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AOO-36873

AIAA 2000-3724Direct Skin Friction Measurement in aRocket-Based-Combined-CycleScramjet CombustorTheodore B. Smith, Joseph A. SchetzVirginia Polytechnic Institute & State UniversityBlacksburg, VAandTrong T. BuiNASA Dryden Flight Research CenterEdwards, CA

36th AIAA/ASME/SAE/ASEEJoint Propulsion Conference & Exhibit

17-19 July 2000Huntsville, Alabama

For permission to copy or to republish, contact the American Institute of Aeronautics and Astronautics,1801 Alexander Bell Drive, Suite 500, Reston, VA, 20191-4344.

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AIAA-2000-3724

DIRECT SKIN FRICTION MEASUREMENT IN AROCKET-BASED-COMBINED-CYCLE SCRAMJET COMBUSTOR

Theodore B. Smith*, Joseph A. Schetz1,Department of Aerospace and Ocean EngineeringVirginia Polytechnic Institute and State University

Blacksburg, VA 24061

and

Trong T. Bui*,NASA Dryden Flight Research Center

Edwards, CA 93523

ABSTRACT NOMENCLATURE

A direct-measuring skin friction gage wasdeveloped for the high-speed, high-temperatureenvironment of the turbulent boundary layer in asupersonic combustor. The design is that of a non-nulling cantilevered beam, the head of which is flushwith the model wall and surrounded by a small (0.005in = 0.0127 cm) gap. Thus, the shear force due to theflow along the wall will deflect the head, inducing ameasurable strain in the beam. Finite element softwarewas used along with simple beam theory to analyze theresponse of the beam to an applied shear load. Semi-conductor (piezo-resistive) strain gages were used todetect this strain at the base of the beam. Cooling waterwas routed both inside the beam and around theexternal housing in order to control the temperature ofthe strain gages. The gage was statically calibratedusing a direct force method. The design was nextverified by testing in a well-documented Mach 2.4cold-flow. Results of the cold-flow tests wererepeatable and within 15% of the value of Cf estimatedfrom simple theory. The gage was then installed andtested in a rocket-based-combined-cycle engine modeloperating in the scramjet mode. The free-jet facilityprovided a Mach 6.4 flow with P0 = 1350 psia (9310kPa) and T0 = 2800 °R (1555 °K). Local walltemperatures were measured between 850 and 900 °R(472-500 °K). Output from the scramjet tests wasrepeatable. Those results and possible refinements tothe gage are discussed.

* Graduate Research Assistant, Student Member AIAAf Holder of the Fred D. Durham Chair, Fellow AIAA* Aerospace Engineer, Member AIAACopyright © 2000 by the American Institute ofAeronautics and Astronautics, Inc. All rights reserved.

Cf Skin friction coefficientD Diameter of floating elementG Gap thickness of floating elementGF Gage factorIf Specific ImpulseL Lip thickness of floating elementM Mach numberP, p Pressureq Dynamic pressureT ShearT TemperatureTCR Temperature coefficient of resistanceTCGF Temperature coefficient of gage factorUe Edge velocityu. Friction velocityV VelocityY Ratio of specific heatsp Density

Subscripts

w Wallx Streamwise componenty Transverse component0 Stagnation condition

INTRODUCTION

Skin friction is an important flow quantity forboth scientific and practical reasons. It is an importantpart of boundary layer correlating parameters, such asfriction velocity:

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Skin friction also helps to define aerodynamicperformance of vehicles, particularly in high-speedflows. One application is the combustor of scramjetengines. Moszee and Snyder' identified skin friction asan important, yet poorly understood, quantity in thedesign and optimization of hypersonic air-breathingengines. Dr. F. S. Billig2 performed calculations for ageneric Mach 10 hypersonic, scramjet-powered vehiclecruising at 94,000 ft (28,650 m). Results are shown inTable 1, below. Small changes in Cf are shown tosubstantially impact the specific impulse (If). Therefore,accurate measurement of skin friction drag is crucial inassessing engine performance.

Q0.0010.0020.0030.0040.0050.0060.0080.010

Ir(lbf*s/lbm)167015401410128011601030790560

Table 1: Impact of Skin Friction onSpecific Impulse1

Skin friction measurement techniques includedirect and indirect methods. Indirect methods have beenthoroughly reviewed by Nitsche, et al3. Indirectmethods infer skin friction from other flow quantities,such as heat flux, using an assumed law, such as theReynolds analogy3. For the highly complex flow in asupersonic combustor, these methods are not wellfounded. Direct methods, on the other hand, assumenothing, and are preferable in such a situation. Winter4

documented early direct measurement designs.Dhawan5 first proved skin friction balances forsupersonic cold-flow in 1953, with later effortsincluding those of Hakkinen6, Roensch and Cadwel!7,Alien8'9, and Voisinet10.

The direct method used was that of a floatingelement mounted on a non-nulling cantilevered beam.A floating element that is flush with the wall of the testsection was mounted to the end of the beam. Thiselement senses the shear force due to the flow anddeflects, causing a measurable strain. The strain is

measured at the base of the beam with piezo-resistivestrain gages. These strain gages give the instrument thesensitivity it needed, being able to measure loads lessthan 50 Pa (0.00725 psi) and greater than 4300 Pa(0.624 psi). Shear levels in the range of 100 to 1000 Pawere expected.

Non-nulling designs are less complex thannulling devices, requiring no feedback loops oradditional mechanisms. Consequently, non-nullingdesigns have better time response than nulling versions.The stiffness of the beam ensures that deflections of thefloating element are small enough that misalignmenterrors are negligible. This stiffness requirement leadsto the necessity of using the very sensitive semi-conductor (piezo-resistive) strain gages.

The purpose of the present work was to designa direct measuring skin friction gage for scramjet andRBCC engine tests. The high temperature environmentcomplicates the design process due to the resultingeffects on the survivability, not to mention theaccuracy, of the sensor. Previous work by DeTurris,Schetz, and Hellbaum", and Chadwick, DeTurris, andSchetz12, has proven the use of piezo-resistive straingages in similar skin friction gages. These were donein scramjet combustor flows, requiring water-coolingsystems to maintain constant temperature of the straingages. Water-cooling was also used to match thefloating element temperature to that of the rest of thewall, and thus reduce errors due to the sensor's effecton the flow. These gages established the use of thenon-nulling, floating-element/cantilevered beam designwith actively cooled strain gages in supersoniccombustion flows.

The present design employs a new beam andstrain gage arrangement with internal cooling.DeTurris, et al11, used a pre-manufactured strain gageassembly with cooling lines around the housing only.Chadwick, et al12, used a quartz beam with strain gagesbonded to the base and cooling lines inside the beamand around the housing. The current design employs astainless steel beam with strain gages bonded to thebase and cooling water routed to and from the inside ofthe beam at the strain gage location, as well as coolingof the housing.

GAGE DESCRIPTION

The genera! gage concept is shown in Figure1. The flow causes a shear that is integrated over thearea of the floating element, inducing a deflection in thebeam. The floating element is mounted to a stainless

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Flow FloatingElement

External Water In // \ " External Water Out

Internal Water Out Internal Water In

Figure 1: Sketch of Skin Friction Gage

steel tube with four semi-conductor strain gages at thebase. The strain gages are arranged in two half-Wheatstone bridges, one for each axis of flow.Hypodermic tubes route water through the base and intothe beam, where a plug blocks the flow just above thelocation of the strain gages. Additional 1/16-in (0.159cm) copper tubes route water around the exterior of thehousing. The cavity between the beam and the housingis filled with silicon oil. This oil provides thermalstability, vibration damping, and protects againstpressure gradient effects across the head.

The material and geometry of the head weredesigned to match as closely as possible those of thesurrounding engine model. This is important to ensurethe same tunnel wall conditions exist at the gagelocation. Changes in material properties and/orgeometry can alter the heat flux and temperaturecharacteristics, causing the gage to measure an incorrectskin friction value. Therefore, the floating element andhousing were both made of copper, with the floatingelement having a length of 1.0 in (2.54 cm).

Stainless steel tubing was selected for thebeam for several reasons. First, it is a high-precision

product readily available from tubing manufacturers.Second, the relatively low (compared to copper)thermal conductivity provides some thermal isolation ofthe strain gages. Thirdly, the interior of the tubingcould be drilled and tapped. This allowed a set-screwto be used as a plug to block the internal water flowfrom travelling to the top of the beam.

Another set-screw was used to thread thefloating element to the beam. Both set-screws werethread-locked with high temperature ceramic adhesive.High temperature ceramic adhesive was also used forbonding the beam into a brass base. Brass was chosenfor the base to provide further thermal isolation.Several holes were drilled through the base to provideaccess for the internal water lines, electricalconnections, and mounting screws. An O-ring groovewas also incorporated to provide a sealing mechanism.

The strain gages chosen were of the semi-conductor (piezo-resistive) type. These offer severaladvantages over the conventional metal-foil type ofstrain gages. First, semi-conductor strain gages havemuch higher gage factors (GF) compared to foil straingages. Typical values of GF for foil gages are around2, whereas semi-conductor strain gages have GF up to150. This allows semi-conductor strain gages tomeasure very small strains, which in turn allows thisskin friction gage to have a very stiff beam. Semi-conductor strain gages can also be made very small; thegages used in this application have a length of 0.060 in(0.152 cm) and a width of 0.008 in (0.0203 cm). Thedisadvantage of using semi-conductor strain gages istheir sensitivity to temperature changes, which makesthe active cooling system of the skin friction gage veryimportant.

Error sources in floating element skin frictiongages were studied parametrically by Alien8'9 Errorsare caused by misalignment of the sensing head withthe surface of the tunnel due to protrusion or recession,and pressure gradients acting on the surface and lip ofthe floating element. Alien investigated the effects ofgap size, lip size, and floating element diameter onthese errors. The results of this study indicated that alarge gap-to-diameter (G/D) ratio actually decreases thesensitivity to protrusion errors. Lip-to-diameter ratio(L/D) was also discussed in terms of boundary layerthickness. Alien learned that the effects of gap and lipsizes become negligible for no protrusion or recessionmisalignment. The pressure gradient can cause errorsin two ways. The first is that the resultant normal forceof a pressure gradient is off-center with respect to thecenter of the beam, inducing a small moment in the

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beam. The second is that the pressure difference actingon the lip can cause a force to be measured in thedirection of the pressure gradient.

The floating element has a diameter of 0.300in (0.762 cm) at the surface. A gap of 0.005 in (0.0127cm) surrounds the floating element. This provides aG/D ratio of 0.0167, which is well above therecommended minimum of 0.005. At a distance 0.030in (0.0762 cm) from the top, the floating elementdiameter decreases to 0.290 in (0.737 cm). This lipgives the floating element a L/D ratio of 0.10, which islarger than that used by DeTurris, et alu and Chadwick,et al12. Most importantly, when constructed, thesegages were assembled with a protrusion or recession ofless than 0.001 in (0,00254 cm). This makes gap size,lip size, and protrusion errors negligible. Filling thegap with silicon oil, which is an incompressible fluidand presents a smooth surface to the flow, makes thepressure gradient error due to force acting on the lipnegligible here.

The silicon oil used to fill the internal volumebetween the beam and the housing has severalimportant functions. First, as mentioned above, the oilprovides a continuous surface to the flow. In addition,it helps to eliminate pressure gradient errors. It alsoprevents flow into and out of the gap, which can disruptthe boundary layer near the gage. Furthermore, the oilprovides thermal stability and isolation for the straingages. Lastly, the fluid provides dynamic damping ofvibrations inherent in test facilities.

The beam was designed using simple beamtheory. The final dimensions included outside diameterof 0.158 in (0.400 cm), internal diameter of 0.102 in(0.260 cm), with the thickness at the flats of 0.109 in(0.278 cm). Given the expected shear level, the strain atthe base of the beam was calculated, accounting for thedifferent cross sectional shapes and materials. Totaldeflection of the head was also calculated, sinceexcessive deflection can cause misalignment errors.Protrusion of the head into the flow due to tilting is alsoa concern when the deflection is large. Therefore, afterthe preliminary configuration was developed, thedesign was then input into a finite element program inorder to verify the expected deformation and strain.

A 3-D solid model of the cantilever beam andthe floating element was created using AutoCAD andimported into COSMOS/Works, a solid modeling/finiteelement analysis program. The static stress analysiswas conducted using the FFE (Fast Finite Element)solver option in COSMOS. This is a fast, robust, and

accurate finite element solver that could providesolutions significantly faster than standard finiteelement direct solvers. The finite element mesh used inthe analysis was generated automatically byCOSMOS/Works using the default high qualitymeshing option. A tangential shear stress of 0.116 psi(800 Pa) was applied to the top surface of the floatingelement, while the base of the beam was held rigidlyfixed.

The FEM-predicted horizontal deflection ofthe beam and the floating element is plotted as colorcontours in Figure 2. Also shown in this figure is thedeformed shape of the beam, magnified by 1800 times.The FEM results agreed well with those obtained fromthe beam theory. The predicted maximum deflection ofthe head was 0.00017 in (0.00043 cm) according to theFEM analysis, and 0.00016 in (0.00041 cm) accordingto beam theory. The resulting maximum protrusion ofthe head into the flow was then only 0.00001 in(0.0000254 cm), which was considered negligible.

UX

r1.7266-004

ft.582e-004

.4386-004

,1.2946-0041.1506-0041.0076-004

8.6276-005

7.1886-005

5.750e-0054.3116-0052.873e-0051.4346-005U. 2326-008

Figure 2: Magnified FEM-predicted Deflection(Color Contours Denote Horizontal Deflection in

Inches)

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EXPERIMENTAL PROCEDURES

The central part of the skin friction sensor isthe strain gage arrangement. For each of the two axesof measurement, two strain gages were applied at thebase of the beam in a half-Wheatstone bridge, one eachon opposing sides. Excitation, bridge completion andbalance, amplification, and filtering were all providedin one unit with Measurements Group 2310 signalconditioning amplifiers. A DC excitation voltage of +5volts was used to power the strain gages, and the signalwas filtered at 10 Hz. For tests in the Virginia Techsupersonic wind tunnel, data was recorded on a PC witha 12-bit A/D card at a rate of 100 Hz. Testing at thefree-jet facility was done at 20 Hz on facilityequipment.

Calibration

The skin friction gage was calibrated byapplying a direct force to the floating element along theaxis of measurement. The gage was first clamped in ahorizontal position, and then weights were hungvertically by string from the surface of the floatingelement. Output in millivolts was recorded from amultimeter, and a linear regression was done to obtain acalibration curve. A sample calibration curve ispresented in Figure 3. Due to slight imperfections inthe machining of the beam and strain gage application,complete orthogonality of the two axes is impossible.Hence, there is a slight output from the "off-axis"Wheatstone bridge. Repeated calibrations showed nochange when the silicon oil was added. Also, afterallowing the strain gages to reach a steady temperature,no change in calibration was seen with internal coolingwater flowing. Calibration of the transverse axis wasaccomplished by rotating the gage 90 degrees.

Cold-Flow Tests

In order to verify the skin friction gage design,tests were conducted in the Virginia Tech 9-in x 9-insupersonic blow-down wind tunnel. This facilityprovided a well understood flow to compare the gageresults against. These tests also demonstrated theversatility of the gage design, being able to operate inmore than one type of facility. Nominal tunnelconditions were set at Mach 2.4, with P0 = 50 psia (345kPa), and T0 = 540 °R (300 °K). Run times weretypically five seconds long. The boundary layercharacteristics of the tunnel floor plate had beenpreviously measured, allowing the skin frictioncoefficient to be calculated by simple methods.

Figure 3: Skin Friction Gage Calibration

The skin friction gage was located in a sectionof the flat wall just downstream of the divergingsection. Cooling to the external lines was not necessaryfor this cold-flow test. Cooling water was supplied tothe internal lines at tap pressure (80 psi, 552 kPa) toprovide thermal stability to the strain gages, thoughlater tests proved that this was also not necessary. Datawas recorded at 100 Hz on a PC via a 12-bit A/D card.The signal was filtered at 10 Hz.

RBCC Freeiet Tests

Tests of the skin friction gage in its intendedenvironment took place in the combustor of a RBCCengine. The engine operated in scramjet mode in theLeg IV freejet facility at the General Applied ScienceLaboratory (GASL). Nominal freestream conditionsincluded M = 6.4, P0= 1350 psia (9310 kPa), and T0 =2800 °R (1555 °K). These conditions were achieved bypreheating the air in a pebble bed, and further heatingthe air by vitiation. Run times were typically 30seconds long, with changes occurring periodicallythroughout the test cycle. The first few seconds of testtime consisted of allowing the forebody cone to traverseuntil the proper inlet condition was achieved. This wasfollowed by several seconds of steady, un-fueled flow,after which fuel was injected from different locations insequenced stages.

The skin friction gage was mounted into anoblong copper plug, which was located along thecenterline of the engine, downstream of several fuelinjectors. Thermocouples near the gage measured walltemperatures reaching 850 to 900 °R (470 to 500 °K).Cooling water was supplied to both the internal andexternal lines from a manifold at 80 psi (552 kPa).

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Data was recorded on facility equipment for about 150seconds at 20 Hz.

EXPERIMENTAL RESULTS

Repeatable results were obtained in both testenvironments. Time response of the gage was not anissue, given the relatively long run times. The gageswere able to capture the significant features of the flow,including the passing of shocks. Another importantfeature of the output is the return of the signal to zero atthe end of each run. This indicates that there was nothermal shift of the output, and that the cooling waseffective.

The skin friction gage directly measures thewall shear in the wind tunnel. For all of the windtunnel tests presented here, a skin friction coefficienthas been calculated by dividing the wall shear by thedynamic pressure, q, of the flow, using

/-> _

"

where q is given by

= l_2

pressure flow in this facility has been studiedextensively, providing good baseline data.

Results of typical Virginia Tech supersonicwind tunnel runs are shown in Figures 4 and 5. Tracesfor both the stream-wise (X-direction) and cross-stream(Y-direction) axes are shown, as well as the totalpressure history. The violent start-up and shutdown ofthe tunnel are clearly registered as spikes in the skinfriction gage output, as shocks and vibration pass. Thereturn of the signals to zero at the end of the run canalso be seen. This reveals that no thermal drift of thestrain gages occurred during the run.

= -yPM2

2

Time (s)

Figure 5: Supersonic Tunnel Run 3

For the Virginia Tech tests, a pressuretransducer near the skin friction gage measured thestatic pressure, P, and both y and M were known. Itshould be noted, however, that for the RBCC tests thedynamic pressure inside the combustor was unknown.Therefore the shear was normalized using the dynamicpressure at the exit of the freejet nozzle. Because ofthe total pressure losses inside the inlet and due to heataddition, the dynamic pressure inside the combustorwould be much lower than the freestream dynamicpressure.

Virginia Tech 9-in x 9-in Supersonic Wind TunnelTests

The skin friction gages were first tested in a well-known supersonic cold-flow. The Virginia Tech 9-in x9-in supersonic blowdown tunnel was used at nominalconditions of M = 2.4, P0 = 50 psia (345 kPa), and T0 =540 °R (300 °K). The boundary layer of the constant-

o

0.004

0.003 -

0.002 -

0.001 -

-0.001

........ 75

f - 2510

Time (s)

Figure 4: Supersonic Tunnel Run 4

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Run3456789

Cfx0.001760.002260.002090.001700.001750.001770.00121

Cfv0.000700.000030.000890.000890.000990.000940.00121

cf0.001890.002260.002270.001920.002010.002000.00173

Table 2: Supersonic Tunnel Results

Table 2 summarizes the results of the VirginiaTech cold supersonic tests. The skin friction coefficientfor this flow has previously been estimated to be0.0020. The results show agreement within 15% of thisnominal value of skin friction coefficient. The flowwas nearly two-dimensional, with only a small cross-stream component of shear.

Windows in the Virginia Tech supersonicwind tunnel made it possible to monitor the gage duringthese tests. Of particular interest was to note theamount of oil that flowed from the gap of the gageduring a run. Since the gages will not work when thereis not enough oil in them, it was important to know howmuch oil the gage might lose in a run. Small streaks ofoil could be seen forming downstream of the gage. Theamount of oil lost by a gage was found to be importantonly after several runs. Tests conducted after asignificant portion of oil had been lost showed muchworse results. This knowledge gained significancewhen installing the gages into the freejet facility, whereaccess was extremely limited, and monitoring of gagebehavior was impossible.

RBCC Freeiet Tests

Tests were next conducted in the Leg IVfreejet facility at GASL in which an RBCC enginemodel was installed. A pebble-bed heater pre-heatedthe air, and the air was then further heated by vitiation,before being expanded in nozzle. This providedfreestream conditions of M - 6.4, P0 - 1350 psi (9310kPa), and T0 = 2800 °R (1555 °K).

Figures 6 through 8 illustrate the results ofthese tests, while Table 3 summarizes the results. Dueto previous damage to the gage, there was only one axisof measurement functioning for these tests. However,since the gage was aligned along the centerline of themodel, the flow should have been nearly two-dimensional.

Centerbodyin position

0.0025 -

0.002 -

c' 0.0015 -

0.001 -

0.0005 -

(fa

i|i

Fueling change

\\\

s J:\,

\

5hut down

^

jjj

0 - ———————————————— I ————— : —————————— I ———————————————— < ———————————————— i

50 60 70 80 90

Time (sec)

Figure 6: RBCC Run 21

0.0035- —

0.003-

0.0025-

0.002-

0.001 5-

0.001 -

0.0005-

Start-up Centerbody Shutdown

/in position ,_ ,. /i Fueling oiianges /\ ^^ \ ̂ /

—— ̂ V

-HH—

>tf\\\\Vv v ,̂

\

—

T

0 T ——————— - - , - - - - - —————— , —————————————

50 60 70 80

Time (sec)

Figure 7: RBCC Run 22

Centerbody Fueling changes Fueled unstartin position \ /^^^^X Shutdowr

0.0025 -

cf 0.0015 -

0.001 -

0.0005 -

0 -4

(U.Anlf**\^V-x,

V A

w,x^ \0 50 60 70 80

Time (sec)

Figure 8: RBCC Run23

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Run212223

Twstart (Pa)485525520

q(Pa)170775166600188630

*-f,start0.002840.003150.00276

Table 3: RBCC Results

These results again show good repeatability.Though not apparent in all of these graphs, the outputreturned to zero at the end of each run, indicating littleor no thermal zero shift. For each run the skin frictioncoefficient starts off at a value of Cf = 0.0027-0.003while the centerbody moves into position and beforefuel is added, and then steadily falls to minimum valuesof Cf = 0.0007-0.0015 after fueling begins. Steadyportions can be seen between fueling changes in someinstances, though lags appear to occur between thenominal change in fueling scheme and the change inoutput.

From previous experience in supersoniccombustion flows, it was expected that the skin frictioncoefficient would increase when fuel was added to theflow. At first glance, this does not appear to be the casein these tests. However, it is important to remember theorigin of the normalizing dynamic pressure. Thedynamic pressure used to convert shear stress into thedimensionless skin friction coefficient was obtained atthe exit of the freejet nozzle, not the combustorboundary layer edge, in these tests. Measurement orcalculation of dynamic pressure inside the combustorwas not possible here. Total pressure losses in the inletwould mean a lower dynamic pressure in the combustorthan in the inlet, which would remain constant duringsteady, unfueled flow. When fuel is injected, though,the accompanying heat addition would cause furthertotal pressure losses, in turn causing the combustordynamic pressure to decrease, so that during the portionof these runs where the shear level decreases, the actualdynamic pressure in the combustor decreasessimultaneously. A factor of 1/6 is a very roughestimate of the ratio of combustor dynamic pressure tofreestream dynamic pressure. Therefore, the actual skinfriction coefficient in the combust might well beincreasing or remaining nearly constant.

MEASUREMENT UNCERTAINTIES

The greatest error source for these experimentswas the temperature sensitivity of the strain gages.Temperature sensitivity of strain gages is describedwith two coefficients- TCR (temperature coefficient ofresistance) and TCGF (temperature coefficient of gagefactor). The change of strain gage resistance with

temperature would cause an output, though no strainwas present. Since TCGF is negative, the change ofgage factor with temperature would cause thesensitivity of the strain gage to decrease, therebyaltering the calibration slope for the skin friction gage.

The water cooling system was intended toprevent heating of the strain gages by maintaining aconstant temperature. The effectiveness of the coolingsystem can be judged by the absence of zero driftduring a given test. For these tests, the signal returnedto the initial value within ±4%.

Pressure gradients were not measured in thesetests. The use of silicon oil in the gap made errors dueto pressure gradients acting on the lip negligible.Assuming a pressure gradient of 1.5 psi/ft (34 kPa/m),the resulting error due to off-center normal force is lessthan 0.25%.

Uncertainties due to the variousinstrumentation components are also small. Thecalibration of the skin friction gages had goodrepeatability and linearity, with virtually no hysteresis.The temperature sensitivity of the strain gagesdominates the uncertainty analysis, resulting in anoverall uncertainty of 10-15%.

CONCLUSIONS

These experiments once again validate the useof the non-nulling, floating-element-cantilever designfor direct skin friction measurement. FEM was used toverify the beam design. The gage was tested in a coldsupersonic flow before being entered into the scramjetcombustor of a RBCC engine.

Results in the cold-flow supersonic tunnelindicated skin friction coefficients within 15% ofprevious estimates. Tests in a RBCC engine installed ina freejet facility showed wall shears in the expectedrange (300-500 Pa). Skin friction coefficient based onfreestream dynamic pressure decreased with fuelinjection (local dynamic pressure was unknown).

Internal and external cooling lines used tocontrol the temperature of the strain gages provedeffective. Future improvements to this cooling systemwill decrease the uncertainties associated with thetemperature sensitive strain gages.

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ACKNOWLEDGEMENTS

The work described above was supported bythe NASA Dryden Flight Research Center underresearch grant number NAG 4-158. The authors wouldlike to thank Scott Thomas of the NASA GlennResearch Center, as well as Dan Cresci and the staff ofthe GASL Leg IV facility for making the RBCCcombustor skin friction tests possible.

REFERENCES

1. Moszee, R., and Snyder, C., "A PropulsionDevelopment Strategy for the National Aero-Space Plane," AIAA-89-2751, 25th JointPropulsion Conference, July 1989.

2. Billig, F. S., private communication, 1998.

3. Nitsche, W., Haberland, C., and Thunker, R.,"Comparative Investigations of Friction DragMeasuring Techniques in ExperimentalAerodynamics," ICAS-84-2.4.1, 14th ICASCongress, Sept. 1984.

4. Winter, K. G., "An Outline of the TechniquesAvailable for the Measurement of SkinFriction in Turbulent Boundary Layers,"Progress in Aerospace Sciences, Vol. 18,1977, pp. 1-57.

5. Dhawan, S., "Direct Measurements of SkinFriction," NACA Report 1121, 1953.

6. Hakkinen, R. J., "Measurements of TurbulentSkin Friction on a Flat Plate at TransonicSpeeds," NACA Technical Note 3486, Sept.1955.

10. Voisinet, R. L. P., "Combined Influence ofRoughness and Mass Transfer on TurbulentSkin Friction at Mach 2.9," AIAA 79-0003,17* Aerospace Sciences Meeting, Jan. 1979.

11. DeTurris, D. J., Schetz, J. A., and Hellbaum,R. F., "Direct Measurements of Skin Frictionin a Scramjet Combustor," AIAA 90-2342,26th Joint Propulsion Conference, July 1990.

12. Chadwick, K. M., DeTurris, D. J., and Schetz,J. A., "Direct Measurements of Skin Frictionin Supersonic Combustion Flow Fields,"ASME 92-GT-320, International Gas Turbineand Aeroengine Congress and Exposition,June 1992.

7. Roensch, R. L., and Cadwell, J. D., "DirectMeasurement of Skin Friction in a HighReynolds Number Supersonic Blow-DownWind Tunnel," Douglas Paper 1728, DouglasAircraft Company.

8. Alien, J. M., "Systematic Study of ErrorSources in Supersonic Skin-Friction BalanceMeasurements," NASA-TN-D-8291, Oct.1976.

9. Alien, J. M., "Improved Sensing Element forSkin-Friction Balance Measurements," AIAAJournal, Vol.18, Nov. 1980, pp. 1342-1345.

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