NASA Contractor Report ]65685

" Nf_S[)-e_-/_%_,C_-3 1176 00167 5348

NASA Contractor Report ]65685

- O NASA-CR_ 165685

MODELANDBOUNDARYAERODYNAMICDATAFROM

HIGH BLOCKAGETWO-DIMENSIONALAIRFOIL TESTS

IN A SHALLOWUNSTREAMLINEDTRANSONIC

FLEXIBLEWALLEDTEST SECTION

S. W. D. Wolf

THE UNIVERSITYOF SOUTHAMPTONSouthampton, England

NASAGrant NSG-7172April 1981

_?.a:,,31 i93i

NationalAeronauticsandSpaceAdministration

LangleyResearchCenterHampton,Virginia23665

CONTENTS

PAGE

I. Transonic Self-Streamlining Wind Tunnel Data i

2. List of Symbols 5

3. List of References 6

Tables 7

Figures 33

AI8I-_I057

I. TRANSONIC SELF-STREAMLINING WIND TUNNEL DATA

In the course of previously reported two-dimensional tests on

a 4 inch (lO.16cm) chord NACA 0012-64 airfoil in the Transonic Self-

Streamlining Wind Tunnel (TSWT)1'2 twenty-four runs were performed

with the flexible floor and ceiling of the test section set 'straight'.

"° These runs provide information on the gross boundary interference

present in a small non-porous test section (with a nominally 6 inches

{15_.24cm}square cross-section), where the model blockage is 8% at

_=O. These tests are discussed here because the associated data may

prove useful in the development of wind tunnel correction techniques.

Four sets of 'straight wall' contours have been obtained

experimentally which give constant wall Mach numbers of 0.3, 0.5, 0.7

or 0.9 in an empty test section. The walls in fact diverge to allow for

boundary layer growth. However, the walls are effectively straight in

the aerodynamic sense with no model present. With a model installed,

the wall boundary layer displacement thickness is altered by the

introduction of longitudinal pressure gradients. The walls are then

no longer effectively 'straight'. Wall adjustments to correct the

contours have not been attempted in this work. This point is raised

as a warning if the 'straight wall' velocity distributions are to be

used for any form of wind tunnel corrections.

V

The 'straight wall' runs are summarised in Table I. Assessment

of wall induced effects on the flow at the model have been made using

the top and bottom wall loadings. These wall loadings are determined

from the imbalance between real measured wall pressures (inside the

test section) and imaginary calculated wall pressures (external to the

test section). The wall induced effects are referred to as 'Residual

Interferences' and have been calculated in terms of induced angle of

" attack (As), induced camber (assessed as an effect on CL and tabulated

as ACL) and induced Mach number perturbation (AM). The force coefficients

CL and CD were determined from integrated model pressures. Notice that

for each run with zero angle of attack, the bottom wall supports a larger

- I-

E (average of the modulus of the pressure coefficient error between

real and imaginary flows along a flexible wall) than the top wall.

Since the imaginary flowfields above and below the test section are

uniform and undisturbed for 'straight wall' cases, the difference

between top and bottom wall E is attributed to the asymmetry of

model position between the flexible walls. This may be due to some

testsectlon centreline displacement or curvature introduced during

the experimental determination of 'straight wall' contours. The

high values of E for the straight wall runs imply high levels of

interference induced by each wall at the model.

The 'straight wail'model pressure distributions are plotted

on Figure I and tabulated in Table 2. Corresponding Mach number

distributions along the flexible walls are shown in Figure 2. It is

evident from the wall data that the peak wall Mach numbers rise

rapidly with increasing freestream Math number (for example, compare

wall data for the _=4°,runs 66, 40 & 42). This effect of flow

compressibility leads to choking of the test section which, of course,

sets an upper limit to the Mach number range of 'straight wall' tests.

The 'straight wall' model data (=, CL and CD) has been corrected

for interference induced by a non-porous test section boundary, using

the conventional technique developed by Allen and Vincenti3. Also, the

corrections due to residual interferences have been applied to the model

data (e and CL). Several options exist regarding the application of the

corrections for these residual interferences. Corrections can be applied

independently to angle of attack, lift and Mach number. Alternatively,

streamwise lift can be corrected for induced camber and Mach number

perturbations while angle of attack is corrected independently. Finally,

the three components of the residual interferences can be related to

corrections to lift while _ and M remain constant. To assist with data

comparisons, only the lift and angle of attack have been corrected for

residual interferences, with no correction applied to the freestream

Math number.

? -- 2

Wall-induced errors at the model are assessed from the wall loadings

made_incompressible by use of linearised theory. In TSWT operation the

residual interferences are made small by wall contouring to streamline

shapes, even up to frees_eamMmch numbers of 0.85. In general, only small

• model corrections can be applied with confidence at transonic speeds. Hence,

in these circumstances, simple incompressible assessment of residual

interferences can be used over a wide range of test Mach number.

The corrected model data is summarised in Figure 3 where the

lift curve slope (dCL/_c) is plotted against freestream Mach number. Results

from 1TSWT straight wall and streamlined wall2 tests are shown together with

theoretical curves derived from linearised theory which are constrained to

pass through the Ibwest Mach number data point on each data set. There is

reasonable agreement between theory and experiment, especially for the

streamlined wall case. The model data corrected for residual errors is

very encouraging considering the gross boundary interference present in

the 'straight wall' tests. The Allen and Vincenti corrections appear too

large, particularly at the higher Mach numbers, illustrating the inaccuracy

of applying only simple corrections to the overall model forces which take

no account of detailed changes in the model flow pattern - for example,

changes in model shock positions.

The straight,wall data was obtained at freestream Mach numbers of

approximately 0.7, 0.5 and 0.3. 'Straight wall' testing above Math No.O'7

was impractical with this model. The walls were set to the appropriate

'straight wall' contours which corresponded closely to the freestream Mach

number of the test. The variation of CL and CD with = for the three Mach

numbers are shown on Figures 4 and 5 respectively together with streamlined

wall data and corrected data where appropriate.

The CL data can be conveniently summarised by the fitting of a

least squares curve to each set of data over the range -8o<_<+8° The

straight line slopes and zero = intercepts are:-

Slope Zero = interceptData dCL/d_ per degree CL

!

Mach No. 0.7 0.5 0.3 0.7 0.5 0.3

T

Straightwalldata .1563 .1197 .1091 -.0898 -.0752 -.0602

Straight wall datacbrrected by Allen .0927 .0875 .0842 -.0519 -.0552 -.0511&IVincenti method.

Straight wall datacorrected for .1203 .0916 .0895 -.0775 -.0672 -.524residual interfer-ences

Streamlined walldata2 .1178 .0965 - -.0753 -.0574 -

i

If the streamlined data is assumed correct as suggested, bycomparison with reference data, then the 'straight wall' residualcorrections seem very good. The ratio of lift curve slopes forstreamlined wall data and residual corrected straight wall data is.97 and 1.05 for Mach numbers 0.7 and 0.5 respectively.

The CD data as shown on Figure 5 relates only to pressuredrag. _ile magnitudes are aerodynamically meaningless, the symmetryof the CD curves about the zero a axis is shown. Future work willinvestigate the momentum defect in the airfoil wake, to assess modeldrag.

p

The 'straight wall' data is presented here in graphical andtabulated form to conclude the summary of current TSWT tests with theNACA 0012-64 airfoil section. Because of the high blockage, thisdata may prove useful to those engaged in the development ofinterference correction methods for transonic wind tunnel testing.

-4-

2. LISTOF SYMBOLS

e Angle of attack

c Model chord

CC Chordwise force coefficient

CL Lift coefficient

CD Pressure drag coefficient

CM Pitching moment about the leading edge

CN Normal force coefficient

Cp Pressure coefficient

E Average of the modulus of the pressure coefficient error

between real and imaginary flows along a flexible wall

! ] Cpr-Ceiln

M .Freestream Mach number

n Number of jacks along a wall

Rc Chbrd's Reynolds number

X Distance from leading edge

SUFFIXES

'" Uncorrected data

'i' imaginary

'r' real

'Top' Top wall

'BOT' Bottom wall

!!

-5-

3. REFERENCES

I. M.J. Goodyer and 'The Development of a Self-StreamliningS.W.D. Wolf Flexible Walled Transonic Test Section'

AIAA Paper 80-0440, March 1980

2. S.W.D. Wolf 'Selected Data from a Transonic FlexibleWalled Test Section' NASACR-159360,September 1980

3. H.J. Allen and 'Wall Interference in a Two-DimensionalW.G. Vincenti Flow Wind Tunnel with Consideration of

the Effect of Compressibility'NACAReport 782, 1944

-6-

TABLE 1 SUMMARY OF TSWT 'STRAIGHT WALL' DATA

Model Data Residual Interferences

i _ _ Aa ACL AM

= _ CL" CD" ETO P EBOT

I 66 I 4° ,706 .5466 .032 +.5 .0649 .O41 .1318 .0665

2 86 3° ,697 .3854 .0027 +.26 .0557 .029 .0897 .O431

3 55 2° .693 .2352 -.004 +.15 .033 .025 .069 .0417

4 54 OO .683 -.Iiii -.0109 -.115 -.0684 .023 .042 .0573

5 68 -2° .701 -.4636 .0013 -.413 -.0491 .033 .0462 .1031

6 67 -4° 8701 -.6624 .0505 -.654 -.0534 •O51 .089 .1742

7 40 4° .520 .4089 -.003 .255 .0629 .O15 .0751 .02368 53 3° .505 .2697 -.006 .302 .0692 .O11 .0665 .0194

9 39 2o .516 .1755 -.0098 .863 .0288 .O13 .0499 .O271

I0 36 0o .505 -.0728 -.O136 -.097 -•0057 .O12 .0290 .0406

II 52 -2° .499 -.3195 -.O136 -.222 -.0424 .O13 .0195 .0609

12 51 -3° .505 -.4415 -.O124 -.298 -.O591 .014 .018 .0724

13 50 -4 ° .504 -.5467 -.0092 -.363 -.0742 .015 .O182 .0857

14 44 IOo .301 .9753 .0565 .719 .1432 .013 .1485 .0473

15 43 8° .298 .8317 .0363 .637 .1186 .O11 .1253 .0385

16 42A 6o .299 .5872 .0133 .411 .0877 .009 .093 i .024

17 I 42 4o .304 .3658 _.0058 .221 .O583 .O08 .0654 .O193'

18'41 2°.296 .1608-.0109 .103 .0237.007.045 .0217

19 ! 40A i O° .293 -.0695 -.0131 -.067 -.0094 .006 .0265 .038520 45 i -2° .297 -.2801 -.0119 -.207 -.0394 .007 .0153 .0573

21 46 -4 ° .296 -.4871 -.0052 -.4 -.0631 .008 .0181 .0856

22 47 -6° .300 -.7399 .0095 -.53 -.1012 .009 .0338 .1078

23 48 -8° .296 -.9106 .0261 -.563 -.1301 1.013 .0378 1396

24 ,49 -I00 .301 -1.052 .0517 -.567 -.1509 i .015 .045 .1688

-7-

TABLE 2

NACA 0012-64

PRESSURE DISTRIBUTIONS

AND FORCES.

-8-

TABLE 2.1

NACA SECTION ANALYSIS0012-64

RUN NO. = 66

I_ING DATA FILE NAME = tWING1.DATINPUT FILE NO. - 19

UF'F'ERSURFACE LOWER SURFACEZCHORD CP LOCAL CP LOCAL

0 -0.0116 -0.01161 -0.6274 0.60422 -0.9894 0.30565 -1,2049 0.09497 -1.2337 -0.03639 -1.2423 -0.1105

15 -1.2114 -0.200320 -1.2715 -0.262425 -1.2749 -0.322829 -1.2822 -0.355735 -1.2805 -0.405640 -1.2323 -0.441644 -1.2444 -0.481250 -1.2099 -0.496655 -1.==74 -0.516360 -1.1239 -0.512964 -0.7273 -0.502070 -0.4999 -0.476975 -0,3835 -0.443580 -0.2949 -0.414285 -0.2085 -0.363390 -0.1150 -0.217595 -0.0223 -0.1353

UPPER 'LOWER TOTAL

CN 0.8644 -0.3169 0.5475CC -0.0314 0.0251 -0.0063CM -0.3214 0.1796 -0.1418

s

AIRFOIL PERFORMANCECL CD CN

0.5466 0.0320 -0,1418

-9-

TABLE2.2--- --w--


RUN NO. = 56:.,

ALPHA = 3.0

MACH NO, =0,697

WING DATA FILE NAHE = _WING1.DATINPUT FILE NO, - 10.

UPPER SURFACE LOWER SURFACE%CHORD CP LOCAL CP LOCAL

0 -0,0163 -0,01631 -0.5707 0.53822 -0.9447 0.23575- -1,1186 0,03527 -1,1501 -0.08819 -1.1677 -0,153215 -1,1414 -0,223720 -1,1519 -0.278325 -1.0677 -0.3276

29 -0.9960 -0,344235 -0.934B -0.388140 -0.7947 -0.409144 -0.7422 -0.433750 -0,7002 -0,437255 -0,6560 -0,441460 -0.6015 -0,430964 -0,5192 -0,399970 -0,4482 -0.375075 -0,3658 -0,349580 -0,2685 -0.°302685 -0,1648 -0.190490 -0.0547 -0,098795 0,0547 -0,0135

UPPER LOWER TOTAL

CN 0,6534 -0.2684 0.3850CC -0.0343 0.0167 -0.0175CH -0.2239 0,1391 -0,0848

AIRFOIL PERFORMANCECL CD CH

0,3854 0,0027 -0.0B48

- i0 -

TABLE 2.3


RUN NO. = 55_i,•

ALPHA = 2.0

MACH NO. =0.6927

WiNG DATA FILE NAME = _STWD.DATINPUT FILE NO. - 14

: UPPER SURFACE LOWER SURFACE%CHORD CP LOCAL CP LOCAL

0 0.0059 0.00591 -0.3569 _ 0.36872 -0.7156 0.05675 -0.8487 -0.11707. -0.8956 -0.21919 -0.8761 -0.277515 -0.8089 -0.323420 -0.7859 -0.369325 -0.7576 -0.402929 " -0.7493 -0.418435 -0.7405 -0.448540 -0.7104 -0.4538.44 -0.7069 -0.471650 -0.6821 -0.468055 -0.6417 -0.458160 -0.5744 -0.43_664 -0.4891- -.0.407270 -0.4180 -0.383575 -0.3469 -0.345580 -0.2672 -0.229885 -0.1767 -0.150190 -0.0690 -0.064695 0.0517 0.0345

UPPER LOWER TOTAL

CN 0.5383 -0.3034 0.2349CC -0.0210 0.0088 -0.0122CM -0.1987 0.1404 -0.0584

" AIRFOIL PERFORMANCECL CD CM

0.2352 -0.0040 -0.0584

TABLE 2.4


RUN NO. = 54?

ALPHA = 0.0

MACH NO. =0.6831

WING DATA FILE NAME = _STWD.DATINPLIT FILE NO. - 15


0 0.0077 0.00771 0.1891 -0.1737•2 -0.1207 -0.49775 -0.2954 -0.59007 -0.3765 -0.63099 -0°4000 -0.6671

15 -0.4307 -0.641720 -0.4488 "0.652625 -0.4615 -0.658029 -0.4745 -0.635035 -0.4907 -0.640440 -0,478k -0.622444 -0.4925 -0.620650 -0.4889 -0.595355 -0.4760 -0.574760 -0,4561 -0.549364 -0.4182 -0.473670 -0.3927 -0.379675 -0.3286 -0.317780 -0.2266 -0.236885 -0.1515 -0.158890 -0.0648 -0.051795 0.0386 0.0634

UPPER LOWER TOTAL

CN 0.3439 -0.4550 -0.1111CC 0.0013 -0.0122 -0.0109CM -0.1475 0.1753 0.0279

i

AIRFOIL PERFORMANCECL CD CM

-0.1111 -0.0109 0.0279

,o; °"

• - 12 -$

-:" ._. :. ,_

'_io"

TABLE 2.5iii|

NACA SECTION ANALYSIS

RUN NO. =681

ALPHA = -2,0

MACH NO. =0,7008

WING DATA FILE NAME = _WING2,DATINPUT FILE NO, - 3

UF'PER SURFACE LOWERSURFACE%CHORD CP LOCAL CP LOCAL

0 -0,0458 -0,0458t 0,5164 -0.60802 0.2198 -0.93895 -0.0035 - -1.16547 -0.1044 -1.0958

9, -0,1618 -1.095815 -0.2418 -1.094020 -0,2818 -1.151425 -0.3183 -1.137529 -0.3560 -1.114935 -0.3803 -_.I11540 -0,4064 -1.090644 -0,4376 -i.I06250 -0,4585 -1.059455 -0.4634 -0.923360 -0.4460 -0.608064 -0,4249 -0.536470 -0.4016 -0.478575 -0.3712 -0.4100SO -0.3345 -0.309185 -0.2265 -0.216090 -0,1080 -0,091795 -0.0106 0.0254

UPPER,. LOWER TOTAL

CN 0.2819 -0,7452 -0,4633CC 0,0162 -0.0311 -0,0149CM -0.1463 0,2707 0.1243

T

" AIRFOIL PERFORMANCECL CD CM

-0.4636 0.0013 0.1243

.°

- 13 -

TABLE 2.6


RUN NO. = 67.- _

ALPHA = -4.0

MACH NO. =0.701

WING DATA FILE NAME = _WING1.DATINPUT FILE NO. - 20

UF'PER SURFACE LOWER SURFACE%CHORD CP LOCAL CP LOCAL

0 -0.0930 -0.09301 0.6434 -0.82952 0.3625 -1.11005 0.1179 -1.36117 0.0000 --1.35219 -0.0728 -1.353815 -0.1786 -1.315820 -0.2359 -1.362525 -0.2931 -1.376329 -0.3509 -1.408735 -0.4047 -1.386140 -0.4464 -1.328844 -0.5020 -1.400050 -0.5593 -1.400055 -0.5749 -1.383860 -0.5870 -1.324564 -0.5942 -1.065770 -0.5809 -0.832875 -0.5507 -0.640580 -0.5178 -0.479285 -0.4785 -0.369090 -0.4097 -0.237195 -0.2427 -0.1593

UPPER LOWER TOTAL

CN 0.3611 -1.0254 -0.6643CC 0.0319 -0.0278 0.0041CM -0.2175 0.4091 0.1916


-0.6624 0.0505 0.1916

TABLE 2.7

NACA SEcr!oN ANALYSIS0012-64

RUN NO. = 40

ALP_IA = 4.0

hi__H NO. =0.5203

WING DATA FILE NAME = _STWD,DATINPUT FILE MOo -- 8

UF'PER SURFACE [.ObJER SURFACE%CHORD CP LOCAL CP LOCAL

0 -0,3297 -0,3297I. -1.3940 0.73452 -1.4600 0,44165 -1,2476 0,2"2817 -1.1235 0,09849 -1.0240 0.032415 -0,8608 -0.051420 -0.7512 -0.101725 -0.7110 -0.147629 -0,6775 -0.!69935 -0,6484 -0.202340 -0.6193 -0,225844 -0,5891 -0,241550 -0,5567 -0,247155 -0,5042 -0,250460 -0.4651 -0,251564 -0.4114 -0.241570 -0,3466 -^_,_75 -0.2828 -0,202380 -0.2046 -0.i77785 -0.1185 -0,149890 -0.0257 -0,068295 0.0682 0.0067

UPPER LOWER TOTAl

CN 0.5298 -0,1221 0.4077CC -0,0495 0.0180 -0.0315CM -0.1683 0.0779 -0.0904


0,4089 -0,0030 -0,0904

- I5 -

TABLE 2.8 ":_::_:!

NACASECTIONhNALYSiS0012-64

RUNNO. = 53 :

ALPHA = 3.0

MACH NO. =0.505

WING DATA FILE NAME = _STWD.DATINPUT FILE NO. - 9

UPF'ER SURFACE LOWER SURFACE%CHORD CP LOCAL CP LOCAL

0 -0.1623 -0.16231 -0.8736 0.54892 -1.0161 0.24765 -0.9156 0.0654

-0.8152 -0.04099 -0.7720 -0.092315 -0_6984 -0.154220 -0.6353 -0.189225 -0°5980 -0.221929 -0.5711 -0.234835 -0.5478 -0.259340 -0.5326 -0.275644 -0.5092 -0.282650 -0.4894 "0.285055 -0.4473 -0.279160 -0.4181 -0.274564 -0.3749 -0.2604 "70 -0.3212 -0.238375 -0.2628 -0.214980 -0.1950 -0.188085 -0.1156 -0.126190 -0.0234 -0.042095 0.0759 0.0200

UPPER LOWER TOTAL

CN 0.4366 -0.1676 0.2690CC -0.0327 0.0123 -0.0204CM -0.1466 0.0874 -0.0593


0.2697 -0.0063 -0.0593

- 16 -

TABLE 2.9


RUN NO. = 39

ALPHA = 2.0

MACH NO. =0.516

WING DATA FILE NAME = _WING1.DATINPUT FILE NO, - 9

UPPER SURFACE LOWER SURFACE%CHORD CP LOCAL CF' LOCAL .

0 -0.0820 -0.08201 -0.5249 0.36092 -0°7240 0.06455 -0.7138 --0.08037 -0.6901 -0.16299 -0.6482 -0.2059

15 -0.6018 -0.245520 -0°5577 -0.268125 -0.5328 --0.293029 -0.5159 -0.299835 -0.4989 -0.315640 -0.4876 -0.326944 -0.4740 -0.331550 -0.4582 -0.324755 -0.4208 -0.314560 -0.3960 --0.304364 -0.3552 -0,285170 -0,3077 -0,261375 -0.2557 -0. _a''80 -0,1957 -0.174285 -0.1233 -0.106390 -0.0351 -0.033995 0.0690 0.0430

UPPER LOWER TOTAL

CN 0.3876 -0.2126 0.1751CC -0.0221 0.0062 -0.0159CM -0.1373 0.0973 -0.0399

AIRFOIL PERFORMANCECL CD CH

0,1755 -0,0098 -0.0399

- 17-

TABLE 2.I0 _ _ :

NACA SECTION ANALYSIS ! _ -0012_64 _

RUN NO. = 36 _"

ALPHA = 0.0!

MACH NO. =0.505

WING DATA FILE NAME = _STWD.DATINFUT FILE NO, - 10

UPPER SURFACE LOWER SURFACEZCHORD CP LOCAL CP LOCAL

0 -0.0418 -0.04181 0.1371 -0.22072 -0.1429 -0.46355 : -0.2683 -0.48007- -0.3137 -0.48799 -0.3253 -0.4972

15 -0.3415 -0.476320 -0.3392 -0.463525 -0.3404 -0.461229 -0.3497 -0.447235 -0.3555 -0.446140 -0.3590 -0.443844 -0.3578 -0.433350 -0.3543 -0.413655 -0.3357 -0.391560 -0.3229 -0.368364 -0.3032 -0.329970 -0.2776 -0.283575 -0.2300 -0.232380 -0.1673 -0.174385 -0.1069 -0.111590 -0.0325 -0.025695 0.0592 0.0709

UPPER LOWER TOTAL

CN 0.2539 -0.3267 -0.0728CC -0.0014 -0.01 _== -0.0136CM -0.1049 0.1223 0.0174


-0.0728 -0.0136 0.0174

- 18 -

TABLE 2.11 °'_ ;_ "; _"_:'

NACA SECTION ANALYSIS.0012-64

.RUN NO. = 52

ALPHA = -2.0

MACH NO. =0.499

WING gATA FILE NANE = %STWD.DATINPUT FILE NO. - 11

UPPER SURFACE LOWER SURFACE_CHORD CP LOCAL CP LOCAL

0 -0.2193 -0.21931 0.5773 -1.0160 "2 0.2785 -1.13085 . 0.1380 -0.97307 -0.0024 -0.86549 -0.0598 -0.830815 -0.1291 -0.732720 -0.1554 -0.670625 -0.1829 -0.638329 -0.2056 -0.595335 -0.2271 -0.578540 -0.2438 -0.563044 -0.2510 -0.535550 -0.2618 -0.499755 -0.2510 -0.466260 -0.2510 -0.431564 -0.2355 -0.380170 -0.2164 -0.323975 -0.1901 -0.265480 -0.1554 -0.190185 -0.1052 -0.114890 -0.0394 -0.015595 0.0430 0.0777

UPPER LOWER TOTAL

CN 0.1405 -0._594 -0.3189CC 0.0122 -0.0369 -0.0247CN -0.0755 0.1518 0.0762

AIRFOIL PERFORMANCECL CD CM _ .

-0.3195 -0.0136 0.0762

- 19 -

TABLE 2.12m,

NACA SECTION A_ALYS_S00112_64 ,'

RUN NO. = 51

.ALPHA= -3',0

MACH NO. =0.5047

HING DATA FILE NAME = _STWD,DATINPUT FILE NO. - 12

a


0 -0.3850 -0.38501 0.7382 -1.50822. 0.4474 _ -1.54955 0.2849 -1.27987 0.1224 -1.09979 0.0506 -1.014915 -0.0389 -0.843020 -0.0789 -0.779425 -0.1142 -0,745329 -0.1460 -0.697035 -0.1731 -0.665240 -0.1954 -0.638244 -0.2108 -0.599350 -0.2249 -0.551055 -0.2202 -0.509860 -0.2249 -0.466364 -0_2155 -0.409770 -0.2002 -0.343875 -0.!778 -0.277980 -0.1495 -0.197885 -0.1083 -0,114290 -0.0495 -0.014195 0.0247 0.0718

UPPER LOWER T0TAL"

CN 0.0976 -0.5378 -0.4402CC 0.0169 -0.0523 -0.0355CH -0.0654 0.1693 0.1039


-0.4415 -0.0124 0.1039

- 20 -

TABLE 2.13

NACA SECTION ANALYSIS0012.64

RUN NOo"= 50

ALPHA = -4.0

MACH NO. =0.5043

WING DATA FILE NAME = _STWD,DATINPUT FILE NO, - 13


0 -0.5778 -0.57781 0.8512 -2.00692 0.5784 -1.92265 0.4010 -1.61707 0.2236 -1.25179 0.1428 -1.112315 0.0386 -0.973020 -0.0117 -0.887525 -0.0550 -0.833729 -0,0937 -0.771635 -0.k241 -0.730640 -0.15!0 -0.695544 -0.1686 -0.648750 -0.1885 -0.592555 -0.1885 -0.542160 -0.1979 -0.49296_ -0.1920 -0.426270 -0.1838 -0.357175 -0.1663 -0.283480 -0.1428 -0.197985 -0.1077 -0.113690 -0,0574 -0,018795 0.0070 0.0562

UPPER LOWER TOTAL

CN 0.0598 -0.6046 -0.5448CC 0.0198 -0.0672 -0.0473CM -0.0556 0.1836 0.1279


-0.5467 -0.0092 0.1279

- 21 -

TABLE 2.14


RUN NO. = 44

ALPHA = 10.0

MACH NO. =0,3011

WING DATA FILE NAME = _WING1.DATINPUT FILE NO. - 3


0 -1.6738 -1.67381 -4.3475 0.99902 -4.2293 0.98425 -2.4619 0.79807 -2.0895 0.62959 -1.8058 0.540915 -1.4659 0.387220 --1.2531 0.292625 -1.1172 0.218729 -1.0137 0.162635 -0.9221 0.106440 -0.8423 0.059144 -0.7714 0.020750 -0.6975 -0.008955 -0.6088 -0.038460 -0o5379 -0.065064 -0.4581 -0.082870 -0.3665 -0.085775 -0.2808 -0.097580 -0.2010 -0.094685 -0.1360 -0.094690 -0.0946 -0.079895 -0.0650 -0.0709

UPPER LOWER TOTAL

CN 0.8656 0.1047 0.9703CC -0.1375 0.0238 -0.1137CM -0.2318 0.0068 -0.2250


0.9753 0.0565 -0.2250

- 22 -

f

TABLE 2.15

NACA SECTION ANALYSIS .0012-64

RUN NO, = 43

ALPHA = 8,0

MACH NO, =0,298

WING DATA FILE NAME = _WING1,DATINPUT FILE NO, - 4


0 -1,1472 -1,14721 -3,3088 1,01442 -2.9586 0.89965. -1,9955 0,67327 -1,7117 0,50429 -1.5_25 0o416615 -1.2559 0,280820 -1.0868 0,202325 -0.9812 0,132829 -0,9027 0,090635 -0°8393 0,042340 -0,7759 0,006044 -0,7215 -0,021150 -0.6642 -0,0483"55 -0,5947 -0.066460 -0,5374 -0.081564 -0.4710 -0.087570 -0°3925 -0°084575 -0,3110 -0,078580 -0,2294 -0,063485 -0.14_9 "0,042390 -0,0634 -0,012195 0,0000 0,0091

q

UPPER LOHER TOTAL

CN 0,7511 0,0776 0,8287CC "0,1028 0,0231 -0,0798CM -0,2157 0,0063 -0,2095

AIRFOIL PERFORHANCECL CD CM

" 0,8317 0,0363 -0,2095 "_

- 23 -

TABLE 2.16


RUN NO, =42

ALPHA = 6,0

;_ACH NO. =0.2987

WING DATA FILE NANE = _WING1,DATINPUT FILE NO, - 5


0 -0,6414 -0,64141 -2,2240 0,9411

-1,9632 0,70445 -1,6155 0,46767 -1,3158 0,3117? -I,12!0 0,236815 -0,9681 0,128920 -0,8482 0,065925 -0°7793 0,012029 -0,7253 -0,021035 -0.6804 -0,062940 -0,6414 -0,086944 -0,6024 -0,110950 -0,5635 -0,122955 -0,5095 -0,134960 -0.4646 -0.143964 -0.4106 -0.140970 -0,3477 -0,134975 -0,2817 -0,I19980 -0.2068 -0.095985 -0.1259 -0.071990 -0,0360 -0,045095 0,0480 -0,0270

UPPER LOWER TOTAL

CN 0.5909 -0,0055 0.5854CC -0.0706 0,0224 -0.0482C_| -0,1771 0,0357 -0,1413


0,5872 0,0133 -0,1413

- 24 -

TABLE 2.17

H c _E_-I A YSIS0012-64

RUN NO, = 42 _

ALPHA = 4.0

MACH NO, =0.3042

_ WING DATA FILE NAME = _WINGI.DAT_'_ INPUT FILE NO. - 6


0 -0.5558 -0.55581 -1.2824 0.7266

-1.2679 0.4_425 -1.0624 0.22297 -0.9553 0.09849 -0._656 0.040515 -0.7_I_ -0.034720 -0,6542 -0.078225 -0.6050 -0.115829 -0.5732 -0.139035 -0o5_71 -0,165040 -0.5240 -0.185344 -0,5008 -0.199750 -0o4748 -0_199755 -0.4313 -0.202660 -0._024 -0.202664 -0.3590 -0._88270 -0,3069 -0.!76675 -0,2518 -0.153480 -0.1882 -0.130385 -0.1129 -0.!04290 -0,0289 -0.078295 0_0695 _ 0.0145

UPF'ER LO_ER TOTAL

CN 0.4576 -0.0931 0.3645CC -0.0457 0.0_44 -0.0313CM -0.1456 0.0614 -0.0841


0.3658 -0.0058 -0.0841

- 25 -J

TABLE 2.18


RUN NO. = 41:

ALPHA = 2.0

MACH NO. =0.2961

WING DATA FILE NAME = _WINGI.DATINF'UT FILE NO. - 7


0 -0.1715 -0.17151 -0.5115 0.3400

•2 -0.6615 0.05215 -0.6309 -0.07967 -0.6033 -0.1501? -0.5635 -0.183815 -0.5206 -0.214420 -0o4808 -0.235825 -0,4594 _ -0.257329 -0.4471 -0.263435 -0.4380 -0.272640 -0.4288 -0.281844 -0.4165 -0.281850 -0.4012 -0.275655 -0.3736 -0.266460 -0.3583 -0.257364 -0.3338 -0.238970 -0.2818 -0.214475 -0.2297 -0.189980 -0,1715 -0_165485 -0.1103 -0.131790 -0.0337 -0.006195 0.0643 0.0704

UPPER LOWER TOTAL

CN 0.3430 -0.1826 0.1604CC -0.0207 0.0042 -0.0165CM -0.1219 0.0823 -0.0396


0.1608 -0.0109 -0.0396

- 26 -

TABLE 2'19


RUN NOo = 40

ALPHA = 0.0

MACH NO. =0.2934

WING DATA FILE NAME = _WING1.DATINPUT FILE NO, - 8

UPF'ER SURFACE LOWER SURFACE%CHORD CP LOCAL CP LOCAL

0 -0.0507 -0°05711 0o1203 -0.23442 -0.1419 -0.43495 -0.2437 -0.44727 -0.2838 -0.44119 -0.2868 -0.441115 -0.3023 -0.419520 -0.2961 -0.410225 -0.296! -0.404029 -0o3053 -0,391735 -0.3115 -0.388640 -0.3115 -0.395544 -0.3115 -0.376350 -0.3053 -0°357855 -0.2899 -0.342460 -0.2776 -0.323964 -0.2560 -0.299270 -0.2313 "0.271475 -0.2066 -0,243780 -0.1820 -0.166685 -0,1234 -0.089490 -0.0031 -0.018595 0.0709 0.0617

UPPER LOWER TOTAL

CN 0.2225 -0.2920 -0.0695CC -0.0017 -0.0114 -0.0131CM -0.0915 0.1097 0.0182

AIRFOIL PERFORMANCE" CL CD CM

-0.0695 -0.0131 0.0182

- 27-

TABLE 2.20


RUN NO, = 45

ALPHA =.-2,0

MACH NO, =0,2972

L_ING DATA FILE NAME = _STWD.DATINPUT FiLE NO, - 3

UPPER SURFACE LOWER SURFACE%CHORD CP LOCAL CP LOCAL.

0 -0.194& -0.19461 0,5343 -0,92362 0,2475 -0.99005 0.1177 -0.83917 -0,0121 -0,74569 -0.0604 -0.712315 -0,1207 -0,630920 -0,1479 -0,585625 -0.1660 -0.558429 -0.1871 -0.522235 -0_2022 -0.504140 -0,2143 -0,495044 -0,2243 -0.464850 -0,2294 -0,431655 -0,2203 -0,401560 -0,2173 -0,374364 -0,2053 -0,335070 -0.1871 -0.289875 -0.1630 -0.241580_ -0,1328 -0,178185 -0,0875 -0,114790 -0,0241 -0,024195 0,0543 0.0664

UPPER LOWER TOTAL

CN 0,1232 -0,4026 -0,2795CC 0,0105 -0.0321 -0,0217CM -0,0649 0,1344 0,0695


-0,2801 -0.0119 0,0695

- 28 -

TABLE 2.21


RUN NO, = 46

ALPHA = -4.0

MACH NO. =0.2959



0 -0.4750 -0.47501 0.8575 -1.8075.2 0,5583 -1,66255 0,3979 -1.33257 0.2375 -1.09199 0.1604 -0.9870

15 0.0617 -0,820520 -0.0123 -0,737225 -0.0339 -0.700229 -0.06_8 -0,657035 -0.0956 -0,620040 -0,1234 -0.589144 -0.1357 -0.555250 -0.1450 -0,505955 -0.1511 -0.459660 -0,1542 -0,428764 -0.1573 -0.370170 -0.1542 -0,305475 -0,1265 -0.2_9880 -0.0956 -0,182085 -0.0771 -0.111090 -0.0216 -0.012395 0.0463 0,0679 _-

UPPER LOWER TOTAL

CN 0.0315 -0.5170 -0.4856CC 0.0191 -0.0583 -0.0392CM -0.0392 0.1569 0.1177


-0,4871 -0.0052 0.1177

- 29 -

.oi..o,.,-

TABL_2.22o-_


RUN NO, = 47

ALPHA = -6.0

HACH NO, =0,3004

WING DATA F!LE NAHE = _STWD,DATINPUT FILE NO. - 5

UPF'ER SURFACE LOWER SURFACEZCHORD CP LOCAL CP LOCAL

0 -0.9486 -0.94861 1.0107 -2,90792 0.8097 -2.49125 0,7654 -2.07757 0,4374 -1.43039 0,3487 -1.3151

15 0.2305 -1,090520 0.1655 -0.975225 0.0827 -0,877729 0.0473 -0,809735 0,0059 -0,768340 -0,0118 -0.706344 -0.0325 -0.656150 -0.0650 -0.605855 -0.0739 -0.546760 -0.0827 -0.496564 -0.0857 -0.416770 -0.0887 -0.357675 -0,076S -0,2837BO -0o0680 -0.209885 -0.0443 -0,130090 -0,0089 -0,041495 0.0296 0,0207

UPPER LOWER TOTAL

CN -0.0623 -0.6746 -0.7369CC 0.0232 -0.0911 -0.0679CH -0.0098 0.1942 0.1845

AIRFOIL PERFORMANCECL C£1 CH

-0,7399 0.0095 0.1845

- 30 -

TABLE 2.23


RUN NO. = 48

ALPHA = -8.0

MACH NO. =0.2959



0 -1.4744 -1.47441 1.0215 -3.97032 0.9485 -3.5812

. 5 0.7585 -2.20107 0.5685 -1.83629 0.4682 -1.684215 0.3405 -1.349820 0.2523 -1.176525 0.1946 -1.057929 0.1398 -0.957635 0.0669 -0.872540 0.0456 -0.805644 0.0091 -0.753950 -0.0061 -0.662755 -0.0426 -0.601960 -0.0578 -0.532064 -0.0851 -0.446970 -0.0790 -0.367875 -0.0730 -0.288880 -0.0851 -0.218985 -0.0578 -0.112590 -0.0304 -0.057895 0.0000 -0.0122

UPPER LOWER TOTAL

CN -0.1001 -0.8053 -0.9054CC 0.0225 -0.1234 -0.1009CM -0.0013 0.2199 0.2185

AIRFOIL PERFORMANCE" CL CD CM

-0.9106 0.0261 0.2185

- 31 -

TABLE2.24


ALPHA = -10.0

MACH NO* =0.3013


UPF'ER SURFACE LOWER SURFACE

ZCHORD CP LOCAL CP LOCAL0 -1.9572 -1.95721 0.9913 -4.9058-4.89392 1.0062 -2 68815 0.9228 '7 0.6906 -2.18809 0.5775 -1.967715 0.4138 -1.565820 0 3275 -1.345525 0.2530 -1._0529 0.1905 -1.074635 0.1340 -0.982440 0.0625 -O.BBit44 0.0387 -0.7918-o.7o5_5o o.oooo -o.61o255 -0.026860 -0.0625 -0.5269-0.437664 -0.0893

-0.0923 -0.345370 -0.262075 -0.1027 -0.211480 -0.113185 -0.1161 -0.178690 -0.1191 -0,142995 -0,1191 -0.1280

UPPER LOWER TOTAL

CN -0.1119 -0.9335 -1.0454CC 0.0241 -0.1559 -0.1318CM -0.0098 0.2460 0.2362


-1.0524 0.0517 0.2362

- 32 -

FIGIIR_ 1

NACA0012-64

SECTIONPRESSURE

DISTRIBUTIONSAND FORCES

- 33 -

_IGT_RF,IoI

NACA 8012-64 Sect i on-1.5 RUN NO AL.Pr',A MACH NO

66 4.o8 8. 766

-1.8 e

0°5'_ UPPER

4.+ LOI.JER

CL CD CM8.5466 8. 8328 -_.14!8

g

I .8 1

FIGURE I.2

NACA 8912-64 See{, i onm

-l.S- RUN NO ALPHA MACH NO56 3.8 8. 697

e's" i" _ UPPER

. LOt,!ER

- CL CD CH8.3854 0.8027 -£1.6848

1.8

'k- 35 ....

IQ: *": •

FIGURE I.3

NACA 0012-64 Sec_. i on

"t's I RU_sNO ALPHA2.0MACHo.693N0 -

-I.0

....... Cp

-O.G+ + + . + +

+ .

. .

Ce _.g

o._ I t ! I I20 _ 60 80 _ 188

×/o

UPPER

+ LOWER

CL CD CH0.23S2 -0.0040 -0.0S84

- 36 -

FIGURE 1.4

NACA 0012-64 Sect, i on-1.s. RUN NO AI_PHA MACH NO

54 8.8 8. 683

4.. "_" . . . . 4- .

8.8 I I I l I20 4e ee ee _ !ce

x/o

0.6" UPPER

+ LOt,/ER

CL CD CH-6. I I 1I -8.0109 0.8279

t.8

- 37 -

FIGURE I.5

NACA 0012-64 Sect, i on-1.s- RUN NO ALPHA MACH NO

68 -2. O O. 7B 1

0.6-_ _ UPPER

_- LOWER

CL CD CM-8.4636 8.8013 8.!24,3

- 38 -

- 39 - _

FIGI_LE i. 7

•NACA 8812-64 Sec_. i on-_.s- RUN NO ALPHA HACH NO

48 4.8 0. $28

13

--1 ,0

, I:1

B

Cp + +_

8.8 I I I I , M. 28 _ 68 88 u 188

. X/O.

4_

8._"'t UPPER

•€. LOWER

CL CD CH0.4889 -8. 8838 -8. 8984

1.8

- 40 -

FIGUREI.8

NACA 8812-64 Sect, i on

-l.s. RUN NO ALPHA MACH NO53 3.8 8. 585

_" _ UPPER

. LO_,4ER

" CL CD CH8. 2697 -8.8863 '8.8593

1.8

- 41 -

FIGITRE i.9

NACA 0012--64 Sec{. ion

-l.s- RUN9NO_ ALPHA MACH NO2.8 0.516 _"

-I .8'

o.s. B UPPER

4. LO_,,'ER

CL CD CMe. 17SS -e.ee98 -e.0399

t.8

: :}

- 42 -

FIGURE I.I0

NACA 8012-64 Sec_ i on-I.s- RUN NO ALPHA HACH NO

36 8. (_ 0. SOS

•-! ,8"

UPPER

+ LOI,.IER.,

- CL CD CM-0.8728 -8.0136 8.0!74

1.8

- 43 -

FIGUP_E!.iI

NACA e812-64. Sec_. i on-t.s RUN NO ALPHA MACH NO

$2 -2. g 0.4.99

4-

-1.8 P 4-

UPPER

4- LOWER

CL CD CM-0.31.9°S -0.0136 0.0762

- 44 -

FIGURE 1.12

4. NACA OOl 2-64 Sec{_ i on-I .S -_" RUN NO ALPHA HACH NO

51 -3.O 8. 585

.

4.

4-4-

4-€.

4-

-O.S _".

4-

4.

Cp 4.

e.8 I ! ! I + 128 4,0 60 88 _ 1884-

e X/C

8._", UPPER

4- LOWER

CI.. CD CM-8.44!5 -8. 8124 8. 1839

1.8

- 45 -

FIGURE 1,13

" NACA 0812-64 Sect, i on-I.s- RUN NO ALPHA MACH NO

58 -4.8 g. SO4

i.

.

g.SUPPER

+ LO\-!ER

CL CD CH --0.5467 -0.0092 0.1279

1.8

- 46 -

" FIGURE I.14

NACA 8t312-64 Sec'c i on. -2.S e

RUN NO ALPHA MACH NO44 18,8 8.381

.B

-I.s

B

-1.8 G

C _P _

-8.Sm

m

8.8 ! I 4. * + + + + . + _ _ _i I I28 _ 68 88 188+

++ X/C

UPPER+

e.s- . * LOWER+ CL CD CM

" + 8.9753 8.8565 -8.2258

|. 8 '+

- 47 -

oFIGURE I.15

-2.s- NACA 8812-64 Sec[ i onRUN NO ALPHA MACH NO

43 8.8 8. 298

"2.8" E!

-1 .S- e

B

-1.8'

p _ N13

-8 .S-

ta

8.8- I _ + + + + + + + T + m28 4- + 48 68 88 188

4.

4- X/C.

UPPER.

_._- + _. LOWER

+ CL CD CN0.8317 0.8363 -8".209S

4-

1.8

- 48 -

FIGURE i.16

-2._- NACA 8812-64 Sec_. i onRUN NO ALPHA MACH NO

42 6.8 8. 299

-2.8-- B

la

-1.8-

P _ B

lae

Br4

B

+ _ +8.8 I + + + + + + + + . . + + _JI ' I 428 48 68 88 B ! 88

,,I.

+ . XJC+ e UPPER

8.s-. . LO_eER

, CL CD CM8.5872 0.0133 -0.1413

1.8

- 49 -

Q

FIGURE I.17

-2.s- NACA 6812-64 Sec_. ion -RUN NO ALPHA MACH NO

42 4-. 9 9. 304

-2.8-

-I ..q;"

m-! .8. FI

t=,l

C _,P

El

__I 13 13 El _I I_

ElEl

4, a, + 4" . 4. 4, 4" 6' 4" . 4""_ ! I ! I .e -I. 4

8.8 . 20 48 68 88 _ 188.

+ x/oUPPER

.e.s- 4, LOWER

I CL CD ON

0.3GS8 -9.90S8 -0.084t

1.8

- 50 -

FIGUI_,£I.18

-2.s NACA 8812-64 Sec L i onRUN NO ALPHA MACH NO

41 2.8 8. 296

-2.8

-I .S-

-I .8

CP Bral_ _

_8._.:a ra

B B B 13 B B '_ r4 B B

+_, _.+ .I- "_ 4" 4" -!, . . 4" . 4" 4" B.I-

8.8 I- I I I _ !4" 28 48 68 88 _. 108

x!c+ _ UPPER

e.s -:-LOWER

CL CD CM0. t608 -0.8189 -0.83,-06

1.8

- 51 -

FIGURE i.19

-2.s NACA B812-64 Sect ionRUN NO ALPHA MACH NO

4B 0.0 0. 298

-2.8

-I ,8"

CP

-0.S"

o.o I 4 I I _ 128 43 68 88 f_ 188

•a X/Oe UPPER

e:s . LOWER

CL CD CM-B.8695 -B.013! 8.0182

1.8--

- 52 -

FIGURE I.20

-2.s NACA 0812-64 SecE ionRUN NO ALPHA MACH NO

45 -2.8 0. 308

-2.8'

-t.B +

Cp +.++

. .

-8._ . . ... .. ++

+

0.8 _ t 'I ! 1 _ I28 4_ 68 80 € 188

a X/OUPPER

e.g- . LOWER

CL cD CM-0.2801 -0.8119 0.069S

1.8"

. _.,,. _'.."_" .r.

z

• . .. ._ .. . -

- 53-

FIGURE 1,21

-=.s- NACA 88I 2-64 Sec_ ionRUN NO ALPHA MACH NO

48 -4.8 0.296

-2.8"

4.

-1.8"

Cp **++ 6 . +

-O._' +4,

..

e.e _ 2_ 4_ 6_ 88 BIba

a UPPER[]

. L_WER8.g'CL CO CM

-0.4871 -_.88S2 B.1177

1.8

- 54 -

FICURE 1,22

-_0so-+ NACA 0012-64 Secl ionRUN NO ALPHA MACH NO

47 =808 8.3_8

4.

F

+

--_08-_ +I

p + ++

•-8.g +.

++

.

b_

2_ _m _ 4_ 68 80 w 1,_8

×/c_ _ UPPER

_._-- + LOWER

It CL CD CH

I:-i. -,. "-6. 7399 8. 8895 8 184S

'1.6

:!

"Z"

- 55 -

FIGURE I°23

-2._ NACA 0012-64 Sec! i onRUN NO ALPHA MACH NO

48 l - 8 . 8 8. 296,I,

-'2.8

,I.

.1 .S'

4-

,I,

-! .8 4,€.

Cp *+,,6

-8._" { 4..

.4-

4.

O.e 218 _ ,_ 68 8_ le9'B

El X/CIB_ UPPER

e.s = _- LOWER

Cl. CD CM-0.9106 8.8261 0.218S

r_1.8 "t_

- 56 -

FIGURE I.244

-2.s. NACA 8812-64 Sec_. i onRUN NO ALPHA MACH NO

49 -18.0 0.301 -.

-2,8' .

+-|.S

+

+

-f.8" .

+

C .P .

-8.g" ..

..

e.eL _ _ _ e8 lee

e X/O, e UPPER

e.B + LOIVER

CL CD CM-1.8S24 810S|7 0.2362

- 57 -t

i

FiGUI_ 2

I-IACHNI_[BER DISTRIBUTIONS ALONG .

THE CENTRELIh'E OF EACH FLEXIBLE WALL

- 58 -

"

F.IGURE 2.1

TSWT MACH NO. DISTRIBUTION...' ..

t·1ALLSRUN NO i~LPHA H.ACH NO

60 4.2 0.700

---------~-------"--t

I1~

t

i

POS!T!ON .cCliNSTRI?:A;1 Dr1/4 CH~R~ PT.CCHv~OS'

-r------rj-----r------,3 4 Ii:'

'---------------

----r------rj-------T----

-

--r-----.-.

TOP WALL

r ----r j r- I 1 : -r-.~ -3 -2 -1 0 , 2I

t~ODEL

BOTTOM WALL

--'.--'---1'

1.1 •

t1.0

0.9I

oJ.

I9.13T

2.7 1C! 1t:J

V1 . ~~ a.G\0 -',-. r

:r. -5u~

,;;:(>Q 1.1 T_I

Ij

1.2 I

19.9TI

0.8I

I0.7I

13.6 1

BOTH WALLS SET uSTRAIGHT-

.."';.

::.:.•:; ';'

... t.:

",;

0\oI

:-i

9 100.!0.9

0.8

0.711

O.!I

e.5 1

FIGURE 2.2

iSWT MACH NO. DISTRIBUTION\ ' ON~ Fl C'XT~l E 1't' LC'f~ I 0 __ \ __ ~ ~~L ~

RUN NO ALPHA MACH NO56 3.0 0.697

TOP WALL

0,./ '---------------------

----...1 ·---r,-----TI----"T'"l----T'----.,..-----TI-----,I-----"r-------,I-.... -3 -2 - t 3 2 3 4 5

IPOSITION DOwNSTREAM C~

fODEL 1/4 CI-lO~D PT. (CHORDS)j .

SOTTOM WALL

--'--'·--·--'T.-.-------y-,....-.. --rl-----,.-----:-~-----_r_----__r----·--r----__,

BOTH ~IALLS SET "STRAIG:;T-

FIGU_ 2,3

TSWT MACH NO. DISTRIBUTIONALONG FLEXIBLEWALLS

RUN NO ALPHA HACH NO55 2._ 0._93

1.9 T TOP 14ALL T,t_'g_

o.71

t 13.6 1

I .-5Z [ Z i i i I I i I i_. -5 -4 -3 -2 -I 8 I 2 3 4 $(,)

:U I/4 CHIRP PT. €CHORDS)MODEL

t...t

c_ I.@ IBOTTOH I_ALL

11.9

8.'J i

o. I :r I l 1" _ i I J i a 1

BOTH,,;.IALLS SET ",'_TR.A.T._HT"

FIGURE 2.4

UAL.LS

I

T

I1

i

- .....-------.l-----~·r

MODE:L

-,.r-----r-l----~------Ir------~ --1..------1-1 0 2:3 4 5

POSrrrON DCtlNSTREAti Of'1/4 C;j-'LR!) PT. CCMCRDS)·

-2-$

BOTTOt1 HALL

TOP ~ALL.

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

r-------,------

t"'l -~. J. J

I~.8 ~

I;

!...

1!1

.'.1~J..

c~

3 ~ .. ~ T

I~.n 1

ic.e 1:~. 7 t

I~.G t0.5 1

r-------r---------y------j----..----

,,,fJ'.~ ., ~,".. ~~.....,,;

:~

BOTH tl.ALLS SET "STf.~AIGHT"

FIGURE 2.5

TSWT MACH NO. DISTRIBUTIONALONG FLEXIBLE WALLS

RUN NO ALPHA t1.4CH NO68 -2.0 0.721

t.L

-~-----~~----~------

BOTTOM WALL

TO? WALL

_._-_._----- ------0.0

1.1 1

.:::1m9.7.1$ ..'. 21.0-,:: r-I-----rol-- -r-I---~I---__Ir-----:-I--- -.----"Tr----r-I-----,1-----.1To -5 -~ -:3 -2 .- t 0 1 2 3 4 Eo~ POSITION DO~NSTREAM Or:; MODEL 1/4 CHORD PT. CCHORDS)

(]9 1. t i

1.3 I0.9 I

0.3

1

BeTH WALLS SET "STRA~GHT·

FIGURE 2.6


RUN NO ALPHA i'1ACH NO67 -4.e 0.701

----------------

MODEL

BOTTOM WALL

TOP WALL

/

---~

tet I1.0

B.g II

9.6 1~ 9.7 tw~!.,IeB.1

2 . ~----r----.....,.-----r------r-----ir------r---- 'r------,..-----.--,-----.,~ f- I I I I i -, I , i-6 -<4 -3 -2 -I 0 2 3 4 6

PvSITIO:-l DOwNSTRE;.tt OF ...1/4 Cl.J.crw fiT. CCHCRDS)

~.- t

. Q 1.1 T

-l I1.9 te.9 t

I

0.8 +I

0.71. 0.6

,...r---~----r·--·---r'-----r-----iTr----~-r------r------r---

BOTH WALLS SET uSTRAIGHT Q

FIGURE 2.7

TSWT MACH NO. DISTRIBUTIONALONG FLEXIBLE wALLS

RUN NO ALPHA 1'~ACH NO40 4.2 2.523

0.8('~l

e.57st"" ~...""l0.":'0:><:..'1

0.52s1

0.5C~J~ 2. 47s1:.:>z

~ ... ~. ·TOP WALL

, i--4 -3

aOTTot1 \JALL

I-2

I-f

Io

t10DEL

I I I234

PCS!T!ON OOUNSTR~M OF1/-4 Cp.ofjJ PT. CCp.oRDSJ

ItI

II6

~-----~

9~57S

2.550

(3.525

0.se"-le.475

r-----r

...•- ..

..

e·

---'T---'-.BOTH WALLS SET ~STRAIGHTP

f-T------rr--- ----'--,

FIGURE 2.8

TSWT MACHNO. DISTRIBUTIONALONG FLEXIBLE_4ALLS

RUN NO ALPHA MACH NO

TOP [email protected]

'3. _?.S" "

i

I EM:TJO

o_ ,_ .47SI =$ ,-- _ i I I I | i i I | |

.j MODEL !/4 C1"-'.3.2DPT.CCF:OR.D$)o<_ @.._z_, BOTTOM WALL

Q...97_.

.525,

@.Se@,

@.47_

r F I 'T T T l i" ! I ,

BOTH WALLS SET .e _. _ "-

FIGURE 2.9

TSWT MACH NO. DISTRIBUTIONALONG FLEXIBLE _/ALLS

RUN NO ALPHA MACH NO- . 39 2.8 8.518

TOP WALL

8._50,

8.52_

•, _o.se_i1I :Jz. i i i i . I _ I I i I i

ig -S -4 -S -2 -1 e l 2 3 4

u_ l l PO_'rT:IONDO_NSTREAH OF

"" 174 C:Jg3gl:)PT ¢C1_9,%3S3._ MODEL ".goo 13,CO_..J BOTTOM %4ALL8.$7S

o.sse

0.525 , / _

o.s_}8

O. 475, " ...

1 i i i i i , i i i

-. BOTH 'dALLS SET _'STRAZS?IT"

FIGURE 2. I0

TS_ITMACH NO. DZSTRZBUTZON.............. ALONGn_T_, F UALLS

RUN NO ALPHA M-_,.

_.a_e,TOP _'ALL I

t ie.s._. T

I

_.._2_ ' Ii

!9..._,

I ":.! I

o', __,,, _T_,L ,koo ._ l i i i i ; i i l i 1

J .T. --._ -4 -::I -2 -I _ I 2 :3 ,4 l;POgIT"ZON DO_ZlqgTR_'_.AM'OF

-_ , FmDEL IQ 0 . ¢ 0_ l T

' BO'}'TOM _ALL i

1

f I" I I t I T T I I" i

_0, H _4ALLS SET "STRAZC_HT:'

I I! i

i i . _ j

FIGURE2.11

TSWT HACH NO. DISTRIBUTIONALONG FLEXIBLE!-,_ALLS

" RUN NO ALPHA HACH NO_2 -2. _ _. 499

:.. o.6_a. TOP WALL T

• 8.57_, I.. _) ,_=__,

_"- T

_. _._as. T

' N " :i

z 1- ] T T I" r l' i i J iI ;.

'° I I oo , : TRo,o.. MODEL I/4 Ct_PD PT.CCHOF,DS)

o' e'6e_"l BOTTOMWALL t..:.:'2!1.52_ ' I

i T I T i _ I I i I 1

BOTH WALLSSET "S'TRA_TGHT=

FIGURE 2.12

TSWT HACH NO. DISTRIBUTION........................ ALONG FLEXIBLE IJALLS .....

RUN NO ALPHA MACH NO• 51 -3.9 0.5_5

0._8,TOP t_ALL

0.576.

0._-25

t_

' _ _.475-, ;2o : , i = , i = z '" 1 z i=: --5 -4 -3 -2 - t 8 t 2 3 4 5I -.._

MODEL t/4c_k_ PT.CQ_RDS)

_.ee_ BOTTOM WALL

_OTH WALL,.?,.SET "3TRAZGHT"

FIGURE 2.13


RUN NO ALPHA MACH NO63 -4.3 0.584

II

t1

F~S!iION pot.n~3T;;~M CF1/4 C;.I,cN) PT 0 {Cr-:O:'DS'

..1I...

I j i I2 3 <4 5

,-

BOTTOM WALL

TOP 'WALL

----~I I I I i I

-04 -3 -2 -1 0 t•

1 HODEL-Jtj

39

•6031

0.575

0.5("-01

0.5'251a0 5001

10.476I

0.6231

2.5751

e.zF.31

C·~-5t~ 0.50.11lJ:en ~ 4-,r::s~...:l

Z ...., -----r-----.,-----y-----~---__T----_,_.----....__---___T----~---___,

(3 -5

~

BOTH t!ALLS SET ltSTRAIGHT·

FIGURE2.14

7SWT HACH NO. DZSTRZ_UTZONA'_,_'_-FLEXZE_LE_JALL$ ........L I,.;!I0

RUN NO ALPHA HACH NO44 1_.0 0.301

e. 4e._, TOP ',/ALL • T

I _-u bJ

_ _._7_ iI ..aZ I i i t i i i t I I I

o

I

•.-_ -4 -3 -2 - l 0 l 2 3 4 S

' _ "........ [ l F,I_SI'.'T_,'I l_l_!l'i'_,_TR-L_H 0FMODEL " I14 Clail.r-iDPT.¢CY,_P.DS_,3

BOTTOH IJALL_.375.

0.3_8.

...... 8.325'

@.27_

l I i I i i I / i i l

' 1BOTH 14A-LSSET "STRAZGHT"

FIGURE 2._15

TSWT HACH NO. DISTRIBUTIONALONG FLEXIBLE%.IALLSRUN NO ALFHA HACH D'O

•1,3 _. _ 8.2.98

8.4_8.TOP WALL

13.37S'

8"32g l13°.2_.=.

u 59.8.273 `L"4 . .-%

_ i _ i I i i i _ i i Ln .r. -3 -4 -3 -2 - t 8 ! 2 3 4 5

(..1

= I MeDEL 1 l/4. ,._-:02-,.PPT.(_1:_33:)

BOTTOM WALL.

8.37_ I....... 1_.3GO.

.275 -_i i i i i i i _ i i I

' _TH "_''__,,.._., SET °STRAZ_-..'HT"

FIGURE 2.16

TSWT MACH NO. DIS'iRI6UTZONALON9 FLEXIBLEI/ALI_$

,'XI;RUN NO _,-.._L^'_,.-_'_ M._CH _0

I_.34_TOP MALL ..

i -, n l i i i I '1 i 1

BOTL-I!_ALI_$ SET ".gTRAZ'_HT_

& ! J t

_ FIGURE 2.17

TSWT _.ACHNu. DISTRIBUTIONALONG FLEXIBLEWALLS

v,, NO ALPHA tIACH NO42 4,. O O. 3_4

._,,;_.TOP WALL

.339.

8.3"L>_,

1_.3:¢

t,..f I

(q_ _.._I ,I _ I I i i t J i l i :-i-- -S -4 -3 -__ _ ! Z 3 4 8

t_

MODEL 114 Ct-;Oi_Fr. cc_J_s_

_.34eI•_ BOTTOM t,'ALL

• _ °:_l

....... I_.-_ l

I i ! i I i _ I i J 1

FIGURE 2.18

TSWTMACHNO.DISTRIBUTION...................... ALOHGFLEXIBLE14ALLS

RUN NO ALPHA "HA,,HNO4! 2.O 0.2_

.s4@,TOP WALL

.4 '.'.4_ £q.@0%

i "" I l i _ i i I i I i i-._ -4 -3 -2 -I 8 I 2 3 4 E

HODEL 1/4 c_p,_ PT.(CV_RDS)

c_2 _,4_l"_" BOTTOM WALL

•_.2._ ]I i i i I I _ l i I l

,.,nTu!--,'ALLSZET °3TRAI_blT _'li,,,';_l • la

J

FIGURE 2. 19

TSWT MACH NO. DISTRIBUTIONALONG FLEXIBLEWALLS

RUN NO ALPHA MACH NO43 _. 8 _. 293

0.340:

z.3_! TOP UALL

oiq

i, i

-.,i g113 "_qlt3"l"",4 _ ° °'*I -, t l l i i I I I J i i

= -_ -4 -3 -2 - I @ I 2 3 4 I_,j

D_,,,b_i'_E.A,IOF"

< l I POS:L'T'T'_N"'_" ;" ":" MODEL 1/4 CPJ2"_PT"CCLJ_'D3)

_e.34_ T BOTTOM _ALL

@.3_2

.3_

! t i i T i : I 1 i 1

_OTH "_' 1_" :3'ZT ,,_',_,'rr..,-:'r,i'_"_i.i,._J . -..,_,l".,Pl_._', i,/ :'-.

.-.

FIGURE 2.20

TSWT HACH NO, D!STRi_.UTZONA'O_'cF_FvTR;F ' "

RUN NO ALPHA HAtH NO45 -2. _ 8.297

s's4s I TOP WALL

- -I .7_. I i i l i i i i i i i

-5 - 4 -3 --'> ~! @ I 2 3 4 ._L_

:_ MODEL !/4c_.2-_PT.c_:Hg?._S)@.84@.

BOTTOM [email protected]@.

... @.31_"

8.2S@

I 'I 't T i i I I i i :

_OTH WALLS SET ,'_",_T'-._'r,

FIGURE 2.21

TSWT MACH NO. DiST21BUTIONALON_ FLEXIBLE "_',_'_L,_o'_

RI I_i • €'_B_,_'_NO ALPHA MA..d HO4G -4. B _. _a4

0.3_B,TOP "ALL

O.B_-B'

0.31B

i e:.n .2.QB _ /_-_- -

l _. i i _ 'i i I' , i i i I•_. -E; -4 -3 -2 - 1 _ l 2 _ 4

MODEL I/4 _-_;4_ PT.CCHO;_._)

C._r_, BOTTOM _ALL

•@._..o.

@.3tO,

.2_3 _.t i i I i i i i i i ,

E_OTH ,, ^1 '

FIGUI_ 2.22

T$_T HACH NO. DISTRIBUTIONALO_G FLEXIB'EI_'ALLS

RUN NO ALPHA HACH NO47 -_. _ B.3_

TOP gALL.37_,

.3_,

!

I _ 1 ! _ i i i i i i i i-Z -4 -3 -2 - ! e ! 2 3 4 5

r._

HODEL t/4 _RD _T.C_-_F,P$_t"

_ _.4_, BOTTOM WALL.375

_.3_,

• - _ ._-2_,

i i i 1 i i = i 1 i i

BOTH WALLS _ET "$TRA2_HT °

FIGURE 2.23

TSWT MACHNO. DISTRIBUTIONALONG FLEXIBLE_ALLS

RUN NO ALPHA HACH I'.'O4.'-8 -8. _ B. 298

0'4_aI TOP WALLQ.37_+

,3501

Z.32gI

' t",%1 I

I I i I I I I I I I

•,2. -Ig -4 -3 -2 -I 8 I 2 3 4 6i t_

MODEL t14 c_,:_a_PT.{Cd_.IRDS}€$

°'_"€_°T BOTTOMWALL

, _'375 l

,. _._3_-_'_l

t I i I i _ i 1 i I 1

........... BOTH _dALL.SsE'r "STRA'rGI-IT" .....

FIGURE 2.24

........................ TSWT_'_"_',.,,,,n NO. DZSTRZBUTZON ........ALON8FLEXZBLE_'" ""

RUN _0 ALPI4A HACH NO49 -I_.8 8.391

_.37g! TOP _ALL

e._z_4

.L__. _7.q.__.;

N_ _ r i i 1 i I 1 | 1 i :t _ --S -,4 -_ -2 - ! 8 1 2 _ ,i,

-_ MODEL i Pc..,.,._,;j _O_':,k,'rl'z?__A_CF!/4 _-,_,.'_ PT.¢r..t_._3S)

BOTTOM WALL

t)._2;_,

_TI-! L'ALLS GET "'-_"_'"' ,'r,,

NACA 0012-64 Section

0.16

@

0.15

0.14

0.13

0.12

0.11

dCL/do<

per degree r0.10 + Walls streamlined

Walls straight• Linearisedtheory

Straightwall data

0.09 corrected by Allen& Vincentl method

r_ Straight wall datacorrected for residual

0.08 interferences

I I I I I

• 0 0.2 0.4 0.6 0.8Mach No

FIG. 3 SUMMARY OF MODEL DATA FROM FLEXIBLEWALLED WIND TUNNELSBELOWMACH 0.8

- 83 -


- 7 "" 1.27x 1M "" 0. ; Rc - 06I i _1 I- I I I U

0.5

0.4 e/ ."0.3

0.2

0.1

C L

o-5 -4 -3 -2 -1 0. 5Angle of Attack (deg)

-0.1----- TSWT streamlined wall data

g TSWT straight wall data

-0.2 '_ TSWT straight wall datacorrected by Allen & Vincentlmethod

-0.3 + TSWT straight wall data+ corrected for residual

A interferences

-0.4

t_ + G

-0.5

-0.6

®

-0.7 I I I t

FIG. 4(a) LIFT CURVE SLOPES; Mco _' 0.7

- 84 -

N A CA 0012- 64 Section

~ 106Mco= 0.5;Rc - 1.02 x0.6 , , , , , , , , , ,

0.5

• 0.4 •

0.3

O

0.2O

4-

0.1 +

CL

0 , - -'1_' Angle of Attack (deg)

!-0.1 /

//

-0.2 / "-*-- TSWT streamlinedwall data.

_/ e TSWT straight wall dataA TSWT straight wall data

-0.3 /+ e methodc°rrectedby Allen & Vincenti

A/ + TSWT straight wall data/ corrected for residual

• -0.4 /+ interferences

A/ o

" -0.5 //

Q

-0.6 l I , I

FIG. 4(b) LIFT CURVESLOPES; Moo "" 0.5

- 85 -

NACA 0012-_ _ctlon

__ 106Moo 0.3 ; Rc 0._ x

1.0 _ ' J ' e

A_0.8-+

f0.6 - ®// -

/"0.4- e_/ -

/

0.2- e/

CL /+0 I 1 I I

- 2 -8 -4 4 8 12

Angle of attack (deg)

-0.2 At -

/° ® TSWT straight wall data-0.4 A

A TSWT straight wall datacorrected by Allen & Vincentlmethod

-0.6 _//* -+-TSWT straight wall data

/

/ corrected for residual..f

A/ e interferences-0.8 +/

Ae

-1.0 -e

-1.2 t I

FIG. 4(c) LIFT CURVESLOPES; Moo"" 0.3

- 86 -

NACA 0012-64 SectionG

Mao-_ 0.7 ;Rc-_ 1.27x106

° 0.06 ' ' ' ' -e- TSWT straightwall data

A TSWT straightwall datacorrected by Allen & Vincentl

0.05 • - method

•--'-- TSWT streamlinedwall data

@0.03

CD &

0.02

0.01

e A0

-5 -4 -3 -2 0 5

Angle of Attack (deg)

" FIG. 5(a) VARIATION OF MODEL PRESSUREDRAG WITH ANGLE OFATTACK ; Mao -_ 0.7

- 87 -


Mao_' 0.5 ; Rc "" 1.02 x 106

0.005

I I I

0 '3 I '1% -5 -4- -2 - 0 1 _ s 4 5 6 7Angle of Attack (deg)

-0.005

A0

-o.ol-0.015

-0.02 i i _ I t i I I I J

--e-- TSWT straight wall data

•, TSWT straight wall data correctedby Allen & Vincentl method

TSWT streamlined wall data

FIG. 5(b) VARIATION OF MODEL PRESSUREDRAG WITH ANGLE OFATTACK ; M_o"" 0.5

I

88 - t


,,, 106Moo "0.3;R c- 0.68x

0.06 _ , , ,

Q

-0.02 i t i I

--e-- TSWT straight wall data

,_ TSWT straight wall data correctedby Allen & Vincenti method

[] TSWTstreamlined wall data

FIG. 5(c) VARIATION OF /V'C2)DELPRESSUREDRAG WITH ANGLE

OF ATTACK ; Moo_ 0.3

- 89 -

1. Report No. 2. GovernmentAccessionNo. 3. Recipient'sCatalc_jNo.

NASACR-165685-4. Title and S'ubtitle 5. Report Date

Model and Boundary Aerodynamic Data FromHigh Blockage April 1981Two-DimensionalAirfoil Tests in a Shallow Unstreamlined 6.PerformingOrganizationCodeTransonic Flexible Walled Test Section

7. Author(s) 8. Performing Organization Report No.

S. W. D. Wolf• 10. Work Unit No.

9. P.er_formingOrganization Name _nd Addr.e_

The University of Southampton 11.Contractor Grant No."

Department of Aeronautics and AstronauticsSouthampton, England NSG-7172

13. Type of Report and Period C_vered

12. Sponsoring Agency Name and Address Contractor ReportNational Aeronautics and Space AdministrationWashington DC 20546 14.SponsoringAgencyCode' 505-31-53-01

15. Supplementary Notes

Langley technical monitor: Charles L. LadsonSemi-annual Progress Report for the period to October 1980. The PrincipalInvestiqator was Dr. M. J. Goodyer.

16. Abstract

This is a semi-annual progress report for the period from May 1980 toOctober 1980 on work undertaken on NASAGrant NSG-7172entitled "The Self-Streamliningof the Test Section of a Transonic WindTunnel." The Principal Investigator isDr. M. J. Goodyer. Included in this report are data fro" an NACA0012-64 airfoilat Machnumberof 0.3, 0.5, and 0.7 with the test section flexible walls set in the"straight" position. Thesedata are presented becausethey may prove useful in thedevelopmentof wind tunnel correction techniques.

. 17. Key Words (SuggeSted by"Author(s)) 18. Distribution Statement :r

Facilities, research and supportF1exi bl e-wal I wind-tunnel Uncla-ssifled - Unlimi ted

I Star Category - 091

19. SecurityClassif.(of thisreport) 20. SecurityClassif.(of this page) 21. No. of Pages 22. Price"

Unclassi fled Unclassi fled 90 A05I

N-30S ForsalebytheNationalTechnicalInformationService,Springfield,Virginia22161

g

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NASA Contractor Report ]65685

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