NASA Technical Paper 1786

Low-Speed Aerodynamic Characteristics of a WPercent-Thick Medium- Speed Airfoil Designed for General Aviation Applications

Robert J. McGhee and William D. Beasley Langley Research Center Ha nzpto 11, Virgiuia

NASA National Aeronautics and Space Administration

Scientific and Technical Information Branch

1980

SUMMARY

An investigation was conducted in the Langley Low-Turbulence Pressure Tunnel to determine the low-speed two-dimensional aerodynamic characteristics of a 17-percent-thick medium-speed airfoil designed for general aviation appli- cations. The results are compared with data for the 17-percent-thick low-speed airfoil and the 13-percent-thick medium-speed airfoil. Theoretical predictions of the drag-rise characteristics for the medium-speed airfoil are also provided. The tests were conducted over a Mach number range from 0.10 to 0.32, a chord Reynolds number range from 2.0 x lo6 to 12.0 x lo6, and an angle-of-attack range from about -8O to 20°.

The results of the investigation indicate that maximum section lift coef- ficients at a Mach number of 0.15 increased from about 1.6 to 2.0 as the Reynolds number increased from about 2.0 x lo6 to 12.0 x lo6. teristics were of the trailing-edge type and were docile at all Reynolds num- bers. The application of a roughness strip near the leading edge of the airfoil decreased the maximum section lift coefficient as much as 0.04 over the test Reynolds number range. Increasing the Mach number from 0.10 to 0.32 at a constant Reynolds number of 6.0 x l o 6 decreased the maximum section lift coef- ficient about 0.03. The magnitude of the quarter-chord pitching-moment coeffi- cient was decreased about 25 percent, and the drag coefficient decreased at all lift coefficients (fixed transition) for the 17-percent-thick medium-speed air- foil compared with the 17-percent-thick low-speed airfoil. The predominant effects of increasing airfoil thickness from 13 percent to 17 percent for the medium-speed airfoils were to decrease the maximum section lift coefficient and to increase the drag coefficient at all lift coefficients (fixed transition).

Stall charac-

INTRODUCTION

Research on advanced-aerodynamics-technology airfoils for general aviation applications has received considerable attention over the last several years at the Langley Research Center. An initial family of low-speed airfoils was devel- oped; this research is summarized in reference 1. Recently, the general avia- tion industry indicated a requirement for airfoils which provide higher cruise Mach numbers than the low-speed airfoils and which still retain good high-lift low-speed characteristics. These medium-speed airfoils have been designed to fill the gap between the low-speed airfoils and the supercritical airfoils for application on light general aviation aircraft. Reference 2 reports the results of a 13-percent-thick medium-speed airfoil designed for a lift coefficient of 0.30 and a Mach number of 0.72.

The present investigation was conducted to determine the low-speed aerody- namic characteristics of a 17-percent-thick medium-speed airfoil designed for a lift coefficient of 0.30, a Reynolds number of 14.0 x lo6, and a Mach number of 0.68. This airfoil is designated as MS(1)-0317. In addition, the results are compared with data for the 17-percent-thick low-speed airfoil (LS (1) -041 7)

and the 13-percent-thick medium-speed airfoil (MS (1 ) -031 3) . Theoretical pre- dictions of the drag-rise characteristics for the medium-speed airfoil are also provided.

The investigation was performed in the Langley Low-Turbulence Pressure Tunnel over a Mach number range from 0.10 to 0.32. The Reynolds number, based on the airfoil chord, varied from about 2.0 x lo6 to 12.0 x l o 6 , and the geometric angle of attack varied from about -8O to 20°.

SYMBOLS

Values are given in both SI and U.S. Customary Units. The measurements and calculations were made in U.S . Customary Units.

cP

C

CC

Cd

c:

Cl

Cm

Cn

h

M

P

R

X

2

Pk - pa7 qa7

pressure coefficient,

airfoil chord, cm (in. 1

section chord-force coefficient,

section profile-drag coefficient, c: d(2) Wake

point-drag coefficient

section lift coefficient,

section pitching-moment coefficient about quarter-chord point,

cn cos O! - cc sin ci

-($ Cp(: - 0.25) d(z) + ($ Cp d(:)

section normal-force coefficient, -@ cP d(:)

vertical distance in wake profile, cm (in.)

free-stream Mach number

static pressure, Pa ( lb/ft2)

dynamic pressure, Pa ( lb/ft2)

Reynolds number based on free-stream conditions and airfoil chord

airfoil abscissa, cm (in.)

"

z airfoil ordinate, cm (in.)

ZC

Zt mean thickness, cm (in.)

a geometric angle of attack, deg

mean camber line ordinate, cm (in.)

Subscripts :

II local point on airfoil

max maximum

W free-stream conditions

Abbreviations:

LS (1 1 low-speed, first series

MS(1) medium-speed, first series

AIRFOIL DESIGNATION

A sketch of the section shape for the 17-percent-thick medium-speed airfoil is shown in figure 1. The airfoil is designated in the form MS(1)-0317. MS(1) indicates medium speed (first series). foil design lift coefficient in tenths (0.30), and the last two digits designate the airfoil maximum thickness in percent chord (17).

The next two digits designate the air-

AIRFOIL DEVELOPMENT

The intention of medium-speed airfoil development was to combine the best features of low-speed and supercritical airfoil technology; this airfoil devel- opment is discussed in detail in reference 2. The design objective of the medium-speed airfoils was to increase the cruise Mach number of the low-speed airfoils while retaining their good high-lift, low-speed characteristics. This 17-percent-thick medium-speed airfoil was designed for a lift coefficient of 0.30, a Reynolds number of 14.0 x lo6, and a Mach number of 0.68. The airfoil shape was changed iteratively until the design pressure distribution was obtained. (See fig. 2 . ) The computer program of reference 3 was used to pre- dict the results of various airfoil modifications.

The design pressure distributions for the 13-percent and 17-percent medium- speed airfoils are compared in figure 2. Note that for the 17-percent-thick airfoil, which has higher induced velocities, the start of the aft upper-surface pressure recovery is located at about 0 . 5 0 ~ ~ compared with about 0.60~ for the 13-percent airfoil. This is required in order to keep the aft pressure gradient gradual enough to avoid separation for the thicker airfoil. The thickness dis-

t r i b u t i o n and camber l i n e for t h e 17-percent medium-speed a i r f o i l are shown i n f i g u r e 3 , and t h e a i r f o i l d e s i g n c o o r d i n a t e s are p r e s e n t e d i n t a b l e I.

MODEL, APPARATUS, AND PROCEDURe

Model

The a i r f o i l model was c o n s t r u c t e d w i t h a metal core around which p las t ic f i l l and t w o t h i n l a y e r s of f i b e r g l a s s were used to form t h e contour of t h e a i r - f o i l . The model had a chord of 61 cm (24 i n . ) and a span o f 91 c m (36 i n . ) and was equipped with both upper- and lower-surface or i f ices l o c a t e d 5 cm (2 i n . ) off t h e midspan. The a i r f o i l s u r f a c e was sanded i n t h e chordwise d i r e c t i o n w i t h N o . 400 d r y s i l i c o n c a r b i d e paper to p r o v i d e a smooth aerodynamic f i n i s h . The model contour accuracy was g e n e r a l l y w i t h i n kO.10 mm (0.004 i n . ) .

Wind Tunnel

The Langley Low-Turbulence P r e s s u r e Tunnel ( r e f . 4 ) is a c l o s e d - t h r o a t , s i n g l e - r e t u r n t u n n e l which can be operated a t s t a g n a t i o n p r e s s u r e s from 1.0 to 10.0 atm (1 atm = 101.3 kPa) w i t h tunnel-empty t e s t - s e c t i o n Mach numbers up to 0.42 and 0.22, r e s p e c t i v e l y . The maximum Reynolds number is about 49.0 x 106 per meter (15.0 x l o 6 per foot) a t a free-stream Mach number of about 0.22. The t u n n e l test s e c t i o n is 91 cm (3 f t ) wide and 229 cm (7.5 f t ) high.

H y d r a u l i c a l l y a c t u a t e d c i r c u l a r p la tes provided p o s i t i o n i n g and a t t a c h m e n t f o r t h e two-dimensional model. The plates are 102 cm (40 in . ) i n d i a m e t e r , rotate w i t h t h e a i r f o i l , and are f l u s h w i t h t h e t u n n e l w a l l . The a i r f o i l ends were a t t a c h e d to r e c t a n g u l a r model-attachment plates ( f i g . 41, and t h e a i r f o i l was mounted so t h a t t h e c e n t e r of r o t a t i o n f o r t h e c i r c u l a r p la tes was a t 0 . 2 5 ~ on t h e model r e f e r e n c e l i n e . The a i r gaps i n t h e t u n n e l w a l l s between t h e rect- angular p la tes and t h e c i r c u l a r p la tes were sealed w i t h metal seals.

Wake Survey R a k e

A f i x e d wake survey rake ( f i g . 5) a t t h e model midspan was mounted from t h e t u n n e l s i d e w a l l and l o c a t e d 1 chord l e n g t h behind t h e t r a i l i n g edge of t h e a i r f o i l . The wake rake used 0.15-cm (0.06-in.) d iameter t o t a l - p r e s s u r e t u b e s and 0.32-cm (0.125-in.) d iameter s t a t i c - p r e s s u r e tubes . The t o t a l - p r e s s u r e t u b e s were f l a t t e n e d to 0.10 c m (0.04 i n . ) f o r 0.61 c m (0.24 i n . ) from t h e t i p of t h e tube. Each s t a t i c - p r e s s u r e t u b e had f o u r f l u s h o r i f i c e s d r i l l e d 90° a p a r t ; t h e s e orifices were located 8 tube d i a m e t e r s from t h e t i p of t h e t u b e and i n t h e p l a n e of measurement f o r t h e t o t a l - p r e s s u r e t u b e s .

I n s t r u m e n t a t i o n

Measurements of t h e s t a t i c p r e s s u r e s on t h e a i r f o i l s u r f a c e s and t h e wake- rake p r e s s u r e s were made by a n a u t o m a t i c pressure-scanning system u s i n g v a r i a b l e - c a p a c i t a n c e p r e c i s i o n t r a n s d u c e r s . Basic t u n n e l p r e s s u r e s were measured w i t h

4

I

precision quartz manometers. digital shaft encoder operated by a pinion gear and rack attached to the cir- cular model-attachment plates. system and recorded on magnetic tape.

Angle of attack was measured with a calibrated

Data were obtained by a high-speed acquisition

TESTS AND METHODS

The airfoil was tested at free-stream Mach numbers from 0 . 1 0 to 0 . 3 2 over an angle-of-attack range from about -8O to 20°. airfoil chord was varied from about 2 . 0 x l o 6 to 1 2 . 0 x 106. tested both in the smooth condition (natural transition) and with roughness located on both upper and lower surfaces at 0 . 0 7 5 ~ . The roughness was sized for each Reynolds number according to the technique described in reference 5 . The roughness was sparsely distributed and consisted of granular-type strips 0 . 1 3 cm (0.05 in.) wide which were attached to the surfaces with clear lacquer.

Reynolds number based on the The airfoil was

The static-pressure measurements at the airfoil surface were reduced to standard pressure coefficients and machine integrated to obtain section normal- force and chord-force coefficients as well as section pitching-moment coeffi- cients about the quarter-chord point. Section profile-drag coefficients were computed from the wake-rake total and static pressures by the method reported in reference 6 .

An estimate of the standard low-speed wind-tunnel boundary corrections (ref. 7 ) amounted to a maximum of about 2 percent of the measured coefficients; these corrections have not been applied to the data. An estimate of the dis- placement of the effective center of a total-pressure tube in a velocity gra- dient on the values of Cd showed these effects to be negligible (ref. 6 ) .

PRESENTATION OF RESULTS

The test conditions are summarized in table 11. The results of this investigation have been reduced to coefficient form and are presented in the following figures:

Figure

Section characteristics for MS(1)-0317 airfoil . . . . . . . . . . . . . 6, 7 Effect of roughness on section characteristics . . . . . . . . . . . . . 8 Effect of Reynolds number on section characteristics; modelsmooth; M = 0 . 1 5 . . . . . . . . . . . . . . . . . . . . . . . 9

Effect of Reynolds number on section characteristics; roughness on; M = 0 . 1 5 . . . . . . . . . . . . . . . . . . . . . . . 10

Effect of Mach number on section characteristics;

Comparison of section characteristics for LS(1)-0417

Comparison of section characteristics for MS(1)-0313

roughness on; R = 6 . 0 x l o 6 . . . . . . . . . . . . . . . . . . . . . 1 1

and MS(1)-0317 airfoils; roughness on; M = 0 . 1 5 . . . . . . . . . . . 12

and MS(1)-0317 airfoils; roughness on; M = 0 . 1 5 . . . . . . . . . . . 13

5

Figure

Effect of angle of attack and Reynolds number on chordwise pressure distributions for MS ( 1 ) -031 7 airfoil; roughness on; M = 0.15 . . . . . . . . . distributions for MS(1)-0317 airfoil; roughness on; R = 6 . 0 x l o 6 . . . . . . . . . . . . . . . . .

Comparison of chordwise pressure distributions for LS(1)-0417 and MS(1)-0317 airfoils; roughness on; M = 0.15; R = 4 . 0 x l o 6 . . . . . . . . . . . . . MS ( 1 ) -031 3 and MS ( 1 1-031 7 airfoils; roughness on;

Variation of maximum lift coefficient with Reynolds number for LS(1)-0417 and MS(1)-0317 airfoils; M = 0 . 1 5 . . . . . . . . . . . . . . . . . . . . . .

Variation of maximum lift coefficient with Reynolds number for MS(1)-0313 and MS(1)-0317 airfoils; M = 0 . 1 5 . . . . . . . . . . . . . . . . . . . . . .

number for MS ( 1 ) -031 3 and MS ( 1 ) -031 7 airfoils; roughness on; R = 6 . 0 x 106 . . . . . . . . . . . .

Variation of drag coefficient with Reynolds number for MS (1) -0317 airfoil; M = 0 .15; c2 = 0 . 3 0 . . . . .

Calculated drag-rise characteristics for medium-speed airfoils; R = 1 4 . 0 x l o 6 ; c2 = 0 . 3 0 . . . . . . .

Effect of Mach number on chordwise pressure

Camparison of chordwise pressure distributions for

I M = 0 .15; R = 4 . 0 x I O 6 . . . . . . . . . . . . .

Variation of maximum lift coefficient with Mach

. . . . . . . . . 14

. . . . . . . . . 15

. . . . . . . . . 16

. . . . . . . . . 17

. . . . . . . . . 18

. . . . . . . . . 19

. . . . . . . . . 20

. . . . . . . . . 21

. . . . . . . . . 22

DISCUSSION OF RESULTS

Section Characteristics

Lift.- Figure 9(a) shows that the lift-curve slope for the 17-percent , medium-speed airfoil in a smooth condition (natural boundary-layer transi-

tion) varied from about 0.11 to 0 .12 per degree for the Reynolds numbers investigated (M = 0 . 1 5 ) . The angle of attack for zero lift coefficient was about -3O. Maximum lift coefficients increased from about 1 . 6 0 to 2 . 0 as the Reynolds number was increased from 2 . 0 x l o 6 to 1 2 . 0 x l o 6 . effect of Reynolds number on maximum lift coefficient occurred for Reynolds numbers below 6 . 0 x l o 6 . The stall characteristics of the airfoil are of the trailing-edge type, as shown by the lift data of figure 9(a) and the pressure data of figure 14. The nature of the stall is docile for all Reynolds numbers tested.

The largest

The addition of a narrow roughness strip at 0 . 0 7 5 ~ (fig. 8 ) resulted in the expected decambering effect because of the increase in boundary-layer thick- ness. The lift coefficient at c1 = Oo decreased about 0 . 0 4 at the lower Reynolds numbers, but only small changes occurred at the higher Reynolds numbers. The roughness strip decreased the 0 .04 for the test Reynolds number range (fig. 19).

c2,max performance of the airfoil as much as

The effects of Mach number on the airfoil lift characteristics at a Reynolds number of 6 . 0 x 706 with roughness located at 0.075~ are shown in fig- ure ll(a). Increasing the Mach number from 0.10 to 0.32 resulted in the expected Prandtl-Glauert increase in lift-curve slope, a decrease in the angle of attack for a stall of about 2.S0, and a decrease in c2,max of about 0.03.

The lift data for the 17-percent-thick low- and medium-speed airfoils are compared in figure 12 for Reynolds numbers from 2.0 x lo6 to 6 . 0 x lo6 and are summarized in figure 18. The design lift coefficients for the low-speed and medium-speed airfoils were 0.40 and 0.30, respectively. The data indicate that the linearity of the lift curve is extended to higher angles of attack for the medium-speed airfoil and that both airfoils develop about the same czImaX at the lower Reynolds numbers. This result is attributed to reduced upper- surface boundary-layer separation for the medium-speed airfoil, as illustrated by the pressure-data comparison of figure 16(b). At the higher Reynolds num- bers (fig. 18), a decrease in CZ,max of about 0.06 is shown for the medium- speed airfoil compared with the low-speed airfoil for airfoils without roughness. Figure 18 also illustrates two interesting features of the Reynolds number effect on czfmax The irregular variation of numbers and the sensitivity of have been improved for the medium-speed airfoil design.

for the 17-percent low- and mediumspeed airfoils. ~2,max with Reynolds number at the lower Reynolds

to roughness for the low-speed airfoil climax

The lift data for the 13-percent-thick and 17-percent-thick medium-speed airfoils are compared in figure 13 for Reynolds numbers from 2.0 x l o 6 to 12.0 x lo6 and are summarized in figures 19 and 20. of increasing the airfoil thickness is to decrease c2,max about 0.10 at Reynolds numbers from 2.0 x lo6 to 6 . 0 x I O 6 . At the higher Reynolds numbers, only small effects of airfoil thickness on Figure 19 also indicates that the sensitivity of what greater for the thicker airfoil. c~,max for both airfoils are shown in figure 20 for a Reynolds number of 6 . 0 x 106. Increasing the Mach number results in similar decreases in €or both airfoils up to about M = 0.28. However, above M = 0.28, the 13-percent air- foil indicates a larger decrease in airfoil.

The predominant effect

~2,max performance are shown. CZ,max

The effects of Mach number on to roughness is some-

CZ,,,~~

clImax, compared with the 17-percent

Pitching moment.- The pitching-moment-coefficient data of figures 8, 9, and 10 illustrate the expected positive increments in Cm due to decreasing the Reynolds number or adding roughness at a constant Reynolds number. This result is typical of the decambering effect associated with boundary-layer thickening for aft-loaded airfoils. increasing the Mach number from 0.10 to 0.32 (fig. ll(c)) shows small effects on the pitching-moment data to about a positive increment in q,, is shown.

At a Reynolds number of 6 .0 x lo6,

a = 8O. At the higher angles of attack,

The pitching-moment data for the 17-percent-thick low- and medium-speed airfoils are compared in figure 12. A reduction in the magnitude of Cm of about 25 percent throughout the c2 range is indicated for the medium-speed airfoil. This result is important because of the expected reduced trim penalties for the medium-speed airfoil at cruise conditions. Comparison of

7

the data for the 13-percent-thick and 17-percent-thick medium-speed airfoils in figure 13 shows essentially no effect of thickness on the pitching-moment characteristics.

Draq.- The design pressure distribution for the mediuwspeed airfoil (fig. 2) shows that a favorable pressure gradient exists only back to about 0.10~ on the upper and lower surfaces at a Mach number of 0.68. The low-speed (M = 0.15) pressure data (fig. 14) show that a pressure peak develops at about 0.06~ on the upper surface of the airfoil at a lift coefficient of about 0.30. Thus, the pressure distributions are not conducive to long runs of laminar flow. Since natural transition usually occurs near the leading edge of airfoils for general aviation aircraft due to roughness of construction or insect remains gathered in flight, the discussion of the drag data is limited to data obtained with fixed transition at 0.075~.

The profile-drag coefficient at design lift (ci = 0.30) decreased from about 0.0116 at R = 2.0 x lo6 to about 0.0090 at R = 12.0 x lo6. (See fig. 10(b) and fig. 21.) This drag reduction is associated with the related decrease in boundary-layer thickness and the accompanying reduction in skin- f r i c t i o n drag. There a r e o n l y small effects of Mach number on Cd ( f i g . 11(b)) over a Mach number range from 0.10 to 0.32.

The drag data for the 17-percent-thick low- and mediuwspeed airfoils are compared in figure 12 for Reynolds numbers from 2.0 x lo6 to 6 . 0 x lo6 with fixed transition at 0.075~. A decrease in drag coefficient at all lift coeffi- cients is shown for the medium-speed airfoil. The small decrease in drag coef- ficients for the medium-speed airfoil at low lift coefficients is associated with the reduced aft upper-surface pressure gradient (fig. 16(a)) and resulting boundary-layer development. The large decrease in drag coefficients at the higher lift coefficients for the medium-speed airfoil is a result of less sepa- ration on the airfoil, as illustrated in figure 16(b).

, I

The drag data for the 13-percent-thick and 17-percent-thick medium-speed

Increasing the airfoil thickness airfoils are compared in figure 13 for Reynolds numbers from 2.0 x lo6 to 12.0 x lo6 with fixed transition at 0.075~. results in the expected increase in drag coefficient throughout the lift coef- ficient range. At the design lift coefficient of 0.30, increases in Cd of about 0.0015 (R = 2.0 x lo6) and 0.0008 (R = 12.0 x lo6) are indicated by increasing the airfoil thickness from 13 to 17 percent.

Theoretically calculated drag-rise characteristics (ref. 3) for the 13- and 17-percent-thick medium-speed airfoils at design conditions are shown in figure 22. Boundary-layer transition was specified at x/c = 0.04 for the calculations to ensure a turbulent boundary-layer development on the airfoils. The estimated drag-rise Mach numbers are about 0.76 and 0.72 for the 13- and 17-percent airfoils, respectively.

Pressure Distributions

The chordwise pressure data of figure 14 illustrate the effects of angle of attack for several Reynolds numbers. As the angle of attack is increased,

8

upper-surface trailing-edge separation is first indicated by the approximate constant-pressure region on the airfoil. Additional increases in angle of attack result in this constant-pressure region moving forward along the air- foil. At maximum lift, trailing-edge separation is present over approximately 20 to 30 percent of the airfoil chord, depending on the Reynolds number. The airfoil stall is of the trailing-edge type, and the stall characteristics are docile at all Reynolds numbers.

The effects of Mach number on the chordwise pressure data at a Reynolds number of 6.0 x lo6 for angles of attack of 8O and 14O are illustrated in figure 15. Increasing the Mach number from 0.10 to 0.32 results in the expected Prandtl-Glauert increase in the value of Cp at CY = 8 O (fig. 15 (a) ) . However, at CY = 14O (fig. 15(b)), this same Mach number increase results in an increase in the extent of upper-surface trailing-edge separation of about 0.05~. This result is attributed to the increased upper-surface pressure gra- dient, which has an adverse effect on the resulting boundary-layer development.

Comparisons of the pressure data for the 17-percent-thick low- and medium- speed airfoils at a Mach number of 0.15 and a Reynolds number of 4.0 x lo6 are shown in figure 16. Note the substantial decrease in the aft upper-surface pressure gradient for the medium-speed airfoil (fig. 16(a)). This reduced pres- sure gradient has a favorable effect on the airfoil boundary-layer development (reduced thickness) and results in a small decrease in drag coefficient at low lift coefficients (see fig. 12). At the higher lift coefficients, this reduced pressure gradient decreases the amount of upper-surf ace trailing-edge separation for the medium-speed airfoil. For example, at a lift coefficient of 1.60 (fig. 16(b)), the medium-speed airfoil exhibits about 0.10~ less separation than the low-speed airfoil.

The pressure data for the 13-percent-thick and 17-percent-thick medium- speed airfoils at a Mach number of 0.15 and a Reynolds number of 4 . 0 x lo6 are compared in figure 17. The pressure data at CC = Oo (fig. 17(a)) illus- trate the increase in upper- and lower-surface velocities due to increased air- foil thickness. Note that the aft upper-surface pressure gradient is about the same for the two airfoils. The effect of thickness on the extent of upper- surface trailing-edge separation is illustrated at CY = 16O in figure 17(c). The 13-percent medium-speed airfoil exhibits about 0.15~ less separation com- pared with the 17-percent airfoil.

CONCLUDING REMARKS

Wind-tunnel tests have been conducted to determine the low-speed two- dimensional aerodynamic characteristics of a 17-percent-thick medium-speed air- foil designed for general aviation applications. The results were compared with those for the 17-percent-thick low-speed airfoil and the 13-percent-thick medium-speed airfoil. Theoretical predictions of the drag-rise characteristics for this airfoil are also provided. The tests were conducted in the Langley Low-Turbulence Pressure Tunnel over a Mach number range from 0.10 to 0.32. The chord Reynolds number was varied from about 2 . 0 x lo6 to 12.0 x lo6. following results were determined from this investigation:

The

9

1. Maximum section lift coefficients at a Mach number of 0.15 increased from about 1.6 to 2.0 as the Reynolds number was increased from about 2.0 x lo6 to 12.0 x 106.

decreased about 25 percent, and the drag coefficient decreased at all lift coef- ficients (fixed transition) for the 17-percent-thick mediuwspeed airfoil com- pared with the 17-percent-thick low-speed airfoil.

6. The predominant effects of increasing airfoil thickness from 13 percent

2. Stall characteristics were of the trailing-edge type and were docile at all Reynolds numbers.

3 . The application of a roughness strip near the leading edge of the air- foil decreased the maximum section lift coefficient as much as 0.04 over the test Reynolds number range.

4. Increasing the Mach number from 0.10 to 0.32 at a constant Reynolds num- ber of about 6.0 x lo6 decreased the maximum section lift coefficient about 0.03.

Langley Research Center National Aeronautics and Space Administration Hampton, VA 23665 November 25, 1980

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I REFERENCES I 1. McGhee, Robert J.; Beasley, William D.; and Whitcomb, Richard T.: NASA Low-

! and Medium-Speed Airfoil Development. NASA TM-78709, 1979. I

1 1

2. McGhee, Robert J.; and Beasley, William D.: Low-Speed Aerodynamic Character- istics of a 13-Percent-Thick Medium-Speed Airfoil Designed for General Aviation Applications. NASA TP-1498, 1979.

I

3. Bauer , Frances; Garabedian, Paul; Korn, David; and Jameson, Antony: Super- critical Wing Sections 11. Mathematical Systems, Springer-Verlag, 1975.

Volume 108 of Lecture Notes in Economics and

4 . Von Doenhoff, Albert E.; and Abbott, Frank T., Jr.: The Langley Two- Dimensional Law-Turbulence Pressure Tunnel. NACA TN 1283, 1947.

1 5 . Braslow, Albert L.; and Knox, Eugene C.: Simplified Method for Determination of Critical Height of Distributed Roughness Particles for Boundary-Layer

i Transition at Mach Numbers From 0 to 5. NACA TN 4363, 1958.

6. Pankhurst, R. C.; and Holder, D. W.: Wind-Tunnel Technique, Sir Isaac I Pitman & Sons, Ltd. (London), 1965. ' 7. Pope, Alan; and Harper, John J.: Low-Speed Wind Tunnel Testing. John Wiley

& Sons, Inc., c.1966.

I

t

11

TABLE I .- MS (1 ) -031 7 AIRFOIL COORDINATES

0.00000 .00200 .00500 .01250 .02500 .03750 .05000 .07500 .10000 .12500 .15000 .17500 .20000 .22500 .25000 .27500 .30000 .32500 .35000 .37500 .40000 .42500 .45000 .47500 .50000 .52500 .55000 .57500 .60000 .62500 ,65000 .67500 .70000 .72500 .75000 .77500 .80000 .82500 .85000 ,87500 .90000 .92500 .95000 .97500

1 .ooooo

0.00099 .01248 .01950 .03099 .04322 .05210 .05893 .06840 .07511 .08033 .08454 .08805 .09096 .09339 .09536 .09694 .09815 .09901 .09952 .09972 .09956 .09909 .09826 .09700 .09535 .09323 .09073 .08777 .08448 .08079 .07672 .07232 .06763 .06269 .05755 .05225 .04687 .04132 .03576 .03013 .02444 .01873 .01302 .00720 .00125

z/c I lower

0.00099 -. 00857 -. 01 366 -.02105 -. 02866 -.03423 -. 03865 -. 04541 -. 05058 -. 05477 - .05817 -. 06099 -.06330 -. 06527 -. 06685 -. 0681 2 -. 06909 -. 06978 -.07021 -. 07036 -. 0701 9 -. 06967 -. 06880 -. 06755 -. 06591 -. 06389 -. 061 38 -. 05845 -. 05501 -. 051 06 -. 04674 -. 0421 4 -.03735 -. 03255 -. 02780 -. 02309 -. 0 1857 -. 01 433 -.01049 -. 0071 9 -. 00460 -. 00289 -.00232 -. 00324 -. 00597

12

TABLE 11.- TEST CONDITIONS

2 x 106 4 x 106 6 x 106 9 x l o 6

X X X X

X

X X X X

X

X

X

1 I 1

12 x l o 6

X

X

M

0.15 .10 .15 .20 .28 .32

I R I Configuration

Smooth Roughness on Roughness on Roughness on Roughness on Roughness on

13

N + 0 w a a

0 4

00

P

CD

2

m

N

0

I I < N

14

I -1

I 1 i

Ai rfoi I - MS ( I 1-03 I3 -- MS ( I 1-03 I 7

M

0.72 0.68

Figure 2.- Calculated pressure distributions for medium-speed airfoils. R = 1 4 . 0 x l o 6 ; c1 = 0.30.

15

.OE

-06 zt/c

.04

.02

0

-.o':, I I .5 I .o x /C

Figure 3.- Thickness distribution and camber line for MS(1)-0317 airfoil.

16

Tunnel sidewalls /- t+-- Diam,= 1.67~ -

d

Airflow t- A

-Circular plate -

I

r - -- -. - - - - - - -TT- -

'r 1. \\\ \\

Airfoil positioning attachment

Top view /_Model attachment plate

Zero incidence reference

Seal detail "z"

Tunnel center line

End view .section A-A

F i g u r e 4.- Typical a i r f o i l model mounted i n wind t u n n e l . c = 61 c m (24 i n . ) .

17

.126c Static- pressure probe

Rad.=0.021c

3.A

. 0 4 2 c

I Static - pressure probes -.,

lu/ - -

-4

(typ.1

.0052 c

Airflow - Tunnel $ - -

Total- pressure probes

(tubes flattened)

1.17c

F i g u r e 5.- Wake survey rake. c = 61 cm ( 2 4 in.).

18

in 0 Iil . .

‘0 u

. . I -

1 c u c u - -

V”

8 u U

I I I I I

E u

In r

0

II

E

-4

m c u

19

. . . . c u r u - - I -

I

u”

TJ Q, 3 c -4 4J c 0 U I

W

a,

m -4 E

Y

in 0 lil el lil 0 L n ’ l 0 0 - - r u r u

I I I I I

E u

20

I

. - 0 0 0 0

-0 u

1 0’ I

I I I I I I -

I

E u

W 0 c

X

0

W

II

p:

h u Y

21

22

8: U

II

a

al Y

23

m u

0 ru

in - 0

Lo

0

in I

0 ..-.

x: a' D

W 0 P

X

0

r*1

II

p:

.. In

0

II

E

7

n m Y

LQ 0 Li, 0 L i , d l 'in 0 Li, 0 Li, cj m'r . . 0 0 - - nr ni - - N a -.. .- I -

I I I I I I

E u" u

u In F 0

0

JJ 4

ro aJ JJ ([I

rl

m m aJ c c P 3

0"

2 rl -4

W L4 -4 4

I-

el 0 I

0

?

n 7 v

v1 E

8 W

m 0 -4 JJ m -4 L! a, L, u 4 L! m c u c 0 -4 JJ u aJ v1

I

F

aJ L4 3 tn .d E

24

. . X

m o m (

U V

0 cu

m -

0 -

m

0

m I

0 - . .

c u m - - u-

c

li, 0 li-l 0 ’ 1 0 . .

I - I

m o m o m

m I

0 - m l cu 0 - - ;

I I I I t

E V

Ll II 5

P f: -4

F

25

u-

W 0

r l h o l i , . . C U C U - -

u”

I li, d l . . I -

I I I I l l

E u

26

. . U u

X a O a , * 7

C .r(

u-

27

in 0 li3 0 Li, 0 Lo d l . . . . r u m - - I -

I

W 0 P

28

u-

r u r u - - u”

U u

I 0 1 I .

I - I I I I I I

E u

29

. . U

V

k- c t

lil 0 lil 0 L n ' l o - - c u c u I I I I I

E V

w 0

30

‘ m 0 h 0 I . . I u r u

V”

U V

0 cu

Lo d

0

m

0

m I

0 &

I 0 h d l . . I -

I

W 0 7

X

0

OI

I t

a .. v)

0

It

c

c - 5

31

. .

. . r u r u - -

3”

U V

. . 0 0 - - 1

4

4

X

0 % . a

II c 0 a u

‘ a , O L C

I I I I I

E V

32

33

0 d

0 0 . .

34

I

1 I I 1 I I I I 1 I

E V

a Q) 3 C -4 JJ C 0 V

I. 03

Q, & 3 m -4 E

35

W 0 F

X

0

w II

a

5 al 3 c

36

t

rm

a e, 3 d

.r( U

u 8

W 0 r

X

0

w II

a A

-Q,

a e, 3 C -4 U C 0 u I. 03

e, LI 3 P -4 ti4

e r 00 0 0

t I- 3

3 3 . .

I

37 I

\D 0 7

X

0

cjr

It

a

38

0 A

m

CD

f

cu

0

cu I

f I

CD I

m I

0 - I I I

W 0 F

39

rm Tm rm

li I -

I c ~

1

i - C

u I I I

-Lu

c

c r

LLLL

1 - - ( c 1 0 ( 0 1

40

.A

W 0 F

X

0

W

II

a

I

03

0) LI 5 P

.A

41

I h 0 h 0 h 0 L i ' l ru m r - -

r u r u r u - - - - I I 1

42

43

W 0 r

X

0

m II

a

44

c u c u (U - - - - I I I

45

t E Z C 0 0

-LLLLJu

o m 0 0 oor

I-

I l l I l l

. . U u

rm rm

I

0 0 0

46

m o Lo 0 m 0 m 0 Lo 0 Lo 01

0 0 0 0 r 0 c u m w Tu c u m m - 0 cu

d - cu cu Lo r

I I I I I I I I I I

E u

47

00000

m r u o r m r u o r t n r u r u m r

u-

r u r u r u - - - - I I 1

II

E

.. c 4J

4J - u m a , - m

m 5 4 0 c h (u a

U-l 0

48

o m ( 0 0 0 o r i i

ui CD 0

U r

o c - - < O O ( . .

49

0 0 0 - d I 4 cu I I 1 I I I I I 1 I

E u

50

0

U m TY a, U

I

m O

I

0

a,

r

51

'r I;

. .

fu aJ 5 c -4 c, c 0 u I

0

aJ LI 5 07

GI

P

-4

- - 0 0 0 0

I

52

.. u ln I- O

0

JJ lu

rn u

JJW u o l u r

'D w l u x 8 : ; Lc

ti J u o -4 c

Lc

4 c

54

. .

mr

0 ru 0

0 Ln 0' ; - e o 0 0 0

55

E u

56

E 0

W 0 7

X

0

hl

II

P;

m d

5 c m

g o u-l II

m c -rl 0

.i . . U u a, m

W 0

57

0

- I'

rr) 9

cu 4

- 9

0

d cu

............... - .......... -. . ,.... I__ . . . . .rrt.??b.. L - ..!.. ... I . . .. frrrrt.. ...... -f._... r r : r r t z :k:: I : :: :!::. .

a,

d-

0

I... .-. . d

I

a, I

E u

58

E 0

a Q) a 3

I l l P . L l l ' l i l l l l l l l l l 1-r 1 - E I U

N I

W 0

Ln 0

d 0

W 0 P

X

0

CJ

I1

&

tu Y

60

W 0 r

X

0

II

cz h e

61

U cu

(D

E 0

62

E u

5

63

E 0

73 0

W 0 r

X

0

CJ

II

p:

r

Y

a 0) a 1 rl u C 0 u I

m 7

al LI 1 m -I+

E

64

L I

-

L

' o h 0 I

6 5

66

W 0

U f - 1 x x

0

u) 0 F

X

0

* II

a

I

* r-

68

-- t I

u l m m f f M M m n l - - I I I I I I I I I I I I

0

69

70

. . I W - - I

I I I

72

13

14

I O 0

I

N

I

00 I

o?

00

(4

N

m m 0

It

o ru

m Y

x /c

(b) c z = 1.60.

F i g u r e 16.- Concluded.

76

I

I

O n I-

cc! I

T 0 I Q 0

5,

09

co

m

-2.8

-2.4

-2.0

- 1.6

- I .2

‘P -.8

-.4

0

.4

.8

1.2 0

x/c

(b) c1 = 6O.

F i g u r e 17.- Cont inued.

78

C

(c) CC = 16O.

Figure 17.- Concluded.

79

2.4

2.2 MS(11-0317

PI ain sym., roughness off FI agged sym., roughness on

2 .o

I .8 CZ,max

I .6

I .4

1.2

I 2 4 6 8 IO I .o IxIO6 R

Figure 18.- Variation of maximum lift coefficient with Reynolds number for LS(1)-0417 and MS(1)-0317 airfoils. M = 0.15.

80

2.4

O I

2.2

2 .o

I .8 CI,max

1.6

I .4

1.2

Figure

I.

19

2

.- Variation of max MS (1 ) -031 3 and

imum m5(1

4 6 R

lift coefficient ) -0317 airfoils.

8 1

with M =

0

Reynolds 0.15.

20x

number

106

for

81

N 0

82

.040

,020

. 010

.om

.OM

.OM

.002

1 . 001 I I

I 2 4 6 8 IO :ox IO6

Roughness Off On

--

R Figure 21.- Variation of drag coefficient with Reynolds number for

MS(1)-0317 airfoil. M = 0.15; cl = 0.30.

a3

M M 4

7 h

4

\ 4

\ 8 A

4 Y

I v, E

l

1 1 I 1 I 1

2 a

D u

8 0

8 0

a

a 4

1. Report No. 2. Government Accession NO. NASA TP-1786

4. Title and Subtitle

LOW-SPEED AERODYNAMIC CHARACTERISTICS OF A 17-PERCENT- THICK MEDIUM-SPEED AIRFOIL DESIGNED FOR GENERAL AVIATION APPLICATIONS

7. Author(s1

Robert J. McGhee and W i l l i a m D. Beasley ,

9. Performing Organization Name and Address

NASA Langley Research Center Hampton, VA 23665

3. Recipient's Catalog No.

5. Repon Date December 1980

6. Performing Organization Code

505-31 -33-05 8. Performing Organization Report No.

- 0 10. Work Unit No.

1 1 , Contract or Grant No. I 12. Sponsoring Agmcy Name and Address

N a t i o n a l Aeronaut ics and Space Adminis t ra t ion Washington, DC 20546

L 13. Type of Report and Period Covered

Technica l Paper

14. Sponsoring Agency Code

- 16. Abstract

Wind-tunnel tests have been conducted to determine t h e low-speed two-dimensional aerodynamic c h a r a c t e r istics of a 17-percent- thick medium-speed a i r f o i l (MS( 1 ) -031 7) designed f o r g e n e r a l a v i a t i o n a p p l i c a t i o n s . The resul ts were compared w i t h d a t a for t h e 17-percent- thick low-speed a i r f o i l (LS(1) -041 7) and t h e 13-percent- thick medium-speed a i r fo i l (MS(1) -031 3 ) . T h e o r e t i c a l p r e d i c t i o n s of t h e d r a g - r i s e char- acteristics of t h i s a i r f o i l are also provided. The tests were conducted i n t h e Langley Low-Turbulence P r e s s u r e Tunnel over a Mach number range from 0.10 to 0.32, a chord Reynolds number range from 2.0 x IO6 to 12.0 x IO6, and an angle-of-at tack range from about -8O to 20°.

17. Key Words (Suggested by Author($))

Low-speed c h a r a c t e r istics Medium-speed a i r fo i l Reynolds number e f f e c t s Mach number e f f e c t s General av i a t ion a i r c r a f t

18. Distribution Statement

FEDD D i s t r i b u t i o n

S u b j e c t Cateqory 02

Avai lab le : NASA's I n d u s t r i a l A p p l i c a t i o n s C e n t e r s

19. Security Clauif. (of this report) 20. Security Clauif. (of this page)

. .

NASA-Langley, 1980

~ ~~

21. No. of Pages

84

22. Price