comparative results from a cfd challenge over a 2d three ... · pitching moment was computed with...

NASA Technical Memorandum 112858

Comparative Results From a CFDChallenge Over a 2D Three-ElementHigh-Lift Airfoil

Steven M. Klausmeyer

Wichita State University, Wichita, Kansas

John C. Lin

Langley Research Center, Hampton, Virginia

May 1997

National Aeronautics and

Space Administration

Langley Research Center

Hampton, Virginia 23681-0001

https://ntrs.nasa.gov/search.jsp?R=19970027380 2020-06-28T13:19:29+00:00Z

Summary

A high-lift workshop was held in May of 1993 at NASA Langley Research Center. A

major part of the workshop centered on a blind test of various computational fluid

dynamics (CFD) methods in which the flow about a two-dimensional (2D) three-element

airfoil was computed without prior knowledge of the experimental data. Comparisons

were made between computation and experiment for (a) lift, drag, and moment, (b) lift and

drag increments due to Reynolds number and flap gap changes, (c) pressure and skin-

friction distributions, and (d) mean velocity profiles.

The results of this 'blind' test revealed:

1. There was good agreement between several codes and the experimental results. In

general, the Reynolds Averaged Navier-Stokes (RANS) methods showed less variability

among codes than did potential/Euler solvers coupled with boundary-layer solution

techniques. However, some of the coupled methods still provided excellent predictions.

2. Drag prediction using coupled methods agreed more closely with experiment than the

RANS methods. Lift was more accurately predicted than drag for both methods.

3. The CFD methods did well in predicting lift and drag changes due to changes in

Reynolds number, however, they did not perform as well when predicting lift and drag

increments due to changing flap gap.

4. Pressures and skin friction compared favorably with experiment for most of the codes.

5. There was a large variability in most of the velocity profile predictions. Computational

results predict a stronger slat wake than measured suggesting a missing component in

turbulence modeling, perhaps curvature effects.

Symbols

C

CD

Cf

q

Cp

M.

n

q_

= cruise or stowed airfoil chord

= drag coefficient

= skin-friction coefficient, where Cf = "c,_etq..

= lift coefficient

= pressure coefficient

= freestream Mach number

= distance normal to airfoil surface

= freestream dynamic pressure

Re

X

0_

A

'_wall

Subscripts

max

= Reynolds number based on cruise chord c

= coordinates along the chord direction

= angle of attack

= differential value

= wall shear stress

= maximum value

Introduction

A High-Lift Workshop/CFD Challenge was held at NASA Langley Research Center in

May of 1993. The primary objectives of the workshop were to define the state-of-the-art in

2D multi-element airfoil prediction techniques, to determine the status of high-lift research

in the U.S., and to discuss high-lift flow physics. Conference participants included NASA

and industry.

The main thrust of the workshop centered on a comparison of several sets of computational

results with 2D experimental data at Reynolds numbers of 5 and 9 million. A total of 15

sets of computational results were submitted in the challenge. All computations were

performed without prior knowledge of the experimental results and each challenge

participant received the experimental data only after sending his/her calculations to Langley.

Experiment

The experiment is described in detail by Chin et. al. (Reference 1). A brief description is

repeated below.

The experimental data set was obtained in the Low-Turbulence Pressure Tunnel (LTPT) at

NASA Langley Research Center in a cooperative effort with McDonnell Douglas Aerospace

West. The data includes force and moment data, pressure and skin-friction distributions,

and mean velocity profiles.

2

Integrationof thechordwisepressuremeasurementsyieldedthe lift data, while integration

of the downstream wake profile yielded the drag data. The skin friction data were obtained

using 0.020 inch diameter Preston tubes. Using the Kline and McClintock method, the

uncertainty in CL, maxwas calculated to be approximately _+0.02 (or less than 1% for a CL. ,,_

value of 4.5), while the uncertainty in C o was calculated to be approximately _+0.0010 for

high-lift models (e.g., 2.5% for a typical C D value of 0.0400). 2 Repeatability studies

confuTned these levels. The maximum error in C t was calculated to be less than 6%. 3

The high-lift model investigated is a McDonnell Douglas Aerospace (MDA) 2D, single-

flap, three-element airfoil as shown in Figure l(a). The 11.55% thick super critical airfoil

model spanned the width of the test section (36 inches) and had a reference (stowed) airfoil

chord, c, of 22 inches. The slat chord is 14.48% and the flap chord is 30% of the stowed

airfoil chord. The airfoil was configured in a typical approach/landing configuration with

slat and flap deflections of 30 ° . Two different flap riggings were used in the test as shown

in Table 1. Figure l(b) defines the nomenclature for gap and overhang. The first flap

rigging was designated as 30P/30N by MDA and Geometry A for the workshop. The

second flap rigging had a larger gap and was designated 30P/30AD (Geometry B).

Computed Results and Comparison to Experiment

Two types of computational techniques were used in the CFD Challenge: 1) Reynolds

Averaged Navier-Stokes (RANS) methods and 2) Potential/Euler methods coupled with

integral boundary-layer techniques. Several turbulence models were represented in the

RANS solutions.

Both structured and unstructured grids were used in the solutions. The structured grids

were of the Chimera and block types while the unstructured grids were composed of

triangular elements.

Challenge participants are listed in Table 2 along with code name, legend key, solution

scheme, grid type, and turbulence model.

The challenge consisted of five computational cases that are list in Table 3. The first three

cases were mandatory for each participant and the last two were optional. These cases

represent four different classes of flow conditions for a high-lift system: 1) attached flow

(Cases1, 2, and4); 2) Flap separationat low angleof attack(Cases1and 4); 3) Ct.=_

(Case3); and 4) Stall (Case 5).

Comparison of Computed and Experimental Data

This section contains the computational results of each high-Eft challenge participant plotted

against the experimental data. To facilitate comparison between calculation and experiment,

the computational results were split into two basic categories according to solution

technique: 1) methods in which an Euler or potential solver was coupled to a boundary-

layer scheme and 2) Reynolds Averaged Navier-Stokes methods.

Force and Moment Predictions

Geomet_ A - Re= 5,000,O00

Computed lift, drag and pitching moment coefficients are plotted against experiment for

Geometry A at a Reynolds number of 5 million in Figure 2. Calculations were performed

at o_'s of 8°, 16 °, and 21 °. Coupled method solutions are shown in Figure 2(a) and RANS

methods are plotted in Figure 2(b). The experimental values are the same in the upper and

lower graphs.

This configuration has flow separation near the flap trailing edge around _---8 ° and the flow

reattaches near _---12 °. The experimental lift curve is nearly linear up to ¢x=14 ° except for a

small dip near _---8 ° where the flap separated. This depression in the lift curve corresponds

to a significant drag increase indicated by the peak in the drag polar near CL=3.1. Above

cz=14 °, the lift curve slope suddenly decreases, possibly due to boundary-layer confluence.

Ct_ = occurred near ct=21 o and was followed by a gradual lift drop-off indicating a fairly

mild stall. Drag was not measured above ct=l 6 °, however, a rapid increase near Ct._

would be expected.

Some of the lift calculations agree quite well with the experiment, but in general the

solutions indicate a higher C L. There is a fairly large variation in lift prediction between the

codes and the variation appears larger for the coupled methods than for the RANS

methods.

Drag calculations by the coupled method solutions agreed with experiment much more

closely than the RANS methods and also showed lower variability between codes. In

general, calculations indicate a higher drag than shown experimentally, particularly for the

RANS methods. Drag calculations by RANS methods are usually very sensitive to the

proximity of the outer boundary to the airfoil and to the outer-boundary conditions,

particularly at large lifts. The outer boundary can influence the force vector angle and a

small angle change can lead to very large drag changes. The effect of outer-boundary

conditions can be seen in the solutions labeled _'yle' and 'kyleff. Much closer agreement

with experiment was obtained by modeling the effect of airfoil circulation on the outer

boundary using the point-vortex method ('kyleff). The coupled methods (except the

'drela' calculations) are not effected by outer boundary conditions which may have reduced

variation in drag predictions.

Pitching moment was computed with reference to the 25% chord location with negative

values indicating a nose down moment. Experimental results show a reduction in nose

down moment as angle of attack increased. This was caused by the large increases in

suction on the slat and main-element leading edge and a relatively constant suction level on

the flap. Predictions from the couple methods showed a larger variation among codes than

the RANS codes, similar to the lift case.

Geometry_ A - Re=9.000.O00 (Increased Reynolds Number)

Computed force and moment coefficients for Geometry A at a larger Reynolds number of 9

million are shown in Figure 3 along with the experimental data. In this case the flap

remained attached at the lower t_'s, thus the lift loss near o_=8 ° and the drag increase near

CL=3.1 for the Re=5,000,000 case are not present. The rift curve remains relatively linear

up to about t_=12 ° followed by a gradual rounding. CL_= Occurs near o_=21 ° followed by

a mild stall.

The calculated results show similar trends as in the Re=5,000,000 case. Computed lift was

higher than that shown experimentally. Variation in lift calculation for the coupled methods

was greater than that of the RANS methods. Drag levels at low angles of attack were better

5

predictedby the coupledmethods. In general, there was a large discrepancy between

calculated drag and the experiment. Pitching moment prediction by the coupled methods

had more variation than for the RANS methods.

Geometry. B - Re = 9,000,000 (Flao GaD Increase)

Geometry B represents an increase in flap gap of 0.23% chord. Increasing the gap

produced a large separation region on the flap around cc=16 ° causing a lift loss, a sudden

large drag increase, and a decrease in nose down pitching moment.

The calculations for this case (Figure 4) show similar trends in lift and pitching moment

predictions but are in much better agreement with the experimental drag than in the previous

two cases (Geometry A, Re=5, 9 million). Interestingly, the computed drag values have

not changed significantly from the previous two cases, however, the experimentally

measured drag increased, narrowing the gap between the computed and experimental

values.

Geomet_ B - Re=9,000,O00 (Detailed Lift/Drag/TVloment)

In the previous three cases, calculations were performed at only three angles of attack

(oc=8 °, 16 °, and 21°). A more detailed prediction of the lift curves, drag polar, and

pitching moment curves are shown in Figure 5 along with the experimental data. The

calculations were conducted with a much closer angle-of-attack increment (ct=4 °, 8 °, 12 °,

14 °, 16 °, 19 °, 21 °, 22 °, and 23°). Very few of the codes predicted a lift break-off at Ct_ _

or the lift decrease near ct=16 ° due to flap separation. The drag increase near CL=3.8 alSO

was not predicted. These results seem to indicate an inability of the codes to predict flow

separation.

Increments in Lift and Drag

One of the primary benefits of using computational methods in developing high-lift multi-

element airfoils is to get some idea of the effects of changing gap/overhang between the

various elements and to determine the effects of Reynolds number on airfoil performance.

6

Determininggap/overhangeffectscansignificantlyreduceconfigurationoptimizationtime

in thewindtunnelby narrowingtheelementpositionmatrix, thussavingtime andmoney.

DeterminingReynoldsnumbereffectsis necessaryfor predictionof theairfoil performance

atflight Reynoldsnumbers.

When calculatinglift and drag changesit is very important to predict the sign and

magnitudeof thechangecorrectly. The sign is critical becauseit determineswhethera

changeimprovesor degradesairfoil performancewhich drives the optimization process.

The magnitude is important because it determines the amount of performance improvement

or degradation associated with a given change.

Geomet_ Changes

The experimentally measured rift and drag increments due to increasing flap gap at

Re=9,000,000 are shown in Figure 6. Flap separation at intermediate angles of attack

caused a lift decrease, however, lift increased at low and high angles of attack where the

flow remained attached. Drag increased across the angle-of-attack range, especially at a's

where the flap separated.

Computed lift increments due to an increase in flap gap are given in Figure 7. At tx=8 °, all

the coupled methods predict a lift loss which agrees with experiment, however, the

magnitude of the change is quite different, with the experimental value hovering near the

zero-increment line and the computed values ranging from -0.02 to -0.07. Six of the

RANS solutions also gave a lift decrease, however, four predicted a lift increase. The

coupled methods have less scatter at the low angle of attack than the RANS methods.

Drag increments due to flap gap change are shown in Figure 8. The experiment indicated a

drag increase, however, the computed results are scattered about the zero-increment line.

Reynolds Number Change

Lift and drag increments for Geometry A due to increasing Reynolds number are shown in

Figure 9. Recall that the flap separated at lower ot's for the Re=5,000,000 condition.

Increasing Reynolds number eliminated the flap separation as indicated by the local peak in

7

AC L and the dip in AC D at o_=8 ° in Figure 9. A large increase in lift was also observed

above ct=15 ° with most of the increase coming from the main element. The wake width

from the main element was smaller at the higher Reynolds number, which may have

allowed a slightly larger suction on the flap which caused a higher loading on the main

element.

Computed increments in lift and drag are shown in Figures 10 and 11 along with

experimental values at 8 ° and 16 °. In general, the computed AC L was lower than the

experimental value. However, most of the calculations fall within the +0.04 to +0.06

variation in the experimental data.

A decrease in drag with increasing Reynolds number was predicted by nearly all of the

codes. At t_=8 °, the computed drag decrease was smaller than that observed

experimentally. The experimental value was large due to reduction of flap separation at

higher Reynolds number. The codes may not have predicted flap separation at the lower

Reynolds number, thus missing this effect. The codes predicted an increase in drag

reduction (AC D more negative) as angle of attack increased. Experimental data is not

available past 15 °, however, the trends seem to indicate that AC D becomes slightly more

positive with increasing angle of attack.

Pressure Distributions

Computed pressure distributions for o_=8.12 ° and 21.29 ° at a Reynolds number of 5

million are used for comparison with experiment.

Pressure Distribution$ at Or=8 o

Computed and experimental pressure distributions for the slat, main element, and flap for

the low ot (8 °) flap separation case are shown in Figures 12, 13, and 14, respectively. In

general, computed suction on the upper surface of the slat (Figure 12) was higher than the

experimental values except for three codes which closely agreed with the experimental

8

values.Theslatproducesathrustforcesincethe it is deflecteddownwardandthe suction

surfacefacesforward. Note thatthe 'kyleff solution shows a larger upper surface suction

than the 'kyle' solution. This difference was largely responsible for the drag differences in

Figure 1. Interestingly, the higher suction levels predicted by the codes suggests that the

computed drags would be less than experiment, not more. However, higher computed

suction levels on the flap pull in the opposite direction, partially canceling the slat

contribution.

In general, the coupled methods showed more scatter between codes than the RANS

methods. The RANS methods also predicted the lower surface/cove pressures more

accurately.

The main-element pressure distribution shown in Figure 13 is typical of multi-elemem

airfoils with a large suction peak near the leading edge followed by a pressure recovery

region and then leveling off into a fiat suction plateau. Note that pressure does not recover

back to freestream levels since the trailing edge is adjacent to the flap suction peak. The

under surface is characterized by high pressure, nearly stagnating flow (Cp near 1).

Computed upper surface suction was generally higher than the experiment except near the

leading edge where calculations were scattered above and below the experimental results.

The stagnation location greatly effects the suction peak. This can be seen in the slat

pressure distributions where the solutions with the higher suction peaks have a more

rearward stagnation point. Small errors in stagnation point prediction can lead to large

suction differences, suggesting that grid clustering around the stagnation point is important.

Separation on the flap is observed in Figure 14 as a slight flattening of the experimental

pressure distribution near the trailing edge. A few of the codes predicted a flattening of the

pressure distribution typical of separation. Nearly all the codes predicted a higher suction

level on the upper flap surface than that found experimentally.

Pressure Distributions at ot=21.29 °

Pressure distributions on the slat, main element, and flap near CL_, (Ct=21.29 °) are shown

in Figures 15, 16, and 17, respectively.

9

Notethevery largesuctionpeakon theslat (Cp=-17experimentally)shownin Figure 15.

For a freestreamMachnumber(M,.) of 0.2, thecriticalCa,is -16.3, thus there is a small

region of supersonic flow near the slat leading edge. Computed pressures on the slat

indicate a higher suction peak than the experiment which would result in a higher Mach

number region and a stronger shock. The coupled method solutions exhibit more

variability than the RANS methods.

The main-element pressures shown in Figure 16 show similar trends as the t_=8.12 ° case.

Suction levels on the flap upper surface (Figure 17) did not change significantly from the

oc=8.12 ° case. Since the flap is located in the downwash from the main element, its local

angle of attack stays nearly constant throughout the angle-of-attack range, thus the

pressures remain fairly constant. Experimental boundary-layer measurements indicated that

the flap was not separated at this angle of attack even though the upper surface pressure

distribution shows a separation-like flattening on the aft 50%. Several of the coupled

methods predicted flap separation as will be seen in the discussion of skin-friction results.

Skin Friction

Computed skin-friction distributions for cz=8.12 ° and 21.29 ° at a Reynolds number of 5

million were chosen for comparison to experiment.

Skin Friction at a=8 °

Computed and experimental skin-friction distributions for the slat, main element, and flap

for the low angle of attack (_=8 °) flap separation case are shown in Figures 18, 19, and

20, respectively.

Computed skin-friction data for the slat (upper surface only) is shown in Figure 18.

Experimental data is not available, since the upper surface boundary layer was nearly all

laminar and the Preston tube skin-friction measurement technique is valid only for turbulent

boundary layers. The calculations show a wide range of skin-friction levels due to widely

varying transition locations. The RANS solutions are split between two distinct skin-

10

friction-coefficient(Cf) levels,onearoundCf=0.01(transitionnearleadingedge)and the

otheraroundC_-0.003(transitionneartrailingedge).Transitionis indicatedby an increase

in skinfriction precededby alow skin-frictionlevel.

Skin-frictionbehavioronthemain-elementuppersurface(Figure19)is typical of a single-

elementairfoil characterizedby largeCf valuesnearthe leadingedge(dueto high velocity

flow andthin boundarylayers)which decreasedownstream(due to slowing flow and

thickerboundarylayers). The largeskin-friction valueat the trailingedgeof the main

elementissomewhatdifferentfrom single-elementairfoils which typicallyshow a drop in

Cfnearthetrailing edgedueto theadversepressuregradient.This occurssincethe trailing

edgeof themainelementis locatednearthe flap suctionpeakwhich keepstrailing edge

velocitieslargeandpreventsa drop-off in Cc The coupledmethodpredictionsshow a

largeamountof scatterbetweencodes,possiblydueto thevariousintegralboundary-layermethods.TheRANSmethodsshowedmuchlessvariability. For mostof thecodes,there

isgoodagreementwith theexperiment.

Boundary-layerseparationon the flap is indicated in Figure 20 as the Cf values approach

zero near the trailing edge. Separation location was estimated from the experimental data ai

about x/_1.06. Computed separation location varied widely among the codes. The large

dips in Cf near x/c=0.92 are the boundary-layer transition locations. Changing the flap

transition location had a large effect on the separation location between the 'amir' and

'amirt' solutions. Separation location moved from 1.07 to 1.10. This seems plausible

since the earlier transition causes a larger energy loss in the boundary layer. Calculated

skin friction on the flap shows more variation between codes than for the main element

possibly due to the influence of the boundary-layer confluence on the flap.

Skin Friction at Or=21 o

Skin-friction distributions for the slat, main element, and flap near CL_= (tX=21.29 °) are

shown in Figures 21, 22, and, 23, respectively. The coupled-method solutions appear to

have less variability than the RANS methods.

The slat skin-friction distribution (upper surface only) is shown in Figure 21. A few

solutions showed a separation bubble near the slat leading edge (x/c=-0.085) which was

11

possiblyinducedby a shock. In general there was a very large variability between the

solutions.

As seen in Figure 22, Cf behavior on the main element is similar to the previous case

although skin friction is slightly higher than the o_=8 ° case near the leading edge and

slightly lower near the trailing edge. The higher values near the leading edge are probably

caused by the highly accelerated flow and the lower values near the wailing edge are

possibly due to the larger adverse pressure gradient.

Experimental skin-friction values for the flap indicate that the flow has reattached at this

angle of attack (Figure 23). A few of the coupled methods predicted flow separation near

the trailing edge.

Velocity Profiles

Velocity profiles were measured at nine locations on the upper surfaces of the roain element

and flap for Geometries A and B. The profiles were obtained using a flattened total

pressure tube and a five-hole probe which were traversed normal to the airfoil surface.

Comparisons to computed results are presented for one location on the main element and

three locations on the flap for Geometry A only.

Velo¢i_ Profiles at or=& 12 o

Velocity profiles near the mid-chord of the main element are shown in Figure 24.

According to the experimental data, the main-element boundary-layer edge is just below

n/c=0.01. The slat wake passes just above the boundary layer and shows very little

velocity deficit. Most of the codes predicted a slightly thicker boundary layer than found

experimentally, this is possibly due to artificial dissipation effects of the RANS methods or

a transition location too far forward. Interestingly, many of the codes predicted a larger

velocity deficit in the slat wake which indicates that the turbulence models may be missing a

key part of the physics. The rapid diffusion of the slat wake may be due to large streamline

curvature effects near the main-element leading edge, however most eddy viscosity models

do not include curvature effects. Another hypothesis is that turbulence production in the

slat cove region may effect the slat wake diffusion.

12

Velocity profilesontheflapnearthe leadingedgearepresentedin Figure 25. Theprofile

in this regionusuallyconsistsof four levels. Going from thesurfaceoutwardthesefour

levelsare: 1) flapboundarylayer, 2) slot flow throughthe flap gap, 3) main element

wake,and4) slatwake. Theflapboundarylayeris extremelythin at this location, thus

experimentaldatacouldnotbe takenin this region. Experimentally,the slot flow extends

from 0.001<n/c<0.01andis ramp shaped indicating strong viscous effects on the slot flow

coming through the flap gap. These viscous effects are quite possibly generated by the

separated and recirculating flow in the flap cove. The main-element wake extends from

n/c---0.01 to n/c=0.025 and is very asymmetric, which is a challenging situation for eddy-

viscosity turbulence models. The calculations show good general agreement with

experiment, however, a few solutions show a large slat wake deficit that did not appear in

the experiment.

At the flap mid-chord (Figure 26), the boundary layer is much thicker, and the main-

element wake is more symmetric and is beginning to merge with the boundary layer. There

are some good agreements between experiment and computed main-element wake location

(e.g., "kyle" and "kyleff'). Variation in the amount of wake deficit prediction is large. A

few of the codes predict a notable slat wake defect also.

Flow separation near the flap trailing edge is indicated by both the experiment and the

calculations in Figure 27. Since the experimental data was obtained using a total pressure

tube, reverse velocities could not be measured. Nevertheless, when the data approaches

zero and becomes quite variable, intermittent separation is indicated. Most of the

calculations also showed a small amount of separation at this location.

Veloci_ profiles at (z=21.29 °

The mid-chord main-element velocity prof'de near Ct._, x is shown in Figure 28. At this

angle of attack, the slat is highly loaded resulting in a strong slat wake at this location. The

slat wake deficit showed better agreement between calculation and experiment for this

condition, however the width and height of slat wake was over predicated.

13

Profilesatthe leadingedgeandmid-chordof the flap are shownin Figures29 and 30.

Thereisconsiderablymorescatteringin thepredictionsfor the flap thanthosefor the main

element,especiallyat themid-chordsectionof theflap.

Velocityprofilesneartheflap trailing edge(Figure 31) show a very largemain-element

wake. In fact theexperimentalprofile containssignificantscatterand very low velocities

indicatingintermittentflow reversalin thewake.

Summary of Results

A workshop on 2D high-lift multi-element airfoils was held in May of 1993 at NASA

Langley Research Center. The primary focus of this workshop was to determine the state-

of-the-art in computational techniques, to discuss industry needs and high-lift flow physics

issues.

A major part of the workshop centered on a blind test of various computational methods in

which the flow about a three-element airfoil was computed without prior knowledge of the

experimental data. Comparisons were made between computation and experiment for

* lift, drag, and moment

• lift and drag increments due to Reynolds number and flap gap changes

• pressure and skin-friction distributions

• mean velocity profiles

The results of this 'blind' test revealed:

1. There was good agreement between several codes and the experimental results. In

general, the Reynolds Averaged Navier-Stokes methods showed less variability than did

potential/Euler solvers coupled with boundary-layer solution techniques. However, some

of the coupled methods still provided excellent predictions.

2. Coupled-methods drag prediction agreed more closely with experiment than the RANS

methods. Lift was more accurately predicted than drag for both methods.

14

3. Thecodesdid reasonablywell in predictinglift and drag changesdueto changesin

Reynoldsnumber,althoughtheyall missedthe dragrisecausedby a mild flow separation

on the flap at low Reynoldsnumber. The codesalso did not perform as well when

predictinglift anddragincrementsdueto changingflapgap.

4. Pressuresandskinfriction comparedfavorablywith experimentfor mostof thecodes.

5. Therewasalargevariabilityin mostof thevelocityprofdepredictions. Computational

resultspredicta strongerslat wake thanmeasured,suggestinga missingcomponentin

turbulencemodeling,perhapscurvatureeffects.

Severalflow-physicsissueswerediscussedduring the workshop. Theneedfor accurate

transitionlocationswasemphasizedsincethisappearsto havea majorinfluenceon overall

airfoil performance.Transitionon the slatuppersurfaceis thoughtto haveamajoreffect

on CL_x perfOrmancesinceit controlsthesizeof the slatwake,which in turn influences

ELmaX"

Three-dimensional testing is really the key item of concern, so more three-dimensional data

is desirable.

The effects of turbulence generation in the slat and main-element coves is virtually

unknown.

15

References

° Chin, V.D.; Peters, D.W.; Spaid, F.W.; and McGhee, R.J.: Flowfield Measurements

about Multi-Element Airfoil at High Reynolds Numbers. AIAA Paper 93-3137,

AIAA 24th Fluid Dynamics Conference, July, 1993.

2. Lin, J. C.; Robinson, S. K.; McGhee, R. J.; and Valarezo, W. O.: Separation Control

on High-Lift Airfoils Via Micro-Vortex Generators. J. Aircr., vol. 31, no. 6, Nov.-

Dec. 1994, pp. 1317-1323.

3. Klausmeyer, S. M.; and Lin, J. C.: An Experimental Investigation of Skin Friction on a

Multi-Element Airfoil. AIAA Paper 94-1870, 12th AIAA Applied Aerodynamics

Conference, June 1994.

Table 1. Slat and Flap Setting

Geometry A

Slat Deflection (_)

Slat Gap, %c

Slat Overhang, %c

Flap Deflection (_Sf)

Flap Gap, %c

Flap Overhang, %c

(30P-30N)

.30 °

Geometry B

(30P-30AD)

_30 °

2.95 2.95

-2.5 -2.5

30 ° 30 °

1.27 1.5

0.25 0.25

16

Table2. CFDChallengeParticipantsSummary

Legend

Key

Kyle

Kyleff

Person Affiliation Program Type Grid

Anderson/ Langley FUN2D RANSBonhaus

Anderson/ Langley FUN2D RANSBonhaus

jones Jones Langley CFL3D RANS

bied Biedron Langley CFL3D RANS

vatsa Vatsa Langley TLNS3D RANS

Unstructured

Modeling Notes

Unstructured

Structured-

Chimera

Structured-

Eiseman

multi -block

Structured-

Eiseman

multi-block

dod Dodbele Langley MCARF Coupled -

mavk Mavriplis/ Langley NSU2D RANS Unstructured

Klausmeyer

stusb Rogers Ames INS2D RANS Structured-

Chimera

stusa Rogers Ames INS2D RANS Structured-

Chimera

stuso Rogers Ames INS2D RANS Structured-

Chimera

drela

hawk

wood

caobb

caoba

Drela M1T MSES Coupled

Hinson/Hawke Learjet MEAFOIL Coupled

amir

amirt

Without point vortex

farfield corrections

With point vortex

farfield corrections

Structured

Baldwin/Barth

turbulence model

Spalart/Allmams

turbulence model

k-_ turbulence

model

Baldwin/Barth

turbulence model

Spalart/AUmaras

turbulence model

Free transition

Fixed transition

Woodson Cessna MCARF Coupled

Cao/Kusunose Boeing INS2D RANS Structured

Cao/Kusunose Boeing INS2D RANS Structured

Amirchoupani Boeing Coupled -

Amirchoupani Boeing Coupled -

17

Table3. CaseSummary

Case Geometry Angleof Attack ReynoldsNumber Comments

3

4*

15"

A (30P-30N)

A

B (30P-30AD)

8.12 ° 5 million flap separation

8.10 ° 9 million

8.10 ° 9 million flap separation

A 16.21 o 5 million

A 16.21 ° 9 million

B 16.24 ° 9 million

A 21.29 ° 5 million Ct_ =

A 21.34 ° 9 million CL._

B 21.31 ° 9 million CL_x

B 4.07 ° to 23.34 ° 9 million

A 23.28 ° 9 million

B 23.34 ° 9 million

flap separation

between 8° to 14°

Stall

Stall

* Optional

18

Main Element

(a) McDonnell Douglas Aerospace three-element airfoil

(b) Nomenclature for multi-element airfoil

Figure 1. Model geometry and nomenclature

]9

6.0

5.0

4.0

CL

3.0

2.0

1.0

Lift

0 5

doddmla

wood

amirt

oxp

(a) Coupled Methods

Drag

I I I I I I I l

10 15 20 25 0.02 0.04 0.06 0.08 0.10 0.12

(X (deg) CD

PitchingMoment

I I I

-0.8 -0.6 -0.4 -0.2

C M

(b) Reynolds Averaged Navier-Stokes Methods

6.0Lift

5.0

CL --{3-- =._bI _lV/ _ _.=

L / --<)- ,_I/ +'_=

2.0 | _ _

t1.0 I I t t I

0 5 10 15 20 25

DragPitchingMoment

o_ (deg) Co C a

Figure 2. Force and moment coefficients for Geometry A, Re=5,000,000

2O

6.0

5.0

C L

4.0

3.0

2.0

1.00

(a) Coupled Methods

Lift

ded

• _ wood

amir f

amlrt

m ex:p

f Drag

I I I I I I I I I I

5 10 15 20 25 0.02 0.04 0.06 0.08 0.10 0.12

0_, (deg)C D

I I I

-0.8 -0.6 -0.4 -0.2

C M

6.0 [ Lift5.0

3-0 / f ---?-- _,,/ f --V--- _..F/ --O- ,,.=


f Drag

2.0 +__[>'-- m

1.0 I I T_ _0 5 10 15 20

tI I I I I I I

25 0.02 0.04 0.06 0.08 0.10 0.12

PitchingMoment

-0.8 -0.6 -0.4 -0,2

(X (deg) CD Cu

Figure 3. Force and moment coefficients for Geometry A, Re=9,000,000

21

C L

6.0

5.0

4.0

3.0

(a) Coupled Methods

Lift

dod

f _ aml

ami

m,,mm, eXp

0I I I I

5 10 15 20 25

OC(deg)

Drag

I

0.02 0.04 0.06 0.08 0.10 0.12

C D

-0.8

PitchingMoment

I

-0.6 -0.4 -0.2

C M

6.0

Lift

5.0


ot_.or /i- _:.

L,/

I _-+-=1.0 I I I I i i

0 5 10 15 20 25G

O_(deg)

Drag

I I i _ I

0.02 0.04 0.06 0.08 0.10 0.12

Co

D

PitchingMoment

i

I I I

-0.8 -0.6 -0.4 -0.2

Cu

Figure 4. Force and moment coefficients for Geometry B, Re=9,000,000

22

5.0

4.0

CL

3.0

2.0

1.00

(a) Coupled Methods

Lift

clod

---{>-- ,,._---_ mir

---_ =,m_

m e_p

I I I I

5 10 15 20

0_, (deg)

m

Drag

I I I I I I

25 0.02 0.04 0.06 0.08 0.10 0.12

C D

PitchingMoment

-0.8 -0.6 -0.4 -0.2

C M

6.0Lift

5.0


Drag

c4Of 3.0

L _r# -----I!-- =,,=

2.0 _ ,==,m

m eXlp

I I I

10 15 20

(_ (deg)

1.0 I0 5

I I I I I I

25 0.02 0.04 0.06 0.08 0.10 0.12

PitchingMoment

C D CM

Figure 5. Force and moment coefficients for Geometry B, Re=9,000,000

23

0.15

I AC L = CL big gap- CL small gap0.10

0.05

AC, o.oo

-0.05

-0.10

-0.15 I0 5 10 15 20 25

_(deg)

0.020

0.015

0.010

0.005

0.000

-0.005

-0.010

-0.015

n

ACD= CD big gap" CD small gap

I I I I I

0 5 10 15 20 25

(X,(deg)

Figure 6. Lift and drag changes due to increase in flap gap, Re=9,000,000(Experiment-- LTPT)

24

0.15

0.10

0.05

0.00

AC L

-0.05

Exp

-----D-'-- doddrela

hawkwood

amir= amirt

Coupled Methods

AC L = C L big gap" C L small gap

0.15 r

0.1o

0.05

0.00

AC L

0 5 10 15O[ (deg)

-0.05

I I

20 25

Reynolds Averaged Navier-Stokes MethodsmExp

caobb

caosa

jones

kylekyleff

stubb

stuko

stusavaLsa

---<]---- bieumavkc

-0.10

-0.15

AC L-- C L big gap" C L small gap

I I I I I

0 5 10 15 20 25(X, (deg)

Figure 7. Lift change due to increase in flap gap, Re=9,000,000

25

Coupled Methods0.030 - Exp

-----D--- dod

drela

0.020- _ .awk /X ;_wood / \ /

----_--- _,r / \ /

coO.OlOr /0.000 - -@ .....

-0.010

-0.020 I I i I i0 5 10 15 20 25

0_, (deg)

0.030

0.020

0.010

zXCD

0.000

-0.010

-0.020

Reynolds Averaged Navier-Stokes Methods

_obbcaosa

jones

kyle

kyle.stubb _

_ _ / \

"0

ACD = CD big gap - CD small gap

I I I I I

0 5 10 15 20 25

(deg)

Figure 8. Drag change due to increase in flap gap, Re=9,000,000

26

0"15F z_C L-" CL Re=9mil" CL Re=5rnil0.10

0.05

0.00

-0.05

-0.10 -

-o.15 I I I I0 5 10 15 20

O_(deg)

I25

0.020 -

0.015

0.010

0.005

0.000

-0.005

-0.010

-0.015

CD Re=9mii" CD Re=5mil

I I I I I

0 5 10 15 20 25

_(deg)

Figure 9. Lift and drag changes due to increase in Reynolds Number,

Geometry A (Experiment -- LTPT)

27

0.20

0.15

0.10

0.05

0.00

-0.05

-0.10

Coupled Methods

drelahawkwood

= amirt ii / ___,,,,,,J,

ACL = CL Re=9 mil - C L Re=5 rail

I I I I ,l

0 5 10 15 20 25

O_

Reynolds Averaged Navier-Stokes Methods0.20

--,,,--- F_xp /1

caobb /_J

cac_a /0.15 _ jones /

_e /----W--kyleff /

o.1o + /

--_-_ _ I V_-"_/_..._

ACt- 0.05 --O---_0.00

_0405

z_'CL - CLRe 9mil " CLRe 5rail

-0.10 I i I I I

0 5 10 15 20 25

(X

Figure 10. Lift change due to Reynolds Number increase, Geometry A

28

0.030

0.020

0.010

AC D

0.000

-0.010

-0.020

Coupled Methods

AC D---- C D Re=9mil " C D Re=5mil

Exp

dod

---_-- drela r_

_ _ hawk

wood /

amir

____= .... a:_rt_ _ _ =r ...... - ........ -/ ............

i I I I I I

0 5 10 15 20 25O(,(deg)

Reynolds Averaged Navier-Stokes Methods0.030 -

Exp

caobb

caosa

jones

0.020 _ kylekyiefl

stubb

stuko

stusa

0.010 _ vatsa

bled

D

0.000

-0.010

AC D-- C D Re=9mil" C D Re=5mil

-0.020 I I I I I

0 5 10 15 20 25

O(,(deg)

Figure 11. Drag change due to increase in Reynolds Number, Geometry A

29

Cp

-4 -

-3 -

-2 --

"1 --

0 -

1 -

Coupled Methods

I S

1 - o-_---o---oo o o-.-/

,%._ ,

_, .."_/..I_i.:'iI &"'-rl J/c_ h'_ ' 0 J£'_

_ ' L-%44"-" _ I "'"

I I I * I

-0.08 *0.04 0.00x/¢

O Exp

dod

...... drela

...... hawk

...... wood

...... amir

...... amirt

I

0.04

-4

-3

-2

Cp-1

0

Reynolds AveragedNavier-Stokes Methods

-- I

I i I i I i

-0.08 -0.04 0.00

x/c

O Exp

caobb

caosa

jones

kyle

kyleff

...... stubb

...... stuko

...... stusa

...... vatsa

...... bied

...... mavkc

I

0.04

Figure 12. Slat pressure distribution -- Geometry A (o_=8.12 °, Re=5 million)

30

Cp

-8 -

-6

-4

-2

2

0.0

Coupled Methods

o Exp

_. dod

..... drela

..... hawk

___,_ - .... wood"_, ..... amir

"-'___ .......

b t

:__.__...__ _ - _---_ ='-G = ,7,

a I a I a I , I

0.2 0.4 0,6 0.8

x/c

i I

1.0

-8

-6

-4

Cp-2

20.0


, I , I , I , I i,

0.2 0.4 0.6 0.8

o Exp

caobb

caosa

jones

kyle

kyleff

stubb

stuko

stusa

vatsa

bled

mavkc

I

1.0

Figure 13. Main-element pressure distribution -- Geometry A

(0_=8.12 °, Re=5 million)

31

-5 Coupled Methods

-4

-3

Cp -2

-1

1

0.80

I

0.85

I

0.90

I I

0.95 1.00

x/¢

o Exp

dod

...... drela

...... hawk

...... wood

....... amir

...... amirt

I I I I

1.05 1.10 1.15 1.20

-5

-4

-3

Cp -2

-1


o Exp

caobb

caosa

jones

kyle

kyleff

...... stubb

...... stuko

...... stusa

...... vatsa

...... bled

mavkc

0.80 0.85 0.90 0.95 1.00 1.05 1.10 1.15 1.20

x/c

Figure 14. Flap pressure distribution -- Geometry A (o_=8.12 °, Re=5 million)

32

Cp

-20 -

-16

-12

-8

-4

o

Cp, crit

Coupled Methods

I , I i I

-0.08 -0.04 0.00

x/c

O Exp

dod

...... drela

...... hawk

...... wood

...... amir

...... amirt

I0.04

-2O

-16

-12

Cp-8

-4

0

- Reynolds Averaged jNavier-Stokes Methods i

.... I ................

- / Cp, crit I

-o Exp

caobb

caosa

jones

kyle

kyleff

...... stubb

...... stuko

...... stusa

...... vatsa

...... bied

...... mavkc

I , I J I , I

-0.08 -0.04 0.00 0.04

x/c

Figure 15. Slat pressure distribution -- Geometry A (o_=21.29 °, Re=5 million)

33

-12

-10

-8

C p -6

-4

-2

20.0

Coupled Methods

'\i_', 0 Exp

', -- • dod

_ ...... drela

'_," ..... hawk

_,_,_, - .... wood

"_, ..... amir

._, - .... amirt

i I I I i I i I

0.2 0.4 0.6 0.8

x/c

I

1.0

C

-12

-10

-8

-6

P-4

-2

20.0


o Exp

-- caobb

-- caosa

-- jones

-- kyle

-- kyleff

..... stubb

..... stuko

..... stusa

..... vatsa

..... bied

..... mavkc

Figure 16. Main-element pressure distribution -- Geometry A(o_= 21.29 °, Re = 5 million)

I

1.0

34

-5 - Coupled Methods

C

-4

-3

-1

m

1 -

0.8(I

0.85

\

\

\

\

\

\

o Exp

dod

...... drela

...... hawk

...... wood

...... arnir

...... amirt

I I I I I I I0.90 0.95 1.00 1.05 1.10 1.15 1.20

x/¢

C

-5

-4

-3

-1


I I I

0.80 0.85 1.15 1.20

o Exp

caobb./2,,

_, caosa

"'"",\ jones

'_ kyle

kyleff

" _ '"-, ...... stubb

...... stuko

" - - _ ...... stusa

_ ...... vatsa

...... bied

...... mavkc

I I I I I I

0.90 0.95 1.00 1.05 1.10

xlc

Figure 17. Flap pressure distribution -- Geometry A (o_=21.29 °, Re=5 million)

35

Of

0.03

0.02

0.01

Coupled MethodsI

!

5

I

II

It

II

I

I_l_

!l'lt

Iiii ,

Tk,lly_

I_ ", _. _ - - -_-_ .......

'it t "_< l

0.00 , I

0.04

IJ ',, f t

I I , I , I

-0.08 -0.04 0.00

x/c

...... drela

...... hawk

...... wood

...... amir

...... amirt

Cf

0.03

0.02

0.01


0.00

0.04

'1

%/ __ ".,

-0.08 -0.04 0.00

x/c

caobb

caosa

jones

kyle

kyleff

...... stubb

...... stuko

...... stusa

...... vatsa

...... mavkc

Figure 18. Slat skin-friction distribution -- Geometry A, upper surface only(0_=8.12 °, Re=5 million)

36

0.04

0.03

Cf 0.02

0.01

0.0

Coupled Methods

I _ 0 Exp...... drela

\

' _ ...... hawk

l!i ...... wood

'1' _ ,_ Upper ...... amir

!lli_ i _ Surface amirt

I t '_" ......

_ \ t /- Lower',, _ . _ _ / Surface

' ....... / " - .... "IT",,Z % / It

0.2 0.4 0.6 0.8

x/c

Of

0.04

0.03

0.02

0.01


UpperSurface

Surface

O Exp

caobb

caosa

jones

kyle

kyleff

...... stubb

...... stuko

stusa

vatsa

...... bied

...... mavkc

, # !

0.0 0.2 0.4 0.6 0.8x/¢

I

1.0

Figure 19. Main-element skin-friction distribution -- Geometry A(0_=8.12 °, Re=5 million)

37

0.03

0.02

Of

0.01

0.00

0.80

0.03

0.02

Of

0.01

I

0.85

Coupled MethodsII

II

II

II

I

I

I

I

I

I i __ O O Exp-up2,:' E×p-low

t _ _ - - - drela

,_ , _ _ hawk

r ' _ _', , , , .//-Upper wood, i, '--\,' , _/ Surface

'i ' _'x _w / amiri ', '. _. . I amirt,_ _ _ ?*_.. , /

,; ";,:'.%../.,' ' ". '_... J-Lower

, ' . ".- ._ _ Surface

i# _TC----'-_"_--_-J--'-_'--_-- - _- - - " ---r..-L-_, c - .{ .i- _ - i i

0.90 0.95 1.00 1.05 1.10 1.15 1.20

x/c

O Exp-up

z, Exp-low

caobb

..... oaosa

jones

kyle

kyleff...... stubb

...... stuko

...... stusa

...... vatsa

...... bied

...... mavkc


O_ Upper

/ Surface

"_'_xx, ,,,,,x-Lower

_ _l'_ Surface ,

0.90 0.95 1.00 1.05 1.100.00 ' I I

0.80 0.85 1.15 1.20

x/c

Figure 20. Flap skin-friction distribution -- Geometry A ((_=8.12 °, Re=5 million)

38

0.09 -

0.08

0.07

0.06

0.05

Cf 0.04

0.03

0.02

O.0f

0.00

-0.01

i

" Coupled MethodsIIPpl

tlII

Vt

_,1). I

l;', |

_'_

_ , ............... -===2- 2'l ---_ 0 0

ii .............

I IL

o Exp

........ drela

........ hawk

........ wood

........ amir

........ amirt

L , 1 , I , I

-0.08 -O04--CX_ 0.00 0.04

Of

0.09 -

0.08 -

0.07 -

0.06 -

0.05

0.04 -

0.03 -

0.02

0.01

0.00 -

-0.01

Reynolds Averaged

Navier-Stokes Methods

0 Exp

caobb

caosa

I_ jones,i kyle'! kyleff

* |\ ii ...... s,ooo>,,_ ._ - ..... stuko

t ...... mavkc

i , I , I L

-0.08 -0.04 0.00 0.04X/C

Figure 21. Slat skin-friction distribution -- Geometry A, upper surface only(c_=21.29 °, Re=5 million)

39

0.0.4

0.03

Cf 0.02

0.01

0.00

IIp

dl I

tll I_

Coupled Methods

O Exp

...... drela

,t Lh_ _ _r- Upper hawk

,I i_,!, , _/ Surface ...... wood# _fl_ _ / / / ......

i '! amir

j N ......dt amirt

,, ,,,,? /..l"i"' J"*'//t i I ¢. / ,""l ' " _' / "'" ,,--Lower

I _. _ _

I k" "'/--. '

0.0 0.2 0.4 0.6 0.8

x/¢

i

1.0

0.04

0.03

Cf 0.02

0.01


UpperSurface

Surface

O Exp

caobb

caosa

jones

kyle

kyleff

...... stubb

...... stuko

stusa

vatsa

...... bied

mavkc

0.000.0 0.2 0.4 0.6 0.8

x/c

i

I

1.0

Figure 22. Main-element skin-friction distribution -- Geometry A (c_=21.29 °, Re=5 million)

4O

0.03 - Coupled Methods

Of

0.02

0.01

0.00

0.80

I

0.85

i Ii I_ 0

I

I I1_I

,_'_ Upper

{ , ,,_- ,i -'1- / Surface/,I I I ,, _ //,, , ',' \ //,,'", ,' \ /

h.t'l " '.F

%1 ,,., ', /- Lower_', ",",, " / Surface

', . 6"":',-,. o -r/

0.90 0.95 1.00 1.05 1.10

x/¢

I

1.15

O Exp-up

...... clrela

...... hawk

...... wood

...... amir

...... amirt

I

1.20

0.03

0.02

Cf

0.01

0.00

0.80


I O

Upper I'

/ Surface illI

1 I

'\"_,,_,N, #- Lower

_i,_" .... /Surface

J _-_-__-_ __ -r.r_ _

i I'_"-_r ..... i ..... , i

0.85 0.90 0.95 1.00 1.05 1.10

xlc

O Exp-up

caobb

caosa

jones

kyle

kyleff

...... stubb

...... stuko

...... stusa

...... vatsa

..... bied

...... mavkc

I I

.15 1.20

Figure 23. Flap skin-friction distribution -- Geometry A (o_=21.29 °, Re=5 million)

41

nlc

0.06

0.04

0.02

0.000.8

O Exp....... bied

....... caobb

....... caosa

....... drela

...... jones

....... mavkc

1.0 1.2

101I

b

ulUO0

0.06

0.04

n/c

0.02

O Exp

....... kyle

....... ffkyle

....... stubb

....... stuko

....... stusa

....... vatsa

0000

0.00 I ,0.8 1.0 1.2 1.4 1.6 1.8

u/U

S

Figure 24. Velocity profiles on the main element -- Geometry A(x/c=.45, (_=8.12 °, Re=5 million)

42

n/c

0.06

0.04

0.02

0 Exp....... bied

....... caobb

....... caosa

....... drela

....... jones

....... mavkc

0.001.2 1.4 1.6 1.8

u/U2.0 2.2

0.06

0.04

0.02

0 Exp

....... kyle

....... kyleff

....... stubb

....... stuko

....... stusa

....... vatsa

0.001.2 1.4 1.6 1.8 2.0

u/Uoo

2.2

Figure 25. Velocity profiles on the flap -- Geometry A(x/c=.898, a=8.12 °, Re=5 million)

43

n/c

n/c

0.12

0.10

0.08

0.06

0 Exp....... bied

....... caobb

....... caosa

....... drela

....... jones....... mavkc

u=II| I

I I!

J _

1msI _

I_t II

4' I %,1|

st _

sSo • 0 _ _I

o.o4I- ,:-"'0 ,'-,"/ [ i S - . . - ___fA"--

0.02 I- .,. -,.0,_,.

0.00 __

0.12

0.10

0.08

0.4 0.6 0.8 1.0 1.2 1.4

u/U

0.06

0.4

0 Exp

....... kyle

....... kyleff

....... stubb

....... stuko

....... stusa

....... vatsa

tlIP

[I!

ISii L

s I TIsSs

• II1_

._. #hiLL

-3_"0 "--,

'" O_o "'l

0.6 0.8 1.0 1.2

u/U1.4

0.04

0.02

0.00

Figure 26. Velocity profiles on the flap -- Geometry A(x/c=1.032, o_=8.12 °, Re=5 million)

44

0.12

0.10

0.08

0.06

0.04

0.02

0 Exp....... bled

....... caobb

....... caosa

0.00

0.0

0.12

0.10

0.08

0.06

0.04

0.02

0.00

0.0

0 Exp....... kyle

....... kyleff

....... stubb

....... stuko

....... stusa

....... vatsa

0.2 0.4 0.6 0.8 1.0

u/U

Figure 27. Velocity profiles on the flap -- Geometry A(x/c=1.112, 0_=8.12°, Re=5 million)

45

0.08

0.06

n/c_ 0.04

o.o2

o Exp....... bied

....... caobb

....... caosa

....... drela

...... jones

....... mavkc

• °

## | t

#.# _I |

#. • _E| •

,_'7, 0 ,'

! t

O! •

mmmalnml _"

0.00 .... n , 1

1.0 1.2 1.4 1.6 1.8, i

2.0

0.08 f i]_,O Exp

0.06 I ....... kyle

L ....... ffkyle -4

1 ....... stuko0.04 r ....... stusa ,_2,,n/c ....... vatsa

0.02f -"""""-"_""_-''__

• ..----_.,t_"O" v- - ""_O"O"o.ooL,,-,=__'o , , , , ,

1.0 1.2 1.4 1.6 1.8, I

2.0

u/U

Figure 28. Velocity profiles on the main element -- Geometry A(x/c=.45, 0c=21.29°, Re=5 million)

46

0.12

0.10

0.08

r_/c 0.06

0.04

0.02

0.12

0.10

0.08

u/U0<3

,t.":'0.06 ,,\_

_'Wh,L__.,_;,-,:-.

.Q,,,

ji_i'_ 0 Exp....... kyle

....... kyleff

....... stubb

._¢° "0 ....... stuko....... stusa

....... vatsa

11

0.04 ',

0.02

0.00

0.8 1.0 1.2 1.4 1.6 1.8 2.0

u/U

Figure 29. Velocity profiles on the flap -- Geometry A(x/c=.898, o_=21.29 °, Re=5 million)

47

0.28

0.24

0.20

0.16

n/c o,_

0.08

0.04

0.00

n/c

0.28

O Exp _b !....... bied /,O ,_

t J #,_L =I

....... caobb , , _,i• e If,,I

....... caosa ' , .IN,%.ql........ drela ,,-_'[ _,','

....... jones "," t_'O' ,O_-_ ''

....... mavkc ""t_,t_. _'° :

i,_ll i I,'=s %%_ ,

.- .... ;..... "-:_'._-o ..u1...-.:...'--- '... ,;--_r_"_._- .....

[ ........... ::-'-_;_HNi_t ..,_- ....... ---'-"-_.-c_-'Ti_ . ,._.,0.0 0.4 0.8 1.2

u/U

0.24 - _!m.0.20 O Exp

....... kyle

....... kyleff0.16 ....... stubb

....... stuko

....... stusa

0.12 ....... vatsa

0.08 i

0,0 0.4 0.8 1.2

u/U(X)

Figure 30. Velocity profiles on the flap -- Geometry A(x/c=1.032, 0_=21.29 °, Re=5 million)

48

0.28

I,,t

,1!

0.24 0 Exp " ",,bled , ,l

....... caobb ,arS _-.

0.20 ....... caosa ,'_" - ,....... drela

I

...... jonesI

O.16 ....... mavkc , I

r --- :::

0.00 '

0.0 0.4 0.8

u/U

0.28

I

1.2

J _J

0.24

0.20

0.16

0.12

0.08

0.04

0.00

0.0

0 Exp

....... kyle

....... kyleff

....... stubb

....... stuko

....... stusa

....... vatsa

0.4 0.8

Figure 31. Velocity profiles on the flap -- Geometry A(x/c=1.112, o_=21.29 °, Re=5 million)

49

REPORT DOCUMENTATION PAGE Form ApprovedOMB No. 0704-O188

Public reporting bun_n for this collection of information is estimated to average 1 hour per response, including the time for reviewing instilJCtions,searching existing data sources,ge_ering and maintaining the data needeci, and completing a_l reviewing the co lectfon of informst_n. Se_l co_rne, nts re_;jarc.l.ingthis burden estimate or any other aspect of this

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May 1997 Technical Memorandum4. TITLE AND SUBTITLE 5. FUNDING NUMBERS

Comparative Results from a CFD Challenge Over a 2D Three-ElementHigh-Lift Airfoil WU 522- I I-51-01

6. AUTHOR(S)

Steven M. Klausmeyer and John C. Lin

7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES)

NASA Langley Research CenterHampton, VA 23681-0001

9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES)

National Aeronautics and Space AdministrationWashington, DC 20546-0001

11. SUPPLEMENTARYNOTES

Klausmeyer: Wichita State University, Wichita, KSLin: Langley Research Center, Hampton, VA

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NASA TM-112858

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13. AB$, HACT (Maximum 200 words)

A high-lift workshop was held in May of 1993 at NASA Langley Research Center. A major part of the workshopcentered on a blind test of various computational fluid dynamics (CFD) methods in which the flow about a two-dimensional (2D) three-element airfoil was computed without prior knowledge of the experimental data.The results of this 'blind' test revealed:

1. The Reynolds Averaged Navier-Stokes (RANS) methods generally showed less variability among codes than didpotential/Euler solvers coupled with boundary-layer solution techniques. However, some of the coupled methodsstill provided excellent predictions.2. Drag prediction using coupled methods agreed more closely with experiment than the RANS methods. Lift wasmore accurately predicted than drag for both methods.

3. The CFD methods did well in predicting lift and drag changes due to changes in Reynolds number, however, theydid not perform as well when predicting lift and drag increments due to changing flap gap.4. Pressures and skin friction compared favorably with experiment for most of the codes.

5. There was a large variability in most of the velocity profile predictions. Computational results predict a strongerslat wake than measured suggesting a missing component in turbulence modeling, perhaps curvature effects.

14. SUBJECT TERMS

High lift; CFD; RANS; Coupled method; Experimental data

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comparative results from a cfd challenge over a 2d three ... · pitching moment was computed with...

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