low-speed aerodynamic characteristics of a 17 percent thick … · 2013-08-31 · low-speed...

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I , NASA TN D-7428

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L O W - S P E E D A E R O D Y N A M I C C H A R A C T E R I S T I C S O F A 1 7 - P E R C E N T - T H I C K A I R F O I L S E C T I O N D E S I G N E D 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

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w 4. Title uxl Subtitle D-7428 LOW-SPEED AERODYNAMIC CHARACTERISTICS OF A 17-PERCENT-THICK AIRFOIL SECTION DESIGNED FOR GENE Ff&- A-WTION APPLICATIONS ~ . . . . . .

7. Authodt)

Robert J. McGhee and William D. Beasley ___._ - .. - . ~ ~

9. Worming Orpization Name nnd Addrsl

NASA Langley Research Center Hampton, Va. 23665

~ .. _--_____I.-- . __ ~~

12. Sponsoring Agency Name and Addrarr

National Aeronautics and Space Administration Washington, D.C. 20546 -

~ . ______. ~_ . ~ _ _ _ _ 1s. s u w m riotes

. . . . .. - ~ - - 6. R e p a t Dam

December 1973 8. F'orkrning Dmniniim cod.

8. Worming Wniut ion R.pwt

L-9132 ~ __ __ _-

10. Work Unit No.

501-06-05-02 _ _ 11. C o n t m or Gnnt No.

.___ _..___m____ _J_E ~ - _-__I -.-

16. Abrtnct

Wind-tunnel tests have been conducted to determine the low-speed two-dimensional

The results were compared with predictions based on a theoret- aerodynamic characteristics of a 17-percent-thick airfoil designed for general aviation applications (GA(W)-l). ical method for calculating the viscous flow about the airfoil. The tests were conducted over a Mach number range from 0.10 to 0.28. Reynolds numbers based on airfoil chord varied from 2.0 x lo6 to 20.0 x lo6. Maximum section lift coefficients greater than 2.0 were obtained and section lift-drag ratio at a lift coefficient of 1.0 (climb condition) varied from about 65 to 85 as the Reynolds number increased from about 2.0 X lo6 to

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17. Key Words (Suggatsd by Author(r1)

Low-speed airfoil section Reynolds number effects Experimental-theoretical comparison GA(W)-l airfoil General aviation aircraft

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Unclassified I I . -I_L

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*For ah by tho Notion01 Tochniclll Information Service,Springfiold, VirOinio 22161

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LOW-SPEED AERODYNAMIC CHARACTERISTICS OF A

17 -PERCENT - THICK AIRFOIL SECTION DE SIGNED

FOR GENERAL AVIATION APPLICATIONS

By Robert J. McGhee and William D. Beasley Langley Research Center

SUMMARY

An investigation w a s conducted in the Langley low-turbulence pressure tunnel to determine the low-speed two-dimensional aerodynamic characteristics of a 17-percent- thick airfoil designed for general aviation applications. The results a r e compared with a typical older NACA 65 se r i e s airfoil section. Also, a comparison between experimental data and predictions, based on a theoretical method for calculating the viscous flow about the airfoil, is presented. The tests were conducted over a Mach number range from 0.10 to 0.28 and an angle-of-attack range from - loo to 24'. Reynolds numbers, based on the airfoil chord, were varied from about 2.0 X lo6 to 20.0 X lo6.

The results of the investigation indicate that maximum section lift coefficients increased rapidly at Reynolds numbers f rom about 2.0 X lo6 to 6.0 X lo6 and attained values greater than 2.0 for the plain airfoil and greater than 3.0 with a 20-percent-chord split f lap deflected 60°. Stall characterist ics were generally gradual and of the trailing- edge type either with o r without the split flap. the section lift-drag ratio increased from about 65 to 85 as the Reynolds number increased from about 2.0 X 106 to 6.0 X lo6. Maximum section lift coefficients were about 30 percent greater than that of a typical older NACA 65 s e r i e s airfoil section and the section lift-drag ratio at a lift coefficient of 0.90 w a s about 50 percent greater. Agreement of experimental results with predictions based on a theoretical method which included viscous effects w a s good for the pressure distributions as long as no boundary-layer flow separation was present, but the theoretical method predicted drag values greatly in excess of the measured values.

At a lift coefficient of 1.0 (climb condition)

; Research

INTRODUCTION

on advanced aerodynamic technology airfoils has been conducted over the last several years at the Langley Research Center. Results of this research have been applied to the design of a 17-percent-thick airfoil suitable for a propeller driven light airplane.

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I -

The subcritical characterist ics of thick supercritical airfoil section research of Some reference 1 indicated performance increases over conventional airfoil sections.

of the features that produce these favorable aerodynamic character is t ics have been applied in the design of a new low-speed airfoil section. This new airfoil is one of several being developed by NASA for light airplanes and has been designated as General Aviation (Whitcomb) -number one airfoil (GA(W)- 1).

dimensional aerodynamic character'istics of the NASA GA(W)- 1 airfoil section. In addition, the results are compared to a comparable NACA 65 series airfoil section. Such sections are presently used on some light airplanes. Also, the experimental resu l t s are compared with results obtained f rom an analytical aerodynamic performance prediction method.

The present investigation was conducted to determine the basic low-speed two-

The investigation was performed in the Langley low-turbulence pressure tunnel over a Mach number range from 0.10 to 0.28. The Reynolds number, based on airfoil chord, varied from about 2.0 X lo6 to 20.0 X lo6. The geometrical angle of attack varied f rom about -loo to 24O.

SYMBOLS

Values a r e given both in SI and the U.S. Customary Units. The measurements and calculations were made in the U.S. Customary Units.

a mean- line designation

cP PL - pco

qco pressure coefficient,

C chord of airfoil, cm (in.)

section chord-force coefficient, CC

max s forward (t/c),,

Cd section profile-drag coefficient determined from wake measurements,

2

'd'

C l

c2, i

cn

h

l/d

M

P

R

t

X

z

CY

P

point drag coefficient,

section lift coefficient,

design section lift coefficient

Cn cos CY - cc sin CY

section pitching-moment coefficient about quarter chord,

Cp d($) - Cp d(5) L.S. U.S.

section normal-force coefficient,

vertical distance in wake profile, cm (in.)

section lift-drag ratio, Cl/Cd

free-s t ream Mach number

s tatfc pr e s s u r e, N/ m2

dynamic pressure, N/m2 (lb/ft2)

(lb/f t2)

Reynolds number based on free-stream conditions and airfoil chord

airfoil thickness, cm (in.)

airfoil abscissa, cm (in.)

airfoil ordinate, cm (in.)

mean line ordinate, c m (in.)

angle of attack of airfoil, angle between chord line and a i r s t r eam axis, deg

density, kg/m3 (slugs/ft3)

3

Ill1 II I l l l l

Subscripts:

L local point on airfoil

max maximum

t thickness

1 tunnel station 1 chord length downstream of model

2 tunnel station downstream of model where static pressure is equal to f ree-s t ream static p re s su re

m undisturbed s t ream conditions

Abbreviations:

GA(W)- 1 General Aviation (Whitcomb)-number one

1. s. lower surface

U.B. upper surface

AIRFOIL DESIGN

The airfoil section (fig. 1) w a s developed by employing some of the favorable characterist ics of the thick supercritical airfoil of reference 1, which indicated performance increases over conventional airfoils at subcritical conditions. In order to expedite the airfoil development, the computer program of reference 2 was used to predict the resuits of various design modifications. The final airfoil shape w a s defined after 17 iterations on the computer. The airfoil is 17 percent thick with a blunt nose and a cusped lower sur - face near the trailing edge. The design cruise lift coefficient was about 0.40 a t a Reynolds number of about 6 X 10 6 . In defining the airfoil emphasis w a s placed on providing good

lift-drag ratios a t cl = 1.0 for improved climb performance, and on providing a maximum lift coefficient of about 2.0. Several key design features of the airfoil are:

1. A large upper surface leading-edge radius (about 0.06~) was used to attenuate the peak negative pressure coefficients and therefore delay airfoil stall to high angles of attack.

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2. The airfoil was contoured to provide an approximate uniform chordwise load distribution near the design lift coefficient of 0.40. To account fo r viscous effects this airfoil incorporated more camber in the rear of the airfoil than the NACA mean camber line (fig. 2).

3. A blunt trailing edge was provided with the upper and lower surface slopes approximately equal to moderate the upper surface pressure recovery and thus postpune the stall.

The airfoil thickness distribution and camber line are presented in figure 2. Table I presents the measured airfoil coordinates.

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APPARATUS AND PROCEDURE

Model Description

The airfoil model was machined from an aluminum billet and had a chord of 58.42 c m (23 in.) and a span of 91.44 c m (36 in.). The airfoil surface was fair and smooth. upper and lower surface orifices located a t the chord stations indicated in table II. A base pressure orifice was included in the blunt trailing edge of the airfoil (x/c = 1.0). In order to provide data for a simple flap deflection, an aluminum wedge was installed on the model to simulate a split flap deflected 60'. Orifices were installed on this simulated flap as indicated in table II.

Figure 3 shows a photograph of the model. The model was equipped with both

I Wind Tunnel

The Langley low-turbulence p res su re tunnel (ref. 3) is a closed-throat single- 2 return tunnel which can be operated a t stagnation pressures from 101.3 to 1013 kN/m

(1 to 10 atm) with tunnel-empty test-section Mach numbers up to 0.46 and 0.23, respec- tively. The maximum unit Reynolds number is about 49 X 106 per meter (15 X 106 p e r foot) a t a Mach number of 0.23. (7.5 ft) high.

The test section is 91.44 cm (3 f t ) wide by 228.6 cm

Circular end plates provided attachment fo r the two-dimensional model. The end plates are 101.6 cm (40 in.) in diameter and a r e flush with the tunnel wall. They are hydraulically rotated to provide for model angle-of-attack changes. The airfoil was mounted so that the center of rotation of the circular plates w a s a t 0 . 2 5 ~ on the model chord line. The air gaps at the tunnel walls were sealed with flexible-sliding metal seals (fig. 4).

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Y

Wake Survey Rake

A fixed wake survey rake (fig. 5) at the model midspan was mounted from the tunnel sidewall and located 1 chord length rearward of the trailing edge of the airfoil. The wake rake utilized 91 total-pressure tubes and five s ta t ic-pressure tubes 0.1524 cm (0.060 in.) in diameter. The total-pressure tubes were flattened to 0.1016 cm (0.040 in.) f o r 0.6096 cm (0.24 in.) f r o m the tip of the tubes. The static pressure tubes had four flush orifices drilled 90' apart and located 8 tube diameters f rom the tip of the tube and in the measurement plane of the total-pressure tubes. Three tunnel sidewall static

total-pressure tubes. One static orifice was located on the center line of the tunnel and the other two orifices were about 0 . 3 5 ~ above and below the center line of the tunnel.

p ressures were also measured from orifices located in the measurement plane of the r

Inst rumentation

Measurements of the static pressures on the airfoil surfaces and the wake rake pressures were made by an automatic pressure-scanning system utilizing variable capacitance type precision transducers. Basic tunnel pressures were measured with precision quartz manometers. Angle of attack w a s measured with a calibrated potenti- ometer operated by a pinion gear and rack attached to the circular plates. Data were obtained by a high-speed data-acquisition system and recorded on magnetic tape.

TESTS AND METHODS

The airfoil was investigated at Mach numbers f rom 0.10 to 0.28 over an angle-of- attack range from about -loo to 24O. Reynolds number based on the airfoil chord was varied from about 2.0 X lo6 to 20.0 X 10 6 , primarily by varying the tunnel stagnation

pressure. The model was tested both with the wake rake installed and removed to determine its influence on the flow over the airfoil. Figure 6 shows typical lift coefficient and pitching-moment-coefficient data and no effects were indicated. The pressure distribution data also indicated no effect of the wake rake on the flow over the airfoil. The airfoil w a s tested both smooth (natural boundary-layer transition) and with roughness located on both upper and lower surfaces a t 0 . 0 8 ~ . The roughness was sized according to reference 4 which indicated a nominal roughness particle height of 0.0107 cm (0.0042 in.) at a Reynolds number of 6 x 106 and 0.0257 cm (0.0101 in.) at a Reynolds number of 2 X 106. The corresponding commercial grit numbers required a r e number 120 and number 60. The transition s t r ips were 0.25 cm (0.10 in.) wide. The roughness was sparsely spaced and attached to the airfoil surface with lacquer. Several different roughness s izes were used fo r the same test conditions and these results are shown in figure 7. For several runs the standard NACA method of applying roughness (number 60 grit wrapped around leading edge on both surfaces back to 0 . 0 8 ~ ) w a s employed (ref. 5). For several test-runs

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oil was spread over the airfoil upper surface to determine if any local flow separation was present. Tufts w2re attached to the airfoil and tunnel sidewalls with plastic tape to determine stall patterns on both the airfoil and adjacent sidewalls.

The static-pressure measurements at the airfoil surface were reduced to standard pressure coefficients and then machine integrated to obtain section normal-force and chord-force coefficients and section pitching-moment coefficients about the quarter chord. Section profile-drag coefficient was computed from the wake rake measurements by the method of reference 6. influence on the static pressures due to the presence of the rake body; therefore, the tunnel sidewall static p res su res were used in computing the section profile-drag coefficients.

The wake rake static-pressure measurements indicated some

An estimate of the standard low-speed wind-tunnel boundarp corrections as calculated by the method of reference 7 is shown in figure 8. These corrections amount to about 2 percent of the measured coefficients and have not been applied to the data. An estimate of the total head tube displacements effects on the values of Cd showed these effects to be negligible.

RESULTS

The results of this investigation have been reduced to coefficient form and a r e presented in the following figures:

Figure

Tuft photographs of NASA GA(W)- 1 airfoil . . . . . . . . . . . . . . . . . . . . . 9

10 Effect of Reynolds number on section characteristics, model smooth . . . . . . . Effect of Mach number on section characteristics, model smooth,

R = 6 X 1 0 6 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11

Effect of Reynolds number on section characteristics, roughness located a t 0 . 0 8 ~ . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12

Effect of Mach number on section characteristics, roughness located at O.O8c, R = 6 x l o 6 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13

Effect of roughness on section characterist ics . . . . . . . . . . . . . . . . . . . Comparison of section characterist ics between NASA GA(W)- 1 and

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NACA 652-415 and 653-418 airfoils . . . . . . . . . . . . . . . . . . . . . . . 15

for various airfoils without flaps . . . . . . . . . . . . . . . . . . . . . . . . 16 Variation of maximum section lift coefficient with Reynolds number

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I I lllIlIlllIllllllllIllllll l l l l l l l l l l l l

Figure

Section characterist ics for 0.20~ simulated split flap deflected 60' . . . . . . . . 17

Effect of angle of attack on chordwise pressure distributions . . . . . . . . . . . 18

Comparison of experimental and theoretical aerodynamic characterist ics . . . . 19

Comparison of experimental and theoretical chordwise pressure distributions . . 20

DISCUSSION O F RESULTS

Experiment a1 Results

Lift.- Figure 10 shows that with the airfoil smooth (natural boundary-layer transition) a lift-curve slope of about 0.12 per degree and a lift coefficient of about 0.52 at a = 0' w a s obtained for all Mach numbers and Reynolds numbers investigated, Maxi- mum lift coefficients increased f rom about 1.64 to about 2.12 as the Reynolds number was increased from about 2 X lo6 to 12 X lo6 at M = 0.15 (fig. 16), with the most rapid increase occurring between Reynolds numbers of 2 x 106 and 6 x 106. Increas- ing the Reynolds number above 12 X lo6 had no additional effect on maximum lift coefficient as shown by figure 1O(b) (M = 0.20).

The GA(W)- 1 airfoil section encounters a gradual type stall (fig. lo), particularly in the lower Reynolds number ranges. Tuft pictures (fig. 9) indicated the stall is of the turbulent o r trailing-edge type. (See also pressure data of fig. 18.)

At a Reynolds number of 6.0 X lo6, increasing the Mach number f rom 0.10 to 0.28 had only a minor effect on the lift characteristics as shown by the results presented in figure ll(a). The stall angle of attack was decreased about 2O and maximum lift coefficient about 5 percent.

The addiiion of roughness at 0 . 0 8 ~ (figs. 12 and 14) did al ter the effective airfoil shape because of changes in boundary-layer thickness, particularly for R = 2.0 X 106 as shown in figure 14(a). For example, the angle of attack for zero lift coefficient changed f rom about - 4 O to -3.6'. No measurable change in lift-curve slope was indicated; therefore, the l if t coefficient at (Y = Oo decreased from about 0.52 to about 0.43. These

effects on the lift characterist ics decreased as the Reynolds number was increased above 2.0 x 106 as might be expected because of the related decrease in boundary-layer thickness. Figure 13(a) indicates that the effects of Mach number with roughness applied to the airfoil were s imilar to those with the model smooth.

Comparisons of the values of ( C Z ) ~ ~ for the NASA GA(W)-1 airfoil with other NACA airfoils without flaps a r e shown in figure 16. Substantial improvements in (c,7kax for the GA(w-')-l airfoil throughout the Reynolds number range are indicated when compared

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to the NACA 4 and 5 digit airfoils and 65 series airfoils. Both the GA(W)-l and 653-418 airfoils have the same design lift coefficient (0.40) and figure 2 indicates both airfoils have roughly the same mean thickness distribution in the region of the structual box ( 0 . 1 5 ~ to 0 .60~) . At a Reynolds number of 6.0 X lo6, a 30-percent improvement in (C2)max is shown for the GA(W)-l airfoil over the comparable 653-418 airfoil. Typical operating ranges of Reynolds numbers for general aviation airplanes are from about 2 X lo6 to 6 x lo6. These improvements result from two primary considerations of the design; first, attenuating of the peak negative pressure coefficients on the upper surface near the leading edge by use of a large leading-edge radius, and second, the attainment of increased aft loading by using the greater aft camber. Figure 15(a) shows a comparison of the lift characterist ics of the NACA 65 series airfoils and NASA GA(W)-l airfoil a t a Reynolds number of about 6 X 106 with roughness located near the leading edge of the airfoils. Even when the large wraparound roughness was employed on the new airfoil, it exhibited superior lift characterist ics to the older NACA 65 series airfoils although it stalled about 4' ear l ier than with the narrow s t r ip roughness now usually employed.

high-lift configuration, a simple 0 . 2 0 ~ split flap deflected 60' w a s installed on the model. The increment in cz a t a Reynolds number of about 6.0 X lo6 between the basic airfoil and flapped airfoil was about 1.46 a t (Y = Oo and about 1.2 a t figs. lob) and 17.) The data of reference 5 indicate an increment in c2 of about 1.40 at CY = Oo and about 1.2 at ( ~ 2 ) ~ ~ f o r the NACA 653-418 airfoil. Comparison of the values of ( ~ 2 ) ~ ~ ~ for the two airfoils with the simulated split flap deflected 60° shows about a 17-percent increase for the GA(W)-l airfoil (3.16 compared to 2.70). Similar improvements a r e also indicated when compared to the NACA 4418 airfoil with simulated split flap deflected 60'. The stall characterist ics of the GA(W)-1 airfoil with the simulated flap were gradual as indicated by the lift characteristics and tuft studies.

In order to obtain some preliminary information on the new airfoil section in a

( c ~ ) ~ ~ . (Compare

Pitching moment.- The pitching-moment-coefficient data (fig. 10) were generally insensitive to F?,eynolds number in the low angle-of-attack range. However, for angles of attack greater than about 4' the low Reynolds number data indicate less negative values of cm. Increasing the Reynolds number, which results in a decrease in boundary- layer thickness, caused negative increments in cm up to airfoil stall. At a Reynolds number of 6.0 X lo6 increasing the Mach number from 0.10 to 0.28 (fig, l l(a)) caused no effect on the pitching-moment data up to about 12'. At higher angles of attack a positive increment in c m is shown.

in cm at Reynolds numbers of 2 X lo6 and 4 X 106. However, at a Reynolds number of 6 X lo6 this increment had essentially disappeared.

The addition of roughness (figs. 12 and 14) at 0 . 0 8 ~ resulted in a positive increment

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Comparison of the pitching-moment data of the GA(W) - 1 airfoil with that of the 653-418 airfoil (fig. 15(a)) at a Reynolds number of about 6 X lo6 indicates a more negative Cm of about 0.04 for the GA(W)-l section. This is expected because of the aft loading of the GA(W)-1 airfoil section as illustrated by the camber distribution of figure 2.

Drag.- The profile drag data of figure 10 generally show, at moderate lift coefficients, the expected decrease in Cd with increases in Reynolds number. This drag reduction is associated with the related decreases in boundary-layer thickness and accompanying reduction in skin friction drag. An increased amount of laminar flow is indicated at a Reynolds number of about 2 X lo6 (fig. lO(a)) by the low values of Cd obtained in the low lift-coefficient range (typical laminar bucket). In practical applications no laminar bucket, such as shown here, should be expected since the design velocity characteristics were not selected for this purpose. Laminar flow designs are generally impractical f o r general aviation airplanes since transition is usually fixed near the leading edge of the airfoil by the roughness of construction or insect remains gathered in flight.

For general aviation application, the drag data of most practical interest are those obtained with a turbulent boundary layer over most of the airfoil chord in the Reynolds number range from about 2 X lo6 to 6 X lo6. Figure 12(b) illustrates the drag data with fixed transition at 0 . 0 8 ~ for this Reynolds number range. The drag coefficient at the design lift coefficient (cl = 0.40) a t a typical cruise Reynolds num- be r of 6 X lo6 is about 0.0108. However, figures 7(b) and 12(b) indicate a large lift- coefficient range where the values of Cd remain approximately constant. This is of particular importance f rom a safety standpoint for light general aviation airplanes where large values of section lift-drag ratio at high lift coefficients result in improved climb performance. Thus, at cl = 1.0 section lift-drag ratios vary from about 65 at R = 2.1 X 106 to about 85 at R = 6.3 X 106 (fig. 12(b)).

A comparison between the section lift-drag characteristics of the GA(W)-1 airfoil For the older type air- and the older NACA 65 se r i e s airfoils is shown in figure 15(b).

foils a coarse size grit was extensively applied (0 .08~) over the airfoil in order to achieve transition in the wind tunnel. This older method of applying large wraparound roughness (NACA standard) results in an increment in cd of about 0.0010 at cl = 0.40 for the GA(W)-l airfoil when compared to the narrow roughness s t r ip now usually employed (NASA standard). However, in order to obtain a direct comparison of the drag coefficients between the airfoils the comparison is made with the older NACA standard method of employing roughness.

On this basis figure 15(b) shows at R =: 6.0 X lo6 that the section drag coefficient, at a cruise lift coefficient of about 0.40, for the 17-percent-thick GA(W)-l airfoil is about

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

0.0010 higher than that fo r the 653-418 airfoil. drag ratios at a lift coefficient of 0.90, the highest' ct for which data were available, indicates about a 50-percent improvement for the GA(W)-l airfoil section (Z/d = 47 fo r 653-418 airfoil compared to Z/d = 70 for GA(W)-l airfoil). The figure also indicates that even greater improvements would probably occur at higher climb lift coefficients.

P r e s s u r e distributions.- The chordwise pressure data of figure 18 illustrate the effects of angle of attack for a Reynolds number of 6.3 X lo6. The data at CY = 0' (cl = 0.47) indicate approximately constant values of Cp from about x / c = 0.05 to x/c = 0.55 for both the upper and lower surfaces of the airfoil (design condition). Upper and lower surface p re s su re coefficients at the airfoil trailing edge are slightly positive. Some upper surface trailing-edge separation is first indicated at an angle of attack of about 8' by the constant pressure region on the upper surface of the airfoil and is also indicated by the nonlinear lift curves above this angle of attack. Increases in angle of attack above 8' resulted in this constant pressure region moving forward along the airfoil and a t maximum lift coefficient (CY = 19.06O) trailing-edge separation was present from about x/c = 0.70 to the airfoil trailing edge. The airfoil stall is of the turbulent, o r trailing edge, type as indicated by figure 18(k) (a = 20.05O) and as observed by means of tuft studies.

However, comparison of the section lift-

Comparison of Experimental and Theoretical Data

Predictions of the aerodynamic characterist ics by the viscous flow method of reference 2 are compared with t h e experimental results at R = 6.3 X lo6 previously mentioned, this viscous flow method w a s employed during the development of the GA(W)- 1 airfoil shape. data well for angles of attack where no boundary-layer flow separation is present (up to about 8'). same lift-coefficient range. Examples of pressure distributions calculated by the theoretical method a r e compared with the experimental p re s su res in figure 20. The agreement between experiment and theory is good over most of the chord length of the airfoil, as long as no boundary-layer flow separation is present.

in figure 19. AS

The theoretical method predicts the lift and pitching-moment

However, the theoretical method overpredicts the drag data throughout this

I

CONCLUDING REMARKS

Wind-tunnel tests have been conducted to determine the low-speed two-dimensional aerodynamic characterist ics of a 17-percent-thick airfoil section designed for general aviation applications. The resul ts were compared with a typical older NACA 65 series airfoil section. Also, the experimental data are compared with predictions based on a theoretical method for calculating the viscous flow about the airfoil. The tests were

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k

ll1ll111 IIIIIII I l l l l l l l l I l l 1

conducted over a Mach number range f rom 0.10 to 0.28. Reynolds number based on the airfoil chord was varied f rom about 2 X lo6 to 20 X lo6. The following results were determined from this investigation:

about 2.0 X lo6 to 6.0 X lo6 and attained values greater than 2.0 fo r the plain airfoil and greater than 3.0 with a 20-percent-chord split flap deflected 60°.

1. Maximum section lift coefficients increased rapidly a t Reynolds number f rom

2. Stall characterist ics were generally gradual and of the trailing-edge type either with o r without the split flap.

f rom about 65 to 85 as the Reynolds number increased from about 2.0 X 106 to 6.0 X 106. 3. Section lift-drag ratio at a lif t coefficient of 1.0 (climb condition) increased

4. Maximum section lift coefficients were about 30 percent grea te r than a typical older NACA 65 series airfoil and the section lift-drag ratio at a lift coefficient of 0.90 was about 50 percent greater.

5. Comparison of experiment with predictions based on a theoretical method which included viscous effects was good for the pressure distributions as long as no boundary- layer flow separation was present, but the predicted drag values were much greater than measured values.

Langley Research Center, National Aeronautics and Space Administration,

Hampton, Va., October 16, 1973.

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REFERENCES

1; Ayers, Theodore G.: Supercritical Aerodynamics: Worthwhile Over a Range of Speeds. Astronaut. & Aeronaut., vol. 10, no. 8, Aug. 1972, pp. 32-36.

I 2. 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.

3. Von Doenhoff, Albert E.; and Abbott, Frank T., Jr.: The Langley Two-Dimensional Low-Turbulence P r e s s u r e Tunnel. NACA TN 1283, 1947.

I ' 4. Braslow, Albert L.; and -ox, Eugene C.: Simplified Method fo r Determination of

Critical Height of Distributed Roughness Particles for Boundary- Layer Transition at Mach Numbers From 0 to 5. 1 NACA TN 4363, 1958.

5. Abbott, Ira H.; and Von Doenhoff, Albert E.: Theory of Wing Sections. Dover Publ., Inc., e. 1959.

6. Baals, Donald D.; and Mourhess, Mary J.: Numerical Evaluation of the Wake-Survey Equations fo r Subsonic Flow Including the Effect of Energy Addition. NACA WR L-5, 1945. (Formerly NACA ARR L5H27.)

7. Allen, H. Julian; and Vincenti, Walter G.: Wall Interference in a Two-Dimensional- Flow Wind Tunnel, With Consideration of the Effects of Compressibility. NACA Rep. 782, 1944. (Supersedes NACA WR A-63.)

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TABLE I.- NASA GA(W)- 1 AIRFOIL COORDINATES

= 58.42 cm (23 in.)]

0.0 .002 .005 .0125 .025 .0375 .05 .075 .loo .125 .150 .175 .20 .25 .30 .35 .40 .45 .50 .55 .575 .60 .625 .65 .675 .700 .725 .750 .775 .800 ,825 .a50 .875 .goo ,925 .950 .975 1.000

(z/c)u, r

0.0 .01300 .02035 .03069 .04165 .04974 .05600 .06561 .07309 .07909 .084 13 .08848 .09209 .09778 .lo169 .lo409 .lo500 .lo456 .lo269 .09917 .09674 .09374 .09013 .08604 .08144 .07639 .07096 .06517 .05913 .05291 .04644 .03983 .033 13 .02639 .01965 .01287 .00604 -.00074

(Z /c ) lOWer

0.0 -.00974 -.01444 -.02052 -.02691 - .03191 -.03569 -.04209 -.04700 -.05087 - .Os426 -.05700 - ,05926 -.06265 -.06448 - .06517 - .06483 - .06344 -.06091 -.05683 -.05396 -.05061 -.04678 -.04265 -.03830 - .03383 -.02930 - .02461 -.02030 -.01587 - .01191 -.00852 -.00565 -.00352 - .00248 - .00257 -.00396 - .00783

14

-6

TABLE 11.- AIRFOIL OFUFICE LOCATIONS

Upper surface Lower surface

x/c 0.0 .00630 .01248 .01730 .02461 .03713 .04961 .06222 .07522 .lo013 .14970 .ZOO04 .24991 .29965 .34991 .39978 .44974 .50004 .5 503 5 .60017 .64996 .70004 .75000 .79987 .a5004 .go004 .95026 .99004

__ Z/C

0.00030 .00228 .03083 .03543 .04143 .04957 .05583 .06 109 .06570 .07313 .OB409 .09209 .09778 .lo170 .lo409 .lo500 .I0457 .lo270 .09917 .09374 .08609 ,07643 .06522 ,05296 .03987 .02643 .01287 .00204 .-

- ..

x/c

0.00678 .01204 .01722 ,02596 .03726 .04970 .06196 .07422 .09957 .14961 .19943 .24965 .30004 .34983 .39991 .45009 .49983 .54970 .59983 .65022 .70022 .75000 .BOO13 . 8 5004 .a9987 .94970 .99004 1.0

Lower surface orifices for 0.20~ simulated split flap deflected 60'

x/c

0.81304 .a2609 .a3478 .a4565 .a5652 .a6739 .a7826 .a8696

-0.03913 -.06087 -.07696 -.09565 -. 11413 -. 13304 -.15217 -.17065

. .

Z/C ~- 0.01635 -.02035 -.02326 -.02683 -.03187 -.03583 -.03909 - .04 200 -.04700 -.05426 -.05930 -.06265 -.06452 -.06517 -.06487 -.06348 - .06 100 -.05691 -.05070 -.04270 -.03387 -.02483 -.01596 -.00857 -.00357 -.00261 -. 006 13 -.00430

__

15

Y

.2

.I

z/c 0

-.I

.I .2 .3 .4 .5 .6 x /c

.7 .8 .9 I .o

Figure 1.- Section shape for NASA GA(W)-l airfoil.

. I 0

,O 8

.O 6

Zt/C .O 4

.o 2

0 Thickness distribution

-

.2 .3 -

. I . 4 .5 x /C

1 1 Camber line - .02 0 .I .2 .3 .4

NASA GA ( W ) - l airfoil NACA 6S3-018 airfoil

.6 .7 .8 .9 I .o

- NASA. GA (w)-l airfoil

- r l l - - - NACA mean line, a=l.O]

.5 .6 .7 .8 .9 */c

Figure 2.- Comparison of camber lines and thickness distribution.

. , . .. , _

17

__t

Airflow L- . A

,

,- Tunnel sidewalls -

Seal detail "z"

T I . 57c

I Tunnel center line

Fnrl u iow sprtinn A-A -..- . ." .. . """. .-.. , . . .

balance attachment

Zero incidence reference

't.- IUL-LULL I I I W U I I L ~ U in wina runnel. mi aimensions in terms of airfoil chord. c = 58.42 cm (23 in.).

19

11111ll1lllll1lllllIl l l l l l l I I

Static- pressure probe --J 7-D--'

.022c

Static -pressure probe

=I

__t

Air f low

Tunnel %--

Total - Dressure probe -3- -3 (tubes flattened)

-L

7

- r 1 .I 9 6 c [ ty P o l

Figure 5.- Drawing of wake rake. All dimensions in terms of airfoil chord. c = 58.42 cm (23 in.).

20

2 .o

I .6

.a Q

.4

0

-.4

CI

-

n

- . I I

- .2 -

Figure 6.- Effect of wake rake on section characteristics. Model smooth; M = 0.20; R E 6 X 106.

21

x

. .

Roughness 0 No. 80 0 No.100

NO, 120

20 24

(a) Variation of cl and c, with a.

Figure 7.- Effect of roughness size on section characteristics. M = 0.20; R = 6 X 106.

22

-03

.02

.o I

0

I

I I I :

1

-1.2 -.8 -.4 0 .4 - .-

Roughness 0 No. 80 0 No. 100 0 No-120

1.2 I .6

(b) Variation of Cd with cz.

Figure 7 . - Concluded.

23

. .

0 No boundary correction i i i I .

it i!

I 0 Boundary correct ion

I .6

I .2

.8

C l

.4

- .4

- .8

.I

0

I

-.I

0 - .2 -12 - 8 -4

i i l

' I 1

Cm

1

i - !

16 4 12 20

(a) Variation of el and cm with a.

Figure 8. - Effect of standard low-speed wind-tunnel boundary corrections on section data. Number 80 roughness located at 0 . 0 8 ~ ; M = 0.15; R 'I 8 x 106.

.03

.02

‘d

.o I

Q 8 -. 4

0

0

.0

No boundary correct ion Boundary correction


Figure 8.- Concluded.

25

C Y = 00

C Y = 14'

a = 40

CY= 16O Figure 9.- Tuft photographs of NASA GA(W)-l airfoil.

a! = 100

C Y = 18'

M = 0.20; R = 2.70 X 106. L-73 - 6898

2.4 . I

0

0

0 A

b

, . . . , . . R 1 . 9 x 1 0 6 3.9 5.7 9.2 2.3

I

I I I

1 i

1

i

I I

I

I

+ - I

1 1 + I I

I

i I

I , , I '

i ;

I I

I

i i

I I i

2 .o

I .6

I .2

.0

=1

.4

0

- .4 I !

- .0

C

Cm

-.I

-.2 -12 -8 -4 4 0

Q ,deg 12 16 20 24

(a) M = 0.15.

Figure 10.- Effect of Reynolds number on section characteristics. Model smooth.

27

IIlll1ll1l11lllI lllllIIlll I l l ll11l1l1ll I l l l l l l l

I I I

R

'd

.04

.03

.02

.o I

0

(a) M = 0.15. Concluded.

I .6

Figure 10. - Continued.

28

2.4 , . . I . . . . I

R 0 2 . 7 ~ 1 0 ~ 0 4.1 0 6.2 A 9.3 n 12.2

2 ,o

I .6 16.5

!

I . I I .

C

- .4 I .

4 . I .

. . I

I

I

I I

- .E I . . . I

:::I: . . . ,

I . . . 1

I . . . I / I .

20 12 2 - 8 - 4 0 4 16

(b) M = 0.20.


.03

.02

'd .o I

R o 2 . 7 ~ 1 0 ~

0.20. Concluded.

10. - Continued.

2.0 2.4

30

. .

0 0

0 A

h

. .

R

6.0 9. I 1.5 3.6

3.5 x IO6 2 .o

I .e

I .2

.E

=1

1

C

- 4

- .8

.I

I

. . . . . . . . I . . . . . . 1;: t . L I

1

-.I

- .2 - 8 12 16 20 24

(c) M = 0.28.


31

___ . .....

R

.03

.02

Cd

.o I

0 ,2 - .8 - .4 0

m tu 'li i ill z

.4 C1

(c) M = 0.28. Concluded.

Figure 10. - Concluded.

32

.8 I .2 I .6 2 .o

7.

... . . . ... .._. .

0 8

Variation of cz and cm with a.

Figure 11.- Effect of Mach number on section characteristics. Model smooth; R = 6 X 106.

33

. .

I I Ill1 II I1 l l l l l l l l l l l I I I

‘d

.04

.03

.02

.o 1

0 - 1.2 -.8 -.4 0 .4 .0 1.2 ‘ 1

I

d

i I

I I 1 t I t- I

I 1 I I I 1

2.0


Figure 1 1. - Concluded.

34

h

.. . . . ...

2.0 '

I .6

I .2

.0

Ca

.4

0

- .4

-.2 I -12 -8

f

I

j i i I

j

i

t i I

-4 0 12

(a) M = '0.10; R = 3.7 x 106.

16 20 24

Figure 12. - Effect of Reynolds number on section characteristics. Number 80 roughness located at 0 . 0 8 ~ .

35

.03

002

.o I

0 - ,1.2 -.8 -.4 .4 C l

(a) M = 0.10; R = 3.7 X lo6. Concluded.


36

2 .o

I .6

I .2

.e c 1

.4

0

- .€

cm -. I

-.2 -

H 0 2.1x IO6 0 4.3 0 6.3

2 - 0 - 4 0 4 8 a,deg

!I 1 I

t

T I I I

i I

P t

i

(b) M = 0.15.


6 20 24

37

:i. . ,

1111111l1l11l1lllllIl Ill I l l1 Ill1 llllllIllll

(b) M = 0.15. Concluded.


38

, . , , : I . .

0

0 3

. - .

M 0.15

.20

I

I

i !

! +

I t !

i i I I

I

t

i I I

I

2 .o

I .6

I .2

.a =1

.4

.2

d

4 8 a ,deg

12 16 0

(a) Variation of c l and cm with a.

Figure 13.- Effect of Mach number on section characteristics. roughness located at 0 . 0 8 ~ . R = 6 X IO6.

Number 80

39

Y

I I I I l11111l11111lllIIllllIlll

-0 E I I

I 1 i d 1 I i I I I

i

M 0.1 5 .20 .28

0 0 0

.05

.04 i

.02

.o I

0 - 0 .4 c1

1.2 I .6 2 .o

(b) Variation of Cd with


40

o o n 2 .o

-.4

-.I

-.2 - t .

I ' 2 .o 16 0 4 0 12 - 8 -4

Figure 14.- Effect of roughness on section characteristics. M = 0.15.

Number 80 roughness located at 0 . 0 8 ~ .

4 1

I ,. . .

I I I1 111llll11l1l1ll1l11 ll1ll1l1l1ll1l1lIl l l l l l l l l l l l l .

d C

-0

1 I -1 - I i i I I I I

2 .o

(a) R = 2 X lo6. Concluded.


42

I 2 .o

I .6

12

=1

0 12 16 20 24

43

.04

.03

-02

.o I

0 -

li I! 't I I I I f

i. L Y

A

d Fh v I

f l

f.

I f I

f

i 1

I

1 L

I .6

(b) R = 4 x 106. Concluded.

Figure 14.- Continued.

44

I

2.41

I

1.61

I .2

.8

C l

.4

0

- .4

- .a

I

C

Cm

-.I

... . . Roughness

0 Off 0 On

. . . , .

- 0 -4 4 0

t I I

8 12 16 20 24

(c) R =: 6 X 106.


45

I I 1 11l11l111lll1l1ll1lll1l1111llllllll Ill1 Ill1 llIllllll Ill1 I l l l l l l l I l l IIIIIIII

005

.04

.03

Cd .02

.o I

.4 -8 =1

1.2 I .6 2 .o

(c) R = 6 X 106. Concluded.


46

NASA GA(W)- I airfoil 0 NASA standard roughness 0 NACA standard roughness

NACA airfoil ,NACA standard roughness (ref.5) 652- 415 653- 418

2 .o

I .6

I .2

.e C l

.4

0

- .4

- .a

,I

0 cm

-.I

- .2 -

I

i i

I , !

I . , . . ,

t : ! :

. . ..

. . P

:it:.-- I .

I . , , I ,

cm data for model smooth I

1' . I . . +\-TI- I l . . . L . . .

. . . I I . .. . . .

. /

-0 -4 0 4 8 Q ,deg

12 16 20 24

(a) Variation of cz and Cm with CY.

Figure 15.- Comparison of section characteristics of NASA GA(W)- 1 airfoil and NACA 652-415 and 653-418 airfoils. M = 0.20; R =: 6 X 106.

47

IIllIIll11111llllIlIlllIll IIIIII llllllIIlllIll IIIIIIII Ill1

.04

.03

.02 'd

.o I

0 - I

NASA GA(W)- I airfoil o N A S A standard roughness 0 NACA standard roughness

NACA airfoil , NACA 652- 415

-- 653- 418

---

( 2 -.e -.4

standard roughness (ref.5 )

(b) Variation of Cd with c z .

Figure 15.- Concluded,

.O

4%

2.2

2.0

I .8

PZh” 1.6

I .4

1.2

I .O 0

NASA G A ( W 1 - I airfoil 0 Roughness off 0 Roughness on

2 4 8 6 R IO 12 1 4 ~ 1 0 ~

Figure 16.- Variation of maximum section lift coefficient with Reynolds number for various airfoils without flaps. M = 0.15.

49

n . .-.w . I .. . . , . . .

3.2

2.8

2.4

2 .o

, . I . , , = - Roughness

Off Off Off

. 08c

R o 2 . 7 ~ 1 0 ~

4.1 0 6.1 A 6.1

~ c , t !

t

I

: 1 I

I t it1 l l i r

c~ 1.6

I .2 ,

i .8

t€J . I . I . . , J

9- l I

.4

0 '

0

-.2

-. 4 -I

I

. !

Cm I

12 16 20

Figure 17.- Section characteristics for 0.20~ simulated split flap deflected 60°. M = 0.20.

50

-9

-8

- 7

-6

-5

cp -4

g i

I I I I I I l l

I I -3 1-

i

I I

cp

0 Upper surface Lower surface

0 !I

I I I I I I I I I I I I I I 1 I I f I I I I I I ! I ~I I , I I ! I ' I I I I I I I I 1 ' 1 I I I ! I I I

0 .I .2 - 3 .4 .5 .6

i

(a) a! = -7.99'.

Figure 18.- Effect of angle of attack on chordwise pressure distributions. Number 80 roughness located at 0 . 0 8 ~ . M = 0.15; R = 6.3 X lo6. (Flagged symbol indicates base pressure orifice.)

51

I .

- -jli 2.0 - -

- 1.6

n I

I

0

0 0 0

n

i I I I I ' I I

__

0

0

I I I I I I I I I I I I I I I I I

0 Upper surfoce 0 Lower surface

1.2 h I 1 I I I I 1 t

(b) (Y = -4.11'.


52

- 2-8E - -

*41 .8 - -

. . . .

. .

. .

0

n

. .

. .

0

n

I I I I I I ~

. -

0

L l

I I I l I l : I

0

U

I I I I I I I I

0

C l

I I I I I I I

o Upper surface Lower surface

0

0

(c) cy = 00.


53

.. ... .,. , . ' 1: . - .. . ,. .

I . . - .

- -2*8F

I I I 1.1 I L ,

a&----

, I I i I I I I I ' I I I I I I I 1 I

-1.6. I . .- .-

.

.- . . . . . .

0 1 ......

. -.

m c . . . . ._ .

. .

.. cP -

-.4 -- Fi

0 E .. . . . . .

E.

... . . .

. -. - - -.

. . . . . . .

. . . . . . .

0

. .....

....

0 ..

. . . . .

. . . .

I L I . I l I I

. . .

. .,. - .

. . . .

.........

0

......

. . . . .

D . . . . .

. .

. . . . . .

- . . . .

_. ... ._ . . . .

. . .

. .

. .Q

I I I I I I I

_ _ .....

. . . .

.

. . . .

.... .-

0

. . u . .

. . . .

1 1.1. I LL.1

. . - 0 Upper surface 0 Lower surfoce

...

- . . . . .

0

O I

...

1.111 1.1 I

. .

0 cl 0

0

' I I1 It I I 1.2 LL1.1, 1.1 I I

x/c

(d) CY = 2.06'.


54

. . .

-2.8

-2 .41

-1.6 u

- 1.2 t cp -.8

-.4 1

0

u

I I I I I I I

0

11

l l l l l l l l

0 Upper surface Lower surface

0

0

I I I I I 1 1 I1

. .

0

0

I I l l 1 1 1 L

0 . I '2 .3 .4 .5 .6 .? .8 x /c

I .o

(e) a = 4.17'.


55

ll1ll111 IllIlllllllIl l l l l l l l I llll1l1l1ll1l111l1l111l11l1ll1 ll! .

. .

0

u

U.U! L

O C

I I I 11.1 I ! 1.1 l.l!J I I l.lJ 11.1.I I1 I t

0 (

" 4

. .

0 ,

c-1 1

0 Upper surface 0 Lower surface

I

'0 .I .2 .3 ,4 .5 .6 .7 -8 . .9 I .o X /C

(f) 01 = 8.02'.


56

(g) a! = 12.04'.


;.' ;

57

2 ..

1111 IIIII lllllll111l I I l11lll11111ll1llllllllllllll I I ,

. . . .

. .

. _ .

. .

. .

0 c

. .

o c

I I I I I I I I ~

-gE- .

. . .

. .-

- .

. .

0

. .

. .

I 0

LII.L.I.I I I

. . . .

. . .

. . .

. . .

ij I

. . .

. . .

0 '

I I I L I I I

.. - . . . . .

. . . .

- . . .

. . .

. .

0

0

I I I I 1 1 I 1

0

. .

n l

I I I I I ! I1

. . . .

. .

. .

0

U

I I I I I I I

. .


0 1

U I

( I I I I I I1

0 .I .2 .3 .4 .5 $6 .7 .0 .9 I .o x/c

(h) CY = 16.04'.


58

.......

(i) (Y = 18.25'.


59

1111 111lll I l l I l l1 Ill11 1l1l1ll1ll1111ll1l11l1ll I I

0 (

u L

, I (.I 1 1 I I

Q l

Q [

I I I I J I I

0

U

I I I I I I I I

0

U

I I I I I I I

0 Upper surface o Lower surface

.3 .4 .5 .6 .7 .0 .9 I .o X/C

(j) a! = 19.06'.


60

.I

-7 F

I

,O

-6 i' o

-5k

0

I [ J

t I 'f'

.3 .4 .5 .6 x /C

(k) CY = 20.05'.



61

. .i

Ill1 l l l l l l IIIIII

3 - - - -

. . .

0 '

... _.

. . . .

O (

u [

I I I I I. I I

.-

. .

. .

. .

....

. .

. Q .

. . . .

0 1

111.11 I I

. . . .

. .

. .

. .

. .

0

.c1 '

. .

.~

. . . . .

0 -

n

.I 1.1 I .I I.

0 0

. . .

Upper surface Lower surface

0 . I .2 .3 ,4 .5 .6 .7 .8 .9 I .o X/C

(1) (11 = 21.14'.


62

- >

. ~ . - I

2.8 . , . 0

. . . . . . . . . . . Ex perlment Theory

. . I ' I

' I

, . , . ,. , '

i i i j j I , 2.4

surface trc separat ic 2 .o

12

.4

0

- .4

-.a

. I

C cm

-.I

-.s -12 - 8 0 12 16 20 24 - 4

(a) Variation of c1 and cm with a.

Figure 19.- Comparison of experimental and theoretical aerodynamic characteristics. Transition fixed at 0 . 0 8 ~ ; M = 0.15; R = 6.3 x 10 6 .

63

1ll1111I111l11 II I Ill

.05

.04

.03

'd

.02

.o I

0 -

0 Experlment Theory

..

,L 2 -.O -,4 I .6 2 .o

(b) Variation of Cd with c i .


64

. . -.

-‘081 Experiment Theory. . ’

-2.41

i

“ I If ‘“JX I. r

- -4 I I

.8

1.2 -- 0 . I

0 0

Upper surface --- Lower surface

.2 .3 .6 .? .8 .9 I ,@

(a) a = -4 .11O.

Figure 20. - Comparison of experimental and theoretical chordwise pressure distributions. Transition fixed at 0 . 0 8 ~ ; M = 0.15; R = 6.3 X lo6. (Flagged symbol indicates base pressure orifice.)

65

, . - . . .: : ”: . .

cP

Experiment Theory Upper surface

--- Lower surface -2.4

.4

3- I L

1 I I 1 I

- 1.-

.5 X /C

.6

(b) c y = oo. Figure 20. - Continued.

66

.7 .0 .9 1.0 :'

(c) (Y = 4.17'.


67

. --.-. - I :

- 3 0 6 1 7 Experiment Theory

.a .9 I .o

~ (d) a = 8.02'.


68

-9

- 8

-7

- 6

- 5

cp - 4

-3

-2

- I

0

I

I

i

Experiment Theory 0 0

a

Ep.

Upper surface _- - Lower surface

L-9132 NASA-Lurgleg. 1973 - 1 i

(e) a = 12.04'.


.6

low-speed aerodynamic characteristics of a 17 percent thick … · 2013-08-31 · low-speed...

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