Effects of forward contour modification on the aerodynamic characteristics of the NACA 641-212 airfoil section

• ......... : :-; : ..... - ...... =::: ::- _: ....... \: :::;:_!2'-_7_ "_-_ . _ - _ .... 2_ _ ........... __ _ 1. Report No. 2. Government Accession No.

3. Recipient's Catalog No.

NASA TM X-529_ 4. Title and Subtitle 5. Report Date September 1975 EFFECTS OF FORWARD CONTOUR MODIFICATION ON THE AERO- 6. Performing Organization Code DYNAMIC CHARACTERISTICS OF THE NACA 641-212 AIRFOIL SECTION 7. Author(s) 8. Performing Organization Report No.

A-6018 Raymond M. Hicks, Joel P. Mendoza, and Angelo Bandettini 10. Work Unit No.

9. Performing Organization Name and Address 505-10-12 Ames Research Center 11. Contract or Grant No.

Moffett Field, California 94035 13. Type of Report and Period Covered 12. Sponsoring Agency Name and Address Technical Memorandum National Aeronautics and Space Administration 14. Sponsoring Agency Code Washington, D.C. 20546 15. Supplementary Notes 16. Abstract Two different forward contour modifications designed to increase the maximum lift coefficient of the NACA 64x-212 airfoil section were evaluated experimentally at low speeds. One modification consisted of a slight droop of the leading edge with an increased leading-edge radius; tile other modification incorporated increased thickness over the forward 35 percent of tile upper surface of tile profile. Both modified airfoil sections were found to provide substantially higher maximum lift coefficients than the 641-212 section. The drooped leading-edge modification incurred a drag penalty of approximately 10 percent at low and moderate lift coefficients and exhibited a greater nosedown pitching moment than the 641-212 profile. The upper-surface modification produced about the same drag level as the 641-212 section at low and moderate lift coefficients and less nosedown pitching moment than the 641-212 profile. Both modified airfoil sections had lower drag coefficients than the 641-212 section at high lift coefficients.

"'17. Key Words (Suggested by Author(s)) 18. Distribution Statement Airfoil Optimization Unclassified - Unlimited Wing STAR Category -- 02 19. Security Classif. (of this report) 20. Security Classif. (of this page) 21. No. of Pages ] 22. Price* Unclassified Unclassified 78 I $4.75 "For sale by the National Technical Information Service, Springfield, Virginia 9,9.16 1 NOMENCLATURE C airfoil chord, cm (in.)

section drag coefficient c d section lift coefficient c l section pitching-moment coefficient referenced to quarter chord C?n pressure coefficient, PL - P,_

%

qoo h tunnel height, m (ft) k roughness diameter, cm (in.)

P static pressure, N/m z (lb/ft _) q dynamic pressure, N/m s (lb/ft 2 ) Re Renolds number based on free-stream conditions and airfoil chord .V airfoil abscissa, cm (in.)

airfoil ordinate, cm (in.)

O_ angle of attack, deg Subscripts 771(./.'¢ maximum L local free-stream conditions iii

EFFECTSOF FORWARD CONTOUR MODIFICATION ON THE AERODYNAMIC

CHARACTERISTICS OF THE NACA 641-212 AIRFOIL SECTION Raymond M. Hicks, Joel P. Mendoza, and Angelo Bandettini Ames Research Center SUMMARY Two different forward contour modifications designed to increase the maximum lift coeffi- cient of the NACA 64t-212 airfoil section were tested at Mach numbers of 0.2, 0.3, and 0.4 and Reynolds numbers of 1 million, 1.5 million, and 1.9 million. The unmodified 641-212 profile was also tested for comparison with the modified sections. One modification consisted of a slight leading-edge droop along with an increased leading-edge radius; the other modification incorporated increased thickness over the forward 35 percent of the upper surface of the airfoil profile.

Lift and pitching moment were determined by integrating surface presstlre measurements and the profile drag was obtained from wake pressures. The models were tested at all the Mach numbers and Reynolds numbers with a narrow strip of roughness located at 12 percent of the chord and without roughness at M = 0.2 and Re = 1.9 million.

Both modified profiles were found to provide substantially higher nlaximum lift coefficients than the NACA 641-212 section. The drooped leading edge incurred a drag penalty of approximately 10 percent at low and moderate lift coefficients and exhibited a greater nosedown pitching moment than the 641-212 profile. Relative to the 641-212 section, the upper-surface modification produced nearly the same drag level at low and moderate lift coefficients and less nosedown pitching moment.

Both modified profiles had lower drag coefficients than the 641-212 section at lift coefficients above 0.8.

INTRODUCTION Airfoils that produce high maximum lift coefficients are desirable for airplanes designed for short takeoff and landing (STOL) capability and for maximum turning performance during maneu- vers. On the other hand, airfoils that produce low drag coefficients are needed for airplanes designed for high-speed applications. After World War II, many general aviation airplanes were designed with emphasis on the high-speed requirements rather than the slow-flight or maneuvering capabilities.

This led many designers to use the NACA laminar-flow sections in the hope of achieving the low drag coefficients exhibited by these airfoil sections during wind-tunnel testing when operated near the design lift coefficient. Such sections rarely achieve the low drag in flight because of manufacturing roughness or poor care of the wing surfaces during service. Furthermore, the lamina>flow sections exhibit a greater decrease in maximum lift coefficient with decreasing Reynolds number below 3 million than do most of the NACA 4- and 5-digit sections. Because of the relatively poor maxi- mum lift characteristics of the 6-series airfoils and the realization on the part of many designers that large amounts of laminar flow are not generally achieved in practice, several efforts have been made to modify the contour of the NACA 6-series sections to achieve greater maximum lift coefficients.

The most widely usedcontourmodificationfor increasing the maximunllift coefficienthas

beena droopof the leadingedgewith or withoutanincreased leading-edge radius(e.g.,refs.1 4).

A relativelylittle usedcontourmodificationfor increasing nlaxinmmlift is to increase the thickness

of the forward sectionof the uppersurfaceof the airfoil. Sucha modificationis suggested in

reference 5.Themainadvantage of theforwardupper-surface modificationis that the maximumlift is increased without incurringthe dragpenaltygenerallyfound with the droopedleadingedge.

Furthermore,the upper-surface modificationproduceslessnosedown pitching moment than a

droopedleadingedge.Because of theseadvantages andbecause of the needto further investigate

upper-surface modifications,a study wasundertaken to compare the two typesof forwardcontour modifications.

Asnotedearlier,the 6-series profileshaverelativelypoor maximumlift characteristics at low

Reynoldsnumbers;hencea 6-series sectionwasdeemedmostappropriatefor this study. Since

NACA63-, 64-,and65-series profilesarebeingusedon manygeneral aviationairplanes, a 64-series sectionwaschosen asrepresentative of this class of airfoils.Thickness ratiosranging from 6 percent to 18percentarein useon general aviationairplanes, hence, a 12percent thick sectionwasselected for this investigation. The conceptdemonstrated hereshouldbe applicable to NACA6-series pro- fileswith thickness ratiosbetween 6 and 18percent.

DESIGNOF AIRFOIL SECTION

The designof profilesfor highmaximum lift coefficientis usuallycarriedout by consideration

of the surface pressure distributionsinceno reliablemethodsareavailable for directcalculation of cI . The procedure used here was to reshape the forward region of the airfoil so that the peak rtl(,1x pressure coefficient and adverse pressure gradient near the leading edge on the upper surface are reduced without significantly changing the basic camber of the 64_-212 profile, thereby retaining the original design lift coefficient. The modified airfoil sections are shown in figure 1 along with the NACA 64_-212 profile. The theoretical pressure distributions tor 3/ -- 0.1 and Re = 1.0× 106 ure presented in figures 2(a) and 2(b). Note the decrease in pressure peak and improved pressure gradi- ent exhibited by both Modification (Mod.) A and Mod. B profiles for c_ -- 6° when compared with the pressure distribution of the 641-212 section (fig. 2(a)). These modified pressure distributions were achieved by repeated iterations utilizing drafting tools and a high-speed computer. Each iteration consisted of drawing a different forward contour for the 641-212 profile and then calcu- lating the pressure distribution and aerodynamic force coefficients for the modified airfoil section by the theory of reference 6 (a CDC 7600 computer was used). Seven iterations were required to develop Mod. A profile and 6 iterations to develop Mod. B. Note that the final pressure distribution for Mod. B at u = 6° exhibits a greater peak pressure coefficient than Mod. A. The test results presented later show a slightly lower maximum lift coefficient for Mod. B than for Mod. A, which is consistent with the relative values of peak pressure coefficients shown here. It might have been possible to further reduce the pressure peak for Mod. B if more iterations had been attempted.

However, because of the time involved in producing the large-scale drawings and generating the "inputs" for the computer for each iteration, the process was terminated after six iterations. In the future, such contour modifications will be carried out very rapidly by the numerical optimization technique described in reference 7.

The maindisadvantage in droopingtheleading edgeto increase the maximumlift coefficientis

illustratedin figure2(b), whichshowstheoreticalpressure distributionsfor _x = 0°. Note the pres-

sure "spike" near the leadingedgeof the lower surfaceof Mod.A. The steepadverse pressure

gradientfollowing the "spike" will cause the boundarylayerto thickenwith anattendantincrease in drag.If theleading edge is droopedenough to producereflexed curvature behindthe leading edge on the lower surface, the pressure "spike" will beaccentuated andthe dragincrease will begreater.

The experimental resultspresented latershowa higherdraglevelfor Mod.A thanfor eitherMod.B

or the 64_-212profile at low lift coefficients,which tendsto supportthe predictedeffect of the lower surfaceadverse gradient.Notethat the lower surfacepressure distributionof Mod.B isvery similarto that of the 641-212 airfoil sectionwhereas the upper-surface pressure distributionshows a

modest"hump" nearthe 10-percent chordstationfollowedby a relatively"flat" adverse pressure

gradient. A gradientof this magnitudeshouldhaveonly a minor effect on the boundary-layer

development andhence little effecton pressure drag.

Thecoordinates for theunmodifiedNACA 641-212sectionand the two modifiedprofilesare

givenin tables1 to 3.

APPARATUS AND TEST PROCEDURE

Models Three airfoil models with the NACA 64_-212, Mod. A, and Mod. B profiles were machined from aluminum billets. Each model had a nominal chord of 15.24 cm (6 in.) and a span of 60.96 cm (24 in.). The models were equipped with 21 upper-surface orifices and 22 lower-surface orifices drilled normal to the surface, which, together with the necessary pressure leads, made it possible to determine the pressure distributions on the model surfaces. The pressure leads from each orifice were set in milled slots in the models, resulting in minute wavyness of tile upper and lower surfaces.

it was felt that such an imperfection was fairly realistic since most manufacturing processes produce some wavyncss of aircraft skins.

Wind Tunnel The tests were conducted in the Ames 2- by 2-Foot Transonic Wind Tunnel, a variable-speed, continuous flow, ventilated wall, variable pressure facility. The tunnel can be used for two- dimensional testing by replacing the ventilated side walls with solid walls where model-supporting thick glass windows arc mounted. The windows can be rotated by a motorized drive system to change the angle of attack. An 82-tube drag rake located 1.75 chords behind the model trailing edge is used to survey the model wake. Figures 3(a) and 3(b) show the Mod. B airfoil installed in tile tunnel along with tile drag rake. Airfoil models are mounted spanning the horizontal dimension of the tunnel test section so that the center of rotation of the side windows is near the 25-percent chord station on the model. The gaps between the ends of the model and the side windows were sealed with silicone rubber adhesive sealant.

Instrumentation

Measurements of the model surfacepressures and the wakerake pressures weremadeby

automaticpressure-scanning systemthat utilizesprecisionpressure transducers. Basictunnelpres-

suresweremeasured with precisionmercurymanometers. Angleof attack wasmeasured with a

potentiometeroperatedby the drivegearfor the rotatingsidewindows.Datawereobtainedby a

high-speed, data-acquisition system andrecorded on papertape.

Tests

The sectionaerodynamic characteristics of the threeairfoilswereobtained at M = 0.2 and 0.3

at Re = 1.0XI06 , 1.5X106 , and 1.9X106 and arm = 0.4 at Re = 1.0Xl06 , 1.9X106 , and 3.0×106 .

(The Reynolds numbers are based on the model chord.) The angles of attack ranged from approxi- mately -3 ° to 18 °, depending on the stalling angle of each model. The models were not tested without the wake rake installed since previous investigations in the 2- by 2-Foot Wind Tunnel have shown that the effect of the wake rake on the model surface pressures is negligible for the rake position used in the present tests. Data were obtained at all test conditions with a 0.159cm (0.0625 in.) wide strip of 0.0064 cm (0.0025 in.) (nominal) diameter glass balls located at the 12-percent chord station in an effort to simulate manufacturing roughness. Data were also taken at M = 0.2 and Re = 1.9X10 6 without roughness.

Pressure coefficients were determined from surface pressure measurements. Section normal force coefficients, chord force coefficients, and pitching-moment coefficients were obtained from an integration of the pressure coefficients. The pitching-moment coefficients were referenced to the quarter-chord point. Section profile drag was calculated from the wake-rake total and static-pressure measurements.

The model angle of attack was corrected for the presence of the tunnel walls by the following equation: Aoe = 6(c/h)c l where As, 6, c/h, and c l are the angle-of-attack correction, correction factor, model chord/tunnel height ratio, and section lift coefficient respectively. The angle-of-attack correction factor (6) is a function of Mach number. The following values were used and the corresponding As, con- verted to units of degrees, was added algebraically to the model geometric angle of attack: M 6 0.2 -0.095 .3 -.150 .4 -.186 (These correction factors (8) were determined during a tunnel calibration conducted by Mr. L. S.

Stivers, Jr.) The Math number corrections due to the presence of the tunnel walls were negligible for the Mach numbers of the present investigation.

RESULTSAND DISCUSSION

Aerodynamic Characteristics

L_/t. - The basic force coefficients for tile three airfoils tested with roughness are presented in figure 4(a) through 4(i). Both Mod. A and B airfoils gave substantially greater maximum lift than the NACA 641-212 section. At Re = 1.0X106 and M = 0.2, both modifications had the same maximum lift coefficient but the stall of Mod. B was somewhat more abrupt than the stall of Mod. A or the 641-212 airfoil (fig. 4(a)). At Re = 1.5XI06 and 1.9×106 , Mod. A has tile highest maximum lift coefficient (figs. 4(b) and (c)). Both modified sections stalled more abruptly than the 641-212. At all test Reynolds numbers at M = 0.3 and 0.4, the two modified airfoils produced nearly equal maximum lift coefficients and showed similar stall characteristics (figs. 4(d)-(i)), except at M = 0.3 and Re = 1.0XI0 6 where Mod. A showed a more gradual stall (fig. 4(d)) than both Mod. B and the 64_-212 airfoil. The two types of forward contour modification considered during this study had little effect on the basic camber distribution of the 641-212 airfoil as evidenced by the nearly constant values of c l at o_ = 0 ° for the three airfoils at all test conditions. As discussed previously, a constant value of design lift coefficient was one of the criteria used to develop these contour modifications. This is important if such modifications are considered for retrofit of existing aircraft.

Summary plots of Clmax versus Reynolds number for the three test Mach numbers are pre- sented in figures 5(a) through 5(c). These figures clearly show the small difference in Clmax for the two contour modifications studied. Further work will be done to optimize the upper-surface modi- fication (Mod. B) in an effort to fllrther improve Clmax since this type of modification does not incur the drag penalty found with Mod. A (see the following discussion). The values of Clmax shown here may be lower than that achieved in actual use on general aviation airplanes since the landing Mach number of most light planes is 0.1 or less and previous NACA data have shown that Clmax can decrease substantially as the Mach number is increased from 0.1 to 0.2 (ref. 8).

Drag. - The profile drag data in figure 4 generally show that the drag level of the Mod. B airfoil is about the same as that of the NACA 641-212 section at low lift coefficients with Mod. A exhibiting higher drag for these conditions. An exception to this result is found at M = 0.2 and Re = 1.0X106 where Mod. B shows slightly more drag than the 641-212 airfoil (fig. 4(a)). However, such a combination of Mach number and Reynolds number is not representative of the cruise condition for most general aviation airplanes and hence is of interest only academically. At all test conditions, both modified airfoils showed lower drag at moderate and high lift coefficients than the 641-212 airfoil (figs. 4(a) (i)). Note that Mod. B had lower drag than Mod. A at high lift coeffi- cients at all test conditions.

The fact that both modified airfoils extend the low drag range of the 641-212 airfoil to higher lift coefficients should be of particular interest to general aviation manufacturers since this means lower drag during climb and hence better climb performance, which is important from a safety standpoint. For example, the data in figure 4(c) show that the section lift/drag ratio of Mod. B varies from 80 at cI = 0.9 to 86 at c l = 1.10, which compares with lift/drag ratios of 67 to 70 for the 64_-212 airfoil over the same lift coefficient range.

PitchiHg mome_lt.--. The pitching moment data in figure 4 show that Mod. A had a greater nosedown pitching moment than the NACA 641-212 airfoil, whereas Mod. B produced less nose- down pitching moment than either of the other two airfoils at all test conditions. These results show another advantage of using an upper-surface contour modification to increase Clmax instead of a drooped leading edge since less nosedown pitching moment means less trim drag. This is an important consideration in choosing a retrofit modification for an existing airplane.

Effect of Surface Roughness The data in figure 6 show the effect of rougtlness on the section characteristics of the three airfoils at M = 0.2 and Re = 1.9×106 . A strip of glass balls, 0.159 cm (0.0625 in.)wide, with a nominal diameter of 0.0064 cm (0.0025 in.) located at the 12-percent chord station was used to achieve a mostly turbulent boundary layer and thereby simulate the surface condition found on most general aviation airplanes in normal service. The major effect of surface roughness was to increase the section profile drag coefficient of all models. The drag increase for the 64 t-212 airfoil was greater than for the other two sections since the 6-series airfoils were designed to achieve large amounts of laminar flow for a range of lift coefficients above and below the design point. Since extensive laminar flow is rarely attained in practice, the data obtained with roughness are more realistic.

Comparison of Experiment with Theory A comparison of experimental and theoretical aerodynamic force data is presented in figure 7 (the viscous theoretical program used here is described in ref. 6). Data are shown for the NACA 64_-212 section at M = 0.2, 0.3, and 0.4 at Re = 1.5X106 and for Mods. A and B atM = 0.2 and Re = 1.5X106. In general, the agreement between experiment and theory is acceptable for the lift and pitching-moment data for angles of attack where the boundary layer is attached, whereas the theory consistently overestimates the profile drag for all three models. The inaccurate drag predic- tions are due in part to the technique used to determine the pressure drag, the integration of the theoretical surface pressures. This procedure is always difficult to use because the pressures are poorly defined near the leading edge of the profile. A new method is currently under development which will use the momentum defect in the wake to calculate profile drag, thereby eliminating the most serious inaccuracy in the theory. Note that the theory predicts the lift and pitching-moment characteristics of the 64,-212 section slightly better than for the two modified airfoils. It is interesting that the agreement between experiment and theory is as good at M = 0.4 as at M = 0.2 or 0.3 since the theory was not intended to be applicable above M = 0.3.

Figure 8 compares experimental and theoretical pressure distributions. Data are presented for the 641-212 airfoil at M = 0.2, 0.3, and 0.4 at Re = 1.5X 106 and for Mods. A and B at M = 0.2 and Re= 1.SX106. Again the agreement between experiment and theory is acceptable at angles of attack where the boundary layer is not separated. As noted with the force data, the agreement between experiment and theory is somewhat better for the (_4_-212 section than for either modi- fied airfoil.

CONCLUSIONS Wind-tunnel tests were conducted to determine the section aerodynamic characteristics of two types of forward contour modifications designed to increase the maximum lift coefficient of the NACA 64_-212 airfoil section. The tznmodified 64 t-212 section was tested for comparison. The experimental data were compared with theoretical predictions. The tests were conducted at M = 0.2 and 0.3 at Re = 1.0XIO 6 , 1.5XIO 6 , and 1.9X]O _ and at /ll = 0.4 at Re = 1.OXIO 6 , 1.9XIO 6, and 3.0XI0 _'. The following results were established: 1. Increasing the upper-surface thickness over the forward 35 percent of the chord was nearly as effective as a drooped leading edge with an increased leading radius for increasing the maxinlum lift coefficient of the 64_-212 section. Both modifications produced about 30 percent more maxi- mum lift than the 64j-212 profile.

2. The forward upper-surface modification did not incur the drag penalty of the drooped leading-edge modification at lift coefficients in the cruise range of most general aviation airplanes.

The drag of the airfoil with upper-surface modification was equal to that of the 64_-212 section at cruise conditions when both airfoils were tested with a mostly turbulent boundary layer.

3. The forward upper-surface modification produced nearly a 25-percent reduction in nose- down pitching moment compared with the unmodified 64_-212 airfoil, whereas the drooped leading-edge modification showed approximately 30 percent more nosedown pitching moment than the 64_-212.

4. Both types of forward contour modification produced less drag at moderate and high lift coefficients than the 641 -212 airfoil, the drag of the profile having the upper-surface modification being less than that of the airfoil with a drooped leading edge at high lift coefficients. The lift/drag ratio of the airfoil with upper-surface modification was about 23 percent greater than that of the 641-212 airfoil at M = 0.2 and Re = 1.gxI06 ate/= 1.10.

5. A comparison of experimental values of lift coefficient, pitching-moment coefficient, and pressure distributions with those calculated by a viscous-flow theory was acceptable at all test Mach numbers for angles of attack where the boundary layer remained attached. However, the drag prediction was poor for all test conditions. A new viscous theory tinder development should improve drag estimates considerably.

Further research will be done, aided by a numerical optimization program, to develop opti- mum forward upper-surface modifications since such contour modifications appear most promising for attaining high maximum lift coefficients, low cruise drag, and low pitching-moment coefficients.

Ames Research Center National Aeronautics and Space Administration Moffett Field, Calif., 94035, April 21, 1975

REFERENCES

1. Kelly,John A.: Effects of Modifications to the Leading-Edge Region on lhe Stalling Characteristics of the NACA 63_-012 Airfoil Section. NACA TN 2228, 1950.

2. Anderson, Seth B.; Matteson, Fredrick H.; and Van Dyke, Rudolph D., Jr.: A Flight Investigation of the Effect of Leading-Edge Camber on the Aerodynamic Characteristics of a Swept-Wing Airplane. NACA RM A52L16a, 1953.

3. Demele, Fred A.; and Sutton, Fred B.: The Effects of Increasing the Leading-Edge Radius and Adding Forward Camber on the Aerodynamic Characteristics of a Wing with 35 ° of Sweepback. NACA RM A50K28a, 1951.

4. Goradia, Suresh H.: and Lyman Victor: Laminar Stall Prediction and Estimation of C . Journal of ' L(max ) Aircraft, vol. 1 I, no. 9, Sept. 1974, pp. 528 -536.

5. Wortman, F. X.: Design of Airfoils with ttigh Lift at Low and Medium Subsonic Mach Numbers. Advisory Group for Aerospace Research and Development. Fluid Dynamics of Aircraft Stalling. AGARD CP 102, 1972.

6. Stevens, W. A.: Goradia, S. H.: and Braden, J. A.: Mathematical Model for Two-Dimensional Multi-Component Airfoils in Viscous Flow. NASA CR-1843, 1971.

7. lticks, Raymond M.: and Vanderplaats, Garret N.: Design of Low-Speed Airfoils by Numerical Optimization.

Paper 750524 presented at SAE Business Aircraft Meeting, Wichita, Kansas, April 1975.

8. Racisz, Stanley F.: Effects of Independent Variation of Mach Number and Reynolds Number on the Maximum Lift Coefficients of Four NACA 6-Series Airfoil Sections. NACA TN 2824, 1952.

TABLE 1.-NACA 641-212 AIRFOIL COORDINATES Upper surface Lower surface x/c y/c x/c y/c 0.00000 0.00000 0.00000 0.00000 .00418 .01025 .00582 -.00925 .00659 .01245 .00841 -.01105 .01147 .01593 .01353 -.01379 .02382 .02218 .02618 -.01846 .04868 .03123 .05132 -.02491 .07364 .03815 .07636 -.02967 .09865 .04386 .10135 -.03352 .14872 .05291 .15128 -.03945 .19886 .05968 .20114 -.04376 .24903 .06470 .25097 -.04680 .29921 .06815 .30079 -.04871 .34941 .07008 .35059 -.04948 .39961 .07052 .40039 -.04910 .44982 .06893 .45018 -.04703 .50000 .06583 .50000 -.04377 .55016 .06151 .54984 -.03961 .60029 .05619 .59971 -.03477 .65039 .05004 .64961 -.02944 .70045 .04322 .69955 -.02378 .75047 .03590 .74953 -.01800 .80045 .02825 .79955 -.01233 .85038 .02054 .84962 -.00708 .90027 .01303 .89973 -.00269 .95013 .00604 .94987 .00028 1.00000 .00000 1.00000 .00000

TABLE 2.- MOD. A AIRFOIL COORDINATES

Upper surface Lower surface

x/c y/c x/c y/c -0.01850 -0.02000 -0.01850 -0.02000 -.01700 -.O0600 -.01700 -.02680 -.01400 .00100 -.01400 -.03110 -.01000 .00650 -.01000 -.03450 -.00500 .01180 -.00500 -.03700 .00000 .01560 .00500 -.03900 .01000 .02200 .01500 -.03980 .02000 .02810 .04000 -.04060 .03000 .03250 .08000 -.04180 .04320 .03940 .12000 -.04300 .07364 .04650 .16000 I -.04410 .09865 .05200 .20000 I -.04540 .14872 .05880 .24000 t -.04680 .19886 .06310 .30079 .04871 .24903 .06640 .35059 1 -.04948 b .29921 .06900 •40039 i -.04910 .34941 .07008 .45018 I -.04703 .39961 .07052 .50000 .04377 .44982 .06893 •54984 -.03961 .50000 .06583 •59971 -.03477 .55016 .06151 •64961 -.02944 .60029 .05619 •69955 -.02378 .65039 .05004 .74953 .01800 .70045 .04322 •79955 -.01233 .75047 .03590 •84962 -.00708 .80045 .02825 .89973 -.00269 .85038 .02054 •94987 .00028 .90027 .01303 1 .00000 .00000 .95013 .006O4 .00000 .00000

TABLE 3.- MOD. B AIRFOIL COORDINATES

Upper surface Lower surface x/c y/c x/c y/c 0.00000 0.00000 0.00000 0.00000 .00418 .01640 .00582 -.00925 .00659 .01970 .00841 -.01105 .01147 .02500 .01353 -.01379 .02382 .03480 .02618 -.01846 .04868 .04820 .05132 -.02491 .07364 .05730 .07636 -.02967 .09865 .06310 .10135 -.03352 .14872 .06820 .15128 -.03945 .19886 .07000 .20114 -.04376 .24903 .07080 .25097 -.04680 .29921 .07100 .30079 -.04871 .34941 .07100 .35059 -.04948 .39961 .07052 .40039 -.04910 .44982 .06893 .45018 -.04703 .50000 .06583 .50000 -.04377 .55016 .06151 .54984 -.03961 .60029 .05619 .59971 -.03477 .65039 .05004 .64961 -.02944 .70045 .04322 .69955 -.02378 .75047 .03590 .74953 -.018OO .80045 .02825 .79955 -.01233 .85038 .02054 .84962 -.00708 .90027 .01303 .89973 -.00269 .95013 .00604 .94987 .00028 1.00000 .00000 1.00000 .00000 II

NACA 641 212

Mod. A

Mod. B

NACA 641-212

Figure 1.- Airfoil sections tested.

-3 I

MOD. A

Cp

-I

I

-2

MOD.B

Cp

-I

f I .

0 .2 .4 .6 .8 1.0

(a) c_=6 °,c l _0.75.

Theoretical pressure distributions for M = 0. l, Re = 1.0× 106 .

Figure 2.-

1 J 1 1

-2

MOD. A

-I

surface )er surface

Cp

_'_-NACA 641-212

-I

MOD.B

Cp

m

0 --

y

0 .2 .4 .6 .8 1.0

(x/c)

(b) o_=O °, Cl _0.15.

Figure 2.- Concluded.

J

._ o _ X © °_ C) © © I rf_ r_ °_

!

1.6

1.4

d

1.2

,g

1,0

xo

.6

o _'-NAoA 64-2,2

fl _ MOD. B

o /

0 4 8 12 16 2_0

co, Deg.

(a) M=O.2, Re= 1.OXIO 6 .

Figure 4.- Effect of airfoil contour modification on section characteristics, roughness at O. 12c.

1.4

1.2

t "1 I.O 1

.8

Cl

.6

.4

.2

0 .01 .02 0

-.12

-'04Cm_78

Cd

(a) M = 0.2, Re = 1.OX 10 6 - Concluded.

Figure 4.- Continued.

1.6

1.4

1.2

b

1.0

Ol

.8

.6

¢

.4

MOD. A

/

.2

I_,_ MOD.B

0 4 8 12 16 20

o::,Deg.

(b) M = 0.2, Re = 1.5X106 .

Figure 4.- Continued.

1.4

1.2--

1.0--

.8

Cl

.6

.2 I

.01 .02

- .12

-04 Cm_08

Cd

(b) M = 0.2, Re = 1.5XIO 6- Concluded.

Figure 4.- Continued.

1.6

1.4

/

1.0

.8

C_

.6

o_NACA 641-212

.4

0 _ MOD. A .2_.

/

16 20

0 4 8 12

(:z:, Deg.

(c) M = 0.2, Re = 1.9X]O 6.

Figure 4.- Continued.

1.6

\

1.2

1.0

I C_ I

.8

.6

I I I

.4

I

!

.2

I

-.12

0 .01 .02 0 -.04 Cm-'08c/4

Cd

(c) M = 0.2, Re = 1.9X106 -Concluded.

Figure 4.- Continued.

1,4"

1.2

I.O

.8

C_

.6

0 _ --I--'_ACA 641-212 __._.---

.4

0 _ MOD. A

.2

0 8 12 16 20

e:, Deg.

(d) M = 0.3, Re = 1.OX 10 6 .

Figure 4.- Continued.

1.4

1.2

_/_ ¢_r ''''_

t/ I t

1.0

.8

I I

J

I I C1, I

.6

I I

I i

I I

.4

I I

.2

I

"0._

-.12

0 .01 Cd .02 0 -.04Cm c_4.08

(d) M = 0.3, Re = I.OX 10 6 - Concluded.

Figure 4.- Continued.

Cl

0 4 8 12 16 20

_, Deg.

(e) M = 0.3, Re = 1.5X106 .

Figure 4.- Continued.

1.4

1.2

1.0

I

.8

i I I t I C_ ,) I

.6

I I I !

I I

.4

q

I I -J

.2

I

.01

.02 0

-.12

-.04Cmc_08

Cd

(e) M = 0.3, Re = 1.5XI06 - Concluded.

Figure 4.- Continued.

1.6

I 1.4.

/i",,o

1.2

/ / /

1.0

\

.8

I:;l

.6

o_-_XoA64,-2,_ /

.4

<) _ MOD. A _-

.2

l'_ (_ MOD. B _....---

o

¢

4. 8 12 16 20

G, Deg.

(f) M=0.3, Re= 1.9X10 6 .

Figure 4.- Continued.

1.6

1.4

¢ " 0

1.2

t I I I

1.0

I

t

|

;>

.8

I I I I I I C_

q

I

.6

I !

I I I I I I

.4

i I

q

I

.2

I

q

.01 .O2

0 -.04Cmc/_.08

Cd

(f) M = 0.3, Re = 1.9X 10 6 - Concluded.

Figure 4.- Continued.

1.4

1.2_

\

1.0

.8

C_

.6

o_---_cA 64,-2,2

.4

0 _ MOD. A

.2

E3_-_- MOD.B

8 12 16 20

0 4

ol::, Deg.

(g) M = 0.4, Re = ].0Xl0 6.

Figure 4.- Continued.

3O

1.6

1.4

1.2

V \ , <)!,

1.0

II l I

.8

I I C_ I

.6

t I I

.4

"7

_r; _?

I I

! '

.2

._ _

I

C

!

<_..,

a Q

--<)

• 02 0

-.04

-.08

Cmc/4

(g) M = 0.4, Re -- 1.OXlO 6 - Concluded.

Figure 4.- Continued.

1.4

1.2

/ /

1.0

'/

.8

,/

C_

.6

,/

0 _-_NACA 641-212 /

.4

0 _ MOD. A

.2

1:3_ MOD.B

0 4 8 12 16 20

_, Deg.

(h) M = 0.4, Re = 1.9Xl0 6.

Figure 4.- Continued.

1.6

1.4

.<

1.2

t N EX |

1.0 I

I I I

.8

I I C_

+

i !

!

,6

I

i

I I

.4

i I

.2

I i I I I %

t'

"0

_>

.01 .02

-.08 -.12

0 Cmc/4.04

Cd

(h) M = 0.4, Re = 1.9X 10 6 - Concluded.

Figure 4.- Continued.

1.6

1.4-

\

\

1.2

1.0

.8

C7,

.6

.4.

0 _ MOD. A

.2

4 8 12 16 20

oC, Deg.

(i) M = 0.4, Re = 3X10 6.

Figure 4.- Continued.

1.6

1.4

f

\ \

1.2

It

1.0

I I I

.8

I I ! I C_ I I I

.6

I !

I I )

?

.4

I I I I

.2

?

I I

?

I I

¢

¢

0 .01

.O2 -.08 -.12

0 Crnc/_.04

Cd

(i) M = 0.4, Re = 3.0X 10 6 -Concluded.

Figure 4.- Concluded.

_-------'_-ACA 641-212 MOD. A MOD. B

1.6

1.4

..____-----

1.0

o

0 1.0 1.2 1.4 1.6 1.8 2.0

RexlO -6

(a) M = 0.2.

Figure 5.- Effect of airfoil contour modification on maximum lift coefficient, roughness at 0.12c _-'NACA 641-21Z_ MOD. A MOD. B

1.6

i ......

F I

i__

1.4

.,.,.---e---- I C_ I max J

1.2

E ._____-----_ -

1.0

>- <

1.0

1.6

1.8

2.O

(b) M = 0.3.

Figure 5.- Continued.

_----_ACA 641-212 MOD. A (_ MOD.B

1.6

.......-,.

.....-.

1.4 F

I ,...- I .---,---.atom,- C_ mox

1.2

1.0

l

¢

0 1.0 1.4 1.8 2.2 2.6 5.O

Re x I0 -6

(c) M = 0.4.

Figure 5.- Concluded.

1.6 r I I

TRANSITION FIXED

K=.0064 cm (.0025 in)

i.4

SMOOTH MODEL

1.2

1.0

.8

C_

.6

t

.4

/

.2

_-----_ACA 641-212

0 4 8 12 16

_, Deg.

(a) NACA 641-212.

Figure 6.- Effect of roughness on section characteristics; M = 0.2, Re = 1.9X 10 6 .

TRANSITION FIXED

K=.0064 cm (.0025 in)

0 SMOOTH MODEL

1.4 --

1.2

f

1.0

.>

.8

C7, ¢.5

.6

/ / / Q

,p

!

.4

I 4,

.2

I I I \

.01 .02 0 -.04 -.08 -.12

Cd

Cmc/4 (a) NACA 641-212 - Concluded.

Figure 6.- Continued.

1.6

o TRANSITION FIXED,

K=.0064 cm (.0025 in) ,,_

0 SMOOTH MODEL

1.4

F

1.2

1.0

.8

C_

.6

/

.4

.2

MOD. A

4. 12 16 20

8o:, Deg.

(b) Mod. A.

Figure 6.- Continued.

o TRANSITION FIXED

K=.0064 cm (.0025 in)

1.6

0 SMOOTH MODEL

1,4 -- 1.2--

1.0

.8--

C_ I

.6

I

.4

I

.2--

I I m -3

.01 -.12 .02 0 -.04 -.08

Cd

Cmc/4 (b) Mod. A- Concluded.

Figure 6.- Continued.

1.6

0 TRANSITION FIXED,

K = .0064 cm (.0025 in)

0 SMOOTH MODEL _M_

1.4.

#,,

1.2

1.0

.8

07,

.6

.4.

.2

MOD.B I

/

16 20

4- 80(,De/2

(c) Mod. B.

Figure 6.- Continued.

0 TRANSITION FIXED

K=.0064 cm (.0025 in)

1.6

0 SMOOTH MODEL

,.4

1.2 /

°

i,l

1.0 t

!

o8 .....................

t C_ C> I

.6

I [ I i

<;._ i

I

.4 +

4) !

I

+ 4_

I

.2 +

I

0 !

b

0 .01

.02 0 -.04 -.08 -.12

Cd

Cmc_

(c) Mod. B -Concluded.

Figure 6.- Concluded.

1.6

EXPERIMEMENT

1.4

THEORY

1.2

t

1.0

® C_

.8

.6

.4

.2

"-N-ACA 64 a-212_'""_--'_-

0 4. 8 12 16

e:, Deg.

(a) M = 0.2, Re = 1.5XlO 6.

Figure 7.- Comparison of experimental and theoretical aerodynamic force coefficients.

Transition fixed at O. 12c.

1.6

L_

o

EXPERIMENT

THEORY

1.4

1.2

1.0

)

.8

® C_

.6

C

.4

.2

(9 ®

0 -.04 -.08

0 .01 .02

Cd

Cmcl 4 (a) M = 0.2, Re = 1.5XlO*- Concluded.

Figure 7.- Continued.

1.6

I 1

I 0 EXPERIMENT I

THEORY

1.4-

1.2

1.0

®

.8

C_

.6

.4

.2

o / f AoA64,-2,2 J

0 4- 8 12 16

, Deg.

(b) M=O.3,Re= 1.5X106.

Figure 7.- Continued.

1.6 _ l

0 EXPERIMENT

THEORY

1.4

1.2

/

!

1.0

®

.8

® c_

.6

E

.2

-.04

0 .01 .02 0

Cd

(b) M = 0.3, Re = 1.5XlO 6 - Concluded.

Figure 7.- Continued.

1.6 L L___

0 EXPERIMENT

THEORY

1.4

1.2

1.0

®

.8

/

CZ

.6

/

.4

.2

, ACA 64 -212_

/

E)

4 8

oc, Deg.

(c) M=0.4, Re= 1.9X10 6.

Figure 7.- Continued.

1.6

L_..___._.L..__

0 EXPERIMENT

THEORY

1.4

1.2

1.0

) ®

.8

®

.6

Cl ¢

.4

.2

C

-.04 -.08

0 .01 .02 0

Cd

Cmc/4 (c) M = 0.4, Re = 1.9X 10 6 - Concluded.

Figure 7.- Continued.

5O

1.6

0 EXPERIMENT

/

THEORY

0 o I

1.4

"3

1.2

1.0

.8

.6

.4

.2

MOD. A

(d) M=0.2, Re = 1.5 x I0 6

4 8 12 16

o[, Deg.

(d) M = 0.2, Re = I.SXIO 6.

Figure 7.- Continued.

1.6

EXPERIMENT

m THEORY

1.4

Q ® Q

1.2

1.0

C)

.8

Cl q.:)

.6

.4

.D

.2 Q

i ® f_ Q Q

-.04. -.08

.02 0

Cmc/4 (d) M = 0.2, Re = 1.5X 106 - Concluded.

Figure 7.- Continued.

1.6 i r

O

EXPERMENT 1

THEORY I ®

1.4

1.2 E)

o C)

1.0

®

.8

C_

.6

.4

.2

MOD.B

0 4 8 12 16

(z:_ Deg.

(e) M = 0.2, Re = 1.5X10 6 .

Figure 7.- Continued.

1.6

0 EXPERMENT

-- THEORY o

®

1.4 -----e_. m

1.2

® ®l

1.0

®

.8

Cl ®

.6

®

I

.4

I

.2

I

®

J

Q

-O8

0 - .04

0 .01 .02

Cd

Cmc/4

(e) M = 0.2, Re = 1.5X 10 6 - Concluded.

Figure 7.- Concluded.

EXP THEORY

UPPER SURFACE

El LOWER

SURFACE

Cp

-.8

-.4

.4

r E

I _ACA 641-212

.8

1.2

0 .2 .4 .6 .8 1.0

X

m

C

(a) M=O.2, Re= 1.5X106,_=-0.13°,Cl=0.115.

Figure 8.- Comparison of experimental and theoretical pressure distributions.

5S

EXP. THEORY

-I.2 j 0 UPPER SURFACE

Cp 0 LOWER SURFACE

I ]

I

0 ® CI_ !

r:l_.N-'r:l-"f - EI-...O_Q..

0 2.._ D.o._.,.o_ '

!

I .4 _,

I _-'_CA 641-212

.8

.8 1.0

0 .2 .4 .6

X

C

(b) M = 0.2, Re = 1.5X106, _= 2.44°,Cl = 0.394.

Figure 8. Continued.

I EXP. THEORY

UPPER SURFACE

[]

LOWER SURFACE

C

_ACA 641-212

/

¢

.8

0 .2

.4 .6 .8

X u C (c) M = 0.2, Re = 1.5X106, a = 5.11°,Cl = 0.678.

Figure 8.- Continued.

F

-5 I

' EXR THEORY

0 _-UPPER SURFACE

-4 -- 0 ..... LOWER SURFACE

-3

CP_2

-I

.B.B_B-I ]-_B-B-I_-B-B-B-[]-B_ i_ _CA 641-212

:5

0 .2 .4 .6 .8 1.0

X C (d) M = 0.2, Re = 1.5X106 , _x= 9.58 °, c I = 1.081.

Figure 8.- Continued.

-I.6

EXP.

THEORY

UPPER SURFACE

[]

LOWER SURFACE

Cp

-.8

-.4

.4

_--I_-NACA 641-212

.8

1.2 I

0 .2 .4 .6 .8 1.0

X

I

C

(e) M = 0.3, Re = 1.5X106 , ot =- 1.9 ° ,c l =-0.059.

Figure 8.- Continued.

EXP. THEORY

UPPER SURFACE

[3

LOWER SURFACE

-.8

Cp

_El-El-El -r

.4

_-----'_NACA 64_-212

.8

I

1.2

0 .2 .4 .6 .8 1.0

X

C

(0 M=O.3,Re= 1.SXlO 6,o_=-0.23°,Cl=0.107.

Figure 8.- Continued.

6O

EXP. THEORY

O

UPPER SURFACE

ID

.... LOWER SURFACE

Cp

.2 .4 .6 .8

1.0

X

c

(g) M = 0.3,Re = 1.5X]06, o_ = 2.16 °, cI = 0.407.

Figure 8.- Continued.

EXP. THEORY

UPPER SURFACE

LOWER SURFACE

Cp

J /El / I

_--"_'_ A 641-212

I

.8 r

1.2

0 .2 4 .6 .8 1.0

X m

c

(h) M = 0.3, Re = 1.5X106 , _ = 4.55 ° , c I = 0.685.

Figure 8.- Continued.

-5

-4

EXP. THEORY

UPPER SURFACE

LOWER SURFACE

Cp

-I

e e ¢%_ "_Q (3 I ...6].... D_E]...t- Q-O -E]-I II_E]_E]_E]_ _'-NACA 641-212_

.2 z .6 .8

1.0

x c (i) M = 0.3, Re = 1.5X l06, o_ = 9.74, cl = 1.048.

Figure 8.- Continued.

EXP. THEORY

0 UPPER SURFACE

LOWER SURFACE

Cp

-.8

.4.

--'_N--A CA 641-212

.8

[2

I.O

0 .2 .4 .6 .8

X C (j) M = 0.4, Re = 1.9X106, o_ =-1.84°,Cl = 0.067.

Figure 8.- Continued.

UPPER SURFACE

THE y l

1.2 -- --- LOWER SURFACE

Cp

-.8

-.4

__ o_o_(

.4

____A 641-212 _

.8

1.2

0 .2 .4 .6 .8 1.0

X R

C

(k) M = 0.4, Re = 1.9X106, o_ =-0.27 ° ' c I = 0.122.

Figure 8.- Continued.

0 0 UPPER SURFACE

-I.6 I EXP. THE RY 1

-I.2 13 LOWER SURFACE

C

I .3

_'_NACA 641-212

.8

' I i

1.2

0 .2 .4. .6 .8 1.0

X

m

C

(1) M = 0.4, Re = 1.9XlO 6, oz - 1.91 °, c l = 0.409.

Figure 8.- Continued.

-4

EXP. THEORY

UPPER SURFACE

r'l

LOWER SURFACE

-5

Cp

-2

L.

p,I, :i_13 ''[ J E!

_NACA 641-212

0 .2 .4 .6 .8 1.0

X

m

(3

(m) M = 0.4, Re = 1.9XlO 6 , ot = 4.16 °, c I = 0.699.

Figure 8.- Continued.

-5

-4

THEORY

XP.

UPPER SURFACE

[] LOWER SURFACE

-3

Cp

-2

-I

_CA 641-212

.2 .4 .6 .8 1.0

X c (n) M = 0.4, Re = 1.9X 10 6 , OL = 8.3 °, c l = 1.03.

Figure 8.- Continued.

0 .2 .4 .6 .8 1.0

X

m

C

(o) M = 0.2, Re = 1.5X106, ot =-0.12 °, cI = 0.103.

Figure 8.- Continued.

Figure 8.- Continued.

-I.6 i _ I_YD TMFn_Y

EXP, THEORY

UPPER SURFACE

n

LOWER SURFACE

Cp

.4

.8

1.2

0 .2 .4 .6 .8 1.0

X I

¢

(r) M = 0.2, Re = 1.5X10 6, t_ =-0.15 ° , c I = 0.119.

Figure 8.- Continued.

.3

THEORY

.-i EXR

-I.2 ---- O

UPPER

SURFACE

I

I

LOWER

I [] SURFACE

Cp

I

!

-.8

I

I

I

- t .)

.4

MOD. A

.8

1.2

0 .2 .4 .6 .8 1.0

X

C

(o) M = 0.2, Re = 1.5XlO 6, ot =-0.12 ° , c I = 0.103.

Figure 8.- Continued.

-I.6

EXR THEORY

0 UPPER SURFACE

rl

LOWER SURFACE

C

C

-.4

"" / [] _ EH:3""

MOD. A

1.0

.2 .4 .6 .8

x

c

(p) M = 0.2, Re = 1.5X106, ct = 5.1 ° , cl = 0.670.

Figure 8.- Continued.

r,7o

EXR THEORY

UPPER SURFACE

rl

LOWER SURFACE

-2 o

MOD. A

0 .2 .4 .6 .8

1.0

(q) M = 0.2, Re = 1.5X106, or= 15.99°,Cl = 1.488.

Figure 8.- Continued.

0 . UPPER SURFACE

EXP, THEORY _--

El LOWER SURFACE

Cp

-.8

-.4

.4

.8

1.2

0 .2 .4 .6 .8 1.0

X

C

(r) M=O.2, Re= 1.5XlO6,a=-O. 15°,cl=O. 119.

Figure 8.- Continued.

-I.6

EXR THEORY

0 UPPER SURFACE

0 .... LOWER SURFACE

-I.2

.8

1.2

0 .2 .4 .6 .8 1.0

X

(s) M = 0.2, Re = I.SX 106 , o_ = 2.47 °, c 1 = 0.393.

Figure 8.- Continued.

-I.8

EXR THEORY

UPPER SURFACE

[]

LOWER SURFACE

-.8

-.4

mJ ].-m...n m

m...w..-

,-o_o.o

.4

I":'1'

d

MOD.B

q

.8

[ I ...........

1.2

0 .2 .4. .6 .8 1.0

X C (t) M = 0.2, Re = 1.5X106, o_ = 5.09, Cl = 0.671.

Figure 8.- Continued.

EXP.

THEORY

UPPER

SURFACE

El

LOWER

SURFACE

-_-_..ee="--

-EI-13-B-E - EI-m-B-E .. 13- El- El- E

/

l

.2 .4 .6 .8 1.0

X m

c

(u)M = 0.2, Re = 1.5X106, o_ = 14.05 ° ' cl = 1.437.

Figure 8.- Concluded.

S_SA-U_S].y, ,075 A-601 8 7)