Experimental Studies of Flow Separation of the NACA 2412 Airfoil at Low Speeds

Airfoil section lift coefficient, section lift qooc c Airfoil section pitching moment coefficient with m section moment respect to .25c location, 2 qooc Ps - Poo Static pressure coefficient, Pt - Poo Total pressure coefficient, H Shape factor (0*/0**) h Razor blade thickness/2 Local static pressure Local total pressure Free stream static pressure Free stream dynamic pressure RN Reynolds number based on wing chord and free-stream conditions Turbulence intensity T u Velocity at the edge of the boundary layer, non- dimensionalized with respect to free-stream velocity u Local velocity non-dimensionalized with respect to

free stream velocity, j Pt - Ps I

qCX) u Non-dimensionalized component of local velocity in x the free stream direction x Strearnwise coordinate z vertical coordinate Angle of attack, degrees ilp Pressure difference between the pressure reading with blade in position and the true undisturbed static pressure Boundary layer thickness

o

u (l--)dz

0* Boundary layer displacement thickness, ~

U o

o

u u -(l--)dz 0** Boundary layer momentum thickness, U U

J

o

~ Shear stress APPARATUS AND PROCEDURE Tests The experimental investigations were carried out in the VlSU 213 cm x 305 cm (7' x 10') low speed wind tunnel fitted with a 213 cm x 91.4 cm (7' x 3') two-dimensional insert em- ploying a NACA 2412 airfoil section having a 61 cm (24") chord and a 91.4 cm (36") span (Fig. 1). The airfoil was fitted with a 1.07 rom (0.042") I.D. stainless steel surface static pressure taps distributed along the mid-span section. The flow field surveys were conducted at angles of attack of 12.4°, 14.4° and 16.4°.which represent pre-stall, stall, and post-stall conditions respectively. Reynolds number of the test was 2.2 x 10 based on the airfoil chord and Mach number was 0.13. Transition was ensured by employing 2.5 rom (0.1") wide strips of #80 carborundum grit at O.OSc on both upper and lower surfaces. In this test series details of flow field were investigated only on the upper surface of the model, and in the wake. At each angle of attack about fourteen chordwise survey stations were selected covering the airfoil upp~rsur face and the wake.

Basic force measurements, surface pressure and local skin friction distributions, flow visualization and hot-film sur- veys were also obtained to supplement the flow field data.

Instrumentation Velocities at heights more than 2.5 rom (.10") above the local surface of the airfoil were obtained by employing a five-tube pressure sensing pitch-yaw probe of 3.175 mm (0.125") diameter (Fig. 2). Velocities very close to the wall and in regions of flow reversal were obtained by a four-tube probe having a pair of pitot and static tubes positioned 180 apart along the tube axis (Fig. 3). The axis of the static tube was located at a height of 0.25 mm (0.01") above the pitot-tube axis. Four- and five-tube probes were mounted in tandem, straddling the model centerline, spaced 7.62 em (3") on either side of the centerline.

Hot-film surveys were conducted to scan the regions of moderate and heavy turbulence employing a Thermo Systems, Inc., 0.05 mm (.002") diameter probe with linearizer (Fig. 4).

Local skin friction was measured by the technique out- lined by East (Ref. 7) employing commercially available razor blades of 0.1 mm (0.0041") thickness. Each blade was trimmed to a 6.4 x 6.4 mm (0.25" x 0.25") square and positioned at the surface static port location where the local skin friction , was to be evaluated. Details of the razor blade dimensions are given in Figure 5.

Unbonded strain gage pressure transducers with a range of ±17.2 kilo-newtons/m (±2.5 psi) were used for all pressure measurements. All pressure measurements were recorded on punch cards.

Methods Lift and moment data were obtained from the tunnel main balance system. The drag was calculated from wake surveys measured at the 0.5c station downstream from the trailing edge.

Flow velocity data were acquired by initially tilting the four- and five-tube probes to align with the local slope of the sur- face. Near-wall velocity data were obtained from the four-tube probe readings. For distances more than 2.5 mm above the sur- face, the five-tube pressure readings were used to obtain total and static pressure, as well as local upwash angle through ap- propriate calibration curves.

Flow reversal was indicated by observing the higher read- ing from the forward- and aft-facing total tubes on the four- tube probe. The data reduction program selected flo~ direction based upon these readings, and utilized the appropriate static pressure tube reading to calculate velocity. Thus the four- tube probe readings were utilized for stations very near the wall, where five-tube data could not be obtained, and for re- gions of reversed flow. Attempts to obtain readings by r9- tating the five-tube probe 180 in yaw for regions of flow re- versal were unsatisfactory. The data usually indicated flow direction opposite to probe direction for both forward and reversed positions. The four-tube probe gave reasonably con- sistent results. The discrepancies between the two instru- ments are attributed to the unsteady nature of the reversed flow, and the high damping characteristics of the five-tube - ---. -. -. . .

probe. Measurements in the wake were made with the probes aligned in the free-stream direction (zero tilt).

Tuft surveys and oil flow methods were employed for ob- servation of the surface flow patterns and determination of the separation point.

Hot-film surveys were made with the traversing mechanism employed for the four- and five-tube surveys. Photos of the velocity fluctuations displayed on the oscilloscope were also recorded.

Local skin friction was measured by positioning the razor blade as shown in Figure 5. This method involves relating the skin friction (.) to the difference between the pressure re- corded by the static hole with the blade in position, and the true undisturbed local surface static pressure (blade removed).

Details of geometrical limitations and calibration are given in Reference 7. Important dimensions are tabulated in Figure 5, for the present experimental set-up.

Data Reduction Force data, with usual wind tunnel boundary corrections, surface pressures, local velocities and flow inclinations were calculated from the measured wind tunnel raw data by com- puter routines developed for the IBM 1130 and 360 computers at WSU. The local velocity is expressed in a non-dimensional form as the ratio of local to free stream velocity. Experi- mental velocity profiles were plotted by a computer routine written for the IBM 1130 computer.

Calibration of the five-tube probe is discussed in de- tail in Reference 8. All the pressure instrumentation employed in the present tests is heavily damped and therefore records time-averaged values.

Typical oscilloscope traces from the hot-film probe were photographically· recorded. Digital volt meter readings of the hot-film probe data were recorded manually. The hot-film was calibrated from time to time during the course of the tests to compensate for wind tunnel temperature variations.

Maximum calibration shifts amounted to 6% of free stream veloc- ity.

The pressure difference 6p, between the surface pressure recorded by the static port with the blade in position and the undisturbed static pressure, is related to the skin friction L by a calibration equation given in Reference 7. The data reduction program utilizes this equation to calculate the local skin friction coefficient.

RESULTS Presentation of Results The results of the present investigation are presented in figures as tabulated below: Table 1 - List of Figures Instrument Figures Type data Airfoil geometry 1 Instrument de- 2 to 5 tails Lift, drag and Force balance and pitching moment wake probe Surface pres- Surface tubes 7 sures Surface flow 8 Tufts Velocity pro- Five-tube and 12.4°,14.4°, files four-tube probes 16.4° Near-wall veloc- Five-tube and 12.4°,14.4°, ity profiles four-tube probes 16.4° Static pressure Five-tube probe 12.4°,14.4°, profiles and surface pres- 11 16.4° sure tubes Static pressure 12.4°,14.4°, Five-tube probe 12 field contours 16.4° Boundary layer 12.4°,14.4°, displacement Five-tube probe 13 16.4° thickness Boundary layer 12.4°,14.4°, momentum thick- Five-tube probe 14 16.4° ness Boundary layer 12.4°,14.4°, Five-tube probe 15 shape factor 16.4° Displacement 12.4°,14.4°, Five-tube probe thickness 16.4° Separation Five-tube probe 16.4° 17 streamline Velocity and 12.4°,14.4°, pressures in Five-tube probe 18 16.4° wake 6a Table 1 - (continued) Instrument CL Figures Type data

--

Total pressure 12.4°,14.4°, Five-tube probe 19 contours in wake 16.4° Hot-film anemo- 12.4°,14.4°, Hot-film field surveys meter 16.4° Skin friction Razor blade 0.2°,12.4°, 14.4°,16.4° Discussion Forces: (Figure 6). These tests were conducted primarily to supplement the flow field data and to provide additional data at low Reynolds number with NASA standard roughness.

Results of the lift, drag and pitching moment measure- ments are shown along with the experimental data at Reynolds 6 6 numbers of 3.1 x-l0 and 5.7 x 10 from Reference 9. It can be seen from Figure 6a that present experimental results of the lift coefficient for the clean model do not agree with the clean model data of Reference 9. A difference in the maximum lift coefficient of 0.15 can be seen between the 6 6 NACA 3.1 x 10 RN and the WSU 2.2 x 10 RN clean data. This difference appears to be rather large for a difference in Reynolds number of 1 x 10 . The reason for this difference is unknown. The differences between NACA data and WSU data with grit are expected since the NACA grit was larger and was ap- plied over a much larger region, causing severe losses in Cn x.max and corresponding increases in drag at high lift coefficients.

The experimental drag and pitching moment data (Figures 6a and 6c) agree reasonably well with the results of Reference 9.

Pressure Distributions: (Figure 7). Surface pressure distributions for an angle of attack range from -3.9° to +12.4° are shown in figure 7a. Figure 7b, 7c and 7d show pressure distributions at the 12.4°, 14.4° and 16.4° conditions selected for detailed flow studies in the present research.

Theoretical pressure distributions from the method of Refer- ence 10 are also given. Separation locations from flow visual- ization studies at these angles of attack are marked on the figures. It is seen that these locations appear quite consis- tent with the beginning of a region of constant pressure for each angle. Constancy of pressure is characteristic of sep- arated flow regions.

Flow visualization studies: (Figure 8). Flow visualiza- tion studies were carried out by attaching tufts to the upper surface of the model. In order to study the influence of the side wall boundary layers on separation patterns, tufts were also applied to the side walls. No evidence of premature side wall separation was observed. Figure Sa shows the tuft photos for a nominal angle of attack range of 0° to 12°. The flow is very steady up to 8°. At 12° the last row of tufts is disturbed with a few tufts exhibiting reversal near the mid-span section. At 14° angle of attack (stall) separation progresses upstream with the last two rows of tufts (aft of O.SO chord) showing reversal. Tufts at the 0.70 chord station are disturbed, with some tufts indicating possible intermittent reversal at near mid-span. The flow is reason- ably two-dimensional (Figure 8b). At post-stall angles of

attack (a = 16°, 18° and 20°) the regions of separation grow

larger and larger and the flow pattern becomes asymmetric.

Thus the extreme post-stall flow pattern appears to have a three-dimensional character. It is interesting to note that the tufts on the side walls are undisturbed.

Limited oil-flow studies were conducted at the pre- stall, stall and post-stall angles of attack, to obtain more detailed definition of separation locations than the tuft studies provide. Since the oil flow is heavily damped, the surface streak patterns tend to represent a mean separation location which is much more difficult to define from tuft patterns. Results of the ilnalysis of combined oil flow and tuft studies are given in Table 2, along with theoretical values calculated by the methods of Reference 10.

Table 2 - Separation Experimental Theoretical Angle of Separation Location Separation Attack from Oil and Tuft Studies Location 12.4° .925c .92c 14.4° .BOc .B2c 16.4° .40c .65c These observations are consistent with surface pressure dis- tributions. These results are in contrast to separation pat- terns for the GA(W}-l airfoil as reported in Reference 1. In the present case, the separation location moved forward 0.40c for a 2° change in angle of attack from 14.4° to 16.4°. In the case of the 17% thick GA(W}-l section, an BOchange in angle of attack moved the separation location only 0.35c.

Velocity plots: (Figures 9 and 10). Computer plots of the measured velocity profiles at the mid-span section are shown in Figures 9a to 9c. The five-tube probe did not in- dicate reversed flow either on the airfoil surface or in the wake. In regions where reversed flow exists either the cali- bration limits of the probe were exceeded or the indicated local dynamic pressure was negative for probe yaw directions of both 0° and 180°.

Velocity profiles obtained from the four-tube probe are shown in Figures lOa to lOb, together with velocities obtained from the five-tube probe for certain z stations. It can be seen that four-tube and five-tube measurements in general agree within ±5% of the free stream velocity. The error appears to be the largest at the O.lOc station (Figure lOb). Discre- pancies between the probe types do not follow any consistent pattern. It is believed that the transition strip contributes to unsteadiness and boundary layer profile distortions at the O.lOc and 0.20c stations.

The flow over the aft portion of an airfoil at stall and post-stall angles of attack is unsteady with intermittent re- versing. The Appendix to this report compares velocity profile measurements as obtained by the two pressure probes used in the present tests and a high-response split-film anemometer used in subsequent separated flow research. These studies show that the turbulent fluctuations near reversal are large, but that the pressure probes indicate approximately the average velocity, and approximately the proper averaqe reversal loca- tion.

Static Pressure Profiles: (Figure 11). Static pressure profiles at various chordwise stations on the airfoil shown in Figures lla to llc were obtained using the five-tube probe.

The surface static pressures as extrapolated from this data show some disagreement with those measured by the surface static pressure ports with no probe present. Special runs to determine the magnitude of the probe interference were made when these effects were observed. These are discussed in appendix B. The runs showed that the probe creates an interference effect which results in slight changes in the separation point and associated changes in pressure levels at the post-stall angle. In some uses the surface static pressure coefficients changes as much as 0.25.

Static pressure contours: (Figure 12). Static pressure contours derived from the pressure dis~ributions obtained at ten chordwise stations and four stations in the wake are shown.

The characteristic high pressure plateau reported in Reference 1 can be seen in Figures 12a and 12b (n = 12.4° and 14.4°).

At the post-stall angle of attack of 16.4° (Figure 12c) however there is no indication of a high pressure region. A vertical pressure gradient from lower surface region to upper surface is also observed in the wake.

(Figures 13 through 17).

Boundary layer characteristics: The displacement and momentum thicknesses show substantial in- creases between pre-stall and post-stall conditions. The rapid growth of the shape factor prior to separation, typical of turbulent separated boundary layers, is clearly seen. A com- parison between measured shape factors at separation and typical value.s of H· (from Ref. 11) is shown in Table 3.

Table 3 - Shape Factors at Separation Separation Point Measured Normally Angle from Tuft and Oil Shape Expected of Attack Flow Value of H Observations Factor H 12.4° .925c 2.12 1.8 to 2.2 .1. 8 .SOc 1. 97 to 2.2 14.4° 16.4° .40c 1. 53 1.8 to 2.2 It is seen that the values of 12.4° and 14.4° are within the normal range of values, while the shape factor at the 16.4° condition is below the normal value. This is believed to be caused by the characteristic post-stall turbulent fluctuations.

Also it should be noted that the flow at this condition is somewhat three-dimensional.

The boundary layer displacement thickness superimposed on the airfoil is shown in Figure 16. It can be seen that the slope of the augmented surface follows the slope of the air- foil surface very closely up to the point of separation and diverges away depending on the depth of the separated layer.

This trend is also exhibited by the separation streamline which is shown in Figure 17 for the angle of attack of 16.4°.

Velocity and pressure distributions in the wake: (Figures 18 and 19). Profiles of velocity, static and total pressure are shown for a vertical traverse range of ±0.2c at each chord- wise station. The progressive growth of the wake width in the longitudinal direction is seen as expected. Static pres- sure profiles exhibit a slight vertical pressure gradient in the direction of the lower surface. Progressive reduction of static pressure gradients in the wake can also be seen.

Total pressure profiles are very regular even at post- stall angles of attack. Contour plots of total pressure (Figures 19a to 19c) are similar at pre-stall, stall and post-stall conditions. Total pressure gradients become smaller at the post-stall angle of attack compared to the pre- stall angle of attack.

Reattachment point in the wake: An examination of the wake velocity profiles (Figures l8a to l8c) indicates the termination of regions of reversal within a relatively short distance downstream from the airfoil trailing edge. This point, which is characterized by a single zero velocity point in the velocity profile, is referred to as the "reattachment point." The points obtained by inspection of the velocity profiles are tabulated in the following table.

Table 4 - Reattachment Point Location Angle of Attack Reattachment Point 1,00 < x/c < 1.05 1.00 < x/c < 1.05 x/c ~ 1.2 These results are consistent with the observations of Reference 1, which also showed that reattachment points for the GA(W)-l airfoil were relatively close to the trailing edge.

Hot-film survey: (Figure 20). Maps of the regions with varying degrees of turbulence are shown. Typical oscilloscope traces of regions of reversal, heavy turbulence, moderate tur- bulence and smooth flow are also shown. Interpretation of the hot-film data for regions of reversal was done in the follow- ing manner. The flow was considered to be reversing whenever the trace indicated zero on the scope. At the outer edge of reversal zones, the flow may be intermittently reversing (less than 50% of the time), in view of the heavy turbulent mixing.

Pressure instrumentation cannot follow the higher flow fre- quencies because of heavy damping. It is interesting to note that regions of reversal measured in this way extend further downstream than the results obtained from the heavily damped pressure probes. Regions of heavy turbulence extend far be- yond 0.50c downstream from the trailing edge for the case of post-stall angle of attack (Figure 20c), whereas for the cases of pre-stall and stall conditions, the regions of heavy tur- bulence terminate within 0.30c downstream (Figures 20a and 20b).

Skin friction distributions: (Figure 21). Local skin fric~ion measurements are compared with theoretical results calculated by the theoretical methods of Reference 7. At low

angle of attack (a = 0.2°, Figure 2la) the upper surface ex-

perimental data show a somewhat higher level of skin friction than theory, while the lower surface data show excellent agree-

ment with theory. At a = 12.4° (Figure 2lb) , the upper sur-

face data show excellent agreement with theory, while the lower surface experimental data are somewhat higher than theory.

At a = 14.4° and 16.4° (Figures 2lc and 2ld) the agreement

between the theory and experiment is good for stations ahead of separation.

CONCLUSIONS 1. Experimental velocity profiles, flow inclinations, static and total pressure distributions have been obtained for the NACA 2412 airfoil, at pre-stall, stall, and post-stall angle of attack conditions.

2. Extensive mapping of the regions with varying degrees of turbulence was done employing a hot-film survey probe.

3. Surface pressure distributions, separation locations, displacement thickness, momentum thickness, shape factor and skin friction show reasonable agreement with theory up to the separation point. Post-separated values are not predicted by present theory.

4. Velocity measurements from the pressure-type probes indicate that the regions of reversed flow terminate at a re- attachment point which is located a relatively short distance (about O.OSc to 0.2c) downstream from the airfoil trailing edge for the test range of angles of attack. The hot-film measurements reveal that intermittent reversal extends somewhat further down- stream than pressure type probe data indicate, but even these re- gions are less than O.Sc in length.

REFERENCES 1. Seetharam, H.C. and "lentz, H.H., Jr.: Experimental Studies of Flow Separation and Stalling on a Two-Dimensional Air- foil at Low Speeds. NASA CR-2560, July 1975.

2. Wentz, W.H., Jr., and Seetharam, H.C.: Development of a Fowler Flap System for a High Performance General Aviation Airfoil. NASA CR-2443, December 1974.

3. Seetharam, H.C. and Wentz, W.H., Jr.: A Low Speed Two- Dimensional Study of Flow Separation on the GA(W)-l Air- foil with 30-Percent Chord Fowler Flap. NASA CR-2844, May 1977.

4. Wentz, W.H., Jr., Seetharam, H.C., and Fiscko, K.A.: Force and Pres- sure Tests of the GA(W)-l Airfoil with a 20% Aileron and Pres- sure Tests with a 30% Fowler Flap. NASA CR-2833, June 1977.

5. Naik, S.N. and Zumwalt, G.W.: An Analytical Model for the S~ud~ of Highly Separated Flow on Low Speed Airfoils, Wlchlta State University, Aeronautical Engineering Dept., AR 77-2, 1977 (in preparation as NASA CR report).

6. r~echtley, E.A.: The International System of Units--Phys- lcal Constants and Conversion Factors (Revised). NASA SP-7012, 1969.

7. East, L.F.: Measurement of Skin Friction at Low Subsonic Speeds by the Razor-Blade Technique. British R&M 3525.

8. Seetharam, H.C., Wentz, W.H., Jr., and Walker, J.K.: Mea- surement of Post-Separated Flowfields on Airfoils. Tech.

Note, AIAA Jour. of Aircraft, Vol. 14, No.1, Jan. 1977.

9. Abbott, I.H. and Von Doenhoff, A.E.: Theory of Wing Sec- tions. Dover Publications, 1959.

10. NASA Langley Research Staff: Viscous Flow Single-Element Airfoil Analysis Computing Routine (Program Airfoil)-, 1976.

11. Schlichting, H.: Boundary Layer Theory, McGraw-Hill Co., Fourth Edition, 1962.

.. ., It II If

~

c==

~- .. - --- ----..-... -.- '-- ..

.._-- ----- .. -.

UPPER SURFACE LOWER SURFACE x/c z/c x/c z/c 0.0000 0.0000 0.0000 0.0000 -.0001 .0028 .0005 -.0028 .0005 .0056 .0015 -.0054 .0012 .0080 .0028 -.0076 .0020 .0098 .0040 -.0092 .0038 .0127 .0062 -.0117 .0060 .0155 .0090 -.0140 .0083 .0180 .0117 -.0160 .0130 .0220 .0170 -.0191 .0178 .0255 .0223 -.0215 .0274 .0312 .0326 -.0254 .0371 .0360 .0429 -.0283 .0469 .0401 .0531 -.0307 .0568 .0438 .0633 -.0327 .0666 .0471 .0734 -.0343 .0766 .0501 .0834 -.0358 .0865 .0529 .0935 -.0369 .0965 .0555 .1035 -.0380 .1165 .0600 .1235 -.0395 .1366 .0638 .1434 -.0407 .1568 .0671 .1633 -.0415 .1769 .0699 .1831 -.0420 .1971 .0723 .2029 -.0423 .2478 .0765 .2522 -.0422 .2985 .0788 .3015 -.0413 .3493 .0792 .3508 -.0398 .4000 .0780 .4000 -.0380 .4503 .0757 .4497 -.0360 .5006 .0724 . .4994 -.0335 .5508 .0683 .5492 -.0308 .6010 .0634 .5990 -.0278 .6512 .0578 .6488 -.0248 .7012 .0516 .6988 -.0216 .7512 .0448 .7488 -.0184 .8012 .0373 .7988 -.0151 .8510 .0293 .8490 -.0118 :) .9008 .0206 .8992 -.0083 .9505 .0113 .9495 -.0049 1. 0001 .0013 .9999 .0013 Figure 1 - Coordinates of 2412 Airfoil.

itch Ports (P &P ) l Total . '3, Pressure ~.I ......... ___ __ Port (P 5) Yaw Ports (P & P ) 3.175 mm (.125") 2 4"\ 12.7 mnt.(.5") R P l mm (.5") R- Tip Details 6.35 nun.

-

~

(.25") q .. D. 3.175 nun (.125")

0.0.

I /

1.905 nun (.075;; 0.0 •

• 457 rom (.018") M~

\

• 30·'

2.362 nun (.09311)~/'~

1.168 mrn (.046") J I-

1.016 Ir.In (.040. )-1

9.525 mm ..

(.375") 0.0.

Probe Wind Shield --~---------- ---- Figure 2 - Five Tube Probe .

. 17

Hemispherical End"

n

Section AA Static Pressure

-- ---~-;~~-~i~!

1------12.80 Probe Details ( • 5 II ) Holes!

JI

co r J. 5D

it Static o -Holes

----- 12. 80 mm---~

r (.5")

8D o - Outside Diameter = 1.09 mm (.043")

1. 28 mm (.05") Hemispherical End

r)

.2~ mm ~\ .67 mm

(.008") (.026") 'Tip Details .77 mm (.03") Figure 3 - Four Tube Probe.

~~ ____ ~~ ______ ~t I~

J I 1.0 .... (.04")

3.2 Jl1m (.125") Dia. r--12. 7 InI" (.50")--..

.....

\.0 1.0 mm (.04") ~~J-.

-"l. _" 1. 6 7 Jl1m (. 0 6 5 .. )

!,-;O 50 mnl' (. 002 .. ) D i a .-'""--

C:=:======::!I=r'

Tip Detail Figure 4 - Hot Film Probe.

Flow ..

Dimensions 1. 07 mm (.042 ) d (.018 ,) e .46 mm .05 rnm (.002") h Surface st~ti;;-pressureTap-\ (.25 ) 1 6.35 mm b 6.35 mm (.25") 20.5 d/h b/h 122.0 l/b 1.0 0.0 e 6x b Criteria from: Ref.7:

~1

b

.. I

1- h > 36

! = 1

b

. /J.x = 0

------- ~--.

---_.-----------_.- Figure 5 - Razor Blade Technique: Details of Dimensions and Positioning.

1.6 ~. 4 1.2 1.0 .8 cR- .6 'SYMBOLS

d WITH GRIT

.• 4

o CLEAN

6 CLEAN

·0 NACA GRIT

§m 12 ~§§ 16 i

-4 (DEGREES) ,;..;.

-...2 a) Lift Figure 6 - Aerodynamic Coefficient Variation of 2412 Airfoil.

,.-;- of , .'

I SYMBOLS R.N.

SOURCE "

o WITH ,GRIT

2.2 x 10 W.S.U.

o CLEAN 2.2 x 10 W.S.U.

6. CLEAN ~,~ 1: x 10 REF. 9 -- --- --'"- .. ----

o NACA GRIT 5.7 x 10 REF. 9

, c~ 0.05 1.,0 ~ ffi

l

N N " t1 I cd -0.1

: . c

m - I r -0,.2 inmnmm

I mmltl HH

, ..

!

1.0 c) Pitching Moment b) Drag Figure 6 -,C9nc1uded.

Alpha Symbol + -3.9° x 2°

~ 4: 3,0

-7 .. 00 ~ 8.3° 41 12.4°.

(Flagged symbols denote lower surface.L -6 .. 00

RN = 2.2 x 1.0

Mach No. = 0.13

-2 .. 00 -1. .. 00 1. .. 00

c

a) Pre-st~ll angles of attack Figure 7 - Pressure Distributions of 2412 Airfoil.

-7.00

o Present Tests

-6.00 - --Theory (Ref. 10) -5.00 S: Separation Point

d

-4.00 \ ~ c p

\

-3.00

q

0'

0\

-2.00

~

0-

0-

0- C).

-1. 00

<9

S

e

S

%6)9

(,)-Q .20 .40 .60 ~_ 0.0

. Sf) Q Q..Q..Q.0-.0..0..0

1.00

Q 0_ J:> .80

~ xI c 1. 00 b) Pre-stall angle of attack, a=12.4° Figure 7 - Continued.

t

= -9.15

I c p min

-7.00

o Present Tests

-6.00 --- Theory (Ref. 10) S: Separation Point -5.00 -4.00 c p -3.00 -2.00 -1. 00 .20 .40 r.. 0 r.'\ 0 0 ~"--

O 00 " \:1- .)::z' - -aD

1. 00

O

0 0 "---- .

0; - XI c 1.00 c) Stall angle of attack, a=14.4° Figure 7 - Continued.

= -11.50

-7.00 o Present Tests .

-6.00 -- -Theory (Ref. 10) S: Separation Point -5.00 -4.00 c p -3.00

o

o

1. 0 L...nootW-..o- d) Post-stall angle of attack, a=16.4° Figure 7 - Concluded.

a} Pre-stall angles of attack Eigure 8 - Tuft Studies.

b) Stall and post-stall angles of attack Figure 8 - Concluded.

.. -

..

'.

NACA 2412 .1 Alpha = 12.4°

RN = 2.2 X 10

N

Mach No. = 0.13

lO , U inf - a) Pre-stall angle of attack, a=12.4° Figure 9 - Experimental Velocity Profiles.

NACA 2412

Alpha = 14.4°

RN = 2.2 X 10

w

Mach No. = 0.13

o U inf _ b) Stall angle of attack, a=14.40 Figure 9 - Continued.

NACA 2412 Alpha::: 16.4°

RN = 2.2 X 10

Mach No. = 0.13

U inf _ Note: Dashed lines indicate Four Tube data c) Post-stall angle of attack, a=16.4° Figure 9 - Concluded.

/j~rimed-: Five Tube Probe Unprimed: Four Tube Probe .010 Xj' Symbol c

----cf:1o--

0.20 El .008 <) 0.30 0.40 t::.

~ 0.50 0.60 ~ .0GEl 0.70 <3> 0.80 0.90 ~ 1. 00 .004 ' -----_-r._~_ .. ___ .002

Q~~~~~~~~~~~~

2.0 0. 1.0 u x a) Pre-stall, angle of attack, a=12.4Q -------.-- Figure 10 - Near Wall Velocity Profiles.

; 32 'Primed: Five Tube Prohe ------ ~ -- -- Unprimed: Four Tube Probe _. __ ._---.

·.010

XI '

Symbol c-- 0.10 0.2Q .008 0.30

<>

I 0.40 ~ 0.50 ~ 0.60 ~ .006 ., 0.70 0 0.80 0.90 (!)

1.00 .00 4 ~ ... ----.- .002

o

.. -- Q.

1.0 2.0 b) Stall angle of attack, a-14.4° -- . __ .... - ._.-.--_.- Figure 10 - Continued.

I 33 Primed: Five Tube Probe Unpri~€d: Four Tuhe Prohe xI .010 SVlT'Dol c 0.10

a

[J 0.20 <) 0.30 .008 0 . .:10 8.

0.50 ~ 0.60 ~ .006 0.70 Q 0.80 .004 0.90 ~ 1.00 .002

o

o 1.0 2.0

----------------------------------------

c) Post-stall angle of attack, ~=16.4° Figure 10 - Concluded.

33a xI Symbol c .10 .20 .30

<>

.40 .50 ~ .60

D

(l .70 <;> .80 .90 1.00 ~ .10 .05 -4 a) Pre-stall angle of attack, a=12.4° Figure 11 - Static Pressure Profiles.

xI

Symbol c .10 0 .20 .30 8 .40 .50 ~ .60

D

8 .70 0 .80 .90 1. 00 ~ .10 .05 o -1 -2 -3 -4 b) Stall angle of attack, a=14.4° Figure 11 - Continued.

xI

Symbol c 0 .10 0 .20 .30

<:>

.40 .50 ~ ~ .60 .70 <) .80

0 .90

~ 1. 00 .1 .05 OL-~~~~ __ -L~~~ __ ~ ______ ~

o -1

-2 -3 -4 c) Post-stall angle of attack, a=16.4° Figure 11 - Concluded.

_---~-- -1.0 ; ....

/ / / -1.25 I

, ,-

--

", -.50· , / -1.50' -.40

-2.0~.1 / /:-J-':-1 5

-.30

. I, ..-

- 2. 5.2Ji I.r--- - - -

~- ..

'( \, I.' ,-:;." ~"'"-----'''---'':~ -3.0 ,I -.10 NACA 2412 Alpha = 12.4° -.05 -.04

RN = 2.2 x 10

-.03 Mach No. = 0.13 -.02 -.01 0.0 c

Ps = .015

a) Pre-stall angle of attack, a=12.4° Figure 12 - Static Pressure Field Contours.

..

01""'" - - I ,. '" I / _1.. SO I I ,-

't /'_7-.00

I I -- -

\

\\ I ,/ <=.0 ., \ _ 2. J

~'\: ~----~------~-

''-~.

-.60 -.65 -.80 ___ -- __ -.50 -1 -.90 -------- • OO~ _----- ,- -1.10 ,," __ ~-------

~/"" ",,---- ----

-1.25 ----.1 " / ",,------, i--;--t / " .... -- -- ......

0 I I , I" -- - ....

1 -. , , I --f ( ," ' .... -.40 l ~ __ ----..::"--.l.....~ -.20 -2.00 /-... ...... ~ -2.10 -.25 , w: lO NACA 2412 -.20

= -.40

Alpha = 16.4°

-.10 6 -.30 RN= 2.2 x 10

Mach No. = 0.13

---------- - . 01 c) Post-stall ansle of attack,a=l6.4° Figure 12 - Concluded.

·03

*

.02 0/ c .01 0 1. 00 ·20 ·40 ·60 ·80 x .

/c .04 [] '.03

*

I 0/ c S 0,.../ .02 ~Jr, "..

.01

cr"1J

-tJ -0- -£)- .

.20 .40 .60 .80

a 1. 00

x/ c .12

-'

Separation Point S: Symbol a.

--

.10 0 12.4° [J 14.4° 16.4°

'0 0

.08 Theory

---

(Ref. 10)

*

0/ c .06

<>

. 04 I

0 /

/ .02 ./ "..

S

<>

--

---

---

6.

---

-

~- 0 .20 .40 .60 .80 1. 00 x/ c Figure 13 - Boundary Layer Displacement Thickness.

.02

**

<5

Ic

S .01

d/~

e -

9-

"_ G- -0- -0--e-

O .20 .40 .60 .80 1. 00

xI

c .02

**

<5

Ic

s .01

~

j 9/

_ [3--0-

"-

._Q._o--8 .

0 .20 .40 .60 .80 1. 00

xI

c .05 s: Separation Point .04 S:lmbo1 a.

0 12.4° .03 14.4° E1

**

<5 16.4°

Ic

Theory

---

(Ref . 10) . 02

<>

/ .01 / S /"

-

-

-

--

--

~-~

o ~~~~~~---r------~------~----~

.20 .40 .60 .80 1. 00 o

xI

c Figure 14 - Boundary Layer Momentum Thickness.

3.0 1-0 / ./ .".

2.0

0\

-

-

--

-

H

-- 0

--- -

S <:)

°

1.0 S : Separation Point 1.00 .20 .40 .60 .80

xI

c [J 3.0

--- ./

./ ./ "./ 2.0 .....

H -

- 0

- -

I

--- 0

·0 S El __ 7 -'- :l.

. ~ El 1.0

o ~----~-------r------~------r-----~

.20 .40 .60 .80 1.00

o

.4.0

<>

3.0

o

,.--

-0--

H

./0

S . .". Symbol L~/<> 2.0

o

12.4°

-----0 <>

G 14.4° <:> <> <> .

16.4°

<:>

Theory 1.0 (Ref. 10)

o ~----~------~------~------~----~

1.00

o .20 .40 .60 .80

xI

c Figure 15 - Boundary Layer Shape Factor.

Symbol ~a ____ 12.4° Figure 16 - Boundary Layer Displacement Thickness Distribution.

p

-

~---------------- __ -=~_~o_--~---~

Figure 17 - Separation Streamline, a=16.40.

z/c .20 .10 -1 -1 1 -1 -1 -1 1 -1 u u' x x a} b) z/c, .20 -1 -1 1 -1 -1 -1 1 -1 1 x c) d) a) Pre-stall angle of attack, a=12~4° ,- Velocity and Pressure. Distributions. in the Wake.

Figure 18 z/c :20 .10 -1 -1 1 -1 -1 -1 c x u Ps x

x / c = 1. 05

b) z/c .

• 2 .10 ":"1 -1 -I -1 -1 -1 1 -1 c c u u x x t P s

c) X/c = 1.20

d) --. - ~ -. - - -_ .. - - - .

b) Stall angle of attack, a=14.4° ". - - Figure 18 - Continued.

.-.

o - Four Tube·

o - Five Tube.

z/c -1 c c u c c U I P P X P s x.

t P t s

-

xI

,a) xI

= 1'.05 = 1.10 b) c .c z/c o 1 ---1, -1' -1 l' -1 c P' t.

d)

XI = 1.20

c) c c) Post-stall angle of attack, a=16.4° ...

Figure 18 - Concluded.

C z/ .' ~.?

C'

~~~~~~~~1~~~~~~~~~~~~:P~t~:=~~1~.=O~~~;~---

=:: a ------..--- .. ___ w ... __ ~ _______ _

r

=====:====-----~-------~~

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

~~~~~~~~~- --' .6 _" '.7

-.05

~

___ -.:::.-=--.&li.-- ---

.... ___ .__ .9 - .... -----_ .

........... ~................. 1.0

.... _------

, 1.'20 1.'40

l~O 0 1.10

1.50 Figur~:19i - Wake Total Pressure Contours.

-.05 ~ , 1.0 1.0 .................... --.

~--.,.,.. ---- ----

I 1. 40 1. 20 1,30 1,50 I~ OO~ 1. PO b) Stall angle of attack, a=14.4° -1. ___ .

Figure ::1~1 - Continued o.

--------- - - - - --

~------

-.2 -= -- -- - -.7 -.05 --------- -.8

~I

-- -- -~9 1.0

--

I I I I I 1.40 1.50 1.30 1.20 1.10 1.00 - --~---- ..... _-- .---.. . -- c) Post-stall angle of attack, a=16.4Q Smooth Moderate Turbulence (!O%<T<50%) Heavy Turbulence (T>SO%l o 01 1 -I I I ~ I Heavy Turbulence - Reversed

mi l Heavy Turbulence

~ ; Moderate Turbulence

0 ' Smooth

RN :: 2. 2 x 10 6

Mach No. = 0.130

a) Pre-s~all angle of attack, a=12.4Q Figure 20 - Hot Film.Velocity Field Survey.

~ , Heavy Turbulence - Reversed

mI I Heavy Turbulence

~ Moderate Turbulence

o Smooth

Ul N

RN = 2.2 x 10

Mach No. = 0.130

b) Stall angle of attack, a=14.4° Figure 26 - Continued.

~ Heavy Turbulence - Reversed

un Heavy Turbulence

Em Moderate Turbulence

o Smooth

U1 W

RN = 2~2 x 10

Mach No. = 0.130 c) Post-stall angle of attack, a=16.4Q Figure : 20 - Concluded.

.014 I (REF. 8 )1 Experiment Theory I I , ----- Upper Surface 0) , ,

cd

Lower Surface .012 '0 \ , I' Separation predicted \ No Separation predicted.

o

\ at .92c.Measured at .925c • .010 I • 010 \ \

o

\ \ \

'\

I \ \ J \ .008 .008 \ ., I· \ \ I \ \ I " \ J \ 0 \ I ',0' \ .006 \ .006 -

V " 0

, \ 0 , \ ':,'C- ~\ , , , f

o

,

'" 0 d

, , .004 .004

" ...... Q 0

, , \.

.........

'-

, , .002 .002

d

o '&., d

c5 &-'a

-0...

o .... _

O.O~----~-------r------~------~------ 0.0 ~-----T------~~----~----~------~ .80 0.0 .20 .40 .60 .60 .80 1.00 0.0 .20 .40 xI c a)i,a = .2° ..

---- --_.-_._--- - ----- -- - - ._- - -- .... -- ---- -- ---- - - Figure 21 - Co~parison of Experimental and Theoretical Skin Friction Distributions.

APPENDIX A Velocity Profiles as Measured by Various Instruments in Unsteady Reversed Flow Fields In the course of the present investigation, various in- strumentation has been developed and used to sense velocity profiles. The flow fields over the airfoil and in the wake are steady for certain angles of attack. For other larger angles of attack, regions of reversed flow accompanied by un- steady flow can occur over the aft portions of the airfoil and in the wake.

Initially a five-tube probe was used for measuring the velocity profiles. However this probe did not accurately sense the regions of reversed flow. Therefore, a four-tube probe was developed and used in the regions where reversed flow existed and regions near the surface of the body. Later the hot film became available and was used primarily to sense the unsteady portions of the flow field and to get the maximum excursions of the velocities. The hot film was also used to sense the regions of flow reversal. As explained in the text, if at any time the flow velocity at some point became zero the flow was deemed to have reversed. Near the end of the pre- sent research a split film anemometer was obtained. This allowed sensing of the velocity both in a positive and negative direction and therefore provided a clearer measurement of the reversed flow regions.

As mentioned in the text any pressure probe device such as the four-or five-tube probe cann~t respond to rapid fluctua- tions of the flow field. These probes therefore measure some integrated average depending on the tube size, oscillation fre- quency, etc. Thus in the regions of unsteady flow some average velocity is measured by the four-tube or the five-tube probe, Al whereas, with the split film actual velocity fluctuations are obtained. If a sufficient sample size can be recorded and processed, a true mean velocity can be determined.

Figure Al shows a comparison of velocity measurements as obtained by the four-tube, five-tube and split film probes at various stations near the aft end of the airfoil. These data were obtained for the 2412 airfoil at an angle of attack of 16.4 degrees. For the split film data, only minimum and maxi- mum values were recorded. These data are shown as the shaded area indicating the maximum and minimum velocities. The dashed line shows the average of the maximum and the minimum and is not the true mean of the time varying velocities. One can see from this figure that the four-tube and five-tube probe measure- ments show an average velocity soemwhere between the maximum and minimum time fluctuating velocities. Thus, one must exer- cise caution in interpreting the data obtained with the various probes. The present comparison does reflect reasonably good agreement as to the flow reversal points measured by split film and pressure probes.

From this series of tests the advantages of using the split film are evident. High rate digital data acquisition and processing techniques are presently being developed to obtain statistical data associated with the velocity fluctuations such as the maximum and minimum velocities, the mean velocity, turbulence levels, standard deviations, etc. These techniques will be applied as standard testing methods at WSU as they are developed and incorporated in the data reduction computer pro- grams.

A2 .20 - .20 - --, " " " .:: : " " " ", : :;::: : Five Tube "" " " " .10 ...

.10: "" " "

:"::: ii: i:" :: :i: 'j

. : :: ~ :. ~ : j .

. . :: ' : . . .

" " Split Film h Four Tu ..L ::j ~ .. :: :::: ...

~ I

-1.0 0 1.0 -1.0 o 2.0

u u x x xI = .80 xI = .90 c c .20 Low Split Film High Split Film : ~ . :: : :: :: Average Between High and Low Split Film

-1.0 o 1.0 2.0

u x xI = 1.00 c Figure Al - Velocity Profiles From Various Instruments in Unsteady Reverse Flow.

A3

Appendix B - Probe Interference Effects on Static Pressure Measurements During the course of the measurements of the flow proper- ties over airfoils it was found that for certain cases the probe and probe mechanism interfered with the static pressure distribution on the airfoil. This Appendix describes the modi- fications and calibrations conducted to remedy the problem.

Since interference of this type had not been encountered on tests of the 17% GA(W)-l airfoil (Ref. Bl), special runs were made with the moderate thickness NACA 2412 and GA(W)-2 airfoil sections to identify the source and magnitude of the interference. These runs showed that the interference was significant only at post-C£, angle of attack (see figures max Bl through B3).

In order to reduce probe mount interference, a new test section ceiling was designed and installed. The new ceiling had a longitudinal slot and structural provisions for mount- ing the probe track and carriage above the test section. The ceiling slot opening was fitted with foam seals to prevent leakage, and a new airfoil-shaped probe strut was designed and fabricated to replace the circular strut used in earlier tests.

Figure B4 shows a sketch of the probe and probe mechanism in the wind tunnel before and after the modifications.

Figure B5 shows that the static pressure distribution after modification with strut installed is essentially the same as without strut. Tests made with the strut plus probe show that the probe influences the surface pressure distribu- tion somewhat, but does not have a radical influence on appar- ent separation point location. Thus most of the interference encountered with the original set-up had been eliminated.

Based upon these calibrations the installation was judged to be acceptable and detailed flow studies at the post-Cn were ~max conducted with this instrumentation configuration.

Bl Also shown on Figure B5 are the surface static pressure coefficient distributions as obtained from extrapolation of the static pressure profiles of Figure 11. Figure~ B6 and B7 are plots of the surface static pressure plots as extrapolated for the pre-stall and stall angles of attack and the static pressure as measured by the surface static pressure ports, without probe.

These give an indication of the interference errors that exist in the data. For the pre-stall and near-stall angles of attack the discrepancies are quite small. For the post-stall case, the data indicate that the separation point is about O.12c further downstream with the probe installed. While this shift in sep- aration location is not unacceptably large, it is responsible for fairly large C charges near the separation point. Pres- p sures aft of separation and near the leading edge show neglig- ble changes.

--Reference Bl. Seetharam, H.C. and Wentz, W.H., Jr.: Experi- mental Studies of Flow Separation and Stalling on a Two-Dimensional Airfoil at Low Speeds.

NACA CR-2560, July 1975.

B2 -7.00 -6.00 o Surface Pressure without probe.

o Surface Pressure with probe at 1.Oc.

-5.00

R. N. = 2.2 X 10

Ha ch No. = O. 13

-4.00 c p -3.00 -2.00 -1. 00 .20 .40 .60 .80 1. 00 Figure Bl - Calibration for Probe Interference, NACA 2412, 0=12.4°.

B3 -8.00 -7.00 o Surface Pressure without probe.

o Surface Pressure with probe at 1.Oc.

-6.00

R. N. = 2.2 X 10

Mach No. = 0.13 -5.00 c P -4.00 -3.00 -2.00 -1. 00

o L-------T--- ____ ,- ______ ~------~------~

o

.20 .40 .60 .80 1.00 Figure B2 - Calibration for Probe Interference, NACA 2412, a=14.4°.

-7.00 o Surface Pressure without probe.

8 Surface Pressure with probe at 1.Oc.

-6.00

R. N. = 2.2 X 10

Mach No. = 0.13 -5.00

o

-4.00

o

c p -3.00

o

El

o

-2.00

o El

[J

o

o

-1. 00

o

o

000000§080000~

[J El [!J O~-------r---------.---- ___ ~ ________ ~ ______ __ .20 .40 .60 .80 1. 00

o

Figure B3 - Calibration for Probe Interference, NACA 2412, a=16.4°.

B5 Probe Travel Track

'---

Circular Strut Probe Hodel ~ Tunnel Floor

""""""""'" ""'" ""'"

a) Before Modification ~~----probe Travel Track ~------r---~------------------~------~ Tunnel Ceiling Airfoil Strut ---.~, , .... ---- Probe

~-------Model

Tunnel Floor

""',""'" '''~ ""'" '" "

b) After Modification Figure B4 - Probe mount and tunnel modifications.

B6 ;---;----r .- - .. --------;------:------.--------- . . , ..'

, , i Unmodified Tunnel Ceiling i i OWing & Strut & Probe I ; ,

L-7--I.---.--.- -.--.+

j Modified Tunnel Ceiling 1._':' I · : r ,. i ' Wing Alone

b?~; -, ~--t-.- ---;-----

[ .... I I .! [) Wing Strut &

l~1.--J.---i-·--· i ____ .. _L __ ._

& Probe 8 Wing & Strut

.. ~. I ' I ,--. ' ___ ~ __

Data Extrapolated Static Pressure

X

, I , , .-. t 'j" 1 - , , , , ---;----;---·T, .. -----.- ..

; i _.; .. . ·1--: .. -; -+-~-'--I'-----i:---'----"-!-----~' -_, -.--+--- "---- .,j..-- - • 1.-- .'---t--~_+--...J----'--_'-...:.-4.-. -- -- -- .- .. - .. , '1 . :: I --·1· - I --._"-- ... ~-------!---;- X I .: !

.j -.-~-----'-- .. - .. --,-.-------- IC. i. I .j , . " ..... __ .... ---+----. - ... ---- .-- --- --- ..... ------ ..... - , . I -------1'"---- - __ - ---- .. .

. 1 Figure B5 - Surface Static Pressure Interference, NACA 2412, a=16.4°.

B7 -7.00 o Surface Pressure without probe.

X Surface Pressure from probe data

extrapolated to the surface.

-4.0 c p

~

o

-2.0

o

~

o~

0~ -1.0

OO~

'O~

O~0V

OO~ o~------~------~------~------~--~~~ .20 .40 .60 .80 1.00

o

Figure B6 - Surface Static Pressure Interference, NACA 2412, a=12.4°.

B8 -8.00 -7.00 -6.00 o Surface Pressure without probe.

-5.00 )( Surface Pressure from probe data extrapolated to the surface.

c p -4.00 -3.00

~

-2.00 -1.00 OL-~--~------~-------r------,- ____ --,

o

.20 .40 .60 .80 1. 00 Figure B7 - Surface Static Pressure Interference, NACA 2412, n=14.40.

B9 :.J " -' . L ; 1IIIIIIIIIIIIIIIII11Ii1Illlilijlllill~llnli~lllllllllllllll· .

3 1176014137195 ---~- __ -----.A __ _ \ \ \ \

•