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Copyright© 1997, American Institute of Aeronautics and Astronautics, Inc.

AIAA Paper 98-0703

TRANSITION ON A THREE-ELEMENT HIGHLIFT CONFIGURATION AT HIGH REYNOLDS

NUMBERS

by

Arild BertelrudAnalytical Services & Materials, Hampton, VA

23666

ABSTRACTAs part of a high-lift flow physics experiment athigh Reynolds numbers to be used for codevalidation, the location and extent oflaminar/turbulent transition has been determinedfor a wide variety of conditions (more than 90combinations of configuration, Reynolds number,Mach number and angle of attack). A databasehas been established to make the data accessible.In the present paper, the methodology developedto determine the transition region extent andsearch for possible separated regions on the modelis described.The transition characteristics foundare described with emphasis on Reynolds numbereffects. Features of importance for the use ofexperimental transition data as input forcomputational codes are also included with adiscussion of the validity of two-dimensionaltesting and matters regarding accuracy andrepeatability.

1. INTRODUCTIONPrediction of the performance and flow physics ofmulti-element, high-lift configurations has beensubject to an extensive effort over the past decades(Dillner et al.*, Valarezo^). Early codes combinedinviscid panel methods with boundary layertechniques for confluent flows and utilizedempirical correlations to arrive at a prediction ofthe flow properties. While most of these codeswere two-dimensional, some of the early codeswere set up to deal with the high-lift problem forinfinite swept wings. This approach did notrepresent the true three-dimensionality since the

Copyright © 1998 by Arild Bertelrud. Publishedby the American Institute of Aeronautics andAstronautics with permission.

spanwise gradients were missing, but some three-dimensional flow features such as attachment-line transition and the possibility ofrelaminarization could be taken into account.

1.1 Code validationA critical element in the development of CFD(Computational Fluid Dynamics) methods is thevalidation of assumptions and computationalmodels through experimental documentation offlow characteristics. As discussed by Cosner^,proper code validation requires an experimentallyexhaustive multitude of investigations.Even in two-dimensional cases, the flow around ahigh-lift configuration is unsteady and three-dimensional with strong interaction between theelements.

Two-dimensional high-lift experiments havesome essential characteristics:-The three-dimensionality is, in general, acombination of two effects:

1. Streamwise displacement of features, orskewing. This essentially means that stagnationpoints, suction peaks and transition/separationare displaced in the streamwise direction due to avarying spanwise load distributions.

2. True three-dimensionality with a spanwiseflow component, due to edge effects, brackets, etc.- Even at Mach=0.2, the flow over one or more ofthe configuration elements may be transonic. Inthe present case the slat is sonic while the typicalMach numbers over the flap are of order 0.3.- Transition on the suction side (as well as the riskof separation bubbles) is highly dependent on thepressure gradients, while the pressure sidetypically exhibits close to zero-pressure gradientflow with the result that laminar flow mayprevail far back on any element.-Extensive cove flows, in the present case on theslat and main elements, cause interpretationproblems due to the effect on transition and asuncertainty in the interpretation to full scale ofthe model's Strouhal numbers and Reynoldsnumber.

A truly two-dimensional high-lift experimentdoes not exist, and it is necessary to make anassessment and identify the conditions where an'approximate' two-dimensionality is attained.Given the reference conditions, the detailed

1


agreement between physical reality and theprediction is essential. This agreement may beconsidered in three levels:

- Pressure distributions- Mean flow field- Turbulence characteristics

Unless all three are validated separately, thecomparison is inadequate and inconclusive.

Agreement in pressure distributions (and hence liftcoefficients) is possible in most cases of interest, i.e. flows without massive separation. However,Mach number, Reynolds number and configurationchange trends may not even be of the correct sign,if the local flow, mean flow and turbulence, is notunderstood.

Experimentally, documentation of the mean flowis feasible but in general carried out for a limitednumber of locations. Transition is an importantparameter since it determines both thesensitivity to separation for the flow over theelements, and has a profound influence on thecharacteristics and strength of attached andseparated shear layers. Compressibility,l o n g i t u d i n a l c u r v a t u r e , l a t e ra lconvergence/divergence and spanwise non-uniformity of transition are important both for theevaluation of the two-dimensionality of theexperiment as well as for the choice of equationsto use and turbulence models to employ.

The turbulence models used in CFD codes quiteoften contain non-physical or hard-to-measureparameters, adding to the complexity ofvalidating the computational methods. Often it isimpossible to determine whether the experimentor the computational trend (or both) are wrong,since the amount of information describing theexperimental flow conditions, in general, is toolimited.

In addition to the ongoing discussion of theapplicability of two-dimensional testing to three-dimensional high-lift configurations, there arefundamental problems regarding Reynolds numberand other scaling effects. Severe limitations existalso in the amount and quality of data that can beobtained in wind tunnel tests.

Over the past decade, several investigationshave been made to establish experimental

databases of use for code validations: Nakayamaet al.4, Chin et al.5, Dominik6, Spaid and Lynch7

with time-averaged flow field data, Klausmeyerand Lin^ with skin friction data. Comprehensiveturbulence flow field data is not in generalavailable. One exception is Pocheron andThibert9, where both attached and reversed flowon a three-element configuration was documented.

1.2 Transition documentationDevelopment of computational codes in recentyears and comparisons with experimental testcases, have demonstrated that a good result inmost cases requires thorough knowledge oftransition location, whether the CFD methodunder scrutiny employs a transition model or not.Surface hot films have been employed to documenttransition in several cases, such as Nakayama eta I.10 and Kreplin and Hohler11.

In the present experiment two principles arefollowed for the transition documentation:- Documentation of transition locations is intendedfor use as input to computational codes. Providinga transition region definition allows the user ofthe data to make an interpretation depending onthe type of modeling used.- Description and discussion concerns the flowproperties of value for a physical interpretationof the flow fields and flow field differencesbetween configurations known to have differencesin performance.

It is also necessary to characterize the generalflow features in terms of the existence ofseparation bubbles, etc. over a reasonableReynolds number regime, since Reynolds numberscaling from model scale to flight conditions isbased on the underlying assumption that the flowfeatures are not drastically altered.

The transition location must be determined for theflow on both sides of each element. Currently thisis commonly used as input in the codes, and is thusa critical element for code validation. Whenthere are disagreements between the measuredand predicted flow, great care must be exercised injudging whether the mean flow representationand/or turbulence modeling causes thediscrepancy.


In the experiment described, transition wasdocumented on all three elements of theconfiguration through the use of 350 surface hotfilms. The statistical characteristics of thedynamic signals were merged with the pressuredistributions and compared with results fromsimple engineering techniques utilizing theexperimental pressure distributions as input.

2. TESTING AT HIGH REYNOLDS NUMBERSThe two scaling parameters of primary interest forthe investigation are the Reynolds number andthe Strouhal number. For atmospheric tunnels, theupper limit in Reynolds number often is 3-4million. For the current 22-inch model,atmospheric runs corresponds to Re=2.6 million. Ifthis model was representative of a 14 ft. chordairplane wing, testing in NASA Langley's LowTurbulence Pressure Tunnel (LTPT) at a full scaleReynolds number of 20 million could beaccomplished but the unsteady shedding from theelements would be significantly different(Bertelrud and Liandrat1 2). As a consequence themodel test might exhibit flow features notexisting in flight and vise versa.

2.1 Test issuesWhile an experiment in LTPT provides a veryhigh Reynolds number, and thus should be reliablefor Reynolds number scaling, several issues needconsideration to assess the experiment as a codevalidation tool. Among these are:

- Model and instrumentation. An existingmodel was used, which means that the numberand location of pressure taps was suited forperformance optimization, but not necessarily fordocumentation of local flow properties.

- Wind tunnel wall effects and three-dimensionality. While sidewall suction canalleviate the problem to some extent (Paschal eta 1.1 3), there is always a three-dimensional walleffect that has to be considered when evaluatingthe data.

Wind tunnel turbulence. The flow isaffected by both wind tunnel turbulence level,length scale and structure; a low turbulence levelin a wind tunnel does not always correspond to alate transition. Both level and anisotropy dependon the streamwise location in the test section, anda high-lift configuration close to maximum liftalso may alter the turbulence characteristics from

the empty-tunnel values. Depending on the scaleof separated regions or vortices, significantinteractions may exist. A comprehensivedocumentation of the flow quality of importancefor high lift configurations includes both meanflow and turbulence characteristics, at thelocation of the model. Most often very limitedinformation is available. For the LTPT, Stainbackand Owen14 measured the turbulence levels andintegral scales both in the test section and at othercritical locations of the wind tunnel circuit. Theresults varied with tunnel pressure and Machnumber, and the test section turbulence lengthscales were found to be of order 3 to 10 ft.,containing measurable turbulence up to 5 kHz, andturbulence levels u'/U = 0.06 to 0.2 % dependingon pressure level. In the present experiment theflow field measurements include Reynolds stressand turbulence documentation.

- Model scale. High-lift configurationscontain several sources of organized vortexshedding, as indicated above. The most dominantmay be assumed to be the transverse vorticitystreets, assumed to emanate from the slat cove andthe main element cove (Kreplin and Hohler11,Savory et al.1^). These scale with the Strouhalnumber rather than the Reynolds number, and inthe current analysis an effort has been made toassess the existence of transverse shedding.

The other type of vorticity of concern isthe longitudinal vortices emanating frombrackets, etc. These interact with the quasi-two-dimensional flow field, and may increase theindicated lift coefficient, since longitudinalvortices, in general, are advantageous (vortexgenerators on wings). However this type ofspanwise non-uniformity is almost impossible tocharacterize in a validation experiment.

- Surface roughness and imperfections.Depending on the unit Reynolds number of thetunnel and the surface finish of the model, thesurface roughness may cause a discrepancy fromfull scale flight conditions. The main roughnesseffect can be anticipated to be caused by the hotfilm sheet with its sensors and leads. Twodifferent issues are of concern: whether or not theroughness causes transition from laminar toturbulent flow, and whether the leads make theturbulent boundary layer a rough surface boundarylayer.


3. FACILITY AND MODEL DESCRIPTIONThe experiment was conducted in NASA Langley'sLTPT over a Reynolds number range from 5 to 15million, based on stowed chord. Figure 1 shows thewind tunnel, which has a 7.5 x 7.5 x 3ft testsection.

The 350 hot films were located in a single row at74 % span. Figure 3 shows how 14 hot films wererun simultaneously. The 294 sensors hooked upcould be tested through three runs.

WAKE SURVEYRAKE STRUT

SIDEW,BLC VENTINGSYSTEM

9 ANTI-TURBULENCESCREENS

Figure 1 Sketch of the NASA LangleyLow Turbulence Pressure Tunnel, LTPT.

Several configurations were investigated. Herethe results observed for one of these, 30P/30N, arediscussed.This configuration is defined by the slatand flap overhangs and gaps; both elements beingdeflected 30 degrees, as illustrated in Figure 2.

Smmdgtel &pe

Figure 3 Sketch of anemometer system hardware

The experimental data has been organized in adatabase. This includes all raw hot film data,18GB (gigabytes) binary, and raw pressure dataalong with reduced data and access codes in aself-contained form as illustrated in Figure 4. Inthe database, data from a wake rake mountedbehind the model has also been included.

-0.2 -0.0

Co nil gu ration:Slat gap

Slat overhangSbt angleFlap gap

Flap overhangFlap angle

30 030 N195%

-Z5 %30degL27*tt 25%

30deg

0.2 0.4 0£ 0.8 I JO

Figure 2 MDA 3-element high-lift model Figure 4Database organizalon

The model has pressure taps mainly in the mid-span region of the three elements, with oneadditional chordwise row on the flap close to thehot film location. Four spanwise rows exist, one onthe slat, one on the main element and two on theflap.

Details on the measurements and the transitionanalysis technique can be found in Bertelrud etal.l^, details regarding the experiment can befound in internal reports and the database whileRumsey et al.1'7 give comparisons betweenNavier-Stokes computations and a subset of thedata.


3.1 Repeatability of reference conditionsThe repeatability of angle of attack, Reynoldsnumber and Mach number is, throughout the entireMach number Reynolds number and angle of attackregime:

Mean Standard dev. Max. diff.Mach 0.0005 ±0.0010 ± 0.003Reynolds[%] -0.02 ± 0.58 ± 2Alphatdeg] 0.0025 ±0.0094 ±0.03

It can be anticipated that the accuracy of theparameters is considerably different. Experiencein other tunnels indicate that there may be subtledifferences in the way the tunnel flow conditionsresult in differences in the performance,especially close to maximum lift.

3.2 Pressure distributionsThe pressure distributions were obtained to verifythe configuration performance, merge with thehot film data and allow identification ofstagnation points and suction peaks.16

-0.20 -0.16 -0.12 -0.08 -O04 -04X) 0.04 0.08 0.12 0.16

Figure 5a Slat pressure distribution. M=0.2 Re=9mill.Figure 5 illustrates typical pressure distributionson the three elements at Mach=0.2 and Re=9million. The plots are done as function ofcoordinate along the surface, which is the naturalcoordinate system for a boundary layer, s/c is non-dimensionalized by stowed chord (22 inches), isnegative on the lower side, and s/c=0 for y/c-0 instowed coordinates. The hot film data wasobtained through use of three repeat runs,typically producing a repeatability in pressurecoefficient, Cp, not exceeding 0.03 deviation inthe suction peak region of the main element andsubstantially lower other places.

Angleof attack

8 deg16deg21 deg

Figure 5b Main element pressure distribution.M=0.2 Re=9mill

Figure 5c Flap pressure distribution M=0.2 Re=9 mill

Figures 5a through 5c illustrates that the slat hasthe highest pressure coefficients, increasing withangle of attack, while the flap has almostconstant suction peak values decreasing withincreasing angle of attack. On the main elementthe location where the slat fits causes a very clearpattern in the pressure distribution which does notnecessarily indicate the existence of a separationbubble.

While the lift coefficient for each element wasdetermined from integration of the pressures, theceiling and floor pressures can be integrated to


provide an overall value (as often doneoriginally). Figure 6 shows that the circulation ofthe high-lift configuration influences the entiretest section.

-02

Figure 6 Tunnel ceiling and floor pressure distributions.vSch=0.2 Re=9mill.

SECONDARYCHORDWISETAP ROW

-190 100

X/C [%]

Figure 7a Flap pressure distribution for 16 degreesangle of attack. Mach=0.2 Re=9 million

Figure 7a illustrates that the pressuredistribution at 74 % span, corresponding to thelocation of the hot film sensors, is somewhatdifferent from the mid span data, and that thereis an appreciable variation in the spanwisedirection.Two-dimensionality of the flow wasmaintained through sidewall suction optimized

for 16 degrees angle of attack. At the highestangles of attack there was an appreciablespanwise variation, as demonstrated for the flapflow in Figure 7b. The figure shows the spanwisepressure distribution at two chordwise positions,an illustration that a pressure row in a particularposition may mask the effects of three-dimensionality.

8 deg.x/c-74*8deg.x/c-100*16degji/c-74M16degjc/c-100H21 dsgj/c-74*21 dea-x/c-IOOH21 dtg-No flm.74*

100

Figure 7b Spanwise pressure distributions on flap.Mach=O.Z Re=9 million

2

o-

-Cp.main.82.5% MAINTAP ROW

4d«g8deg12degISdeg19dog21 deg22 dag23 deg

Spanwise location Z/B [%]

0 20 40 60 80 100"igure 7c Spanwise pressure distribution on main element.82.5% stowed chord) Mach=0.2, Re = 9 millionThe spanwise non-uniformity is considerablysmaller on the slat and the main element, asillustrated in Figure 7c. Only for angles of attackhigher than 16 degrees is the variation noticableon the main element.


To locate stagnation points and suction peaks, it isnecessary to interpret the measured pressures asvelocities and Mach numbers as described byBertelrud and Graves^ to obtain smooth curvesusable for engineering estimates. Thecomputational model of the high-lift elementsusually have an order of magnitude betterresolution than the experimentally determinedpressure distributions. Since boundary layers andseparations are sensitive to the pressuredistributions and the changes in pressuregradient, a smooth interpolation routine isneeded.

33 Mach number distributionsAt any given point, the local Mach number isdetermined from the ratio of total pressure andstatic pressure. Upstream of the model, the totalpressure is approximately constant. However, dueto the finite size of the wind tunnel, there will, ingeneral, be a varying static pressure.

Although the reference Mach number isapproximately 0.2, the local Mach numbers on theelements can be much higher, and can reach 0.75on the slat at 16 degrees angle-of-attack. Formaximum lift coefficients, the highest Machnumber is close to unity.

4. DOCUMENTATION OF TRANSITIONREGIONS

In this section elements of the technique used todetermine the transition region are described.They are well-defined statistical entities, incontrast to the estimation of the transition regionitself, which is an interpretation of availableinformation.

4.1 Standard deviationThe standard deviation of the hot film signal iscommonly used for transition detection. Anincrease in amplitude ratio of two to three existsbetween laminar and turbulent flow, whiletransitional flows usually exhibit much higherlevels. The interpretation of standard deviationsis influenced by the mean level of the signal. Thenonlinearity of the hot film response results inhigher sensitivity to fluctuations when the meanshear level is low.

4.2 SkewnessSkewness is a measure of asymmetry of the signal.Typically pure laminar and turbulent flowsexhibit nearly zero skewness. Transitional flowwill correspond to one sign in the beginning oftransition, zero close to where the intermittencyfactor is 0.5, and then a reversal in sign.

4.3 FlatnessFlatness has a value close to 3, for laminar andturbulent flows, while transitional flowscorrespond to large values. Most often two peakscan be foundbut if transition occurs over a shortdistance, it may not be possible to distinguish twopeaks of flatness.

4.4 ScalesThe autocorrelation function is used to provideinformation on the Taylor microscale and theintegral length scale (TennekesandLumley1^). Inthe analysis the scales are used as a rough guideto identify sensors that might exhibit a problemor a region of strong correlation.

4.5 Phase relationsIn the stagnation region of an airfoil, the phaserelations between pairs of films may be used toestimate the location of the stagnation point,since films on either side will exhibit 180 degreephase reversal (Nakayama et al.10). Thistechnique is not always applicable on a multi-element airfoil since the free stream may containstrong quasi-steady flow structures.

4.6 Cross-correlationsThe cross-correlation function was used todetermine the time-delay for maximumcorrelation between pairs of sensors. This wasused to identify stagnation or reattachmentregions, but also to identify the coherent structuresin the turbulent part of the boundary layers. Themaximum correlation between signals and how farfrom zero time delay this maximum correlation isshifted is identified.

Considerable care must be taken in drawingconclusions from cross-correlations between surfacehot film sensors. Since each individual film has adifferent transfer function, the accuracy of crosscorrelations is limited. Reattachment regions, aswell as separation regions, are notoriouslyunsteady, at a shedding frequency that is at bestsingle-valued.


4.7 StationarityStationarity is determined through examinationof all data sets available, and verify whether ornot the statistical values change. Since the lengthof each 1024 sample data set is only 20.5 msec,frequencies below 100 Hz may be present (andobserved) in the data but are not resolved.

5. TRANSITION LOCATIONThe transition region is determined from acombination of the standard deviation, theskewness and the flatness of the hot film signals.The scales are used as a guide along with samplesof the signals to make sure that any noise ordisturbance that might be present is notinterpreted as transition.

Two ground rules for the analysis are:1.) each element is handled separately2.) a merged data file is used per angle-of-attack,containing all chordwise pressure information andall validated hot film statistical data.

The pressure distribution is included in thecomparisons since it provides a good indication ofwhere transition can be expected to occur, and,more importantly, indicates the sensitivity of thehot films. Another feature of the hot film signalsis their behavior in a stagnation point region,where the mean shear approaches zero. In some ofthe figures a higher standard deviation is visiblein the stagnation region. This is mainly due to thenon-linear effects of the hot film. The skewnessand flatness are used to verify that this indeed isnot a transitional signal but a stagnation pointfeature.

The most conclusive test once general regions havebeen determined based on the statisticalparameters, is to examine the signals directly,evaluate the cross-correlation function and makedecisions based on direct signal interpretation.

As explained in Ref.16, the characteristics of hotfilm anemometry, require transition determinationto be based on mean voltage, standard deviations,skewness and flatness. The use of severalstatistical moments alleviated interpretationproblems due to the non-linearity of hot filmsignals, and also enabled distinguishing betweenstagnation regions, transition and separatedregions.

0.10

0.08

0.06

0.04

0.02

0.000.00 0.01 0.02 0.03 0.04 0.05 0.06 0.07 0.08 0.09 0.10

Hgure 8a Transition determination: Standard deviation.M=0.2 degRe=9mffl

0.00 Q02 0.04 0.06 0.08Figure 8b Transition determination: Skewness.Mach=0.2 Re= 9 million

o.io

One example is given for transition on the flapsuction side. Figure 8a through 8c show thedistributions of the statistical moments used todetermine the transition location for each of theelements. Figure 9 illustrates how the transitionregion indication differs depending on the toolused for its determination.

8


0.00 0.01 0.02 0.03 0.04 O.OS 0.06 0.07 0.08 0.09 0.10

Figure 8c Transition determination-Flatness.Mach=0.2 Re=9 million

are accurate to 1-2% chord, except two regionswith very sparse sensors. The naturalunsteadiness of the flow implies that a much moreaccurate determination is not possible, no matterthe sensor separation or averaging length.

6. ANALYSIS OF MERGED HOT FILM ANDPRESSURE DATA

The statistical analysis is based on sample sets of1024 samples length each, averaged overapproximately 15 sets per hot film for each of the294 hot films for each of the (typically) 9 anglesof attack, i.e. roughly 40,000 data sets perconfiguration.

In this section only the flow at Mach=0.2 andRe=9 million is discussed.

SUCTION PEAK

S/0.0 «% | | CL

— — — - Stagnation pointSuction peakTransition startsTransition ends

— "•• ~ COVEReattachment forward~ COVEReattachment rear

Figure 9 Translion region characteristicsMach=0.2 Re=9mill FLAP-suction side

Figures lOa through lOc show the transitionlocation for the three elements as function of angleof attack at a Reynolds number of 9 million. Theslat flow is mainly laminar, while the mainelement has laminar flow over the main part ofthe pressure side. On the suction side, transitionoccurred close to the leading edge, with theexception of 8 degrees angle of attack.

The accuracy of the feature locations is dictatedboth by the separation between adjacent sensors(minimum 0.1 in. or 0.45 % chord) and the type offeature. Stagnation points are in general accurateto 1-2 % chord, suction peaks to 1 % (both based onpressure tap separation), while transition regions

8 10 12 14 16 18 ZO 22 24

Angle of attack [deg]

Figure lOa Transition Mach=0.2 Re=9 million - SLAT

6.1 Slat flowFigure lOa shows the coordinate s/c on theordinate. The apex of the slat is at s/c=0. Theupper trailing edge is located at s/c=0.16 on thetop of the figure. The lower side stretches fromapex down to the cusp at s/c=-0.06. The cove regioncovers from s/c=-0.06 back to the trailing edge ats/c= -0.20. Since the s/c coordinate system wrapsaround the slat s/c=0.16 and s/c= -0.20 are thesame point.


The slat flow stagnation point starts out at theapex (i.e. close to y/c=0, where s/c=0). As theangle of attack increases, the stagnation pointmoves down and back, until stagnation occursalmost on the lower edge of the slat. Meanwhile,there is no suction peak for 4 degrees angle ofattack there is a smooth acceleration all the wayto the trailing edge on the suction side. From 8degrees on, a suction peak appears close to theapex. This suction peak moves very little - of theorder 1 % chord - as the angle of attack isincreased to 23 degrees. The transition regionmoves forward from being at the rear upper edgefor 4 degrees angle of attack, until it occurs almostimmediately behind the suction peak at highangles of attack. It appears that the beginning toend of transition in general is 2 % chord, except for4 degrees where transition starts close to the uppertrailing edge and probably sheds as anintermittent shear layer, possibly causingspanwise non uniformities.

For 4 degrees the standard deviation indicates anon-laminar flow only at the very last sensors.The standard deviation is seen to exhibit a cleartransitional region only at 8 degrees, where itcovers 5-6 sensors. The skewness shows theexpected two-sided deviation while the flatnessover the same sensors gives two same-sign peaks,but with a longer transition region.

The upper side transition for angles of attackhigher than 8 degrees is almost undetectable,indicating that it happens over an extremelyshort distance. This may be explained by the veryhigh adverse pressure gradients where transitionis nearly instantaneous either through attachedbypass transition or through a closed separationbubble.

6.2 Main flowThe flow features found on the main element areillustrated in Figure lOb.

The pressure distributions have virtually nomovement of the low pressure region, only atsuccessively higher pressure coefficient values asthe angle of attack is increased. The suction peakstays close to s/c=0.01, but the values areconsiderably lower than on the slat. The surfacecurvature region close to s/c=0.1 provides a

pressure distribution resembling the shapenormally cause by a reattaching laminarseparation bubble (at high Reynolds numbers).However, for most of the angles of attack, therewas no evidence of reversed flow in the hot filmdata, and transition occurs almost immediatelybehind the suction peak. For 4 degrees, theflatness factor indicates a clear transitionalregion from s/c=0.01 to 0.06. At higher angles ofattack it becomes very short, and is centeredaround s/c=0.02.

— — — — Stagnation pointSuction peakTransition startsTransition endsMinimum pressureTransition startsTransition ends

4 6

Figure lOb Transition Mach=0.2 Re=9 million - MAIN

8 10 12 14 16 18 20 22 24Angle of attack [deg]

On the lower side of the main element, thestagnation point moves back somewhat. A majorfeature is the existence of slight suction peak mid-chord at low angles of attack and an associatedregion of favorable pressure gradient. The suctionpeak moves back and the favorable gradientbecomes weaker as the angle-of-attack increases.The hot film statistics indicate that this causesthe start of transition to move from s/c=-0.33 to -0.4 at the two lowest angles, and jump back forhigher angles. In fact, it appears that the flowover the lower trailing edge of the main elementis transitional for all angles of attack above 8degrees. This may have an influence on the mainelement cove flow and the subsequent flap flow.

10


6.3 Flap flowThe flow features found on the flap are illustratedin Figure lOc.

0.08 -,

0.074

0.06-

S/C 0.05 -

0.04-

0.03 <

0.02-

0.01 -

0.00-

-0.01 -

-0.02 -

1 ——

) ——

—— t

—— {

>Sj

-*4

~ ~ ~ - Stagnation point~~^~~~ Suction peak - main row— O — Transition start— • — Transition end— — — - Suction peak - extra row

HH

r*

\ ——

) ——

—— < k^

J.

~~

"N

- — •

*•-

v_>^>«

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GSiMCaiVI

-qr^•>»,

r»-

ft

4 6 8 10 12 14 16 18 20 22 24

Angle of attack [deg]Figure lOc Transition Mach=0.2 Re=9 million FLAP

The suction peak for the main row of pressure tapsmoves slowly from s/c=0.02 to 0.03 as the angle ofattack is increased, while it is further forward forthe extra tap row. Up to 19 degrees angle ofattack, the changes are most;y in scaling. Forhigher angles of attack it appear that the flow isseparated or near separation. The standarddeviation distribution shows transition in theregion s/c 0.04 to 0.06, except for 23 degrees angleof attack where it is further forward. The highlevel in the rear (s/c>0.20) for 8 degrees isuncertain. For 23 degrees the high level mayindicate simply low shear attached or reversedlow. Neither the skewness nor the flatnessindicates anything other than low shear,attached flow.

On the lower part of the flap, the hot film signalsindicate that the flow is laminar; the non-stationarity found in the rear is associated with afairly low frequent oscillation that may be causedby the main cove vortex shedding.

7. SEPARATION REGIONS AND COVE FLOWSLocal separations exist on the configuration forseveral test conditions, and reversed flow isassociated with the coves for all conditions. In thepresent section, a couple of these cases of reversedflow will be examined.

7.1 Main element separationAn interesting difference between 5 million and 9million is the 'double' intensity peaks, as shownin Figure lla and lib. At Re=5 million the hotfilms indicate a double peak for 4, 19 and 21degrees a double peak. At Re=9 million, there aresingle peaks indicating attached transitioninstead.O.IOi

0.08

0.06

0.04

o.oz

0.00

4 deg8 deg12 deg16 deg19 deg21 deg

0.00 0.01 O.OZ 0.03 0.04 0.05 0.06 0.07 0.08 0.09 0.10

Figure lla Distribution of standard deviation on mainelement for several angles of attack. M=0.2 Re=5 mill.

0.12

0.10

0.08

0.06'

0.04-

0.02

0.00'

4 deg8 deg12 deg16 deg19 deg21 deg

0.00 0.01 0.02 0.03 0.04 0.05 0.06 0.07 0.08 0.09 0.10

Figure lib Distribution of standard deviation on main elementfor several angles of attack. M=0.2 Re=9 mill.

11


In the region at 4 degrees angle of attack, as shownin Figure 12a, there are two regions of non-correlated sensors, if the cross-correlation functionis evaluated for adjacent sensors. The time shiftchanges sign in between the two, taken to be anindication of reversed flow. The interpretation isthat the 5 million case for 4, 19 and 21 degreesangle of attack has a closed laminar separationbubble - an argument not contradicted by themeasured pressure distribution or time traces fromthe sensors covering the region. Figure 12b showstime traces from the region of reattaching flow. Atthe front films (158 and 157) the flow is stillseparated. Further back the along the surface(from film 156 on) the flow has reattached. Asimilar pattern (not shown) exists at the region ofseparation.

Figure 12aData fromSCP /SON, M=0.2, Re=5 mi 11. alpha=4 degMaximumcrosscorrelation on main nose.

7.2 Slat cove flowUtilizing the same correlation technique, it ispossible to identify probable regions ofreattaching flow in the slat cove. The flow ishighly unsteady, and reattachment can be expected to move back and forth in thereattachment region. Figure 13 shows the slatregion at Re=9 million with this feature indicatedalong with the stagnation point and the start oftransition. Throughout the angle of attack range,the flow leaving the lower cusp can be assumed toundergo transition as a free shear layer beforereattaching in the cove. The region ofreattachment in the cove part of the slat islocated at the lower trailing edge for the 4degrees angle of attack. Since the upper side flow

is only (early) transitional the slat wake can beassumed to be asymmetric with strong low-frequency components.

64 128 132Samples (256 sample s=5.12 msec)

256

Figure 12b Signal traces (zero-shifted)corresponding to Figure 12a. M=Q2 Re=5mill.Alpha = 4 deg.

As the angle of attack is increased, the upper flowbecomes turbulent at the trailing edge, and thecove reattachment moves forward so that even theflow from the lower side can be assumed to beattached turbulent. The stagnation point movesback until it sits almost at the cove edge.

y/cTRANSITIONBEGMNNGASFUNCTONOfANGU--OF-ATTACK

I MAIN

POSSBLEREATTACHWNTREGION(NTEXraETED)

I \ STAGNATION POINTI ' „ HOVE«NTWrTH

6 12 16 " ANGU OF ATTACKX/C

-0.10 O.OS

Figure 13 SLAT flow features. M=02 Re=9 mill.

12


0.07

0.06

0.05

0.04

0.03

0.02

0.01•O.ZO -0.19 -0.18 -0.17 -0.16 -0.15 O.14 -0.13 -0.12 -0.11 -0.1 0

Figure 14 Distribution of statistical properties in the cove region.Mach=0.2 Re=9mill Alpha=19 deg.

In Figure 14 the corresponding distribution ofstatistical properties is illustrated. Notice thatthe change in standard deviation does notcorrespond to the flatness minimum. There is noevident way of showing the flow reversal basedon these single-film statistical properties.

8. COMPARISON WITH ENGINEERINGTECHNIQUES

Simple engineering techniques can be used tocompare the experimentally determinedtransition regions with simple formulae. Figure 15illustrates how Bezier curves have been createdbased on the experimental pressures. A simplecompressible integral method (Mann andWhitten^O) has been used to compute the laminarboundary layer properties. Based on these,different empirical formulae (MIchel21) fortransition locations were applied, as exemplifiedin Figure 15 for flow on the main element.

-0.05 -0.04 -0.03 -0.02 -0.0! -0.00 0.01 0.02 OO3 0.04 0.05Figure 15Suction peak regjon of main element Results from engineering criteriaand thehctfilme)q3eriment.Mach=Q.2, Re=9mill,Alpha = 1

9. REYNOLDS NUMBER EFFECTThe Reynolds number effect was documented overfour Reynolds numbers: 5,9,12 and 15 millionsbased on configuration chord, albeit with thehighest Reynolds number data being taken atMach=0.18 instead of 0.2. Also, the highestReynolds number data was only taken at threeangles of attack: 4,8 and 16 degrees.

9.1 Slat flowAt high angles of attack, a very moderate changeof the pressure distribution is found, and a minortransition adjustment occurs. In most of the cases,there is a trend of of transition moving furtherback as the Reynolds number is decreased.

Figure 16 shows the changes in pressuredistributions on the slat for Re=5, 9, 12 and 15mill, at 8 degrees angle of attack. The two issuesdemonstrated in the figure are that for 8 degreesangle of attack the slat pressure distribution doesnot change from Re=9 up to 15 million andtransition moves forward between 5 and 9 million;i.e. for increasing Reynolds number above Re=9million only the length of the transition regiongets shorter.

13


-0.30

-ace -0.04 -ooz 0.00 ao2 0.04 o.os o.oo 0.10 0.12 0.14 o.ie

Figure 16 Reynolds nurrber effect on pressure distribution. SlatM=a2 Alpha =8 deg.

9.2 Main flowFigure 17a shows the transition location on theupper side of the main element as function ofReynolds number at three angles of attack. Thereis a distinct movement forward and shortening ofthe transition region with increasing Reynoldsnumber.The most dominant difference in flow conditions atRe=5 million compared with Re=9 million is themovement of transition on the lower side towardsthe main element cusp/cove, as illustrated inFigure 17b.

• 4deg a m 8 deg a - 16 deg

s/c(

(

-0.1

-0.2 s

^SfcfcTRANSmOTi—— ' —— «-

SUCTION '̂BAK ~*

STAONATJoil POINT

(

1KSUCTION P

STAGNAT

?^KDNPONT

^^~"

0.1

I

< ^$1

STAGNATE

«%^

^*~^

J POINT

"^

10 1!f~' • • • • „ > " " ' '1S~'5 ><> 1Re [million] Re [million] Re (million]

Figure 17a Flow characteristics on the main element,upper side, as function of Reynolds numberfor three angles of attack.

D Transition start, Re-5 million• Transition end, Re-5 million0 Transition start, Re-9 million• Transition end, Re-9 millionA Transition start, Re-15 million2 Transition end, Re-15 minion

-0.7512 16 20 24

Angle of attack [deg]

Figure 17b Transition locations on main element,lower side as function of angle of attackat three Reynolds numbers.

The pressure distribution of the main for Re=5million is barely noticeably different from 9million The transition on the suction side wasfound to have moved somewhat back by 0.05 in s/c,while the transition on the lower side is moveddownstream by 25 % chord, from -0.4 to -0.65 %.Notice that for 8 degrees the end of transition stilloccurs in front of the cove 'entry7.

Comparing with the distributions of standarddeviations obtained at 9 million Reynolds number,there is a marked change, with two distinct peaksexisting around s/c=0.015 and 0.045. As discussedin section 7.1, it is suggested that the two peakscorrespond to a reattaching (closed,short) laminarseparation bubble that does not exist (or is notdetectable) at Re=9 million.

9.3 Flap flowPredominant features are the relatively longtransitional region on the suction side,approximately 4 % chord, and also the largedifference between the mid-span and theauxiliary tap row suction peaks. This isillustrated in Figure 18. The pressure distributionis not noticably altered, while the beginning oftransition on the suction side has moved back byroughly 0.5 to 1 %. In the region 12 to 16 degreesangle of attack transition start did not move atall.

14


The end of transition moved from 7% chord atsmall angles-of-attack and 6% at the high angles,by 1% when the Reynolds number is reduced from 9million to 5 million,o.io-

s/c0.08

0.06-

0.04

0.02

0.00

Re-5 millionTransition symbolsas Figure 17b

Re-9 million

*•-

-•-

_—~~—» _MAIN_ROW ^

"STFWROW" —

Angle of attack [deg]8 12 16 20 24

Figure 18 Transition locations on flap, upper side asfunction of angle of attack at three Reynolds numbers:Re=5 (squares), 9 (circles) and 15 (triangles) million.

10. ASSESSMENT OF EXPERIMENT VALIDITYIn this section some of the conditions that mayaffect the validity and usefulness of the data arediscussed. It is important to distinguish betweenlocal and global disturbances: local disturbancesalter the flow only at the disturbance location andwill not have any measurable effects downstreamor to the sides, global disturbances (e.g. trippingtransition) will have effects that may jeopardizethe test as a code validation case.

10.1 Film sheet effectThe hot films were deposited on separatepolyamide film sheets wrapped around eachelement. The sheets were 0.002 in. thick and theglue layer thickness was approximately 0.002 in.,i.e. a total of 0.004 in., or roughly 0.02 % chord.

A number of concerns exist:1. Surface roughness. The leads were deposited onthe sheets, and in certain cases their roughnessmay affect the turbulent boundary layer.2. Sheet bubbling occurred at least during the lastpart of the test over the flap region.3. The sheet edge was smoothed initially, butthere may be some effects both initially and asthe test progressed.

10.2 Three-dimensionalityAs described earlier, it is necessary to distinguishbetween at least two different three-dimensionaleffects, spanwise non-uniformity and three-dimensional flow with appreciable crossflow.

The main information available to assess thethree-dimensional effects is the spanwisepressure distributions on the elements along withpressure distributions from the secondarychordwise row on the flap.

The pressure distribution on the slat is uniform.There is no indication its non-uniformity should bestrong enough to influence the main element as faras transition location is concerned.

On the main element (Figure 7c) up to 16 degreesangle of attack, only a slight variation existsbetween 70 and 100 % span. At higher angles, thevariation increases until there is a variation ofalmost 0.6 in pressure coefficient compared to theleft side.

Figure 7b shows the spanwise pressuredistributions for the flap as function of angle ofattack, at Mach=0.2, Re=9 mill. As expected, thespanwise non-uniformity increases with angle ofattack, and is more pronounced on the flap,compared with conditions further forward.

A good indicator on the overall effects on theconfiguration can be found in the 100% (TrailingEdge) pressure coefficients which is a measure ofthe pressure recovery. Inviscid flow should have atrailing edge pressure equal to the stagnationpressure. In reality, for practical airfoils, a valueof around 0.15 -0.2 for the pressure coefficient canbe considered to indicate attached flow. However,if the pressure coefficient goes to zero or becomesnegative, there is a high probability that theflow is separated over the trailing edge. Based onthis it appears that the flow may be marginallyseparated in the rear corner of the flap close to 100% span. Successively, the separated regionappears to grow for increasing angles of attack,until it is felt at the midspan for 22 degrees angleof attack.

15


10.3 Side wall mass-flowThe sidewall suction is performed throu use of asuction box and perforated tunnel sidewalls. Thesuction setting was done to minimize span wisepressure gradients at 16 degrees angle of attack.

10.4 Comparisons with previous testsThe main difference is a difference in referencepressure, corresponding to a shift in Cp values of0.02 to 0.04. At higher angles of attack thedifference increases somewhat. The largestdifferences were observed in the suction peakvalues.The repeatability within the test wasfound to be better than the repeatability from testto test.

11. CONCLUDING REMARKSThe paper contains a discussion of the flow on allthree elements and Reynolds number effectsassociated with transition and separation. Thepurpose is to describe in physical terms a summaryof the near-wall flow conditions, and indicatehow to utilize the information for code validationpurposes.

The assessment of the transition data is that upthrough 16 degrees angle of attack the flow issufficiently two-dimensional to utilize theinformation for code validation purposes. As theangle of attck is further increased, in particularthe flap information becomes tainted by three-dimensionality. Using the experimentalstagnation points and suction peaks as flowfeature locators (rather than the strictgeometrical coordinates), the transition regionscan be identified.

The general flow characteristics of the model aresimilar to earlier tests with the pressuredistributions accurately repated. However thereis an influence of the hot film sheet that has to beconsidered in using the data. The number of hotfilms used had to be very large to obtain transitiondata of acceptable resolution since thetransitional regons were not known a priori.

Several conclusions and comments can be made:-The transition regions have been

established, based on multiple transition criteria.-Localized reattachment regions in the

slat cove and the main cove have been identified.-Even seemingly small difference in

pressure distributions, can be accompanied bysignificant and major differences in viscous flowcharacteristics, such as separations, transitionregion extent and physical characteristics.

-The presence of the hot films increasedthe span wise non-uniformities, especially at thehighest angles of attack on the flap.

The intention is to compare results from a numberof transition criteria to the entire transitiondatabase with the short-term goal of being able toincorporate a simple engineering technique intothe framework of Navier-Stokes codes.

ACKNOWLEDGEMENTSThe work was carried out under NASA contract

number NAS1-96014. The author thanks Eric V.Roback, Sheryl Johnson, John. B. Andersjr., C.B.McGinley and J.H. Morrison for their help anduseful discussions. The author also acknowledgeshelpful comments by S.X. Ying and F. W. Spaid ofMcDonnell Douglas.

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2. Valarezo,W.O.: "Topics in high-liftaerodynamics." AIAA Paper 93-3136, Julyl993.

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7. Spaid, F.W. and Lynch, F.T.: "High Reynoldsnumber, multi-element airfoil f lowfieldmeasurements." AIAA Paper 96-0682 Jan 1996.

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10. Nakayama,A., StackJ.P., LinJ.C. &Valarezo,W.O.: "Surface Hot-Film Techniques forMeasurements of Transition, Separation, andReattachment Fonts." AIAA Paper 93-2918,AIAA 24th Fluid Dynamics Conference, July 6-9,1993,Orlando, FL.

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15. Savory,E., Toy, N., Tahouri,B. and Dalley,S.: "The flow regimes in the Cove regions Between aSlat and a Wing and the Flap on a Multi-elementAirfoil." in Engineering turbulence - Modelingand Experiments (Rodi and Ganic,Eds.), Elsevier,1990.

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18. Bertelrud, A. and Graves, S.: "Analysis ofFlight Flow Characteristics in the Leading EdgeRegion of Swept Wings at Subsonic and TransonicSpeeds" AIAA Paper 94-2152, 7th Biennial AIAAFlight Test Conference, Colorado Springs, CO, June20-23,1994

19. Tennekes,H. and LumleyJ.L.: "A First Coursein Turbulence." The MIT Press, Cambridge, Mass,U.S.A. and London,England, 1972.

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21. MIchel,R.: ONERA Report 1/1578-A (1959)

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