AIAA 92-0267 Experimental Studies of a Two-Element Airfoil with Large Separation K. Biber and G.W. Zumwalt The Wichita State University Wichita, KS

30th Aerospace Sciences Meeting & Exhibit

January 6-9/1992 / Reno, NV For permission to copy or republish. contact the American Institute of Aeronautics and Astronautics 370 L'Enfant Promenade, S.W., Washington, D.C. 20024

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AIAA-92-0267

EXPERIMENTAL STUDIES OF A TWO-ELEMENT AIRFOIL WITH LARGE SEPARATION

Kasim Biber' and Glen W. Zurnwalt** Wichita State University

Wichita, KS 67208

Abstract

An experimental program was undertaken to study a two-element airfoil with moderate through large sep- aration. The GA(W)-2 airfoil with 25% single-slotted flap was the test model at a Reynolds number of 2.2~10' and Mach number of 0.13. The test matrix consisted of single-element airfoil and 30' and 40" flap deflection angles with optimum and narrow gaps at near- and post-stall angles of attack.

Airfoil performance was studied through the measure- ments of aerodynamic forces and surface pressures. Flap-deflected configurations indicated a large degree of hysteresis with angle of attack changes. Flow field static pressures were measured by a specially designed splitter plate and disk probe and presented by contours. Total pressures were measured by a double-pitot tube. - Flow field sensitivity to these probes were determined by means of surface pressure and force measurements. Two-component mean velocity and turbulence quanti- ties were measured by an end flow X-probe hot film anemometer.

The experimental program revealed the difficulty of measuring static pressures in separated flows. The pres- ence of a probe, no matter how small, changes the 00w geometry in the interacting parts of the flow field. The results of this experimental study could assist and guide the analytical investigations for a better prediction of such separated flows on multi-element airfoils.

Momenclature

C

C, =Drag coefficient C, =Lift coefficient

Ci,,, =Maximum lift coefficient C,,, =Moment coefficient

=Mean chord, 2-ft. for the GA(W)-2 model

* Graduate Student, Student Member AIAA t ** Distinq. Prof. of Aerospace Engineering,

Assoc. Fellow AIAA

Copyright 0 American Instltude of Aeronautla and A8rronautics, In<., 1901. AI1 right. r.,.&.

C,

C,,, G/c =Gap-to-chord ratio

P =Static pressure

p~ =Total pressure qm =Dynamic pressure, 24 psf. indicated for the

Re u, v =Fluctuating velocity components u2, v 2 =Reynolds normal stress components

uv =Reynolds shear stress U, V =Mean velocity components

UrC, =Tunnel reference velocity V,,, =Total mean velocity, d m x, y

X, Y

a =Angle sf attack

6 =Flap deflection angle

=Static pressure coefficient, (p - pm)/qm

=Total pressure coefficient, ( p ~ - pw)/qw

present tests

=Reynolds number based on chord length

=Airfoil coordinates, x is along the chord line

=Tunnel coordinates, X is along the tunnel cen- terline (positive downstream)

SilbscriDts ref =Tunnel reference value 00 =Free-stream quantity

I. INTRODUCTION

The flow mechanism on multi-element airfoils dif- fers from that on single-element airfoils by the com- plexity of the flow field introduced by the flap deploy- ment. The presence of confluent boundary layers, cove separation bubble, off-the surface pressure recovery, strong viscous-inviscid interactions and pressure gradi- ents across the wake characterize the distinguished fea- tures of the flow around multi-element airfoils. When the flow separates from the upper surface, the unsteady vortical motion inside the separated wake makes the Bow even more complicated to deal with. Fig.1 shows such separated flow field with its interactive parts. Suf- ficient knowledge of this flow mechanism is essential in multi-element airfoil design and performance predic- tions. This is why there have been numerous computa-

tional and experimental investigations of multi-element airfoils since they were first used in high-lift generation.

A vast majority of multi-element research has been conducted for the pre- or near-stall angles of attack con- ditions without separated flow. Even if there is separa- tion, it is on the flap at low angles of attack. The work of Ref.(l) appears to be the most detailed and extensive experimental data of this kind available for the com- putational codes, such as the one reported in Ref.(2), developed on multi-element airfoils. But no field pres- sure data were taken. The flow field in the vicinity of a single-slotted flap has been investigated in Ref.(3) for attached and in Ref.(4) for separated flow cases. The measurements of attached flow field quantities such as mean velocity and turbulence on multi-element airfoils have been reported in Refs.(5-7). In Refs.(S and 9), the high-lit airfoil performance data at angles of attack up to post-stall and flow field pressure and velocity data by using pitot-static pressure probe up to stall conditions have been reported.

Fig.1: Theoret ical model for largely sepa- r a t e d flow

Needless to say, there is a lack of experimental data for realistic multi-element airfoil configurations. This data should include the airfoil performance within its angle of attack limitations and flow field data by means of pressure, mean velocity and turbulence.

The purpose of the experimental investigation pre- sented in :.his paper was to provide the data base along with sufficient information to assist in the development of computational methods for the prediction of a highly separated flow field on a flapped-type airfoil. The pri- mary tasks can be outlined as follows:

To study the effects of flap setting on the airfoil per- formance by means of force and surface pressure mea- surements

To deterniine the hysteresis effects on the aerody- namic coefficients and surface pressure distributions.

To verify the flow sensitivity to the interference of the measuring probe.

To check the credibilty of the splitter plate by means of data comparison with a disk probe.

To obtain the pitot pressure profiles at selected chord- wise stations for the determination of the reversed flow regions.

* To measure flow field pressures and present them by pressure contours.

To measure mean velocity components and turbu- lence quantities at selected chordwise stations.

To discuss the flow field on the flapped-type airfoil based on the measured quantities and observed physical events.

11. EXPERIMENTAL A R R A N G E M E N T

W i n d Tunnel and Model

The test was conducted in the Wichita State Uni- versity (WSU) 7- by 10-foot wind tunnel, which is a closed return tunnel with atmospheric test section static pressure and top tunnel speed of 180 miles per hour. The test model was mounted between two inserts, providing a 7- by 3-foot two-dimensional test section. This test section has been calibrated and its results have been reported in Refs.(8 and 9).

The test model used was the 13% thick GA(W)-2 (General Aviation) airfoil equipped with 25% single- slotted flap. The airfoil model had a reference chord of 2-foot at flap nested position and span of %foot. The main wing terminates at 87.5% chord of the airfoil and has a cove region where the flap retracts for single- element configurations.

d

The model was attached to the tunnel balance sys- tem through a set of 42-inch in diameter alcminum end plate disks. The model pivot location was at 50% chord station. Along the mid-span of the main wing and flap, there were 69 pressure ports available for the surface pressure measurements.

Unlike the flap, the bonnday layer transition on the main wing was fixed by two layers of 2.4 mm wide trip strips at 5% upper surface and 10% lower surface. By doing so, maximum possible turbulent flow would be present on the model upper surface.

- Test ConflRurationa

With the inclusion of the single-element airfoil, five different airfoil configurations, each one corresponding to .i gap, were selected for the present experiments. Table 1 shows the selected test matrix.

- 2

The selection of the test matrix was based on the previous flap optimization studies conducted at WSU for the same GA(W)-2 airfoil model (Ref.9). Each f lap deflected case (30" and 40') had an optimum and a narrow gap. The optimum gaps produce the highest Cfma=. In this case, the main wing wake does not merge with the boundary layer on the flap upper surface until downstream of the flap trailing edge, but stays adja- cent. The narrow gaps, on the other hand, produce confluent boundary layers for a short distance down- stream from the flap leading edge. Flap overlap with the main wing was very small and not considered as a major flap setting parameter.

'd

Table 1: The test matrix

Cases I A B C D E

a

6

G/c

100 00 00 00 00

15' 12.8' 12.8' 12.8' 12.8'

20' 15.5' 15.5' 15.5" 15.5" 0' 30' 30" 40' 40"

0 0.03 0.02 0.015 0.0115

the pressure in prescribed directions although the flow is multi-directional, especially in the shear layer recom- bination region. The splitter plate meaures static pres- sures regardless of the flow direction in the plane of the plate, by pressure taps on the plate surface. Therefore, the new field pressure measuring instrument was de- veloped for the present experimental program. During the prelimenary design process, the major issues were the plate size and its support system. The wind tun- nel blockage and three-dimensional effects, especially at post-stall conditions were Considered as major limi- tations. Tests showed the plate would flutter to some degree a t high angles of attack.

.

Pig.2: Geometr ic details of s p l i t t e r p l a t e (All dimensions i n inches)

The plate with its dimensions is shown in Fig.2. It was supported by the strut that has been used for flow field surveys, as shown in Fig.3. It had an elliptically shaped leading edge and rounded bottom edge in order to obtain a better flow quality on the surface. The slope of the bottom edge is straight rather than curved to get closer to the model curvature at all selected angles of attack. The plate was made of aluminum and had 42 pressure taps on one side. These taps were opened to aluminum pressure tubes buried inside the 0.25-inch plate. The ends of pressure tubes were extended verti- cally as much aa 20 inches, and connected to the tunnel pressure transducers.

The long length of the tubing would tend to smooth out any pressure fluctuations, so the pressure readings must be regarded as time-averaged values.

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The flow quality on the plate surface was checked by oil Row visualization without and with the two- dimensional inserts in the empty test section. The oil flow showed a small separation bubble at the leading edge, hut the elliptic shape minimized the bubble size. Also, some vortices developed on the bottom edge due to its inclination with respect to the main flow. But these events did not indicate a significant change in the pressure data obtained for plate calibration.

The splitter plate swept the flow field above and downstream of the model. Starting well above and up- stream of the model, the survey was made by mov- ing the plate in j-iuch horizontal (downstream) steps. This made the front two columns of pressure taps over- lap the locations of the last two columns of the previ- ous position. After moving downstream of the model, the plate was moved forward and lowered at most 3 inches. Downstream steps were made again. At least 0.25-inch clearance from the model surface was main- tained. When sizeable pressure differences were sensed at the same position by the plate overlap at two dif- ferent steps, the data from the innermost taps on the plate were considered to be the most accurate. Other- wise, data from overlapped taps were averaged.

Plate flutter limited the portion of the field which could he measured. It occured at the most extended, that is, the lowest, strut position when in the wake of the model at high angle of attack. Flutter was moni- tored visually from the tunnel ceiling window, and no data were used when flutter waa present.

nror,w).r *,wall

Pig.3: Probe positioning in the flow field

Pressure Probes

A disk probe, 1.5-inch in diameter and O.?-inch thick (Fig.4-a! was used to measure field static pres- sure for data comparirou with the splitter palate. It has one pressur? tap on the center of each side. Its circumferance is rounded t o obtain better Row quality

LI on the surface. Before it was used, the probe wan cal- ibrated t o make sure both sides read the same static pressure when aligned with the free-stream.

m- 8;

a-) Disk Probe b-) Double-Pitot Tube

Fig.4: Geometric details of pressure probes LJ

(All dimensions in inches)

A pitot pressure probe with forward and backward facing pressure taps was used to measure total pressures at survey stations. The geometric details of the probe are shown in Fig.4-b.

Hot Film Probe

Flow field velocity and turbulence quantities were measured by an end flow X-probe hot film with con- stant temperature anemometer system. The hot film probe was a custom made TSI model 1241. It con- sisted of two sensors 90' to each other and 45' to the probe axis. The probe axis was positioned parallel to the flow, as shown in Fig.5. Each sensor was made of platinum coated film on a quartz wire and had diame- ter of sensing area, 0.002 inches (0.05 mm) and length of sensing area, 0.04 inches (1 mm).

A 90a angle adapter (TSI model 1157) was used to alig:n the probe axis with the main flow. The adapter w-. supported by a dual sensor probe support. Two coaxial cables, one for each sensor, were utilized to c a n y the sensor signal from the dual support to the thermal anemometer in the tunnel control room..

.c/

4

The signals were processed through the Intelligent Flow Analyzer (IFA) IO0 thermal anemometer. The anemometer is capable of expanding to 16 channels and is computer controllable. Sensor overheat can be mon- itored remotely or manually. An overheat ratio of 1.5:1 was used for the hot film probe. The IFA 100 signal conditioner features D.C. suppression, gain and filter settings to condition the anolog signals. A multichan- nel digitizer, IFA 200 waa used for high speed analog to digital conversion of anemometer signals. The IFA 200 was programmed for sampling rate of 20 kHz per channel with simultaneous analog to digital conversion.

[ J Lr I

a3 X I C nwJu--

-

Fig.5: Geometric detai ls of hot fllm probe (All dimensions in inches)

Frequency response of rapid veIocity fluctuations was optimized by square wave testing at maximum free- stream velocity. This test was done every time the sys- tem was activated for measurements.

The conditioned signals were visualized by a digital oscilloscope over the period of data acqutition. The signals were interpereted as the turbulence traces of the flow field and printed out by a plotter.

Hot 51m cal ibrat ion

The hot film probe was calibrated prior to its use for data acquisition. The relationship between the digi- tized anolog data and velocity wan constructed through calibration in a facility using a TSI 1125 calibrator.

The calibration consisted of two stages. First, speed calibration data were acquired for the variation of air flow with sensor voltages when the probe axis was parallel to the air flow. Then the sensors were yawed between -/+ 25 degrees in 5-degree increments to find the axial flow cooling effect in the calibration. The speed and yaw calibration data were fed into the computer to find a relationship between the voltages induced from the sensor cooling effect and velocity of the flow. The calibration curve was non-linear, a de-

IV. RESULTS AND DISCUSSION

The data obtained by traversing along the vertical survey stations will be presented as profiles a t these survey stations. Composite plots of profiles overlayed on the model shape are provided for a better visual data comparison of various measurement stations. The coordinate X is defined to be the axis along the tunnel centerline (positive downstream] passing through the model pivoting location (the half-chord) while Y is the coordinate vertical to X-axis and tangent to the model leading edge at each angle of attack setting. (X = 0 at the leading edge).

Force and Moment

The aerodynamic forces, lift, drag and moment, were measured for all test configurations and reduced to coefficient form as Cr, C, and C,,, respectively. A comparison was made with the previous WSU results reported in Ref.9 and very small shifts in the plots of these coefficients were seen. However, the angle of at- tack range of the present tests was from 8O to 22', which is higher than the range selected in Ref.9.

1 , , , , , , , , , , , , ,,.A I ' " ' I

5 10 15 20 25 -10 -5 k L E OF ATlACK (DECREE)

Fig.6: F l a p se t t i ng effect on the airfoil per- fo rmance

Results of those tests with optimum gaps are pre- sented in Fig.6. It is clear that deflecting the flap from nested position to 40' increases the C,, but decreases the stall angle of attack. For 40" of flap deflection, there

also present in the Ref.9 data. Flow visualization stud- ies showed that this was due to an upwash effect from

is an increase in l i t slope just prior to stalling. Thi.. 0 W a s

" suable feature to operate the hot film sensor?, at low the slot flow to the wing trailing edge. The termination of the wing trailing edge at 0.875 chord rather then 1.0 speed conditions.

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chord (12.5% cutout) reduced the so-called 'Coanda effect" of the slot flow. Instead it helped the slot flow act as an ejector on the wing trailing edge to diminish the effects of adverse pressure gradients. This meant an addition to the lift besides that which resulted from higher circulation around the airfoil.

After the stall, the flap-deflected configurations show a sharp decrease in Cl and C, and increase in C,+ Previously published test results have quit tak- ing data after this abrupt stall. The present test re- sults indicated nearly constant aerodynamic coefficients after this stall, as shown in Fig.6. The higher the flap deflection, the larger the drop and the wider the flat line appears to be. Increasing the angle of attack caused another sharp drop in the aerodynamic coeffi- cients. This interesting feature of flapdeflected config- urations was always present and interpreted as a two- stage stall: stall on the flap and then stall on the wing. Hysteresis studies give a better understanding of this phenomenon.

Hysteresis Effects

The airfoil performance was determined for in- creasing and decreasing angles of attack for all airfoil configurations to check if there is any hysteresis loop on the recovery of post-stdl conditions.

- ; O -'5 0 5 1 0 15 20 25 L H U E O f AlTACS (DECREE)

Big.7: Absence of the hysteresis effect for test case-A.

Fig.? shows this type of study on the single- element airfoil. The aerodynamic coefficients recover from the post-stall condition witbout any loss of per- formance.

Fig.8 shows the same type of measurement for the 30" flap with optimum gap. As can be seen, about 20%

0 of the Cf,,,is lost as the angle of attack is decreased below 14'. The loss of lift continues until the angle is decreased to 7'. There appears to be a downward shift in the lift curve between these two angles. It is known that the primary function of the flap is the opposite of what is happening in the lift curve. That is, the flap increases the C,,=,by shifting the lift curve in the neg- ative angle of attack direction. But the present results show that whatever performance is gained by deflecting the flap is diminished or eliminated within the hystere- sis loop. The hysteresis values were repeatable.

1

r

-10 -5 0 5 10 I5 20 25

a W

ANCLE OF ATTLCK (DECREE)

Big.8: Hysteresis a e c t on the aerodynamic coe5cierits for test case-33.

Surface pressures were measured for the study of hysteresis effects on the flapextended configurations. The selected angles of attack included the beginning, center and termination of the hysteresis loop. At each angle of attack, the surface pressure distribution was obtained twice: the first one was for increasing angle of attack, and the second one was for the decreasing angle of attack within the whole range between -8" and 22". Figs.9, 10 and 11 show examples of the results. As can be seen from these figures, there is no hysteresis effect at 14O and 15.5' angles of attack, but below these angles, corresponding to the hysteresis loop in the l i t plots, there is discrepancy in the surface pressure distri- butions. The highest discrepancy occurs around 12.8' angle of attack, which is near the Ct,,,angle.

Separation eitber on the wing or on the flap plays an important role in explaining both two-stage stall and hysteresis loop. The flow separation on the airfoil elements can be observed by determining the regions of constant surface pressures. Before the hysteresis loop, at 7' angle of attack, there is separation only on about 60% of the flap. As the angle of actack increases, the

6

flow becomes attached both on the 6ap and on the main v wing until the Clma=angle. After the first stall, the flow

is separated from about 65% chord to the main wing trailing edge, but completely separahed from the flap (Fig.10). The flatness on the aerodynamic coefficients between 14" and 15.5" angles of attack corresponds to increasing separation on the main element, as seen in Figs.10 and 11.

_ _ - - - - _ _ - - - - -'-I - - - -

.oo .2s .so .75 9 . 0 0 .7s 1 .00 C W O I D V I E L DIsrmcL ./a

Fig.10: Hysteresis effect on the surface pres- sure distribution for test case-B at a= 14''.

As the angle of attack decreases below 14O, the flow on the flap remains separated and attached again at the end of the hysteresis loop. The flow on the main wing

L partially reattaches beyond the Cj,,,angle of attack. But suction pressures on the main wing are lower than

expected (see Fig.9). This is because the flap does not provide the necessary downwash for the wing as it does while increasing the angle of attack,

.7s 1 1 0

Fig.11: Hysteresis effect on the surface pres- sure distribution for test case-B at a= 15.5'.

The lower line of the hysteresis loop can be at- tained by another method. If the wind tunnel velocity is increased to test speed with the model at a low angle of attack, and then the angle of attack is increased, the upper Ci values result. But if the high angle, say, 12.8' is set before increasing the air speed, the lower leg of the hysteresis plot occnrs. Increasing speed for a fixed angle of attack causes flow separation on the flap at almost all speeds. The probe interference tests shown in Figs. 13 and 14 were conducted in this manner and show totally separated 6ap flow. Those who test multi- element airfoils should be aware of this phenomenon.

Slot Flow Effects on the Hvsteresis LOOP The experimental data obtained for hysteresis

studies show that the loss of the airfoil performance for flapextended configurations are a consequence of the change in Row mechanism around the flap. The flow around the 6ap is very complex due to the interactions between different layers of the flow. The boundary layer 60w over the wing is thick and turbulent over the most of the wing. It terminates at wing trailing edge and extends over the 6ap as a wake with adverse pressure gradient at low angles of attack. On the other hand, the boundary layer over the 6ap is thin and laminar and tends to stay so due to the favorable pressure gradient and convex curvature. This favorable pressure gradi- ent is provided by the flow p,xsi;ig through the gap. The flow is highly accelerated. potential flow with low pressures compared to the neighboring Rows. This low

7 ,

pressure region creates strong pressure gradients across the flow just above the flap.

The slot flow centerline tends to follow the Rap curvature under the influence of strong pressure gradi- ents in the wing wake. The inclination of the centerline from the main flow primarily depends on the Eap de- flection angle. As the angle of attack increases, the angle of centerline with the main flow also increases, influenced by the wing wake over the boundary layer on the flap. Since the wing wake is stronger, it pushes the flap boundary layer dowaward towards the surface, limiting its ability for separation. This type of inter- action continues until the stall-angle of attack. At the stall angle, the separating streamline at the wing trail- ing edge czn not resist the adverse pressure gradients and starts moving just upstream of the trailing edge. When this happens, the boundary layer on the flap be- comes free of strong wake influence. It can now become thicker and stronger. It can even be separated with a slight increase of angle of attack. This is when the first abrupt stall occurs on the flap as explained earlier. In fact, the flap boundary layer always had the potential of separation before the wing boundary layer due to the flap deflection. But with the aid of the slot flow, the wing wake suppresses this potential until the stall.

After the first stall, the separating streamline moves further upstream on the main wing. This is where there appeara to be nearly constant aerodynamic coefficients as mentioned earlier. This steplike change is short and tenniuates by the stall of the main ele- ment. Once both airfoil elements stall, there occurs a large separated region on the airfoil. The separated wing wake mixes with the separated boundary layer on the flap so that their separate identity no longer exists. This causes a big deficiency in the aerodynamic coef- ficients. The airfoil configuration needs to recover this deficiencics by pitching the nose down. The question here is whether the airfoil can gain the same Cr,,, that it had just before the first stall. There has not been a specific answer to this question in the current litera- ture. One very obvious reason for this is the limitation on the top angle of attack. But a complete predictiblity of airfoil performance requires the knowledge of stall re- covery and hence tests of this kind.

Although the main wing recovers its stall and r e gains its part of lift, the flap stays stalled for a wide range of angle of attack. Flat surface pressures on the flap show this. It seems that the flap loses its participa- tion in the lift generation due to its uncambered effect after the stall. Also once the flap becomes separated, its separated wake forms a bubble shape region easily; that is, the flap shear layer combines with the sepa- rating streamline from the flap trailing edge at some

distance downstream. Decreasing angle of attack does not have much effect on the shape of the bubble, at least not as much as the flap deflection angle has on its formation. The rate at which this bubble decays is much slower for decreasing angle of attack than it is for increasing angle of attack. This is simply be- cause once the wing wake and the flap boundary layer are both separated, they exchange momentum between each other for energy balance. The high momentum parts of wing wake are carried through the low mo- mentum parts including the flap flow. A huge mixing takes place at different layers of flow. Even though the geometric requirement of the conditions in which this mixing is created may be reversed precisely, dynamic changes of the wake are irreversable in nature. This process reduces the favorable effect of slot Row on the flap upper surface and leaves the flap boundary layer separated until low angles of attack.

J

The hysteresis effects on the aerodynamic coeffi- cients show that the hysteresis loop for the 40' flap case with optimum gap is the first and 30' flap case with optimum gap is the second largest of all tested configurations. This means that the flap setting plays an important role in the irreversibility of the flow. High flap deflections have higher potential of holding a sep- arated bubble for longer time whenever its conditions are present. This potential becomes more readily avail- able if there is a gap large enough to provide longer slot flow. In other words, the slot flow angle with the main flow and its length on the flap determines the level of hysteresis.

This hysteresis of two-element airfoils may not be present in three-dimensional testing or in actual flight conditions. This may be present only for two- dimensional wind tunnel testing of largely separated flows. The tip vortex of a wing probably assists reat- tachment and the effect ripples along the span. This three-dimensionality of the separated flow is denied within the two-dimensional test section.

L/

Probe maverse Effects

The need for this type of study arose from the fact that any probe used to map flow field properties dis- turbs the surrounding flow to some degree, depending on its size and support system. This probe could be either a relatively smaller size, like hot film probe or a larger size, like the splitter plate. In any case, the level of disturbance needed to be determined for the highly separated flows to estabilish the level of confidence in the flow field measurements.

Probe traverse effccts were first determined for the model's aerodynamic coeffi.:ients. The force measure-

ments were done while the probe traverse was posi- tioned at one of the survey stations, as shown in Fig.3. For the splitter plate, X and Y locations of the pres- sure port no.32 (see Fig.2) was used as the position indicator. Angles of attack were initially set to the pre- scribed angles for this study and then the data were recorded at three ten second intervals at each probe position. The average of these data were compared to those obtained when there was no probe traverse in the flow field. The comparative results showed no signifi- cant change in the aerodynamic coefficients. Data re- peated itself at all probe traverse positions. This meant that the probe traverse does not affect the airfoil per- formance but may change the flow structure on and around the model. This became more clear after the measurements of surface pressures.

L'

2,o CHORDVISE DISTMES. x/o

Fig.12: Probe traverse effect on the surface pressure distribution for test case-A at a= 15"

Figs.12, 13 and 14 show examples of t,he surface pressure distributions while either one of the pressure probes are at the prescribed locations of the survey sta- tions. As can be seen from these figures, the flow field is not affected significantly by the probe presence at any of the survey stations for test cases A and B at near- stall angles of attack. Of the three probes, the double- pitot tube is the smallest and is thought to cause the least disturbance. The splitter plate has the largest size and causes the most disturbance. Therefore. the double-pitot tube was exluded from the comparative study of pressure distribution for the flapextended con- figurations.

Fig.14 shows that the probe disturbs the Bow field at post-stall angle of attack and moves the separation point up or downstream depending on its location. The level of disturbance decreases as the probe is positioned

in the wake region behind the model trailing edge. Be- cause of its size, the splitter plate disturbs the flow field more then the disk probe, as anticipated.

2- .oo .25 .so .75 1.00 .75 7.00

CYOSDWISF c4tllLIICF. 4-

Fig.13: Probe traverse effect on the surface pressure distribution for test case-B at a= 12.8".

I I - 1 4 - I I I - 1 2 - I ' -10-

I I -e- I 1

I

I

I

.OO .25 .so .75 1.00 .7s 7 . 0 0 fH010115L 0nl-L .,.

Fig.14: Probe traverse effect on the surface pressure distribution for test case-B at a= 15.5".

The flow sensitivity to probe presence at post-stall conditions primarily results from the strong interac- tions between different parts of the separated flow field. Introduction of the probe to any part of the separated flow field promotes the unsteady flow character and moves the whole separated wake bubble back and for- ward very quickly. When surveying the flow field, the probe enters the separated wake from its shear layer. The probe affects, not only the strong viscous-ihviscid interaction, but also the vortical motion originating from it.

9

Data Comparison of S u l i t t e r P l a t e and

Disk Probe

The pressure data of the splitter plate were com- pared to those of of the disk probe. The purpose of this comparative study was to determine the credibility of the splitter plate for field pressure measurements.

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . p.... ......... j ................... ..........

. . . . . . . . . . . . . . ....... .yPw -:12.sDEC..

....... ..... .!. . . . . . . . . i ........ : . . . . . . . . I . . ...... .j : 0 s : ,.m :la 1.41 :1* :

-1 1 -o,, . . . . . . . . . . . . . . . . . . . . . . . . . . . . . I. . . . . . . . . i ........ j .......... c p . . . . . .: i s x € _ f :

. . . . . . . . . ........ ........ ........ ........ ..... -0.6 j ; . . . . . . . . . , (. I I .'.'t 0.00 0.25 0.50 0.75 1.00 125 1.50 1.75

A X ! N OISTUCE. X/c

Fig.16: Pressure data comparison of sp l i t t e r p l a t e and disk probe for test case-B at iz.ao.

........................................................... s l . o , ' ' ' .

u' ........................... .......... . . . . . . . . . 0.3 : . U H * - : I S 5 DEG.. + mr rmec 0.6 ' 4 ........ i imr* Y? . . .ra:o?. - i . o n E c i ... .: ........ - s p u r n W E ..........

0

......... ........ ........ ........ ....... . . . . . -0.6 ......... (. I , ,. ,. ...I

AYUL [HSTWCE. X/c

4 : 0.00 0.25 0.50 0.75 1.M) 125 1.50 1.75

Fig.18 Pressure data compar i son of sp l i t t e r p l a t e and disk probe for test case-B at a= 15.5'.

The probe positioning was the same as the one described earlier for determining the probe traverse ef- fects. Both the plate and disk probe surveyed the same stations, but the disk probe picked more data points

4 closer to the model because of its relatively smaller size.

Figs.15 and 16 show the results for test case-B. The data obtained from both instruments agree reasonably well at all stations at near-stall angle of attack (Fig.15). But when the model was set to the post-stall angle of attack, the disk probe pressures are slightly less then the splitter plate data in the most downward positions of these probes (Fig.16).

Due to its relatively bigger size, the splitter plate experiences a thicker boundary layer over its surface than does the disk probe. Therefore the pressures on the splitter plate are likely to be higher than those on the disk probe. Also, three-dimensional effects may flutter the splitter plate in its most extended positions.

P i t o t Pressure Profiles

Flow field total pressure profiles were measured at the survey stations to determine the reverse flow re- gions. The forward (upstream) and backward (down- steram) facing tubes of the measuring probe made the simultaneous total pressure measurements at two sides.

Total pressure profiles were presented in the coef- ficient form as C,, = ( p ~ - p , ) / q - , . The forward and backward pressures reach the free-stream pressure at the vertical survey lines. d

............................................................ s j . . . . . . . . ........ j . . . . . . . . : ....................

NPHI -:12.a DEG.: + F W A R D Ne€ - B T W m TUBE :DELTA - 20.0 CEO.:

yl 0.6 . . . . . . . . !cl!c,..?,?:?! ..... + ........ .........

. - I I . -0.4 . . . . . . . . (.. ...... .:. ........ :.. ....... j ...... ..;. ...... ..,- ....... :

! s c U E > : .LP, :

. . . . . . . . . . . . . . . . . . ........ ........ ...... -0.6 j ! I. I ........,........ ' ' 'I 0.00 0.25 0.50 0.75 1.00 1.25 1.50 1.75

AYW OLSTWE. X/c

Fig.17: Pitot pressure proflles for test case-B at a= 12.8'.

Figs.17 and l a show the pitot pressure profiles for 30' flap case with the optimum gap at near- and post- stall angles of attack. The data show a tendency of increasing both side total pressures ss moved to down- stream. The forward tube readings are positive and the hackward tube readings are negative in the regions of

10

forward flows. But as the flow begins to reverse, the for- ward tube readings are decreasing while the backward tube readings are increasing. In large separated flow regions (Fig.lB), the forward pressures become more negative than the backward ones. After the maximum flow reversal, this flow tendency happens in the other way until the free-stream conditions are reached again.

celerates without any major steepness and reaches the free-stream pressure in relatively short distance down- stream. As can be seen from the figure, the free-stream contour is reached at a downward-rearward location. This means that the Rap pulls the wing wake down- ward vertically to increase the airfoil performance at its optimum setting.

............................................................ Q l . O l . . ' ' :

. . . . . . . . ......... t

........ . . . ..... ........

. . -0.4 . . . . . . . . ;... . . . . ..:.. ...... .:. . ..... ..+. ....... !.. ...... .:. . cp;. ' 'I

i & A : -0.6 1 ; . . . . . . . . . ,. ...... ..,...... .. ', ...... ..,. ...........................

0.00 0.25 0.50 0.75 1.M) 1.25 1.50 1.75 UUL MSTANCE. x/c

Fig.18: Pitot pressure proflles for t es t case-B v at a= 15.5".

Flow Field S ta t i c Pressure Measurements

Flow field static pressures were measured by using the splitter plate described earlier. During the move- ment of the splitter plate in the flow field, there was a gap in the pressure data from the airfoil upper surface. This gap wm filled by the data obtained from the disk probe. Even though data comparison of the splitter plate and disk probe did not show a significant differ- ence, the disk probe was thought to yield better results close to the airfoil. This was because of the possible in- terference between the relatively large size splitter plate and the boundary layer on the airfoil upper surface.

Results are presented by constant pressure con- tours. These contours were connected with the upper surface pressures in the airfoil bounded regions and left isolated in the far wake regions. They approach to the airfoil as nearly orthogonal for the attached flows and parallel for the separated flows.

Figs.19 and 20 show the results for the 30' flap case with optinum gap. At 12.8' angle of attack, a gradual change of the streamwise field pressures about the lifting surfaces indicates it:) favorable effect for the

v flow to be attached all over the upper surface. The flow accelerated at the leading portion of the airfoil de-

Y ................................................................

. . . . . . . . . . . . . . . ; ....... i i ~ ; ; ....... . . . . . . . . ........ ...............

. . . . . . . 1 ........ ; ....... j . . . . . . . j ................. ; ....... I ....... 1

0.00 025 0.50 0.75 l.W 125 1.50 1.75 2.W NUL D S T W E . X/S

Fig.19: Flow 5eld pressure con tour s for test case-B at a= 12.8'.

............................................................. s l .o l ' ' ' . ;

0.00 0.25 050 0.75 l.W 135 1.M 1.75 2.W NUL wr*Na x/c

Pig.20: Flow fleld pressure con tour s for test case-B at a= 15.5'.

At 15.5' post-stall angle of attack (Fig.20), there are isobars along with the flap and wing upper surface. These are the signs of stall both on the wing and on the flap. As the flow becomes separated from the surface, the constant pressure regions are also formed in the flow field of the lifting surfaces and the far wake.

11

Mean velocity and turbulence data were obtained by using an end flow X-probe. The measurements were done at 11 stations for 12.8' and 12 stations for 15.5' angles of attack. All hot film data were non- dimensioned with the free-stream velocity and scaled with the flow field coordinates to be able to present them in composite plots.

Mean Velocitv Resul ts

The mean velocity data are presented by vectoral plots to show the velocity deficits and flow directions at the survey stations. Also, the asymmetric wake flow and slot flow are well represented in these plots.

Fig.21: Mean velocity data for test case-B at a= 12.8'.

Fig.21 shows the results for test cases-B at near- atall angle of attack. As can be seen from the figure, the highly accelerated flow at the 6rst station gradu- ally decelerates as it moves downstream. The velocity profile at the flap trailing edge experiences the largest deficit. The deficit ia recovered at the following down- stream wake stations.

At post-stall angle of attack, the mean velocity data were not displayed correctly in the separated wake regions. This was simply because the hot film probe does not respond to the flow reversals. A split film probe would provide this type of data if it were used. However the results near the separating line are still valid. The double-pitot tube showed the region of flow reversals as mentioned earlicr.

The velocity vectors showed the flow tendency to enclose the separated wake region behind the flap trail-

ing edge at 15.5' angle of attack. That is, the flow from the upper surface inclines downward and the flow from the lower surface inclines upward until they com- bine at some region in the wake. This recombination is interpreted as the termination of the flow reversal. It theoretically corresponds to a free-stagnation point as considered in the separated wake modelling of com- putational investigators. But the overall results of this experimental program reveal that there is no such sin- gular points describing the shear layer recombination. Large separations at post-stall flow conditions make the flow quite unsteady and unpredictable. There are vor- tices of varying strength in the separated wake. They move in clockwise and counterclockwise directions de- pending on which shear layer they are originated from. The vortices from the upper surface in general play a dominant role over those from the lower surface because of their longer length of shear layer.

The accelerated slot flow is well defined by the ve- locity vectors in the gap. The size of this flow is di- rectly determined by the gap. The flow direction of the jet-like flow appears to be upward, different from the surrounding flow vectors. In general, this Bow closely follows the flap upper surface curvature until its ter- mination. But here the slotted flap with almost zero overlap does not show this trend of multi-element air- foils. Instead, there is increasing upward flow as the flap is deflected from 30" to 40°, clearly showing the signs of the vorsical motion on the main wing trailing edge.

\ur

L/

Turbulence Resu l t s

Flow field normal stresses are presented in the composite profiles of U and V components at the sur- vey stations. Fig.22 shows the results. As can be seen, the flow iregularity is more pronounced in the normal stress profiles then it is in the mean velocity data. U- component normal stress is larger then V-componenet at stations near the airfoil model, as expected. But at the stations in the far wake regions, V-component becomes larger, which is believed to be due to the in- creased flow irregularity. This irregularity increases even more with increasing angle of attack or flap de- flection angle.

Flow field turbulent intensities are presented for only profiles of U-component at the survey stations in Fig.23. The +.rends of turbulent intensity profiles are apparantly similar to those of velocity profiles. The maxima and minima of velocity deficits at the survey stations seem to correspond to those of turbulent in- tensities.

ii

............................................................ s ' . o l . . . . ; . . . . . . . . .........

........ . . . . . . . . . . . . . .

........ i ........ . . . . . . . . . j . . . . .

: 1 : I ' i ~ o ~ d ~ o & m n i s : I 21 urc!, ,........I v 1 Urc( ) I 10 -0.6 ......... ,.........I ......... I ........,........ ...... 0.00 0.25 0.50 0.75 1.00

AXIAL DlSTANCf

, .... 1.25 1.50 l.h : x/c

Fig.22: Reynolds normal stress profiles for test case-B at a= 12.8".

............................................................ 1.01"'

, u' 0.8 ...... ..; ........................... . . i . ........ ;.. ....... ;... ...... ; b 1 :UIU -112.8 DEG.;

:Dam - 24.0 0EG.j * u - m P M I ( m * m B n ........ ... ..... $ 0.6 jc/c ?.?! > ........ I

E

, ........ , ........ ., 0.00 0.25 0.50 0.75 1.00 1.25 1.50 1.75

MIN DDTANCE. X/c

Fig.23: Turbulent intensity profiles for test case-B at a= 12.8O.

Flow field shear stresses are presented with pro- files at the survey stations. Fig.24 shows the results. The shear stress is known to be the correlation be- tween the fluctuating U and V velocity components. The flow field complexity is quite visible in the shear stress profiles. In the far wake regions, there are multi- ple maxima and mir.ima in the profiles, corresponding, .. although ver j ronghly, to the mean velocity gradients (Ref.7).

............................................................

' . O l " " i u' 0.8 . . . . . . . . . j . . . . . . . . . . . . . . . . . . j . . . . . . . . . > . . . . . . . . i . . . . . . . . j ..........

i .wIU 4 2 . 8 DEG.. : M T I - m.0 DEC.'

.......... b Y ] - 0.6 . . . . . . . . ic/~...i."?? ..... : ......... 0

........ I ........ ........ ......... ........ . . . . . . . . I ........ I 0.W 0.25 0.50 0.75 1.00 1.25 1.50 1.75

A X I N 0IST.WCE. x/c

Fig.24: Reynolds shear stress profiles for test case-B at a= 12.8'.

V. CONCLUSIONS

Detailed studies of the GA(W)-Z airfoil with 25% slotted flap with moderate through large sepu a t ' ion have been made in this experimental program. Dur- ing these studies, five different airfoil configurations, with the inclusive of the single-element airfoil case, have been tested in a low speed tunnel. The following con- cluding remarks can be made as a result of this exper- imental study:

The airfoil performance data showed that the stall phenomenon occured in two stages for the flapextended configurations. The Brst stall occured on the flap and the second one on the main wing. Both appeared to be rather abrupt. This feature of multi-element airfoils has probably not been reported by the previous inves- tigators, because they have quit taking data after the first stall.

Lift charecteristics of the airfoil model with 40' flap deflection showed a peciluar jump just before the first stall. This was believed to be a result of the flap being slotted.

Airfoil performance was determined in a loop of in- creasing and then decreasing angles of attack. The data showed a large degree of hysteresis loop for the flap-extended configurations. The hysteresis effects ap-

13

peared to be resulting from the complex Row mecha- nism around the slotted Rap and from the nature of two-dimensional testing with low aspect ratio. Sepa- rated Row on a two-element airfoil is seen to depend on the Row history.

The surface pressure measurements conducted for the hysteresis study also confirmed that the Row first sep- arates from the flap and then progresses on the main element wing, contrary to what some other investiga- tors of this field claim.

At post-stall conditions, Butter of the pressure mea- suring instrument became a problem. However data comparison of the splitter plate with the disk probe showed small differences in the regicns of accelerated Row.

Probe traverse effects were determined while either one of the pressure measuring instruments was in the Row field. The model force and moment measurements were not affected by the traversing probe. But some surface pressures at post-stall conditions showed some discrepancies which were the largest in the accelerated Row regions. This was mainly because of the travers- ing right above the model centerline where the surface pressure taps were located.

The reversed Bow region was determined by the total pressure measurements of the doublepitot tube.

Field pressure contours were orthogonal ‘io the sur- face in the attached Bow regions and parallel to the surface in the separated regions. In the near wake or f a r wake regions, there appeared to form the islands of constant pressures depending on the scale of Row sep- aration. The data did not clearly indicate the regions of recompression or recombination as claimed by some theoretical investigators.

The end Bow X-probe hot film anemometer was em- ployed for the measurements of mean ve1ocit.y and tur- bulence quantities at the survey stations.

REFERENCES L,

(1) Braden, J.A., Whipkey, R.R, Jones, G.S. and Lilley, D.E., ‘Experimental Study of the Separating Conflu- ent Boundary Layer”, vol.1-Summary, NASA CR-3655, vol.11-Experimental Data, NASA CR-16608, 1083.

(2) Stevens, W.A., Goradia, S.H. and Braden, J.A., “Matematical Model for Two-Dimensional Multi Ele- ment Airfoils in Viscous Flow”, NA.SA CR-1843, 1971.

(3) Olson, L.E. and Orloff, K.L., “On the Structure of Turbulent Wakes and Merging Shear Layers of Multi- Element Airfoils”, AIAA paper 81-1238, 1981.

(4) Adair, D. and Clifton, H.W., ‘Turbulent Separated Flow in the Vicinity of a Single-slotted Airfoil Flap”, AIAA paper 88-0613, 1988.

(5) Van Den Berg, B. and Oskam, B., ‘Boundary Layer Measurements on a Two-dimensional Wing with Flap and a Comparison with Calculations”, AGARD CP- 271, 1979.

(6) Brune, G.W. and Sikawi, D.A., “Experimental In- vestigation of the Confluent Boundary Layer of a Multi- Element Low Speed Airfoil”, AIAA Paper 83-0566, 1983.

(7) Nakayama, A,, Kreplin, H.-P. and Morgan, H.L., “An Experimental Investigation of Flow Field about a Multi-Element Airfoil”, AIAA paper 88-2035, 1988.

(8) Wentz, W.H., jr. and Seetharam, H.C., ‘Develop- ment of a Fowler Flap System for a High Performance General Aviation Airfoil”, NASA CR-2443, 1974.

(9) Weuts, W.H., jr. and Fisko, K.A., ‘Pressure Dis- tributions for the GA(W)-2 Airfoil with 20% Aileron, 25% Slotted Flap and 30% Fowler Flap”, NASA CR 2948, 1978.

( IO) Biber, K., “Experimental Studies of a TWO- Element Airfoil with Large Separation”, Ph.D. Disser- tation, Wichita State University, Wichita, XS, Novem- ber 1991.

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