High Reynolds Number Transonic Tests of an NACA 0012 0.3 ...€¦ · APPARATUS The test program was conducted in the 8 in by 24 in (20.32 cm by 60.96 cm) two dimensional test section

.-

NASA Technical Memorandum 100527

,

High Reynolds Number Transonic Tests of an NACA 0012 Airfoil in the Langley 0.3-Meter Transonic Cryogenic Tunnel

Charles L Ladson and Acquilla S. Rill

(PAS&-TI-100527) I I X E RElfBOLDS IiDEBEB Y88-1E571 f 3 A l S C I I C 'IESIS 01 d IOaCB CO1; AIBPOU I1J

r u b m m ( N A S A ) 150 p CSCL O 1 A Unclas IEB I A P G L E Y 0-3-EBIEB I E B Y S G Y I C CBXCGBNXC

G3/02 0130053

DECEMBER 1987

National Aeronautics and Space Administration

Hampton, Virginia 23665 ---

https://ntrs.nasa.gov/search.jsp?R=19880009187 2020-05-14T05:05:09+00:00Z

SUMMARY

Tests were conducted in the two dimensional test section of the Langley 0.3 meter Transonic

Cryogenic Tunnel on an NACA 0012 airfoil to obtain aerodynamic data as a part of the

Advanced Technology Airfoil Test (ATAT) program. The test program covered a Mach number

range of 0.30 to 0.82 and a Reynolds number range of 3.0 x 10' to 45.0 x 10'. The stagnation

pressure was varied between 1.2 and 6.0 atmospheres and the stagnation temperature was varied

between 300 K and 90 K to obtain these test conditions.

Plots of the spanwise variation of drag coefficient as a function of normal force coefficient and

the variation of the basic aerodynamic characteristics with angle of attack are shown. The data

are presented uncorrected for wall interference effects and without analysis. This data adds to

the existing data base to be used for tunnel comparison purposes as well as assessment of wall

interference and correction procedures.

INTRODUCTION

A comprehensive test program has been undertaken in the Langley Research Center 0.3 meter

transonic cryogenic tunnel (0.3-m TCT) to obtain the static aerodynamic characteristics of a

series of two dimensional airfoils at transonic speeds and flight equivalent Reynolds numbers.

This program, known as the Advanced Technology Airfoil Test program (ATAT), was sponsored

by the Langley Aircraft Energy Efficiency Project Office and the Transonic Aerodynamics

Division. A description and status report on this program was first presented in reference 1.

The primary objectives of this program, as stated in this reference, were to provide the U. S.

transport industry the opportunity to test and compare their most advanced airfoils to the latest

I

NASA design at high Reynolds numbers in the same facility and to provide experience in the

design and construction of models for testing in a cryogenic environment as well as testing

techniques unique to this environment. Included in this test program were four series of tests: (1)

Correlation airfoils, (2) Advanced NASA supercritical airfoil, (3) Industry airfoils, and (4)

Foreign technology airfoils.

This report presents results obtained from tests of the NACA 0012 airfoil which was tested as a

part of the Correlation series of airfoils. Tests were conducted at Mach numbers from 0.30 to

about 0.82 at chord Reynolds numbers from 3 x 10' to 45 x 10'. To obtain these conditions, the

stagnation pressure varied from 1.2 to 6.0 atmospheres and the stagnation temperature from 300

K to about 90 K. Model surface pressures, wake survey pressures, and tunnel wall pressures

were obtained during the test program.

Plots and tabulated values of model surface pressure distributions as well as tabulated values of

the integrated force and moment coefficients obtained in this test program are presented in

reference 2. This paper contains plots of the spanwise variation of drag coefficient as a function

of normal force coefficient and the variation of basic aerodynamic coefficients with angle of

attack. Normal force and pitching moment coefficients were obtained from integration of the

surface pressure distributions and the drag coefficient was obtained from integration of the wake

survey data. The data are uncorrected for wall effects and are presented without analysis. In this

format, the data should be useful for tunnel comparison purposes as well as extending the

existing data base for use in evaluating wall interference and correction procedures.

SYMBOLS

All measurements and calculations were made in U.S. Customary Units; however, the

2

L

measurements are presented with the International System of Units (SI) in parentheses.

a

b

C

‘d

‘d,corrl

‘d,corr2

‘d,corrS

Cd,C01T4

‘d,corrS

cm

cn

M

pt

R

Tt

Y

angle of attack,deg

airfoil model span, 8.00 in (20.32 cm)

airfoil model chord, 6.00 in (1 5.24 cm)

section drag coefficient ( c ~ , ~ ~ ~ ~ if valid data exist, otherwise, c ~ , ~ ~ ~ ~ )

section drag force coefficient from integration of wake total head profile for probe located at y/(b/2) = 0.125

section drag force coefficient from integration of wake total head profile for probe located at y/(b/2) = 0.0

section drag force coefficient from integration of wake total head profile for probe located at y/(b/2) = -0.125



section pitching moment coefficient about model 25-percent chord point from integration of model surface pressure coefficients

section normal force coefficient from integration of model surface pressure coefficients

average free stream Mach number

free stream stagnation pressure

Reynolds number based on model chord

free stream stagnation temperature, K

spanwise distance from tunnel centerline, (positive toward right hand side)

APPARATUS .

The test program was conducted in the 8 in by 24 in (20.32 cm by 60.96 cm) two dimensional test

section of the Langley 0.3-meter transonic cryogenic tunnel. This facility is a continuous flow,

3

fan driven, pressure tunnel which uses nitrogen gas as the test medium. The test section is

rectangular in cross section, has solid sidewalls, and slotted top and bottom walls. Two slots are

located in each of these walls with a spacing of 4.0 in (10.16 cm) and a total open area ratio of

0.05. Further details of this tunnel and test section are contained in reference 3. - d

Data were recorded on a shared, remote central data acquisition system. This system is referred

to as the original data acquisition system for the two dimensional test section of the 0.3-m TCT

in reference 3 and further details of the capability of the system are given in this reference.

MODEL

The model tested was a two dimensional model of the NACA 0012 airfoil with a chord of 6.00 in

(15.24 cm) and a span of 8.00 in (20.32 cm). The equation for the ordinates of this airfoil shape

was first presented in reference 4 and the ordinates for this model were obtained from the

computer program of reference 5.

The model was constructed of A286 stain-dss steel which is an acceptable material for cryogenic

test conditions. To locate all pressure instrumentation tubes internally in the model, it was

constructed in two halves, the tubing installed, and then these two halves were bonded together.

By locating the tubes internally, the model surface can be maintained in a very smooth condition

which is required for high Reynolds number testing. One longitudinal row of pressure orifices

was located on each surface of the model near the mid-span location and three spanwise rows

were located on the upper surface. The ordinates for each orifice location are presented in the

tabulated data. An orifice diameter of 0.010 in (0.25 mm) was used for all locations.

.

Upon completion of model construction and after several cryogenic temperature cycles, the model

4

ordinates were measured and compared to the design values. The maximum deviation from

design was found to be 0.003 in (0.076 mm) with most of the model surface being within the

manufacturing tolerance of 0.001 in (0.025 mm).

c

TESTS

The tests of this airfoil covered a Mach number range from 0.30 to 0.82 and a Reynolds number

range of 3.0 x 10' to 45.0 x lo6. To obtain these test conditions, the tunnel stagnation pressure

varied between 1.2 atm and 6.0 atm and the stagnation temperature varied between 300 K and 90

K. Measurements of model surface pressure, wake total head surveys, sidewall pressures at the

wake survey station, tunnel floor and ceiling pressures, and various tunnel parameters were made

during the tests.

For these tests, the wake surveys were made at a station 1.2 model chords downstream of the

model trailing edge. Data were obtained with both free and fixed locations of boundary layer

transition. For the fixed transition tests, data were obtained only at Reynolds numbers of 3.0 x

10' and 9.0 x 10'. The angle of attack range for all tests was from zero degrees to that for drag

divergence. Limited data were also obtained at negative angles of attack as a check for

asymmetry which could result from flow angularity, model inaccuracies, or incorrect angle of

attack setting.

For the fixed transition data, transition strips were located on both the upper and lower surfaces

of the model at the 5-percent chord station. These strips were about 1-percent chord in length

and consisted of #240 carborundum grit applied with an acrylic spray.

5

TEST PROCEDURES

Model Pressures

All model surface and tunnel floor and ceiling pressures were measured using 48-port scanni-

valves connected to high precision variable capacitance type pressure transducers. Three samples

of data, taken at 0.25 second intervals, were recorded for each port. These readings were

averaged during the data reduction process. To provide for lag in the response time of the

pressure tubing and transducer, a delay time was invoked after each port change before the data

were recorded. This delay time varied from 2.0 seconds near the model leading edge where large

changes in pressure occurred between ports to 0.5 seconds for the aft portion of the model.

Wake Pressures

Surveys of the total pressure in the model wake were made using a traversing rake cantilevered

from one of the tunnel sidewalls. This rake contained five individual probes such that a spanwise

survey could also be obtained. The drive mechanism consisted of an electric stepper motor which

was synchronized with the scannivalve system so that the rake would advance one step in distance

with each scannivalve step. For these tests, approximately 75 steps were made as the rake

traversed the model wake region. Tunnel sidewall static pressure data was obtained at the same

tunnel station as the rake probes for use in data reduction. Individual pressure transducers were

connected to each of the probes as well as the sidewall pressure taps to keep response time to a

minimum, For this test program, the survey probes were located approximately 1.1 model chords

downstream of the model trailing edge.

6

Corrections

c

All data and tunnel parameters were corrected for real gas effects by the method of reference 6.

No corrections have been made for wall interference effects on the data presented herein;

however, these corrections have been made for some of this data and the results presented in

reference 7. These corrections reduce the test angle of attack for all conditions and the free

stream Mach number, mainly at supercritical conditions.

PRESENTATION OF RESULTS

The results of this investigation are presented in this paper as plots of integrated force and

moment coefficients obtained from surface pressure distribution data and wake surveys. These

data are presented for reference purposes only and no analysis of the results is presented. An

index to the various figures for both free and fixed transition data follows.

I. Effects of normal force coefficient on the spanwise variation of drag coefficient for various Mach and Reynolds numbers. Transition free.

Mach number Figure

0.30 0.40 0.50 0.60 0.65 0.70 0.74 0.76 0.78 0.80 0.82

7

1 2 3 4 5 6 7 8 9 10 11

II.

111.

IV.

V.

Effects of normal force coefficient on the spanwise variation of drag coefficient for various Mach and Reynolds numbers. Transition fixed.

0.30 0.50 0.70 0.74 0.76 0.78 0.80

12 13 14 15 16 17 18

Variation of basic aerodynamic characteristics with angle of attack for various Mach and Reynolds numbers. Transition free.

0.30 0.40 0.50 0.60 0.65 0.70 0.74 0.76 0.78 0.80 0.82

19 20 21 22 23 24 25 26 27 28 29

Effects of fixing transition on the variation of basic aerodynamic characteristics with angle of attack for various Mach and Reynolds numbers.

0.30 0.50 0.70 0.74 0.76 0.78 0.80

30 31 32 33 34 35 36

Comparison of the effects of Reynolds number on the basic aerodynamic characteristics for various Mach numbers. Transition free.

0.30 0.40 0.50 0.60 0.65 0.70 0.74 0.76 0.78 0.80 0.82

37 38 39 40 41 42 43 44 45 46 47

8

VI. Comparison of the effects of Reynolds number on the basic aerodynamic characteristics for various Mach numbers. Transition fixed.

0.30 0.50 0.70

48 49 50

9

REFERENCES

1. Ladson, Charles L. and Ray, Edward J.: Status of Advanced Airfoil Tests in the Langley 0.3-Meter Transonic Cryogenic Tunnel. NASA CP-2208, 1981.

2. Ladson, Charles L.; Hill, Acquilla S.; and Johnson, William G., Jr.: Pressure Distributions From High Reynolds Number Tests of an NACA 0012 Airfoil in the Langley 0.3-Meter Transonic Cryogenic Tunnel. NASA TM- 100526, 1987.

3. Ladson, Charles L. and Ray, Edward E.: Evolution, Calibration, and Operational Characteristics of the Two-Dimensional Test Section of the Langley 0.3 Meter Transonic Cryogenic Tunnel. NASA TP-2749, 1987.

4. Jacobs, Eastman N.; Ward, Kenneth E.; and Pinkerton, Robert M.: The Characteristics of 78 Related Airfoil Sections From Tests in the Variable-Density Wind Tunnel. NACA Rep. 460, 1933.

5. Ladson, Charles L.; and Brooks, Cuyler W., Jr.: Development of A Computer Program To Obtain Ordinates for NACA 4-Digit, 4-Digit Modified, 5-Digit, and 16-Series Airfoils. NASA TM X-3284, 1975.

6. Adcock, Jerry B.: Real-Gas Effects Associated With One-Dimensional Transonic Flow of Cryogenic Nitrogen. NASA TN D-8274, 1976.

7. Gumbert, Clyde R., and Newman, Perry A.: Validation of A Wall Interference AssessrnentlCorrection Procedure for Airfoil Tests in the Langley 0.3-m Transonic Cryogenic Tunnel. AIAA-84-2151, 1984.

10

0 -.3860 0 -.1950 0 -.0109 A .1694 h ,3544 b 5351 n .7187 0 ,7983 0 .8859 0 .96oi @ 1.0282

1.0656 0 1.1106 A 1.1575

(a) R = 3.0 x 106

Figure 1.- Effects of normal force Coefficient on the spanwise variation of drag coefficient for various Reynolds numbers at a Mach number of 0.30. Transition free.

11

43

0 0 0 A h 0 a 0 0 n e

e .048

.040

.032

.008

0

-.0113 -.3902 -.2020 -.0163 .1112 .3597 .5488 .7303 .8181 .9054 .9867 1.0672 1.1374

-1 .o -.8 -.6 -. 4 -.2

Y/b/2

(b) R = 6.0 x lo6

Figure 1.- Continued.

0 .2 4

0 0 0 A h n n

0 n

0

e ffl e A

Cn

- -4005 -.2061 -.0174 .1747 .3703 .5641

.7418

.8358

.9301

1.0086

1.0779 1.1558 1.1596 1.2011

.048

.040

.032

Cci corr .024

.016

.008

0 - 1 .o -. 8 -.6 -. 4 -. 2 0 .2 .4

Y/b/2

(c) R = 9.0 x lo6


13

0 -.4025 0 -.2134 O -.0181 A .1724 h .3676 [5 .5604 0 .7459 0 .8319 0 .9107 n .993i

1.0613 BI 11.1123

1.1520 A 1.1873

(d) R = 15.0 x lo6

Figure 1.- Concluded.

14

.048

.040

.032

C, corr -024

.016

.008

0

C n

0 -.0213 0 -.4073 O -.2075 A -.0104 h .1863 b .3708 0 3609 0 .7502 Q .8444 A .9236 @ .9925 a 1.0220

-1 .o -.8 -.6 -. 4 -. 2 0 .2 .4 Y/b/2

(a) R = 3.0 x lo6

Figure 2.- Effects of normal force coefficient on the spanwise variation of drag coefficient for various Reynolds numbers at a Mach number of 0.40. Transition free.

1s

0 -.4077 0 -.2065 0 -.0141 A .1819 h .3772 b .5706 n ,7582 0 .8518 Q ,9462

- n 1.0092 @ 1.0486 a 1.0620 e 1.0558

(b) R = 6.0 x lo6


16

.048

.040

.032

C, corr .024

.016

. ooa

0

-.4138 -.2160

-.0196 ,1783

.3763

3727

.7649

.8530

.9445 1.0173

1.0406

1.0566 1.0573

-1 .o -.a -.6 -. 4 -. 2 0 .2 .4

Y/b/2

(c) R = 9.0 x 106


17

0 -.4153 -.2146

O -.0145 A ,1844 h ,3827 b .5780

n .76n 0 .8592

0 .9440 n 1.0220

(3 1.0409 ffl 1.0519 e 1.0531

C i corr

.

Id) R = 15.0 x lo6


18

.048

.040

.032

Cci corr -024

.016

.008

0

0

0 A

h c\

0 n

0 n

cn

-.0123

-.4301 -.2235 - ,0144 .1918

,3971

,5989 .7880 .8623

.9125

-1.

(a) R = 3.0 x lo6

Figure 3.- Effects of normal force coeffwient on the spanwise variation of drag coefficient for various Reynolds numbers at a Mach number of 0.50. Transition free.

19

4'

0 -.0099 0 -A300 0 -.2165 A -.0135 h .1936 n .39% n ,6084 0 ,7925 0 .8767 n .93i9 @ .9649

,9754

(b) R = 6.0 x lo6

F i n 3.- Continued.

20

-.4352

-.0176

-.0144 .1946 .4009

.6100

.8030 ,8789

.9647

.9788

(c) R = 9.0 x lo6


21

0 -.4351 -.2232

O -.0158

A .1905 b .4017 h .6073

0 .8728 6 .9314

n ,7999

n .9553

,9515

(d) R = 15.0 x 106

Fiure 3.- Continued.

. 22

3 3

0 0 0 A h 0 n

0 n

0

e

cn

-.0116

-.4385 -.2284 -.0163

.1961

.4107

.6226

.8162

.8193

,8992

,9446

.9594

-1 .o -. 8 -.6 -. 4 -. 2 0 .2 .4

Y / W

(e) R = 30.0 x lo6


23

Cn

0 -.0234

0 4 5 7 5 0 -.2334

A -.0184

h .2063 0 .4240

n .ti337 0 ,7290

@ ,8118 , A .8520

.048

.040

.032

C, corr .024

.016

.008

0 -1 .o -.8 -.6 -.4 -. 2 0 .2 .4

Y/b/2

(a) R = 3.0 x 106


24

47

.04E

.04C

.032

Cci corr .024

-01 6

.008

0 -

En

0 -.0203

0 -.4595 0 -.2302 A -.0112 h .2089 b ,4287

Q .6564 . 0 ,7479

9 .8513 n .8987

e .9191

0 -.8 -.6 -.4 -. 2 0 .2 .4

Y/b/2

(b) R = 6.0 x 106


cn

0 -.4654 0 -.2340 0 -.0140 A .2078 h .4315 c\ .6587 Q .7503 0 .8306

. Q .8636 d .8831

.048

.040

.032

C, corr .024

.016

. OOE

C -1

(c) R = 9.0 x lo6


26

cn

0 -.4537 I3 -.2357 0 -.0163 A .2045 b .4287 [5 .6581 0 .7398 0 .E466 0 .8440

.048

.040

.032

C, corr .024

-01 6

.008

0 -1 .o -.a -.6 -. 4 -. 2 0 .2 .4

Y / W

(d) R = 15.0 x lo6


27

0

0 A b

b 0 0 0

0

cor r

-1 .o -.a -.6 -.4 -.2 Y/b/2

(e) R = 30.0 x lo6


28 ...-

Cn

-.0129 -.4768 -.2448 -.0144 .2113

.4393

.6644

.7631

.8412

.8648

0 .2 .4

Cn

0 -.0178 0 -.4768 Q -.2443 A -.0176 b .2163 b .4464 0 ,5534 0 .6636 0 .7426 0 ,7984

(a) R = 3.0 x lo6

Figure 5.- Effects of normal force coefficient on the sprnwise variation of drag coefficient for various Reynolds numbers at a Mach number of 0.65. Transition free.

29

0 -.0179 0 -.4770 0 -.2386 a -.Oil1 b .2180 D .4471 0 ,5623 0 ,6908 0 .7882 0 .8317

.8648

j (b) R = 6.0 x lo6

Fiiun 5.- Continued.

30

0 0 0 a n 0 0 0 0 0

(c) R = 9.0 x 106


C n

-.4821 -.2457 -.0176 .2136 .4393

.5561 ,6613

.7131

.7538

.7815

31

cn

0 -.4740

0 -.2436 0 -.0163 A .2186

b A 7 4 [5 .5614 0 .6643

0 .7420 0 .7564

0 .7741

.048

.040

-1 .o -.a -.6 -. 4 -. 2 0 .2

Y/b/2

.032

.016

.008

.4

(d) R = 15.0 x 106


32

.048

.040

.032

C, corr .024

.016

.008

0 -

Cn

0 -.0229

0 -SO33

., 0 -.2534 A -.0146 b -.221a D .4603 0 ,5750

0 ,6970

o .7a23 o ~ a 9 0

0 -.8 -.6 -. 4 -.2 0 .2 .4

Y/b/2

(e) R = 30.0 x lo6


' 33

.048

.040

.032

C, corr .024

.016

.008

a

Cn

0 -.0167

O -.3913

0 -.2611 A -.0219

b .2204

D ,3382 0 .4613

0 .5868

0 .6680

4 -1


34

(b) R = 6.0 x lo6


35

0 -.0211 0 -.3847 0 -.2531 A -.0139 b .2324 D .3562 n .4879 0 .6019 0 .6053 0 .6889 @ .7417

.7907

Cn

O -.4078 0 -.2687 0 -.1435 A -.0187 n .loo9 b .2222 0 .3613 0 .4944 0 .6067 n .6656

-1 .o -.a -.6 -.4 -.2 0 .2 .4

Y / W

(c) R = 9.0 x lo6


36

.048

.040

.032

C, corr *024

.016

.008

0 -1

0 0 0 a b

0 0 0 0

0 8 B

Cn

-.3782 - .2548 -A341 -.0149 .1040 .2226 .3493 .4739 .5690 ,6535 .6731 .6942

-.a -. 6 -. 4 -. 2 0 .2 .4 Y / W

(d) R = 15.0 x 106


37

.04E

.04C

.032

C, corr -024

-01 6

.008

0 -1

0 0 0 A b n 0 0 0

0

Cn

- ,0244 - .4053 -.2702 -.1392

-.0171

,2326 .3599 .4974

,6147

.6811

(e) R = 30.0 x 106


38

.048

.040

.032

.008

0

0 0 0 A

0 n n 0 0 0 @

C n

-.4189 -.2763 -.1388 -.0098 .1154 .2477 .3787 .5074 .5785 .6476 .7277

-1 .o -.8 -.6 -.4 -.2 0 .2 .4

Y/b/2

(a) R = 6.0 x lo6

Figure 7.- Effects of normal force coefficient on the spanwisc variation of drag coefficient for various Reynolds numbers at a Mach number of 0.74. Transition free.

39

0 -.3988 0 -.1589 0 -.0191 A .1246 b .2504 0 .3906 0 .4867 0 .5400 0 .5729 0 .5919

-1 .o -.8 -.6 -.4 -.2 0 .2 .4

Y/b/2

(b) R = 9.0 x lo6


40

cn

0 -.4019 O -.2766 0 -A471 A -.0187

b .lo88 b .2495

n .3686 0 .4756 0 .5486 0 .5711

(c) R = 15.0 x lo6


41

37 Cn

0 -.0274 O -.4287 0 -.2891 A -.1517 b -.0170 n ,1154 0 .2560 0 .2587

. 0 .4037 0 .3886

5024 @ 3893 e .6010 A .6215

-1 .o -. 8 -.6 -. 4 -. 2 0 .2 .4

Y/b/2

(d) R = 30.0 x lo6

Fiiure 7.- Concluded.

I 42

Cn

0 -.0201 D -.4206

O -.2861

A -.0107 b .1208

b .2626 0 .3895

0 .4975 0 .5520 0 .6187

.04€

.04C

.032

C, corr .024

.016

.008

0 -1 .o -.8 -. 6 -. 4 -.2 0 .2 .4

Y / W

(a) R = 6.0 x 106


43 I

0 -.3924 O -.3019 0 -.la1

A -.0238 b .1181

[5 .2521

0 ,3721

0 .4592

0 A945

0 ,5180

-1 .o -.a -.6 -.4 -. 2 0 .2 .4

Y/b/2

(b) R = 9.0 x lo6


44

0 0 0 a b b n 0 0

0 a3

C n

-.4275

-.2887 -.1434 -.0155

.1159 ,2448 .3577

.4545

.4985

.5276 ,5546

(c) R = 15.0 x IO6


45

0

0 0 h 0 D

n 0 0

n a3

cn

-.0329 -.4324 -.2936 -.1595 -.0202 .1206 .2606 .4023 .5187 .5358 .5562

.048

.040

.032

.024

.016

.008

0 -1 .o -.8 -.6 -. 4 -.2 0 .2 .4

Y/b/2

(d) R = 30.0 x lo6


46

Cn

0 -.0278 O -.1552

0 -.2970

A -.0153 b .1204

0 ,2670

n .3775

0 .3764

0 .5066

0 .5489

@ ,6015

B -.4254

.048

.040

.032

.008

0 -1 .o -. 8 -.6 -.4 -.2 0 .2 .4

Y/b/2

(e) R = 45.0 x l o 6


47

Cn

0 -.0151 0 -.4119 0 -.2950 D -.1678 0 -.0220 n .ii6i 0 .2595 0 .3893 0 .4597 0 3169

,5824

(a) R = 6.0 x lo6


48

.048

.040

.032

C ~ I corr -024

.016

.008

0

Cn

0 -.4044 O -.3130 0 -.1851 A -.0326 h .1277 0 .2466 0 .3543 0 .4255 0 ,4434 0 .4645

- 1 .o -.8 -.6 -.4 -. 2 0 .2 .4 I

Y/b/2

(b) R = 9.0 x lo6


49

0 0 0 n b

D 0 0 0

0 (3

Cn

-.a181 -.2975

-A551 -.0179

.1190

.2493

.3544

.4329

.4609

.4797

3139

(c) R = 15.0 x lo6


50

.048

.040

.032

C, corr .024

.016

.008

0

O -.0308 O -.4256 0 -.3062 A -.1696 0 -.0267 D .1218 0 .2680 0 .4039 0 .4751 0 .4849

-1 .O -.a -.6 -. 4 -.2 0 .2 .4

Y / W

(d) R = 30.0 x lo6


51

O -.0196 0 -.4319 0 -.2908 A -.1667 b -.0208 I3 .1332 0 ,2750 0 .3933 0 .4914

-1 .o -.a -.6 -. 4 -.2 0 .2 .4

Y/b/2

(e) R = 45.0 x lo6


52

.041

.04C

0 0 0 A b

D 0 0 0

-.0246

-.2976 -.0127 .1242

.2648

.3658

.4051

.4611

.5331

-1. .2

(a) R = 6.0 x lo6

Figure 10.- Effccts of normal force coefficient on the spanwise variation of drag coefficient for various Reynolds numbers at a Mach number of 0.80. Transition free.

.4

53

.048

-040

.032

C, corr .024

-01 6

.008

0

'0 -.2856 O -.1646 0 -.0288 a .io91 b .2380 D .3332 0 .3823 0 .4026 0 ,4382

-1 .o -.a -.6 -. 4 -. 2

Y/b/2 0 .2 .4

(b) R = 9.0 x lo6


54

.048

.040

.032

C, corr .024

.016

.008

0

cn

0 -.2873 0 -.1539 Q -.0147 A .1182

b .2395 b .3466 0 ,3804 0 .4023 0 .4390

-1 .o -.a -.6 -. 4 -. 2 0 .2 .4

Y/b/2

(c) R = 15.0 x lo6 Figure 10.- Continued.

55

0 0

0 h b n D 0 0

cn

-.0426 -.3030 -.1615 -.0234 .1151 .2615 .3835 .4180 .4169

.048

.040

.032

C, corr .024

.016

.008

0 -1 .o -.a -.6 -.4 -.2 0 .2 .4

Y / W

(d) R = 30.0 x lo6


56

0

0 0 A b

C n

- .0204

-.1514

-.0155

,1351

.2863

.048

.040

.032

C, corr -024

.016

.008

0 -1 .o -.8 -.6 -.4 -. 2 1

Y/b/2 .2 .4

(e) R = 45.0 x 106


57

--

.048

.040

.032

C, corr .024

.016

C

0 0 0 A b D

D * o

0

C n

-.0132

-.2751 -A513 - .0070 .1185 .2414

.2944

,3563

.43w

-.8 -.6 -. 4 -. 2 0 .2 .4 -1 .o Y / W

(a) R = 6.0 x lo6


58

. OI a

.040

.032

‘ti corr -024

.016

.008

0

0 0 0 a h n 0 0 0

0

Cn

-.0229 -.2661 -.1354 -.0138 .1073 .2118 .2892 ,3206 ,3455 .4018

-1 .o -.8 -. 6 -. 4 -. 2 0 .2 .4

Y/b/2

(b) R = 9.0 x lo6


59

0 -.2685 O -A421 0 -.0200 A .lo19 b .2169 0 ,3005 0 .3287 0 .3558 0 ,3909

(c) R = 15.0 x 106

Fisure 11.- Continued.

60

O -.0226 O -.2847

0 -.0133 h .0909 b ,2542 b .3361 0 .3463

.04€

.04C

.032

C, corr -024

.016

.008

0 -1

(d) R = 30.0 x 106


61

Cn

0 -.0036

-.1472

0 -.0043 A .1145

.048

.040

.032

C, corr .024

.016

.008

0 ~ ~~

-1 .o -.8 -.6 -. 4 -. 2 0 .2 .4

Y/b/2

(e) R = 45.0 x lo6


62

0 -.3773

0 -.2879

0 -.1%0

-.0115 b ,1744

0 .3607 0 3435 0 .7263 0 .a147 0 .goo0

a 1.0661

Q 1.1387

A 1.2006

e .9a74

(a) R = 3.0 x lo6

Figure 12.- Effects of normal force coefficient on the spanwise variation of drag coefficient for various Reynolds numbers at a Mach number of 0.30. Transition futcd.

63

O -.3876 0 -.2095

, 0 -.0088 A .1914 b .3851 D .5764 0 .7709 0 .8640 0 .9535 0 1.0438 e 1.1215 B 1.1935

1.2447 4 1.2538

-1 .o -. 8 -.6 -. 4 -. 2 0 .2 .4

Y/b/2

(b) R = 9.0 x lo6

@re 12.- Concluded.

64

0 -.4190 a - . m i

. 0 -.0083 h .1995 b .4047 n .6iiz 0 ,8186 0 .8940 0 3377

-1 .o -.a -.6 -.4 -. 2 0 .2 .4

Y/b/2

(a) R = 3.0 x lo6

Figure 13.- Effects of normal force coefficient on the spanwise variation of drag coefficient for various Reynolds numbers at o Mach number of 0.50. Transition fixed.

65

O -.4256

D -.2160

I 0 -.0028 15 ,2109 b .4213

[5 .6325

0 .8393 0 .9120

. 0 .%72

' 0 .9809

.048

.040

.032

.008

0 -1 .o -.8 -.6 -. 4 -.2 0 .2 .4

Y/b/2

(b) R = 9.0 x lo6


66

0 -.3838

O -.2536

0 -.0095

h .2378 b .3661 0 .5043 0 ,6230

Cci corr

I I -1 .o -.8 -.6 -.4 -. 2 0 .2 .4

Y/b/2

( I ) R = 3.0 x lo6

Figure 14.- Effects of normal force coefficient on the epanwise variation of drag coefficient for various Reynolds numbers at a Mach number of 0.70. Transition fixed.

67 I

0 -.4175 0 -.2589

0 -.0059

h .2545

b . a 2

0 A245

0 ,6354

0 .6428

(b) R = 9.0 x lo6


68

0 -.4423 0 -.2847 Q -.1405 A -.0019 b .1346 n 0 .4360 0 .5351

Figure 15.- Effects of normal force coefficient on the spanwise variation of drag coefficient for a Reynolds number of 9.0 x lo6 and a Mach number of 0.74. Transition fixed.

69

O -.1526

O -.3047

0 - .a74 A -.0022

b ,1428

e .2947

0 A369

1 9 A935

0 A933

Figure 16.- Effects of normal force coefficient on the spanwise variation of drag coefficient for I Reynolds number of 9.0 x lo6 and a Mach number of 0.76. Transition fixed.

70

Cn

O -.1748 0 -.3209 0 -.0028 0 ,1627 B 3162 D .4155

.048

.04C

.032

‘d corr

.016

.008

0 -1 .o -.8 -.6 -. 4 -. 2 0 .2 .4

Y/b/2

Figure 17.- Effects of normal force coefficient on the spanwise variation of drag coefficient for a Reynolds number of 9.0 x lo6 and a Mach number of 0.78. Transition fixed.

71

.048

.040

.032

C, corr .024

.016

.008

0

C n

0 -.1788 0 -.0045 0 A714 A .3112 fh .3531

-1 .o -.8 -.6 -.4 -.2 0 .2 .4

Y/b/2

Figure 18.- Effects of normal force coeffiient on the spanwise variation of drag coefficient for a Reynolds number of 9.0 x lo6 and a Mach number of 0.80. Transition fixed.

r 72

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148

(D

Report Documentation Page 1 Report No

NASA TM-100527 _____

2 Government Accession No 3 Recipient's Cdtalog No

- -____ ---____

Charles L . I i idson and AcquiLIa S . H i L 1

4 Title and Subtitle

High Reynolds Number Transon ic T e s t s on a n NACA 0012 A i r f o i l i n t h e Langley 0.3-Meter T ranson ic Cryogenic Tunne 1

I 10 Work Unit No.

-

5. Report Date

December 1987 6 Performing Organization Code

505-61-01-07 ___-- --i - _ _ __ _____ ~

9 Performirig Orgarmation Name and Address

_____-__I_---

12 Sponsoring Agmcy Name and Address

N a t i o n a l Aeronau t i c s and Space A d m i n i s t r a t i o n Washington, DC 20546

____ __ __-._ [ 11 Contract or Grant No NASA i . s n p l e y Research Cen te r

13 Type of Report and Period Covered

Tcchni r a 1 Nc~morandurr, - -__-__

14 Sponsoring Agency Code

19. Security Classif. (of this reportt 20. Security Classif. (of this page1 21. No. of pages

U n c l a s s i f i e d U n c l a s s i f i e d 149

- 16. Abstract

22. Price

A0 7

T e s t s were c o n d u c t e d i n t h e two d imens iona l t e s t s e c t i o n of t h e LangLey 0.3-meter T ranson ic Cryogenic Tunnel on an NACA 0012 a i r f o i l t o o b t a i n aerodynamic d a t a as a p a r t of t h e Advanced Technology A i r f o i l T e s t (ATAT) !)rngram. a Reynolds number range of 3 . 0 x 10 t o 45.0 x 10 . The s t a g n a t i o n p r e s s u r e was v a r i e d between 1.2 and 6.0 atmospheres and t h e s t a g n a t i o n t empera tu re w a s v a r i e d between 300 K and 90 K t o o b t a i n t h e s e t e s t c o n d i t i o n s . P l o t s of t h e spanwise v a r i a t i o n of d r a g c o e f f i c i e n t as a f u n c t i o n of normal f o r c e c o e f f i c i e n t and the v a r i a t i o n of t h e b a s i c aerodynamic c h a r a c t e r i s t i c s w i t h a n g l e o f a t t a c k a r e shown. The d a t a are p r e s e n t e d u n c o r r e c t e d f o r w a l l i n t e r f e r e n c e e f I e c t s and w i t h o u t a n a l y s i s .

The t e s t program covered 8 Mach number gange of 0.30 to 0 . 8 2 and

17. Key Words (Suggested by Authorls)) I 18. Distribution Statement

U n c l a s s i f i e d - Unlimited A i r f o i l s High Reynolds numbers Cryogenic wind t u n n e l s S u b j e c t Category - 02

b

i

1

d

NASA FORM 1626 OCT 86

Documents

High Reynolds Number Transonic Tests of an NACA 0012 0.3 ...€¦ · APPARATUS The test program was conducted in the 8 in by 24 in (20.32 cm by 60.96 cm) two dimensional test section