^ ;_. ;-

A S A TECHNICAL

MEEV9ORAND U

I

X

wo

NASA TM X-2532

s; / r~.->

! Wfr

MS; -;ii .M/ iYii-i

AERODYNAMIC CHARACTERISTICS OF

TWO NASA SUPERCRITICAL AIRFOILS

WITH DIFFERENT MAXIMUM THICKNESSES

by Charles D. Harris

Langley Research Center

Hampton, Va. 23365

NATIONAL AERONAUTIC S AND SPAC E ADMINISTRATION • WASHINGTON, D. C. • APRIL 1972

€0NFIDENTiAt-

-4.
/ •;. • .,] •.-.::•. .-;:..-.^v5 ,,;.-, n V;•- :j v:r.t.;r;:

:i ,:',_ ' ,.* : "'..'»,') in'"."" ".i t •::•••".•. Ci '.: oT^-7

.::•. "'.'.. , , :;-'o -:?; ;~.'.\i *::"'.;'.•;:: '.''.., y

i ".. . \~ c"?:"::~ j-j f;i.;s "• " :": .' r.-:r- *^
J

THIS PAGE IS UNCLASSIFIED

ERRATA

NASA Technical Memorandum X-2532

AERODYNAMIC CHARACTERISTICS OF TWO NASA SUPERCRITICAL

AIRFOILS WITH DIFFERENT MAXIMUM THICKNESSES

By Charles D. Harris

April 1972

Pages 23 to 35: Section drag coefficients, c^, for the 10-percent-thick airfoils shown in

figures 7, 8, and 9 are incorrect because of an erroneous instrumentation sensitivity

constant. Correct drag levels are roughly 8 percent greater than indicated. Discussion and conclusions pertaining to these figures are unaffected.

Issued August 1973

THIS PAGE IS UNCLASSIFIED
1. Report No.

NASATMX-2532

2. Government Accession No. 3. Recipient's Catalog No.

4. Title and Subtitle

AERODYNAMIC CHARACTERISTICS OF TWO NASA SUPERCRITICAL AIRFOILS WITH DIFFERENT MAXIMUM

THICKNESSES (U>

5. Report Date

April 1972

6. Performing Organization Code

7. Author(s)

Charles D. Harris

8. Performing Organization Report No.

L-8173

9. Performing Organization Name and Address

NASA Langley Research Center

Hampton, Va. 23365

10. Work Unit No.

742-73-01-14

11. Contract or Grant No.

12. Sponsoring Agency Name and Address

National Aeronautics and Space Administration

Washington, D.C. 20546

13. Type of Report and Period Covered

Technical Memorandum

14. Sponsoring Agency Code

15. Supplementary Notes

16. Abstract

Wind-tunnel tests have been conducted at Mach numbers from 0.60 to 0.81 on two supercritical airfoils with maximum thicknesses of 10 and 11 percent of the chord. The results

demonstrate the influence of variations in certain airfoil shape parameters and provide a.

source of systematic experimental data for the supercritical airfoil.

CLASSIFICATION CHANGE

UNCLASSIFIED

Station, NASA

"information Facility

.CLASSIFIED

SUBJECT J£**6SWERAL

*j&"->S#

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

Supercritical airfoils

Transonic aerodynamics

19. Security Oassif. (of this report) 20. Security Classif. (of this page)

Unclassified

This mater^cDWtairis^formatioriaffgfling the nationaj,^e*ense of the Unjje>*-States

IB m^^fTthe esgflffjgelaws, Tit^a??5!sX.. Sees. TQ^SSPlsk. the transmission orTevelation of wmch in any mann^rnoan unauthorized p^on is prohibited by law.
AERODYNAMIC CHARACTERISTICS OF TWO NASA SUPERCRITICAL

AIRFOILS WITH DIFFERENT MAXIMUM THICKNESSES*

•

By Charles D. Harris

Langley Research Center

SUMMARY

Wind-tunnel tests have been conducted at Mach numbers from 0.60 to 0.81 to determine the aerodynamic characteristics of two NASA supercritical airfoils with maximum

thicknesses of 10 and 11 percent of the chord. Other geometric dissimilarities were

present which would prevent a direct comparison of the two airfoils, but the results provide a source of systematic experimental data for the supercritical airfoil.

For the thinner airfoil, stall (onset of trailing-edge separation) begins at an approximately 0.1 higher normal-force coefficient at the higher test Mach numbers, and the drag

divergence Mach number at a normal-force coefficient of 0.7 was 0.01 higher. Both

effects are associated with lower induced velocities over the thinner airfoil.

INTRODUCTION

In principle, the NASA supercritical airfoil (refs. 1 to 4) is shaped to reduce the

drag associated with energy losses due to shock waves and flow separation. It was conceived to have extensive regions of local supersonic flow over the upper surface but have

only weak shock waves near the trailing edge; thus, satisfactory performance characteristics were maintained well beyond the critical Mach number (the free-stream Mach number at which the local velocity becomes sonic at some point on the airfoil).

This report documents results of an early phase, of the supercritical airfoil wind- ~

tunnel development program involving comparisons of two supercritical airfoils with

maximum thicknesses of 10 and 11 percent of the chord. Although maximum thickness

was the primary variable between the two airfoils, other geometric dissimilarities were

present which would prevent a direct comparison. The results are useful, however, in

demonstrating the effects of variations in certain airfoil shape parameters and provide a

further source of systematic experimental data for the supercritical airfoil.

The wind-tunnel results presented herein for Mach numbers from 0.60 to 0.81 were

obtained in the Langley 8-foot transonic pressure tunnel. Normal-force, drag, and
pitching-moment coefficients -y=r* 'fetermined from static-pressure measurements along

the surface of the airfoil and ttsl-;,ressure measurements in the wake of the model.

SYMBOLS

pressure coefficien:.

c

p,sonic pressure coefficier.: c-v/responding to local Mach number of 1.0

c chord of airfoil, cesirrietftrs (inches)

cd section drag coefficient, \ cd —

cd point drag coefficient (ret. 5)

cm section pitching-moment coefficient about the quarter-chord point,

i.s. U.K.

"-n

M

m

P

q

R

• \ A vr ^ A

section normal-force coefficient, / Cn — - / Cp ~

Z_ j *• C l_~j. *> C

I.S. U.S.

Mach number

dv

slope of airfoil surface, -i

dx

static pressure, newtons per meter2 (pounds per foot2)

dynamic pressure, newtons per meter2 (pounds per foot2)

Reynolds number based on chord of 10-percent-thick airfoil, 63.5 centimeters

(25.0 inches)

airfoil leading-edge radius, centimeters (inches)

airfoil thickness, centimeters (inches)
agSiSS^a^

y

ordinate along airfoil reference line measured from airfoil leading edge,

centimeters (inches)

ordinate normal to airfoil reference line, centimeters (inches)

z vertical distance in wake profile measured from top of rake, centimeters

(inches)

,- a angle of attack of airfoil reference line, degrees

Subscripts:

I conditions at local point on airfoil

max maximum

te trailing edge

00

 conditions in undisturbed stream

Abbreviations:

l.s. airfoil lower surface

u.s. airfoil upper surface

APPARATUS AND TECHNIQUES

The apparatus and testing techniques used during this investigation were the same

as those described in references 1 and 2. These descriptions are repeated herein for

convenience. A number of previous papers have discussed the design philosophy of the

*. supercritical airfoil concept (refs. 1 and 2, for example), and these discussions are not

repeated herein.

Wind Tunnel

The investigation was conducted in the Langley 8-foot transonic pressure tunnel.

This tunnel is a single-return, rectangular wind tunnel with controls that allow for the

independent variation of Much number, stagnation pressure, temperature, and dewpoint

. 6). The upper and lower walls of the test section are axially slotted to permit
testing throve the transonic speed range. The total slot width at the position of the

model aver*.^d about 5 percent of the width of the upper and lower walls.

The f,}jcj Side walls and slotted upper and lower walls make this tunnel well suited

to the inves?;5rnQOn of two-dimensional models since the side walls act as end plates

while the sl'/*s permit development of the flow field in the vertical direction.

Models

Airfoil number designations used are those assigned as part of the numbering system for the overall supercritical airfoil development program.

Models; f,f two supercritical airfoils (fig. 1), each having a thickness-chord ratio at

the blunt trajij,,^ ecjge of Q.Q! (trailing-edge thickness 1 percent of the chord), were used

in this invej-nation. The trailing-edge thickness was somewhat greater than that shown

to be desirahio in reference 2, but this should not invalidate conclusions drawn from the

data present! herein. Coordinates and surface slopes of the two airfoils are presented

in tables I a/i/i jj and ^e chordwise distribution of surface slopes is presented graphically in figu/ & 2.

The 11 percent-thick ((t/c)max = O.ll) airfoil (airfoil 5) was derived from an earlier airfoil v/m, sharp trailing edge (airfoil 4) by rotating the rear lower surface downward about HIM G4-percent-chord line to the desired trailing-edge thickness. (Airfoils 4

and 5 are cdu,,,.lred and discussed in ref. 2.)

The 10 porcent-thick ((t/c)max = O.lOj airfoil (airfoil 9), considered to be more

representative of the midsemispan section thicknesses used on present-day transports,

had a smaller loading-edge radius and increased rear camber (fig. 1). Figure 2(a) indicates slightly induced curvature around the 50-percent-chord station and increased curvature over approximately the rearmost 25 percent of the upper surface of the airfoil

for the 10-per rent-thick airfoil. (For small values of slope, curvature may be closely

approximate^ |,y (ne second derivative of the airfoil contour, d2y/dx2.) .

The air foil models were mounted in an inverted position and completely spanned the

width of the tunnel. Angle of attack was changed manually by rotating the model about

pivots in the Innnel sidewalls. Sketches of airfoil 9 and the profile drag rake are presented in figure 3, and a photograph of one of the airfoils and the profile drag rake

mounted Ln thP (uimei is shown as figure 4.

Boundary-Layer Transition

Bas-e-d CM\ (he technique discussed in reference 7, transition strips were applied

along the 25-\>owent-chord line on both the upper and lower surfaces to simulate full-scale

Reynolds nusnisM- boundary-layer characteristics at the trailing edge as well as full-scale

W
shock-wave positions. The simulation is limited on the upper surface to those conditions

in which the shock wave occurs behind the transition, that is, to the higher test Mach num-

• bers. Full-scale simulation on the lower surface would be valid through the Mach number range of the investigation since laminar flow can be maintained ahead of the trip for

all conditions. The transition strips consisted of 0.25-cm-wide (0.10 in.) bands of No. 90

carborundum grains.

Caution should be exercised when comparing the present results with results from

earlier supercritical airfoil investigations since transition grit size and locations used

during earlier phases of the supercritical development program differed from those just

[

 described.

Measurements

Surface pressure measurements.- Normal forces and pitching moments acting on

the airfoils were determined from surface static-pressure measurements. The surface

pressures were obtained from chordwise rows of orifices located approximately 0.28c

from the tunnel center line on airfoil 5 and 0.32c from the tunnel center line on airfoil 9.

The maximum range of the transducers in the differential pressure-scanning values used

to measure the static pressure at the surface was ±68.9 kN/m2 (±10 Ib/in2). The orifices

were concentrated near the leading and trailing edges of the .airfoils to define the severe

pressure gradients in these regions. In addition, a rearward-facing orifice was included

in the trailing edge of the 10-percent-thick airfoil (airfoil 9).

Wake measurements.- Drag forces acting on the airfoils, as measured by the

momentum deficiency within the wake, were derived from vertical variations of the total

and static pressures measured across the wake with the profile drag rake shown in figure 3(b). The rake was positioned in the vertical center-line plane of the tunnel, approximately 1 chord length rearward of the trailing edge of the airfoil. The total-pressure

tubes were flattened horizontally and closely spaced vertically (0.36 percent of the airfoil

chord) in the region of the wake associated with skin-friction boundary-layer losses.

^ Outside this region, the tube vertical spacing progressively widened until, in the region

above the wing where only shock losses were anticipated, the total-pressure tubes were

spaced about 7.3 percent chord apart. Static-pressure tubes were distributed as shown

* in figure 3(b). The rake was attached to the conventional center-line sting mount of the

tunnel which permitted it to be moved vertically to center the close concentration of tubes

on the boundary-layer wake.

Total and static pressures across the wake were also measured with the use of

differential pressure-scanning values. The maximum range of the transducer in the

valve connected to total-pressure tubes intended to measure boundary-layer losses was
i/i t Ib/in2); the corresponding maximum range for measuring shock losses

and ' steW ^''»6ure was ±6-

9

 kN/m2 (±J Ib/in2).

Reduction of Data

fyli'itfailon of cn and cm.- Section normal-force and pitching- moment coeffi-

• te "//*--/<: obtained by numerical integration (based on the trapezoidal method) of the .

~ .;//;<';'•• pressure coefficient measured at each orifice multiplied by an appropriate

weigMJ/'K f«";t/>r (incremental area).

/'yjr;u)iiUon of C(j.- To obtain section drag coefficients from the total and static

c ur.H h'^iind the model, point drag coefficients for each of the total-pressure meapress'."

rittlt\n w<;re computed by using the procedure of reference 5. These point drag val-

>4, jH/t:iln summed by numerical integration across the wake based on the trapezoidal m<r»>i'"1.

Wind- Tunnel Wall Effects

'/')»« niiijor interference effect of the wind-tunnel walls was an upflow at the inverted

ode) 'I'''|M UP^OW» proportional to the normal-force coefficient, results in the measured

> ;.j,. jingle of attack being significantly greater than the aerodynamic angle of attack

at the MKl|<)r normal-f°rce coefficients. The mean value of this upflow (in degrees) at

the mi'J'''101'*' °'

 the model ma

y

 be

 estimated by the theory of reference 8 to be approxi-

. , ;j nui(!H the section normal-force coefficient. Based on experience in other

, dj;/(t,,,nlonal tests in the Langley 8-foot transonic pressure tunnel, however, such

a corr«?«'"<m ^8 believed to be unrealistically large. Because of this uncertainty, the

Kt.(tul geometric angles of attack are used in the results presented herein.

'j'h« theory of reference 8 also indicates that tunnel blockage effects would be small;

consef<"t)"".V.

 IU) corrections have been applied to the data to account for blockage effects.

TEST CONDITIONS

wore conducted at Mach numbers from 0.60 to 0.81 at a stagnation pressure

of 0 10 J;' MN/ui^ (1 atm). Wind-tunnel Reynolds numbers corresponding to these conditions a»Mt lui.-u'd on the chord of airfoil 9 (63.5 cm (25.0 in.)) varied as shown in figure 5.

The st«ii'ul<'

on

 'emP

erature

 of the tunnel air was automatically controlled at approximately

322 K ( I'W° ^ and tlie air was dried unti*

 tne

 dewpoint in the test section was reduced

suffic)"i'ily

 u> rtvoid

 condensation effects.
PRESENTATION OF RESULTS

Section force and moment coefficients of the 11-percent- thick airfoil (airfoil 5) are

presented over an extensive angle -of -attack range in figure 6. Data are not available for

the 10-percent-thick airfoil. with the blunt trailing ed^e (airfoil 9 as described herein)

over an angle -of -attack range as extensive as that for airfoil 5. Data are available

(ref. 2) however, on an airfoil with the same section coordinates as airfoil 9 but with' a

cavity in the trailing edge (airfoil 9a). Since the effect of the cavity on the aerodynamic

characteristics is not large, the aerodynamic characteristics of airfoil 9a presented in

figure 7 provide a good approximation of the characteristics of airfoil 9. Figure 8 presents a direct comparison between the aerodynamic characteristics of airfoils 5

(11 percent thick) and 9 (10 percent thick) over ar. abbreviated angle-of-attack range

around those at which the design normal-force coefficient for airfoil 9 of 0.7 occurs.

The drag- rise characteristics are summarized f;r A normal- force coefficient of 0.7 in

figure 9 and chordwise pressure profiles of the t^v Airfoils are compared in figures 10

and 11.

DISCUSSION

Aside from the difference in maximum thkJrx'SS, other geometric dissimilarities

were present which would prevent a direct, detail?-: comparison of the two airfoils solely

on the basis of differences in maximum thickness.. However, general observations concerning the results can be made.

Section Characteristics

Immediately apparent in figure 8 is a subs^rdal increase in normal-force coefficient for the 10-percent-thick airfoil when comn^w with the 11-percent- thick airfoil at

the same angle of attack. The slope of the normil-corce curve was not greatly affected.

More significantly, however, at the higher test Mien numbers (M = 0.78 to 0.81), the

normal-force coefficients at which the thinner axr; ?il begins to stall (manifested by an

abrupt increase in the drag level) is approximate 0.1 beyond those at which the

11-percent-thick airfoil begins to stall; thus, a .cssgn normal-force coefficient of 0.7 is

permitted for the 10-percent-thick airfoil 'commrsi with about 0.6 for the 11-percentthick airfoil. (See also figs. 6 and 7.) Althouci. r^'erred to as stall, separation onset

would be a more descriptive term since there i~ vt a sudden large loss of lift but a gradual stall beginning at the trailing edge (rear sercr'.cion) and reflected as merely a change

in the slope of the curve of normal force as a ir:. -on of angle of attack. The abrupt

increase in drag at a particular Mach number -v- iiso due in part to increased wave
losses. (Compare, for example, the pressure profiles for a = 1.0° and 1.5° at

M = 0.79, figs. 10(o) and 10(p).)

Figure 9 shows a drag divergence Mach number roughly 0.01 higher for the thinner airfoil at cn = 0.7. Such an increase in the drag divergence Mach number is consistent with the rule of thumb for conventional airfoils that changes in maximum thickness ratio of 0.01 result in changes in the drag divergence Mach number of approximately

0.01. The higher drag divergence Mach number at cn = 0.7 and the extension of the

normal-force coefficient at which stall begins for a particular Mach number are related

to the lower induced or perturbation velocities (less negative pressure coefficients) over

the thinner airfoil as illustrated in figure 11.

The peak in the drag-rise curve (fig. 9) at M = 0.78 was due to the second region

of supersonic flow on the upper surface (inherent in the supercritical airfoil concept at

intermediate off-design conditions, ref. 4) developing to such an extent that a second

shock wave was formed. (See, for example, figs. 10(k) and 10(1).) A reduction of this

peak would be achieved by decreasing the trailing-edge thickness (ref. 2).

Pressure Distributions

Lower surface.- The differences in the section characteristics of the two airfoils

may be largely identified with the noticeably lower induced velocities (less negative pressure coefficients) over approximately the forward 60 percent of the lower surface on the

10-percent-thick airfoil (fig. 10). These lower induced velocities may be attributed to

the reduced curvature of the thinner airfoil in this region (fig. 2).

Although the level of compression on the rear lower surface (referred to as the

lower surface cusp) was increased slightly for the 10-percent-thick airfoil, differences

in the pressure coefficients in this region were not as noticeable as on the forward region.

This was probably because the dissimilarities in curvature over this region were not as

pronounced as over the forward region and also because the flow in this region is not as

sensitive to small curvature variations.

Upper surface. - The effects on the rear upper surface pressure distribution (fig. 10)

associated with the dissimilarities in curvature (fig. 2) rearward of approximately the

60-percent-chord station (reduced curvature around the 50-percent-chord station and

increased curvature over approximately the rearmost 25 percent of the chord for the

10-percent-thick airfoil) are increased magnitudes of the second velocity peak (more

negative pressure coefficients) and a tendency for the forward shock wave to stabilize at

a slightly more forward location at intermediate off-design conditions for the thinner airfoil (figs. 10(1), 10(n), and 10(o), for example). Although the effects associated with these

changes in upper surface curvature cannot be completely separated from the influence of
the other geometric differences present, the results are consistent with the trend of the

data presented in reference 4 on the effects of changes in rear upper surface curvature.

The upper surface leading-edge velocity peak was considerably higher on the

10-percent-thick airfoil when compared with the 11-percent-thick airfoil at the same

angle of attack. If compared at equal values of normal-force coefficient rather than

angle of attack, good agreement in leading-edge peak velocities was indicated. (See

typical comparison in fig. 11.)

CONCLUDING REMARKS

Wind-tunnel tests have been conducted at Mach numbers from 0.60 to 0.81 to determine the aerodynamic characteristics of two NASA supercritical airfoils with maximum

thicknesses of 10 and 11 percent of the chord. This report documents the results of

these tests to provide a source of systematic experimental data for the supercritical airfoil. Although maximum thickness was the primary variable, dissimilarities were present which would prevent a direct comparison between the two airfoils. However, general

observations concerning the results can be made. For the thinner airfoil, stall (onset of

trailing-edge separation) begins at an approximately 0.1 higher normal-force coefficient

at the higher test Mach numbers, and the drag divergence Mach number at a normal-force

coefficient of 0.7 was 0.01 higher. Both effects are associated with lower induced velocities over the thinner airfoil.

Langley Research Center,

National Aeronautics and Space Administration,

Hampton, Va., March 16, 1972.

P^$P^PI^f!i!ffl p^
REFERENCES

t

1. Whitcomb, Richard T.; and Clark, Larry R.: An Airfoil Shape for Efficient Flight at

Supercritical Mach Numbers. NASA TM X-1109, 1965.

2. Harris, Charles D.: Wind-Tunnel Investigation of Effects of Trailing-Edge Geometry

on a NASA Supercritical Airfoil Section. NASA TM X-2336, 1971.

3. Blackwell, James A., Jr.: Aerodynamic Characteristics of an 11-Percent-Thick Symmetrical Supercritical Airfoil at Mach Numbers Between 0.30 and 0.85. NASA

TMX-1831, 1969.

4. Harris, Charles D.; and Blackwell, James A., Jr.: Wind-Tunnel Investigation of Effects

of Rear Upper Surface Modification on an NASA Supercritical Airfoil. NASA

TM X-2454, 1972.

5. Baals, Donald D.; and Mourhess, Mary J.: Numerical Evaluation of the Wake-Survey

Equations for Subsonic Flow Including the Effect of Energy Addition. NACA

WR L-5, 1945. (Formerly NACA ARR L5H27.)

6. Schaefer, William T., Jr.: Characteristics of Major Active Wind Tunnels at the

Langley Research Center. NASA TM X-1130, 1965.

7. Blackwell, James A., Jr.: Preliminary Study of Effects of Reynolds Number and

Boundary-Layer Transition Location on Shock-Induced Separation. NASA

TND-5003, 1969.

8. Davis, Don D., Jr.; and Moore, Dewey: Analytical Study of Blockage- and LiftInterference Corrections for Slotted Tunnels Obtained by the Substitution of an

Equivalent Homogeneous Boundary for the Discrete Slots. NACA RM L53E07b,

1953.

10
TABLE I.- SECTION COORDINATES OF 10-PERCENT-THICK SUPERCRITICAL

AIRFOIL 9 WITH 1-PERCENT-THICK TRAILING EDGE

fc = 63.5 cm (25 in.)l

x/c

0.0075

.0125

.0250

.0375

.050

.075

.100

.125

.150

.175

.200

.250

.300

.350

.400

.450

.500

.550

.575

.600

.625

.650

.675

.700

.725

.750

.775

.800

.825

.850

.875

.900

.925

.950

.975

.990

1.000

Calculated

y/c

Upper

0.0162

.0196

.0250

.0287

.0316

.0359

.0390

.0414

.0434

.0449

.0462

.0480

.0492

.0498

.0500

.0498

.0493

.0485

.0479

.0472

.0465

.0456

.0445

.0433

.0418

.0402

.0382

.0359

.0332

.0299

.0261

.0217

.0164

.0102

.0028

-.0021

-.0057

Lower

-0.0162

-.0198

-.0257

-.0297

-.0328

-.0373

-.0406

-.0431

-.0450

-.0465

-.0476

-.0492

-.0499

-.0500

-.0495

-.0486

-.0469

-.0442

-.0422

-.0396

-.0362

-.0316

-.0259

-.0202

-.0151

-.0105

-.0066

-.0033

-.0007

.0011

.0020

.0019

.0004

-.0027

-.0079

-.0123

-.0157

m

Upper

0.850

.570

.345

.255

.204

.145

.110

.086

.069

.056

.045

.029

.017

.008

.000

-.007

-.013

-.021

-.025

-.029

-.034

-.039

-.046

-.053

-.062

-.073

-.085

-.100

-.118

-.140

-.165

-.194

-.229

-.269

-.316

-.348

- 37 » u t \0j

Lower

-0.874

-.610

-.374

-.276

-.219

-.153

-.114

-.087

-.067

-.052

-.040

-.021

-.008

.003

.014

.025

.042

.070

.091

.120

.157

.206

.239

.216

.193

.169

.145

.118

.089

.056

.017

-.031

-.090

-.164

-.256

-.321

- 37 • v 10 \J

Experimental

y/c

Upper

0.0160

.0196

.0250

.0286

.0314

.0358

.0389

.0415

.0433

.0448

.0461

.0479

.0491

.0498

.0500

.0499

.0494

.0485

.0480

.0474

.0465

.0456

.0445

.0433

.0419

.0401

.0382

.0359

.0332

.0300

.0264

.0220

.0167

.0103

.0035

-.0016

Leading-edge radius: 0.0212c

Lower

-0.0165

-.0201

-.0259

-.0299

-.0329

-.0374

-.0407

-.0432

-.0451

-.0465

-.0476

-.0491

-.0498

-.0500

-.0494

-.0485

-.0468

-.0440

-.0420

-.0393

-.0357

-.0310

-.0255

-.0200

-.0152

-.0109

-.0072 -

-.0041

-.0014

.0005

.0016

.0016

.0004

-.0026

-.0073

-.0120

- 0• W1 i S V7 1

r- t

11
• '

«*®^^K*ra^».

TABLE H.- SECTION COORDINATES FOR 11-PERCENT-THICK SUPERCRITTCI^I,

AIRFOIL 5 WITH 1-PERCENT-TfflCK TRAILING EDGE

[c = 63.0 cm (24.8 in.)]

•>

?

x/c

0.0065

.0125

.0250

.0375

.050

.075

.100

.125

.150

.175

.200

.250

.300

.350

.400

.450

.500

.550

.575

.600

.625

.650

.675

.700

.725

.750

.775

.800

.825

.850

.875

.900

.925

.950

.975

1.000

y/c

Upper

0.0158

.0203

.0267

.0302

.0334

.0381

.0416

.0444

.0466

.0484

.0499

.0521

.0536

.0545

.0548

.0549

.0544

.0534

.0529

.0519

.0512

.0502

.0490

.0477

.0461

.0443

.0422

.0398

.0370

.0338

.0300

.0256

.0204

.0144

.0074

-.0008

Lower

-0.0157

-.0206

-.0271

-.0316

-.0351

-.0403

-.0440

-.0469

-.0491

-.0508

-.0521

-.0539

-.0548

-.0549

-.0541

-.0524

-.0497

-.0455

-.0426

-.0389

-.0342

-.0285

-.0224

-.0165

-.0112

-.0065

-.0024

.0011

.0039

.0059

.0070

.0069

.0056

.0024

-.0028

-.0108

m

Upper

1.000

.612

.376

.280

.225

.161

.124

.098

.080

.065

.054

.036

.023

.012

.001

-.006

-.014

-.023

-.028

-.032

-.038

-.044

-.051

-.058

-.067

-.077

-.089

-.104

-.120

-.140

-.164

-.191

-.223

-.260

-.302

-.352

Lower

-1.02-4

-.6815

-.4-LL

- 3 * ~

-.2-SB

-.1T-4

-.LJiD

- 1 ~n

-!o~s

-.OifO

_-.(}«*."

-.02:5

-.0" D

.003

.02.5

.0-45

.0*7

.LT-4

.121

.If. 5

.211

.2-t4

,2-t-9

.22~5

.201

.177

.Lr.2

.125

.095

.0*3

.023

-.026

-.037

-.154

-.250

-.231

Leading-edge radius: 0.0223c

'

•t

1

>

:•

(

•t'

H

&

K;

12

E

?^3;3TM!^^?S^?;p$!rr(l^tj^^

 tiT*^J
OF

PAGE 15

8

8

co

0)

o

-4->

O)

>J

co

0)

S-.

s,

13
ORIGINAE PAGE IS

OF POOR QUALITY!

1.0

.8

.6

m

.4

1

\

t

V

\

\\

\\

\

\ X

X<^ „

•^

^

- - ,.

— ^-:

_ .

-.

••

•

- •-. ;

J .01 .OZ ,(J6 .04 .Ob r,;

x/c

-

<VOmox

0.10 (airfoil 9)

.1 1 (airfoil 5)

i —

.08

...

.09 .1

m 0

-.2

-.3

-.4

J !

•4 -5 .G

x/c

.7 .8 .9 1.0

(a) Upper surface.

Figure 2.- Chordwise distribution of utvfoil surface slopes.

a

li

14
ORIGINAL PAGE IS

OF POOR QUALITY

-.2

m -.

-A

-.5

-.6,

I!

0.10 (airfoil 9)

.11 (airfoil 5)

0 .01 .02 .03 .04 .05 .06 .07 .08 .09 .10

x/c

m 0

-.1

-.2

-.3

-.4

.4 .5

x/c

.6 .7

(b) Lower surface.

Figure 2.- Concluded.

.8 .9 1.0

15
\Angle-of-ottack plate

A

L

_ . Tunnel center I ine

3.430c

T

.320c

low of orifices

-Trailing edge

Top view

-Model pivot

-Tunnel wall

r~

Zero incidence reference

.. Tunnel center line

-c=63.5cm(25.0in.)

i

;

 i

Tunnel !

window '

End view,section A-A

(a) Airfoil 9 mounted in tunnel.

Figure 3.- Apparatus. Dimensions in terms of chord of airfoil 9 (c = 63.5 cm (25 in.)).

16

-ffl^P!-.t*^^> . *. ---
Airflow

Static-pressure tube

.020 c

Totai- pressure tube

Diameter = 0.002

Static -pressure tube — ' —

Tubes flattened and staggered I

.240c

.I20c

.960c

(b) Profile drag rake.

Figure 3.- Concluded.

17
s

-7^P^

'- ''" ' ! '

 /

co

CD

tr>

18
8.0

7.8

7.6

1 7.4

c

ttF 7.2

7.0

6.8

6.6

X I O

/

/

5

/

'

/

/

/

/

/

/

/

/

/

/

/

s

/

i 1

.60 .64 .68 .72

M

.76 .80 ,84

Figure 5.- Variation with Mach number of test wind-tunnel Reynolds number

based on chord of airfoil 9 (63.5 cm (25 in.)).

19
ORIGINAL PAGE IS

OF POOR QUALITY,

0 .2 .4 .6 .8 .0 1.2 1.4 1.6

Figure 6.- Variation of section drag coefficient, angle of attack, and section pitchingmoment coefficient with section normal-force coefficient at various Mach numbers

for 11-per cent-thick supercritical airfoil (airfoil 5).

20
1=0.60

a.deg o

0

M=0.65

1 = 0.70

M=

1 = 0.76

'=0.78

= 0.79

M=0.80

M=0.£

r

•4

.81

M=0.60

r

.78

.79

.80

OEIGINAi; PAGE K

OF POOR QUALITY;

.70

IP

.76

.2 .4 .6 1.0 1.2

Figure 6.- Continued.

.8

c

n

1.4 1.6

21
f^*4^^$$^'**i

*

s

!

s

^

1H*^

.04

0

0

0

0

0

0

0

cm

o

0

-.04

O8

-.12

-.16

.20

-.24

C

22

ORIGINAL PAGE IS

OE POOR QUALITY

M =C

M=_C

M = (

M=C

M =

M.

M=C

M r

M=C

).60

).65

' .

D.7C

'.74

D.76

D.7J

).79

i.80C

).8I

Q—

3

Q

i

—a- fT} F

Q

-o —

O

Q-,

— <3—

D — E

-a —

-o— ,

•^

-Q"

)— C

— CK

^

"^o-

•Q

^

r-0"

— -i

^

N

D

^

-a

.81

*>**•

^

l_r

3^_

>a

N

s^

.80

JS

3—-

— ^

-^

S.

s

a

.79

^

x^

1

.78

) M =

— — —

^— ^

X

N>

o.e

a6

3=

^

.76

D

5

y

 .;

.74

\

j

'0

) .2 .4 .6 .8 1.0 1.2 1.4 1.6

&

Figure 6.- Concluded.

~xxnfrm.n. LJ

l l

1 1

1 1

i I

1

\ I

!

! f

1 1

1 1

*' %I

1 41

t 1 r 1

1 I

1 1

'- 1

:: 1
ORIGINAL PAGE IS

OF POOR QUALITY!

.UD

.04

.03

.02

.01

0

0

0

0

0

0

0

0

0

M=0

M<

M=C

M=C

M = '

M =

M=C

M=<

M=C

.60

).65

).70

).74

D.7€

0.7J

).79

D.8C

).8I

3

'

.

G

Gu

D —

^

•*

o —

^

•

-^

-H

\

C

^v "(*

-(Ty —

>

-,

> —

y

r

j.

o — ~

,

4

^

^x

J^

s

0

V

Gf

y

/

y> .8

.81

M

/

^

x

^

/

X

-^

f

(p. 7

0

=o.e

/

c/

/

1/1

/

D

^

.78

9

0

0

i

/

<^

x

/

.74

.76

.65

f

1

1

/

r

/

/

1

.70

f

• 2 .4 .6 .8 1.0 1.2 1.4 1.

Cn

Figure 7.- Variation of section drag coefficient, angle of attack, and

section pitching-moment coefficient at various Mach numbers for

a 10-percent-thick supercritical airfoil with 1.0-percent-thick

trailing edge with cavity (airfoil 9a). Data from reference 2.

23

*

*

^^
a,deg

i -- ••

r

fei

Figure 7.- Continued.

24

B9?P^!^^?3S^!®WP^^^wS5^»!!i*SSR^^P^^1^^ il^Wl^pJ^S^^^Ss^^^^^B^^S^y
-'-.'

.04

0

0

0

o

0

o

0

cm

0

0

-.08

-.1 2

.16

-.20

-—

M=C

M=C

M=C

M

=C

m —

M -

M.

M=C

.60

).65

).70

).74

0.7(

O.7

D.7<

0.8C

).8I

3

3

)

(

G

a

^ —

> — .

1

 — ^ —

— ~

1

— • — .

— c

— .

% r v

r~<

o_

\

\

Q^

&*

rr

r

y^

>~~~~~-

^

3-0

r^

^>

V.

.81

f

s

ti.

.80

-er"

— — -"

N

N

-o

.79

o w

f>— -

\

x

.78

=0.(

-<J

a,

^

\

\v

50

.65

^

%

^

.76

O

.74

.70

/

0 .2 .4 .6 .8 1.0 1.2 1.4 1.6

Figure 7.- Concluded.

<«

fTi <?hTiT 1 TII^III « r25
-.12

-.20

.016

.012

c

d

.008

.004,

**.

Figure 8

26

.•^OQNJQB^eN^IAlJ"'

-J /

r/

S-*

\

r

J

S

f-f

S*

r(

/*

=&

— C

r

/

c

S

==*

r^:

/

/

\z-~-

----

/

*FV\J

t—l

(_J

.

(t/c)max, percent

o 10

a 1

. -

4

3

a.deg

2

1

> .6 .7 .8 .9 1.0 I.I

c

n

(a) M = 0.60.

.- Comparison of supercritical airfoils with maximum thicknesschord ratios of 10 and 11 percent.

zn"n?xrru

*._

;piJB»liii!!ffl^^
^.^vsMttss*^^ m

-'

cm -.16

~

.016

I .012

1

 cd

I .008

1 nnd.

— —

L

r

J

X

— D

i— t

x°

— §

-(

_ — •

f

/

/

^d

sJ

u

/

^

O^"

n/

^J

o

x

/•^

^

^

^

/ \

^

si

^

i

X

3

(t/c)max percent

0 10

a l l

y

)

D

4

3

a, cleg

2

1

1

i

w

^

1

1

%'&•

i'i

1-

S|

£>

l|

rl ? •» i" *

^ i*

1'^

B

II

V •*!

!l

14s

rl

H

ij&?,:••*

3j/ .- • 5*

1^.1

^

 t:

\ .-

U

.7 .8

Cn

.9 1.0 I.I

(b) M = 0.65.

Figure 8.- Continued.

27
QN^I&ENT-tA^P

-.12

cm -.16

-.20

.020

.016

.012

.008

.004

(Vc)mox, percent

0 10

n i l

-&

a

aa.deg

.5 .6 .7 .8 .9 1.0 I.I

(c) M = 0.70.

Figure 8.- Continued.

28
-.1 2

cm -.16

-.20

.020

.016

012

.008

.004

0 10

D M

,5

^

2 a,deg

.8 .9 .0

(d) M = 0.74.

Figure 8.- Continued.

29
-.16

cm -.20

-.24

.020

.016

cd .012

.008

.004

(V0max,percent

o 10

a 1 1

.5 .6 .7 .8 .9

a,deg

0

p

.0 I.I

(e) M = 0.76.

Figure 8.- Continued.

30
-.16

cm-.20

-.24

.020

.016

c d .Q|2

.008

« *.

•flffff TFTIJl'JJ 1'IKIL

r.i

O—

p-~^__ b^

(t/c

)max, percent

o 10

a l l

rr

'•

^

r

6

^-^"^

' ~

^

j

—

©

-;

f

^

V "

^-^

^

^

/

^L—

^T1

0

'

f

/ .

1

^^.

/

/

/

/

/

?

0

"N

> .6 .7 .8 .9 1.0 1.

cn

3

2

a.deg

1

0

-

1 f

i-

»

i

f

?'

•>?^

;

1

[.

t

r

r

i

I

I

i

?

:

If

>

(-•

;

I

i.

:

'•

r

1

f

(f) M = 0.78.

Figure 8.- Continued.

I

31
*!Bp§*^il&gfPSBg&^^

-.16

Cm ?O

^m •*-v-

/

-.24

.024

.020

.016

cd

.012

.008

.On4

—MeeOl^lDBNTrA^-*

^i?=

(t/cW)x> percent

c 10

a l l

-f^

D-

•• — t ^°

1 <

 — ~— .

^

v-" J"

~\ , 1

 _

r-r^

•— ^— '

-^"

ris<

•~^-^.

x^

"^

/

X

'

^

,-X

x-U

^-^

J

/

/

/

f*' ^

r~

S

-^

/

/

/

/

^>—

r/

^^

/

/

/

p^-N

/-v

3

2

1 a,deg

0

L

c

S

s --

1!

I

 -

b'

t

jt _

>

r

t

?

ii

1

i

S;

" .

i

fL

[

L1

?

i

1

t'

I

(

1

P

1

F

i

^

1

1

&

k

.6 .7

Cn

.8 .9

(g) M = 0.79.

Figure 8.- Continued.

1.0

I

32
^^d^^^t^^^^^St&i^i^^^^l^a^^S^

-.16

cm -.20

-.24

.028

.024

.020

.016

.012

.008

.004

o 10

a 1 1

P

er

 '

cent

.5 .6 .7

Cn

.8 .9 1.0

(h) M = 0.80.

Figure 8.- Continued.

33

s?ajpss3p3^^
-.1

-.20

-.24

.024

.020

.016

c

d

.012

.008

f

(t/c)max,Percent

o 10

a l l

p

a,deg

.5 .6 .7 .8 .9 1.0

(i) M = 0.81.

Figure 8.- Concluded.

34
cd

.018

.016

.014

.012

.010

(

f

/

c

)max>Percent

10

II

.008

.006

Figure 9.- Comparison of variation of section drag coefficient with Mach number

at section normal-force coefficient of 0.7 for supercritical airfoils with maximum thickness-chord ratios of 10 and 11 percent.

t

(j

i

?

I

A •.

ft

r

35

^giPJJiy^
||!|pasata^aSteiaiS8S8fiai8a^

o

00

CD

O CD

"V1 ~

-

*

/7\

f

O

i

,

rP

CJ

O ( D

ii

" rLnJ n

1

•^^

-H C

Oj (I)

CVJ « «

•S S

h 0,

• !H 2

O <U ^

o a ^3

(^ I 3 r?

CM 00 <j- o «- 03 CVJ II ^Q

- i ' • - .« a -H •

,?0 • - § °

 • _ • O 2 o CQ

< >

o

CO

Q.

0

1

3P

O CD

CM -

.

36

G

O

O

o

o

o

.-.

o

^^

o

o

©.

o

o

0

\

<$r

n

El

El

B

_

El

B

_

a

a

P- i

D

H [

D

D

fe

D

n n

o

oII

0

X.

-* - v-'

_ J .+3 O

II 5 g

"M /S

00

 ^ • 1 1

3 CQ

co co

CQ <U

CD a o

CO 2

X £•§

^j- o

0

o

OJ a>

» . fn

g) .

0

C\J 00 ^ O ^" 00 CM

— — ' '

Q.

0

flTWMTT^ £gffiT=.

*.w^'^'>K^^^^^^^^^i^:fa^^^^^^t^^f^^»^^^^''^'^3^-{
0) <

0

0 0 & 0 1w t/ 1> a 1 £3o a

O

^

(

0°

(2 /

r

onQnOO

/-\ O

(-}

\~) ooD

O

Paaaa1

-hb

Da

u

 ia

ta. n

• onXo<

o

CsiI

CVJ

00

00

O

00

<\J

CV1

oLOII .

5

DM' 1

O

c. O0k1

Ccu O>^

C

(

0•7

/•

j

0

cL

O

qr

n

oonoo©_ 0Pi oD ^

O

r.

fr

^DDEDB_ ar \ c

aD

°E

D

u -Da ^

faD-i 13 r•

nc

0•

__• o11Xo^O

oco•oII 00

oOio

j1<u^a

CM

Q.

O

37
ORIGINAL PAGE IS

OF POOR QUALITY;

9) O>

0 0

£ o

1j 1in in

a> O>

Q. 5

Q. 0

ID -J

0 D

°r

(

-f f

r

f

G

©

_

>°

f

O

0

/"}

o

/-\

o

c\

O

0

o'

D

 U

D

Q

a

Q

TD

D

D

D

Q

D

Fb

a

n_

•

0

UX

<

cv

Q

CVl

IP CVJ 00

*'

I

00

00

CVJ

1

< ? o

Q.

O _^

o

0

CO

CL

O

1

1

1

1

1 %

c

o

o

O

o

o

o

©

_

o

o

-\

)

o

(&*-

a

B

.

E

C

E

C

B

^ C

H

a t

Q

1 j*1 1 — 1 f~\

O

0

II

x

o

o

c

CO

o

II

00

3

U

O

O

(U

(4

§,

^JcJ

1

O

oJ

i

CD_

1

OJ

_*

1

00

I*

1

<-

r

Q.

o

00 00

38

PplpAA^pii^^
J^st!lleKSKSi

*

0

Q> 0)

0 0

5"*- ** o~

0) 0>

aQ-. S O

0 0

3

 r

— fi***®8*TM^ ORIGINAL PAGE IS

OF POOR QUALITY

o

r o CD

f cvJ -'

? G e

o

u

0°'

1

©,

0

)

o

o

o

o

0

o

D

Q

a

a

B_

a

B

.

Q

JQ

u

D

3

Q

— i

r

6

X

o

C\J 00 ^- Q ^j. CO o

1

 Q.

o

_<-v

'

C

J

/)_

cC

0

C

9

,

f

o

>

0

o

o

o

o

0

o

\

o

o

&r

B

.

Q

G

i

a

]

D

T

]

Q

D

LJ f

E

3

 a

Q

Cl rn

g

c>

"x

a

.

00

<

x

<tf

•

»

O*

in "g

J° » ^

* AJ r^

. Q KO

.. u

q § -.

0 S

n a>

CO §

s

 fe

•

x

M-

•

CM

•

<~«

39
iO) 0>

•E °

V) CO

o> <5

Q. 5

o a

0°

U

D CO> CVJ 0

NJ ~r ~r"

(

/

C

c>

f

(.

o

>-\ ~j

o

o

0

o

)

o

E

r

E

E

C

o

 L

(-

r

 Q

D

u

Q

a

-j

d

n

3

o ' '*• o <t OD c\

I* • '

CD

ID

SB.

f.

r

O

<

•GP

fO

i

o

'c

o

cC

o

1

1 c

1 0

o

3

O

O

o

o

D

O

(T

Q

D

a

D

a

D

a q

•-U (

E

-

^ r

a

n

rr\

g

dnX

0

•*-•

CVJ

0

g

d

it

5

. i*

S§

o

a.

CD

OJ

cvi

I

00

*

00

CVJ

O

ex

o

40
0

> 0>

S'o

1W <r 1t

8. 1 S3

.0 a

8

l

r

Q

< G©_ 8

f

oQ)•\J0^o

/"vo1 D

O[cI

G

L

1]DD•n^D

a

D

u

ar

%a-ii

• ouX0<^f

O

cvj1

(£>• T

CM . — '. \

00•11

«-1* a

.

0

qCMi

COi

CM•1

00r

•1Q

.

o

00

00

CM

CM

(€?

/^N

p.

((

OO

<n_

CL

Oh1

(

©

.

O)-\j3O0O•\)0

0

.<§»*

DQB_ aS_ aanDE

aa

 a

U [ . Q n

hQ'n

(— i f~\

o• o

• IIXo§

00

CM

CM

E- OII

•oo>3IOId»H

(1)

bCs

41
PAGE

0> 0)

0 0

€ -S

3 ZJ

u. >-

d> 0>

Q. $

0. O

O D

0°

)

)

D

O

(

(

c

f

o

)

3

3

O

s

j

)

i

O

0 ^

D

Q

Q

D

4:

D

a

d

E

a

a

CD

X

OJ

I

CD

1

OJ

1

00

r

sjr

a.

O

00

oo

•

oo

Q.

O

00

CJ

OJ

/

o

1

o

0

ex

O

•

O

-\

j

O

n '

0

O

O

o

O

O

^

a

a

G

r

a

a

a

Q

0

u

G

D Q

a—

n f i

o

o"

II

1

;§.

n

ob

es

II

o

co

O

o

o X

CM

42
ORIGINAL PAGE p:

OF POOR QUALITY

0) <0

0 0

o o

H~ *|

w- k_

trt m

. k.

o3 o>

53

o °

$

(

o

(

©_

o

)

o

•

0

Q

0

•N

0

c r\

J

a

Q

Q

a

B -

a

i

?

 C

0^

1

1

D

m

a

f

Q

a

a

D

6n

X

o

2

O COC O CO <fr O <v)- CO C

• • • * » *

O

^ CO CO

CO - -

I I 1

00 ^f O <3"

• •

1 1

a.

O

oo

<

X

OJ

O

0

§

,

'

0

o

c

o

CL

O

1

^

3

O

O

O

1

 C

O

(

f

I

E

f—>

(^

]

a

3

]

3 Q f

3

)

LJ

a

n

a

n

on..

o

6

n

X

0

CO

CO

CO

o

in

co

o

n

s

tl

o

O

43
0

>o

,pv_

</>k_0)§

: ID0

Dvii

<uooH—w.

IS) k_0)o.3D

(1:>•1

0J•1

o

/*^

d?

r)V3

0

/•

O

/")G/(

r-\

Dr

o

•\)-^o/^oo• »; f

•*.

^~\

Qa.anaa_ Qa_ ' ti r•

>

ot

n

Q

C

D

3

u

a

a

sj

aDD-| 1-•

a0

•onXo<C

V

CD

f.

i'

<X

CM

a.

O

(i

©_

_

oo^5D

(

•}

O©

.C

0

(/) CL ^

f"

)?

/-\oQOf

-C

^i o)

[

($• aCDa- aa]-i^,

D'a

LJa

nt-

Q

__f

-i

O• OIIXo^

s*\

O _J

O T3

• 0)

>H 3 n .S•y ?J GO .. o ^ '. d 2 II a,

CO

§>

*-•.!

CVJ

q

cvi C

D CVJ T

CDr

CD CVJ

Q.

O

44
CVJ

I

CD

-

I

CM

-f

1

00

f

1

00

-rT cvj

O

§

w

o

0

a

>*—

to

<5

o

D

:> CD

sj* — '

I

c

1

o

•)

)

1

1

r

o

(

O

o

Q

0

..

' c

[

[

c

Q

C

[

C

°

l

]

:

]

ib

m

D

U

E

9

. _

a

D

G

D

n

CM GO ^f O *f

• • .

1 1

a.

0

O

O

1

o

c

o

o

r

o

CO

a.

0

Q

)

o

0

(

1

o

•)

r-\

D

Q

B_

a

B.

a

Q

D

E

a

3

u

C

«

\

D

n

6

n

X

o

oo cv

Q

O

o"

n

X

o

CO

CD

<X

00

CD

o

in

o

II

T3

a»

o

U

i

•

o

»-i

0)"

CVJ

e^o

a.

O

45
ORIGINAL PAGE IS

OF POOR QUALM

<1> 0)

0 0

1

ID W

 1>

1 1

£3

0 G

fe

0

(^

o

/-

c

n

9

~\

a

o

)

(-)

o

(^

<

e, -

D

m

)Q

Q

?

TV

3

°

v

)

Q

a

a

a

Q

a

a

-i

•

OII

•5

D CO <NJ 00 <j- Q ,-j. 00 cv

V .-' 7 • r

1 Q.

o

00

CVJ

5

j

1

Ci) cvj

O

G

O

§

.

t

a

S

o

<

o

'c.

o

CL

O

?Q

D

r

D

C

f

•**!

>

-\

O

'C

n

 G

E

3

n

a

h

r

i

i

D

D

i

'a [

c

D

I

)

1

D

B

-

r;

<d

hS

,.

a

r aD

0

n

x

i^

0

(TO

•

•

«"f

•

OJ

*

J

•

o

n

00

c-(

o

II

0)

3

O

O

a.

o

46
0) 0)

8 8

•si in o>

i_ in) <O

a. 5

Q. 0

ID _l

O G

> Cf

j

1

(

G

O

$

o

e.

3

O

D

O

G

a

a

a

a

UQ

D

0

"

(

}

a

i

) CVJ CO ^ O

•' "' '' k . "

0

- 0

c

6

0

9

,

*~

O

o

o

^\

o

)

o

'c

o

CL

O

-\

O

1 C

e.1

i

r

1

.

e

- ,

j

c

<$"

a

D

a

Q

i

3

n

u

r

a

a

a

a

d

.§

^ oo c\

,

• ..

a

— ra_

a

3 — C ]

0

6

nX

o

CD

CVJ

I

CVJ

1

00

."

1

00

00

CD

o

X

CVJ

%>

g

-d

II

§

CD

C\J

,

fP

ex

o

47
0> 0)

0 0

•g £

(O (O

1. 1 5:3

o n

i

r\

0

Q

0

o

f\

0

(~\

^J

o

n

*0

<

c

)

o

G

H

Q

r-

[

E

I:

o

l

o

1

]

Q

n

E

D

D

CD

a

OII

X

o

£

O CO (M 00 <j- o ^- 00 O

• ••••. * • . • • -

M T T 1 1

00

D CiD C

(

f

<s

C

/

(*

vJ

O

©,....

o

o

0

r\

>

>

1

c0

G

G

U

•c

)

)

5

n

a.

<?

Q.

)

3 °

t

r

<&

Q

n

a

a

CD

a

Dac

c

\

a

'a c

3

u

C

3 O

<;

%

a

i

r

rn

C0

0

d

u

ZZ

c

o•

(JO

CD

*

vl

•

CVJ

•

u

VJ

o

ift 73

-

1

' I

II .g

O CH

o

... O

°C3- 1^

O 2

11

 0

a

- H

<u

I

Q.

O

48

^
i i

4> <U

iJ ° 2 °

?

w v 1>

1 1

33

0 D

©

_

9

6

_

C

r\

o

r\

O

Q

1

h

D

•s

O

©

_

G

a

D

a

a

V

o

ot

a

m

LJ

CD

D

Q

a

6MX

o

g

D CD C\J 00 <fr O ^f 00 0

i i 7' '' iQ.

O

CO

O

(

(

/

\

C

r

£

o

c

L

.O

O

(/)

CL

0

ol

5

">

)

fe

O

(*\

Q

r

\

9

_

s 1

^1

1

1

1

C

[

\ r3 •

r

E

a,

Q

^

3-

)

]

j

3

b Q

D

U |

Q

n

fe

Q

nn_

o

o*nX

o

<

/^>

O ( Q CV J 00«4 - O ^ 00^ 0

~ T Til - • -

oo

o

II

3

oT

o

O

I

o

f-4

<U

E

s

CVJ

? . j

f

tQ.

O

49
a) <

u

o o € € 3 a 1Q. 0

 1

ID -1

0 D

CfC(

/-

1

\JG

/-j OO1

(

>

i>. i o>o

QnQnQ— , aB_, PHD

^>

^a 3

a

a ~f a m

Q

aB_. ai

• ong§

CVJI

toI

OJ•I

00r

a.

O

CO

q

c\ji

CD• T

OJ• T

CO

00

Q.

O

00

CVJ o

Oo

OJ

(/(Cc0el1 oe_ oA$

£

(~\

)D, 5-\j)

3

/-

\_'cwco\>i

.

o

 GE

^

3aa3]%

a

n

LJG a .. n

%B_ El

O.. C

D

o• OIIX <— o

t-• oII •aCD

•.dOO0)(H

00OJ

CVJ

50

^-^^^
CD 0)

S 8

1-

</> (

1A

<u o>

Q. 5

Q. 0

ID -J

0 D

C

o

0

n

O

e_

O

0

n

<£

O

o

/-

0

D

P

S.

P

E

r-

§C

0^

'°

E

Q

CJ

E

D

%

Q

«

O

ii

X

0

§

CO

o

X

CD 00 00 CM

CVJ

I

I

Q.

o

i

(

D

e

i

O

9

_

O

e

.

o

o

D

b~\J

6

_

o

'c.

o

CL

O

1

1

1

1

1

1

1

1

©_

3

\ •

D^

D

Q

Q

Q

Q

Q

S

.

QG

Q

Q a

LJ

G

A

D CO CM 00 tf O <d * • . * ^-* ^ \j T - i r

\

El

n. ri

0

d

"x

§

/-\

h oq c\

• ^ „

o

IO

II

a

6T

c-

•

o

II

S

TJ

CD

3

I

O

0)

00

CVJ

o.

o

51
•••&'-• »j&£*Q&^fes^^^^ - • gUinSrfiKa'FHWgMa'

C

O

•••MB—^-—-,,.—,.- ORIGINAL PAGE IS

^^^@tff®

S4^^BW5B

 OE POOR QUALITY

0

0> Q)

0 0

1to c/ I)

»_ i_

a> o>

Q. 3

O D

)

j

^

o

o

o

o

o

Q

§

%

•>

G

£

1L

CD ^ 00 ^

C.

O

'E

o

V)

CL

O

k

(

Q

^

O

0

3

^

•\

0

"" H

Q

Q

11

1

3

0°'

a

n

u

Q

Q

o"

II

X

o

^ o <j oo o

Q.

o

 e

D E

C

C

a

~

B

&

)

1

3

3

a a

3

r

LJ

a

i

D

-Q-O

O

6

X

o

'o

/-*

00

CD

X

CVJ

•

O O 3

i »a ^ §• •

.„ 0

3 § d ° iH

II o

CD ""—'i £ ^^ ^j* ^^

• "*»*''

CD

X

CO

/-N

!

>-

I

j

!'*

!:-•'

|;V

t .

&"K

K

t;"TS'

i

1

52

O (£)

cvi —

i i

CM 00 <fr

T" i* r

a.

O

0 00 CVJ
ORIGINA

L PAGE IS

OF POOR QUALITY

Q>0s§t_Q) Q

.

IDO

5J

<o0oH—k_col_<ua

cr

)

oj•

e_^ o©_Gco(<ccG/•1ka)•v)-)i!>Dr

<s

c

O

DaaaUfDOr

</

"i

a

C

CD

5

QQ 3 ^

%a

r•

aD

oIIo<c\

OcJi

ID• T

00

Q.

O

00r

o.

o

00

00

oo

O -H

e, oc/(<(0oG o /

}\)-i)\>

o

'cof Q. O

(

I D

&

D3DQDI]

•QC

D

u

E

n

PfeDo r

~

i

c5Xo\O

ff\

UJ••

r-f-

••

r^i

oo

c>II oUio

CVJ

..f

53
0) 0>

0 0

S o

1

«/) (

 1

A

8. 1 £3

O 0

(

>

C\

o

/-s

O

s-}

o

<^\

o

Q

o

(->

1

^

(-}

(

o

a

a

a

B.

a

B.

#u

r>

o

6

'

~!

D

Q

D

H

a

a

a

a

a

n

*

o

II

X

o

<

O CO OJ 00 ^- o <t 00 <\

cvi 7 7 ' O

O

CVI

1

:A

CD CVJ CO ^j-

« • •

7 T ' i

o

a

0 00

00

CVJ

0

o

O

0

O

O

Q

O

\

o

'c

o

CO

Q.

O

1

nc

0

I

(^

3

PI

Q

a

D

3

a

a

3

] a

u

a

a

^

O

o*

"x

• •»— •

03

CO

2s "O

0)

T^ 'Q

ii -3

o

. o

... o

o

oo '.

o 2

II

s?5^v«;ft?3iwi:^^ ^'''^l. TM":s
ORIGINAE PAGE IS

OE POOR QUALITY*

-1.0

"0 .1 .2 .3 4 .5 .6 .7 .8 .9 1.0

x/c

Figure 11.- Chordwise pressure profiles at M = 0.60.

NASA-Uuigley, 1972 1 L-8173 55

IMP^U^iilJIM^
iSBPiiaagsiBMjp^
MA^

'The aeronautical and space activities of the United States shall be

conducted so as to contribute . . . to. the expansion oj'human knowledge of phenomena in the atmosphere and space. The Administration

shall provide for the widest practicable and appropriate dissemination

of information concerning its activities and the results thereof."

— NATIONAL AERONAUTICS AND SPACE ACT OF 1958

NASA SCIENTIFIC AND TECHNICAL PUBLICATIONS

TECHNICAL REPORTS: Scientific and

technical information considered important,

complete, and a lasting contribution to existing

knowledge.

TECHNICAL NOTES: Information less broad

•in H<>p e but nevertheless of importance as a

loiitributio n to existing knowledge.

T HC H N 1C A L M EMOR ANDUMS:

Information receiving limited distribution

'xv.uise of preliminary data, security classification; or other reasons.

CONTRACTOR REPORTS: Scientific and

uvhnical information generated under a NASA

oinir.ict or grant and considered an important

contribution to existing knowledge.

TECHNICAL TRANSLATIONS: Information

published in a foreign language considered

to merit NASA distribution in English.

SPECIAL PUBLICATIONS: Information

derived from or of value to NASA activities.

Publications include conference proceedings,

monographs, data compilations, handbooks,

source-books, and special bibliographies.

TECHNOLOGY UTILIZATION

PUBLICATIONS: Information on technology

used by NASA that may be of particular

interest in commercial and other non-aerospace

applications. Publications include Tech Briefs,

Technology Utilization Reports and Notes,

and Technology Surveys.

Details on the availability of these publications may be obtained from:

SCIENTIFIC AND TECHNICAL INFORMATION OFFICE

NATIONAL AERONAUTICS AND SPACE ADMINISTRATION

Washington, D.C. 20546

r

r

L

•|;

.*.

^Klusy^.a^u^.HJi.Ll^KJ'i.