A recontoured, upper surface designed to increase the maximum lift coefficient of a modified NACA 65 (0.82) (9.9) airfoil section

NASA Technical Memorandum 85855

'"

NASA-TM-85855 19840009076

'FOB. REYEro:.N~

I

I

A Recontoured, Upper Surface

Designed to Increase the

Maximum Lift Coefficient of a

Modified NACA 65 (0.82) (9.9)

Airfoil Section

Raymond M. Hicks

February 1984 LANGLC:Y RESEARl:I-' CENTER LIBRARY, NASA HAMPTON, VIRGINIA

NI\S/\

National Aeronautics and Space Administration NASA Technical Memorandum 85885

A· Recontoured, Upper Surface

Designed to Increase the

Maximum Lift Coefficient of a

Modified NACA 65 (0.82) (9.9)

Airfoil Section '

Raymond M. Hicks, Ames Research Center, Moffett Field,' California

NI\S/\

National Aeronautics and Space Administration Ames Research Center Moffett Field, California 94035 NOMENCLATURE p - P L 00 c pressure coefficient.

p qoo airfoil chord. m (in.)

c section drag coefficient section lift coefficient CQ., section pitching moment coefficient referenced to quarter chord c m 6* H shape factor. 8 upper surface thickness parameter K M Mach number 2 2 p static pressure. N/m (lb/ft ) 2 2 q dynamic pressure. N/m (lb/ft ) Re Reynolds number based on free-stream conditions and airfoil chord distance along airfoil surface. m s velocity. m/sec u x airfoil abscissa. m (in.)

airfoil ordinate. m (in.)

y a angle of attack. deg displacement thickness. m 8 momentum thickness. m Subscripts max maximum min minimum L local 00 free-stream conditions iii SUMMARY A wind-tunnel test was conducted to assess the effectiveness of a recontoured upper surface to increase the maximum lift coefficient of a modified NACA 65 (0.82) (9.9) airfoil section. The recontoured airfoil was slightly thicker when compared to the original airfoil. The modification was confined to the forward 50% of the chord. The recontoured and original airfoils were tested at Mach numbers of 0.3 6 6 and 0.4 and at Reynolds numbers of 2.3x10 and 4.3xl0 • The recontoured airfoil showed a higher maximum lift coefficient, lower drag coefficients, and similar pitching moment characteristics when compared with the original airfoil section at all test conditions. At lift coefficients near the design value, the recontoured airfoil had only slightly less drag than the original profile did, whereas at high lift coefficients the drag of the recontoured airfoil was substantially lower.

The improvements found for the recontoured airfoil of the pres~nt study are similar to those found during previous investigations of recontoured 6-series airfoils with less camber.

INTRODUCTION Previous experimental evaluations of a recontoured upper surface of the NACA 64 -212 and the NACA 632-215 airfoil sections showed that substantial increases in the maximum lift coefficients were achieved at Mach numbers of 0.2, 0.3, and 0.4 6 6 and at Reynolds numbers between 1.9x10 and 2.5x10 (refs. 1 and 2). The improve- ments in maximum lift of these airfoil sections were achieved by reducing the adverse pressure gradients near the leading edge along the upper surface at high angles of attack. Both of the above sections had low camber which is typical of many high-performance single and twin engine business aircraft designed over the last 40 years.

The current investigation was conducted to assess the effects of a recontoured upper surface of a highly cambered 6-series airfoil typical of propeller and com- pressor blade sections. The airfoil chosen for this study was an NACA 65 (0.82) (9.9) airfoil with a circular arc camber which was the profile used at the 83% radial station of the compressor blades in the 40- by 80- by 120-Foot Wind Tunnel at Ames Research Center. The wind-tunnel results reported herein provide an evaluation of the design methodology of references 1 and 2 at lower thickness-chord ratios and greater camber.

The type of recontouring studied here is useful for the retrofitting of exist- ing lifting surfaces when the permissible change to the airfoil's profile is limited to relatively small changes brought about by manufacturing constraints.

THEORETICAL AIRFOIL RECONTOURING The technique used to recontour the modified NACA 65 (0.82)(9.9) airfoil section is similar to the method described in references 1 and 2. The upper sur- 1 2 8X face thickness was increased by adding the function Kx / (1 - x)/e to the ordi- nates of the upper surface of the normalized airfoil. The parameter, K, was increased incrementally and the pressure distribution was analyzed for each value of K by using Program H (ref. 3). The four parameters considered during the design process were the boundary layer shape factor H, the separation parameter [(S/u)(du/ds)], the peak pressure at the leading edge at high angle of attack, and the pressure gradient which follows the peak pressure. The most useful parameter, and the one which dictated the amount of thickness to be added to the airfoil in this investigation,was the peak pressure at the leading edge. The other three parameters were monitored to ensure consistency with the trend indicated by the peak pressure; i.e., if all four parameters indicated that the amount of thickness added was beneficial, the design was accepted. A plot of the peak pressure (abso- lute value of Cp . ) vs K is shown in figure 1. Note that the peak pressure mln decreases rapidly as K increases, until a plateau is reached. The value of K at the beginning of the plateau determined the amount of thickness to be added to the upper surface of the airfoil. This design method is based on the assumption that C£ is inversely proportional to the magnitude of the peak pressure near max coefficient shown in figure 1 is the the leading edge for attached flow. The lift over 95% of the upper surface of the highest value predicted to have flow attached to maximize C£ directly by using original airfoil. It would be more effective max a numerical optimization algorithm coupled to a Navier-Stokes code; however, such a technique does not exist, so it is neC~SSqry to rely on simpler methods such as the one used here.

The recontoured and original airfoil sections are shown in figure 2. The coor- dinates are given in table 1.

MODE;LS Two airfoil models with the modified NACA 65 (0.82)(9.9) and the recontoured profiles were cast of aluminum epoxy material with the pressure tubing laid into the material during casting. Pressure orifices were subsequently drilled normal to the surface to meet the tubing at 25 locations on the upper surface and at 18 loca- tions on the lower surface. Each model had a l5.24-cm (6-in.) span and chord. A model photograph is shown in figure 3., WTNDTUNNEL The tests were conducted in the Ohio State University 6- by 22-Inch Transonic Airfoil Tunnel (ref. 4). The tunnel is a blow,down facility with perforated floor and ceiling and has a Mach number range from 0.2 to 1.07 and a Reynolds number range 6 6 from 2x10 to 34x10 depending on Mach number. The tunnel has separate plenum cham- bers above and below the test section. An installation photograph showing a model installed in the tunnel is shown in figure 4.

TEST CONDITIONS The section aerodynamics characteristics of the two airfoils were obtained at 6 6 Mach numbers of 0.3 and 0.4 and Reynolds numbers of 2.3xl0 and 4.3xl0 • The angles 0 0 of attack ranged from approximately _4 .to 17 depending on the angle of stall for each model. Data were obtained at all test conditions· with free transition because the full-scale Reynolds number was attained during testing, and because of diffi- culty in simulating a realistic in-service surface condition on a wind-tunnel model.

Pressure coefficients were determined from surface pressure measurements. Sec- tion normal force, chord force, and pitching moment coefficients were calculated by integrating surface pressure coefficients. The pitching moment coefficients were referenced to the quarter chord point. Section profile drag was calculated from pressures measured by a traversing total pressure probe and a separate static pres-- sure probe.

The model angle of attack can be corrected for the presence of the tunnel walls by use of the following equation:

a = a 0 Sc

true geom. -. ~ where atrue is the corrected angle of attack and a • is the angle of attack geom set during testing. The angle of attack used in this report is the geometric angle.

RESULTS AND DISCUSSION Force Coefficients The aerodynamic force coefficients for both airfoil sections are presented in figures 5 and 6. The recontoured airfoil exhibits a higher maximum lift coefficient than the original 6-series section at all test conditions. The increase ranges from a 10% improvement at Mach 0.4 and a Reynolds number of 2.2xl0 to. 27% improvement at Mach 0.3 and a Reynolds number of 4.3xl0 • The lift-curve slope appears to be slightly greater for the recontoured airfoil than for the original airfoil at all test conditions. The stall characteristics of the recontoured profile are somewhat more gradual than for the modified 6-series section at all test conditipns. Note that the maximum lift coefficient of the recontoured airfoil is more sensitive to Reynolds number and Mach number than it is for the modified 6-series airfoil (fig. 7).

The profile drag data of figures 5 and 6 show a somewhat lower level of drag for the recontoured airfoil at moderate lift coefficients except at Mach 0.4 and a Reynolds number of 2.2xl0 (fig. 6(a». At higher lift coefficients the drag of the recontoured airfoil is considerably less than that of the modified 6-series section at all test conditions. The lower drag of the recontoured profile at moderate lift coefficients is opposite to the trend observed when a similar recontouring was applied to lower cambered 6-series airfoils (refs. 1 and 2). In those studies the recontouring caused a. loss in laminar flow over the upper surface at lift coeffi- cients near the design value which resulted in slightly higher profile drag. The lower drag of the recontoured section at high lift coefficients is consistent with the result found with the earlier 6-series modifications reported in references 1 and 2 and is due to improved pressure recovery over the aft region of the airfoil.

The data of figures 5 and 6 show that both airfoils have similar pitching moment characteristics. The aerodynamic center position is nearly the same for both airfoils. This result is somewhat different than that, reported in references 1 and 2 wh~re the modified airfoils w~re found tb have a slightly more forward posi- tion of the aerodynamic center than the original 6-series .. airfoils had.

Pressure distributions for both airfoils are shown in figures 8 through 11.

All pressure distributions show a negative trailing edge pressure coefficient at all angles of attack which is apparently caused by the fairly large trailing edge blunt- ness of both airfoils .. Pressure distributions are shown for both Mach numbers at the higher Reynolds number only because the general shape of the curves are a weak function of Reynolds number.

The main effects of the recontouring are production of a less favorable pressure gradient over the forward 50% chord at the lower tift c,qefficients (compare figs. 8(a) and 10(a)), and a reduction of the peak negative pressure at the leading edge at high lift coefficients (compare figs. 8(d) and 10(d)). The reduction in favorable pressure gradient at low-lift coefficients apparently did not cause pre- mature transition on the recontoured airfoil as indicated by the drag curves shown earlier. The reduction in adverse pressure gradient at high-lift coefficients resulted in better pressure recovery near the trailing edge of the recontqured pro- file (figs. 8(e) and10(e)) which explains the higher maximum lift coefficients and greater angle of stall for the recontoured airfoil.

An analysis of compressor flow indicates that the recontoured airfoil will reduce the stall speed of the compressor by approximately 20% compared with the original 6-series airfoil (Borst, Henry V.; Private communication, May 1983).

CONCLUSIONS A wind-tunnel test was conducted to evaluate a recontoured upper surface designed to improve the maximum ~ift coefficient and stalling angle of a modified NACA 65 (0.82)(9.9) airfoil section and to improve the speed margin before stall of a compressor using this airfoil. The test conditions were M = 0.3 and 0.4, 6 6 x10 Re = 2.3 , and 4.3x10 • The following results were achieved: 1. Increasing the upper surface thickness over the forward 50% of the chord of the modified NACA 65 (0.82)(9.9) airfoil increased the maximum lift coefficient 6 6 by 10% at M = 0.4 and Re = 2.3x10 and by 27% at M = 0.3 and Re = 4.3x10 • 2. The recontouring had a negligible effect on the pitching moment character- istics of the modified 6-series airfoil.

3. The recontouring produced slightly lower drag at lift coefficients near the design value at most test conditions.

4. The recontoured airfoil had substantially lower drag than the modified 6-series airfoil did at high~lift coefficients.

5. The stall of the recontoured airfoil was somewhat more gradual than that of the 6-series airfoii.

6. The improvements found for the recontoured airfoil in the present study are similar to those found during previous investigations of recontoured 6-series air- foils with less camber.

7. A compressor flow analysis indicates that the recontoured airfoil will give a 20% lower speed before stall when compared to the original NACA 6-series airfoil.

REFERENCES 1. Hicks, Raymond M.; Mendoza, Joel P.;and B~ndettini, Ange~o: Effects of Forward Contour Modification on the Aerodynamic Characteristics of the NACA 64 -212 Airfoil Section. NASA TMX-3293, 1975.

2. Hicks, Raymond M.; and Schairer, Edward.T.: Effects of Upper Surface Modifica'- tion on the Aerodynamic Characteristics<ofthe NACA632-2l5 Airfoil Section.

NASA TM-78503,1979.

3. Bauer, Frances; Garabedian, Paul; Korn, David; and Jameson, Antony: Supercriti- cal Wing Sections II, Lecture Notes in Economics and Mathematical Systems.

Springer-Verlag, 1975.

4. Lee, J. D.; Gregorek, G. M.; and Korkan, K. D.: Testing Techniques and Inter- ference Evaluation in the Ohio State University Transonic Airfoil Facility.

AIM Paper 78-1118, AIM 11th Fluid and Plasmadynamics Conference, 1978.

TABLE 1.- AIRFOIL COORDINATES Modified NACA 65 (0.82)(9.9) airfoil coordinates x/c x/C Yu/C YL/C 0.00000 0.00000 0.99931 -0.00394 0.00354 0.00844 0.99527 -0.00327 0.00591 0.01038 0.94893 0.00180 0.01048 0.01362 0.89842 0.00491 0.02231 0.01975 0.84782 0.00634 0.04651 0.02978 0.79756 0.00667 0.07101 0.03827 0.74745 0.00606 0.09558 0.04590 0.69757 0.00492 0.14546 0.05889 0.64779 0.00321 0.19556 0.06935 0.59843 0.00122 0.24592 0.07799 0.54910 -0.00099 0.29655 0.08459 0.49991 -0.00312 0.34722 0.08937 0.45069 -0.00499 0.39805 0.09224 0.40164 -0.00650 0.44899 0.09344 0.35246 -0.00791 0.49985 0.09246 0.30315 -0.00941 0.55060 0.08935 0.25373 -0.01079 0.60127 0.08457 0.20421 -0.01206 0.07810 0.15434 0.65185 -0.01316 0.70210 0.07021 0.10423 -0.01355 0.75233 0.06090 0.07874 -0.01337 0.80223 0.05063 0.05340 -0.01259 0.85197 0.03931 0.02748 -0.01093 0.90159 0.02738 0.01440 -0.00912 0.95118 0.01522 0.00908 -0.00775 0.99675 0.00471 0.00639 -0.00655 1. 00065 0.00391 0 .. 00000 0.00000 Recontoured Airfoil Coordinates 0.00000 0.00000 0.99931 -0.00394 0.00354 0.01363 0.99527 -0.00327 0.00591 0.01694 0.94893 0.00180 0.01048 0.02200 0.89842 0.00491 0.02231 0.03074 0.84782 0.00634 0.04651 0.04254 0.79756 0.00667 0.07101 0.05089 0.74745 0.00606 0.09558 0.05762 0.69757 0.00492 0.06805 0.64779 0.00321 0.14546 0.19556 0.07605 0.59843 0.00122 0.24592 0.08269 0.54910 -0.00099 0.29655 0.08701 0.49991 -0.00312 0.34722 0.09153 0.45069 -0.00499 0.39805 0.09366 0.40164 -0.00650 0.44899 0.09436 0.35246 -0.00791 0.49985 0.09304 0.30315 -0.00941 0.25373 0.55060 0.08972 -0.01079 0.6012.7 0.08479 0.20421 -0.01206 0.65185 0.07823 0.15434 -0.01316 " 0.70210 0.07029 0.10423 -0.01355 0.75233 0.06095 0.07874 -0.01337 0.80223 0.05066 0.05340 -0.01259 0.85197 0.03932 0.02748 -0.01093 0.90159 0.02739 0.01440 -0.00912 0.95118 0.01523 0.00908 -0 ~ 00775 0.99675 0.00471 0.00639 -0.00655 0.00000 0.00000 1. 00065 0.00391 5.2 4.8 4.4 4.0 :2 ::E 3.6 Co u 3.2 2.8 2.4

o .02 .04 .06 .08 .10

K Figure 1.- Effect of forward airfoil thickness on peak pressure coefficient M = 0.39, C~ = 1.35.

-- RECONTOURED AIRFOIL ------ ORIGINAL AIRFOIL

.1r~· ~

y/c _.:C=---~--,===

1.0 .9 .8 .7 o .1 .2 .3 .4 .5 .6

x/c Figure 2.- Airfoil sections tested.

It.:S'f ~., ,~" Figure 3.- Wind tunnel model.

Figure 4.- Airfoil model in the Ohio State University 6- by 22-Inch Wind Tunnel.

0-- RECONTOURED AIRFOIL [}--- ORIGINAL AIRFOIL 2.0 1.6 ..

.8 .4.

0'-'-------'----'------'---'----------'--'---------' -.1 -.2 .06 0 -4 0 4 8 12 16 0 .02 ex, deg C / mC (a) Re = 2.3 X 10 0-- RECONTOURED AIRFOIL [}--- ORIGINAL AIRFOIL 2.0 1.6 _--0---0 1.2 .8 .4

o '-- '-- __ -'-- __ -'----- -'-----_-----.1

-.4 0 4 8 12 16 0 .02 .04 .06 0 -.1 -.2 ex, deg Cd C mC/ (b) Re = 4.3 X 10 Figure 5.- Aerodynamic characteristics of the original and recontoured airfoils, M = 0.30.

0-- RECONTOURED AIRFOIL 0--- ORIGINAL AIRFOIL 2.0

T

1.6 ...0---------, ,- ;:f '/ I .8

-.1 -.2 .06 0 8 12 o

.02 16 0 4 C ex,deg m Cf4 (a) Re = 2.2 X 10 0-- RECONTOURED AIRFOIL 0--- ORIGINAL AIRFOIL 2.0 1.6 1.2 .8 .4 0'-- __ -'----- __ .1-- __ -L..- __ -'---_----' -.1 -.2 .06 0 .02 .04 -4 0 4 8 12 16 0 C ex, deg Cd mCf4 (b) Re = 4.3 X 10 Figure 6.- Aerodynamic characteristics of the original and recontoured airfoils, M = 0.40.

RECONTOURED AIRFOIL --- ORIGINAL AIRFOIL 1.8 1.6 1.2 1.0 L..- __ ..l.- __ .....I- __ -L __ -L __ -.l.. __ --''-- __ J

o 2.0 2.4 2.8 3.2 3.6 4.0 4.4

Re (10)-6 Figure 7.- Effect of Reynolds number and Mach number on the maximum lift coefficient of the original and recontoured airfoils.

-1.0 0 0

°

°

-.5 0

°

C

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p

<>

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<>

<>

.5 o UPPER SURfACE

o LOWER SURFACE

(a) C = 0.52, 0' = 0°

Q 1 0 " . 0 .2 .4 .6 .8 1.0 x/c -1.0 o o o -.5 o o o

o <>

o f-------------- ----.:0'='l" :H

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.5 o UPPER SURFACE

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(h) C = 0.71, a' = 2° Q 1.5 .2 .4 .6 .8 1.0 x/c

Figure 8.- Pressure distributions for the original airfoil; M = 0.30, Re 4.3xl0 •

-1.5 o UPPER SURFACE

<> LOWER SURFACE

-1.0 D D D D D D D

CltJ D D

D D D D -.5 .

D D C p D D D 0 0 0 0

<9

.5 (c) C = 0.91, ex = 4" Q 1.0 1.0 .2 .4 .6 .8 x/c -5 o UPPER SURFACE

<> LOWER SURFACE

-4 -3 D -2 D D D D D D D 0 DOD D -1 D o o D o o

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DO 0f-------------------'='-l:..ID

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/'.. <> <> <>

000 v

<f> (d) C = 1.26, ex = 8°

Q 1 i'(.'.:r----'------l----'----..L-----'

.8 1.0 o .2 .4 .6

x/c Figure 8.- Continued.

-6 o UPPER SURFACE

<> LOWER SURFACE

-5 -4 o o o o o 0 o 0 o 0 -1 o o o o o 0 DDDDrn

o 1-----------------------1

o 0

o

<> 0 0 0

0 <>

. # (e) C = 1.36, CY = 10°

Q 1'v'-----L-----L----'-- __ ..L- __ ----'

o .2 .4 .6

1.0 .8 x/c Figure 8.- Continued.

-6 o UPPER SURFACE

<> LOWER SURFACE

-5 -4 ~3 C p -2 o o 0 o 0 -1 0 0 o o 0 O~------------------rl

o

o o

000 0 .0 (f) C = 1.37, a = 12() Q 1 .

O· .2 .4 .6 .8 1.0 x/c Figure 8.- Concluded.

-1.0 o -.5 o o <) <) <) <) <) <) .5 1.0 .

o UPPER SURFACE

<> LOWER SURFACE

(a) C = 0.53, ex = 0° Q 1.5 '-- -'-- __ -L --'---- -'-- __ ---' .2 .4 .6 .8 1.0 x/c -1.0 0 0 [] <) .<)<)<) <) <) C <) <) <) p <)

<>

<>

o UPPER SURFACE

<> LOWER SURFACE

(b) CQ = 0.74, ex = 2° 1.5 .2 .4 0 .6 .8 1.0 x/c

Figure 9.- Pressure distributions for the original airfoil; M = 0.40; Re 4.3xl0 •

1.5 .2 .4 0 .6 .8 1.0 x/c Figure 9.- Continued.

-5 o UPPER SURFACE

o LOWER SURFACE

-3 o -2 o o o 0 000 0 o 0 -1 o ,0 o 0 o 00 0

o I-- "_O---.:O::::..:::J" %

o 0

o

o

(>

<9

1 "" (d) C = 1.26, ~ = 8° Q .6 .8 1.0 x/c Figure 9.- Continued.

-5f <> UPPER SURFACE

o LOWER SURFACE -4 EhJ -3 o o -2 o o 0 . 0 -1 o 0 o o o o 0 o 0 . 0000 0 0 0 0

#

(e) C = 1.36, ex = 10° Q .2 .4 .6 .8 1.0 x/c Figure 9.- Continued.

-5

o UPPER SURFACE

o LOWER SURFACE -4 C PCRIT OJ -3 0 o o -2 o o 0 o 0 -1 o 0 o o

00 <>

o 0 o 0000

o I---------------------(.)i

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1#

2L...-----'--------'----.l-------L _

o .2 .4 .6 .8 1.0

x/c Figure 9.- Concluded.

-1.0 0 0 0 0

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o UPPER SURFACE

<> LOWER SURFACE

(a) C = 0.54, a = 0° Q

1.0 a

.2 .4 .6 .8 1.0 x/c -1.0

&DO

0 0 0 0 0 0 0 0 0

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C <> <>

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p

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.5 o UPPER SURFACE 1.0 .

"

<> LOWER SURFACE

1.5 a .2 .4

.6 1.0 .8 x/c

Figure 10.- Pressure distributions for the recontoured airfoil; M = 0.30 Re 4.2xl0 •

t -1.5 o UPPER SURFACE

<> LOWER SURFACE

EBJ· o 0 o -1.0 o 00000 o 0 o 0 o o -.5 o o o

o

o 01-----'-----------------IL:.l-:1.l <>

o 0 00<> <>

<>

<9

o

1.5 .2 .4 .6 .8 1.0 x/c Figure 10.- Continued.

-4 o UPPER SURFACE

<> LOWER SURFACE

-3 .

o o o -2 o o o 00000 o 0 -1 o 0 o o o o

o o

o

o 000 0 000

.2 .4 .6 .8 1.0 x/c Figure lO.-Continued.

-5 o UPPER SURFACE

<> LOWER SURFACE

-4 -3 -2 o C p o o 0 o 0 o 0 -1 [j 0 o []

o

o o

o o

o

o

... L .. --- __ -----l...-..... ..,..-......L--. --' ~ 1 .2 .4 .6 .8 1.0 x/c Figure 10.- Continued.

-6 o UPPER SURFACE

o LOWER SURFACE

-5 -4 [:J -3 C p -2 [) [) -1 [.J

o

OED

o

o

0 0

&'

(f) C = 1.62, IX = 12°

Q .2 .4 .6 .8 1.0 x/c Figure Continued.

10.- -----------------------_._-------------------- -9 o UPPER SURFACE

<> LOWER SURFACE

[] [] [] [] [] [] [] [] -1 fJ [] [] [] []

<>

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(g) CQ = 1.69, a = 14° .6 .8 1.0 x/c Figure 10.- Continued.

o UPPER SURFACE

<> LOWER SURFACE

C o PCRIT -5 C p -3 -1 00 0 0

o

o

o 0 0

(h) CQ ::;: 1.74, c< ::;: 16° 3 0~------:------.1:--_-....L __ ---l. __ .------J .2 .4 .6 .8 1.0 x/c Figure 10.- Concluded.

-1.0 0 0

° 0

-.5

eP

°

o· <P % <:>

<>

C p

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.5 .

o UPPER SURFACE (a) C = 0.55,0' = 0°

o LOWER SURFACE

Q 1.0 L- ---'--- __ ----L_~ _ _l_ _'__ __ ----'

1.0 .8 .6 .4 .2 o

x/c -1.0 c:JD0 0 0 0

°

-.5

°

<><><>

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C

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p

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.5 1.0 o UPPER SURFACE

<> LOWER SURFACE

(b) C = 0.76,0' = 2°

Q 1.5 0 .2 .4 .6 .8 1.0 x/c Figure 11.- Pressure distributions for the recontoured airfoil, M = 0.40, Re = 4. 3x10 • ·1.5 o UPPER SURFACE ~

o LOWER SURFACE

-1.0 o 0 0 o 0 0 0 -.5 C p 0 0 0 0 0 000 0

<9

.5 1.0 (e) C = 0.97,0' = 4" Q 1.5 O~--_L -'-----_----------~-----'-- ---L .2 .4 .6 .8 1.0 xie Figure 11.- Continued.

-4 Cl UPPER SURFACE

<> LOWER SURFACE

o o o -2 o 0 []

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I (d) C = 1.36, (X = 8°

Q .6 .8 1.0 x/c Figure 11.- Continued.

-5 o UPPER SURFACE <) LOWER SURFACE -4 C PCRll -3 -2 o ~ 0 DOD o 0 -1 o 0 o o [] o

o

. 0 0 Ol----------------~--=-~ <) <) <)

° <)

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0°<>

&

.0 (e) C = 1.5, ex = 10° Q 1.0 0 .2 ~4 .6 .8 x/c Figure 11.- Continued.

-6 D.

o UPPER SURFACE

o LOWER SURFACE

-5 -4 DC PCRIT -3 D D D C P -2 D o o D D D .~ -1 o D D D D D 0 DDDDrn d

o

0 0 0'

1~

(f) C = 1.54, cu= 12° Q .2 .4 .6 .8 1.0 x/c Figure ll.- Continued.

-6 o UPPER SURFACE

<> LOWER SURFACE

[J -5 [J -3 [J [J [J [J -1 ,,-------------------------------------,.------------- -6 o UPPER SURFACE

o LOWER SURFACE

-5 o ··4 C °PCRIT -3 o o -1 0f------------------nJ (h) C = 1.47, C\' = 16° Q .2 .4 .6 .8 l.O x/c Figure 11.- Concluded.

1. Report No. 3. Recipient's Catalog No.

I 2. Government Accession No.

NASA TM-85855 5. Report Date 4. Title and Subtitle SURFACE DESIGNED A RECONTOURED, UPPER TO INCREASE Februarv lQRll 6. Performing Organization Code THE MAXIMUM LIFT COEFFICIENT OF A MODIFIED NACA 65 (0.82)(9.9) AIRFOIL SECTION 7. Author(s) 8. Performing Organization Report No.

Raymond M. Hicks A-Q5lll 10. Work Unit No.

9. Performing Organization Name and Address T3334Y Ames Research Center 11. Contract or Grant No.

Moffett Field, CA 94035 13. Type of Report and Period Covered Ic' 12. Sponsoring Agency Name and Address TeGhnical Memorandum National Aeronautics and Space Administration 14. Sponsoring Agency Code Washington, DC 20546 505-31-21 15. Supplementary Notes Point of Contact: Raymond Hicks, MS 227-8, Ames Research Center, Moffett Field, CA 94035, (415) 965-6396 or FTS 448-6396 16. Abstract surface was designed to increase the maximum lift A recontoured upper section which was coefficient of a modified NACA 65 (0.82)(9.9) airfoil numbers of' 2.3xl0 and tested at Mach numbers of 0.3 and 0.4 and Reynolds for comparison with the 4.3xl0 • The original 6-series section was tested recontoured section.

found to have a higher maximum lift coeffi- The recontoured profile was The recontoured test conditions than the original airfoil.

cient at all characteristics and nearly the same pitching moment airfoil showed less drag all test conditions.

as the original 6-series airfoil at recontoured airfoil of the present study The improvements found for the investigations of recontoured are similar to those found during previous 6-series airfoils with less camber.

17. Key Words (Suggested by Author(s)) 18. Distribution Statement Airfoil design Unlimited Propeller design Compressor design Subject Category 02

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19. Security Class;f. (of this report) 22. Price" ,.20. Security Classif. (of this pagel 21. No. of Pages Unclassified Unclassified 40 A03 "For sale by the National Technical Information Service, Springfield, Virginia 22161 IIIII1I1II1 ~~rlrllll~lIrll~r~II~11 ~Iilr~~~ 111111/1111 , 3 1176 00513 4573