" 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
https://ntrs.nasa.gov/search.jsp?R=19810012558 2018-06-17T16:13:56+00:00Z
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--
NACA SECTION ANALYSIS0012-64
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
NACA SECTION ANALYSIS0012-64
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
NACA SECTION ANALYSIS0012-64
RUN NO. = 54?
ALPHA = 0.0
MACH NO. =0.6831
WING DATA FILE NAME = _STWD.DATINPLIT FILE NO. - 15
UPPER SURFACE LOWER SURFACE%CHORD CP LOCAL CP LOCAL
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
NACA SECTION ANALYSIS0012-64
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
AIRFOIL PERFORMANCECL CD CM
-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
AIRFOIL PERFORMANCECL CD CM
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
AIRFOIL PERFORMANCECL CD CM
0.2697 -0.0063 -0.0593
- 16 -
TABLE 2.9
NACA SECTION ANALYSIS0012-64
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
AIRFOIL PERFORMANCECL CD CM
-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
UPPER SURFACE LOWER SURFACEZCHORD CP LOCAL CP LOCAL
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
AIRFOIL PERFORMANCECL CD CM
-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
UPPER SURFACE LOWER SURFACE%CHORD CP LOCAL CP LOCAL
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
AIRFOIL PERFORMANCECL CD CM
-0.5467 -0.0092 0.1279
- 21 -
TABLE 2.14
NACA SECTION ANALYSIS0012-64
RUN NO. = 44
ALPHA = 10.0
MACH NO. =0,3011
WING DATA FILE NAME = _WING1.DATINPUT FILE NO. - 3
UPPER SURFACE LOWER SURFACE%CHORD CP LOCAL CP LOCAL
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
AIRFOIL PERFORMANCECL CD CM
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
UPPER SURFACE LOWER SURFACEZCHORD CP LOCAL CP LOCAL
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
NACA SECTION ANALYSIS0012-64
RUN NO, =42
ALPHA = 6,0
;_ACH NO. =0.2987
WING DATA FILE NANE = _WING1,DATINPUT FILE NO, - 5
UPPER SURFACE LOWER SURFACE%CHORD CP LOCAL CP LOCAL
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
AIRFOIL PERFORMANCECL CD CM
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
UPPER SURFACE LOWER SURFACE%CHORD CP LOCAL CP LOCAL
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
AIRFOIL PERFORMANCECL CD CM
0.3658 -0.0058 -0.0841
- 25 -J
TABLE 2.18
NACA SECTION ANALYSIS0012-64
RUN NO. = 41:
ALPHA = 2.0
MACH NO. =0.2961
WING DATA FILE NAME = _WINGI.DATINF'UT FILE NO. - 7
UPPER SURFACE LOWER SURFACE%CHORD CP LOCAL CP LOCAL
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
AIRFOIL PERFORMANCECL CD CM
0.1608 -0.0109 -0.0396
- 26 -
TABLE 2'19
NACA SECTION ANALYSIS0012-64
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
NACA SECTION ANALYSIS0012-64
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
AIRFOIL PERFORMANCECL CD CM
-0,2801 -0.0119 0,0695
- 28 -
TABLE 2.21
NACA SECTION ANALYSIS0012-64
RUN NO, = 46
ALPHA = -4.0
MACH NO. =0.2959
WING DATA FILE NAME = _STWD.DATINPUT FILE NO. - 4
UPPER SURFACE LOWER SURFACE%CHORD CP LOCAL CP LOCAL
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
AIRFOIL PERFORMANCECL CD CM
-0,4871 -0.0052 0.1177
- 29 -
.oi..o,.,-
TABL_2.22o-_
NACA SECTION ANALYSIS0012-64
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
NACA SECTION ANALYSIS0012-64
RUN NO. = 48
ALPHA = -8.0
MACH NO. =0.2959
WING DATA FILE NAME = _STWD.DATINPUT FILE NO. - 6
UPPER SURFACE LOWER SURFACE%CHORD CP LOCAL CP LOCAL
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
NACA SECTION ANALYSIS0012-64
ALPHA = -10.0
MACH NO* =0.3013
WING DATA FILE NAME = _STWD.DATINPUT FILE NO. - 7
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
AIRFOIL PERFORMANCECL CD CM
-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
TSWT MACH NO. DISTRIBUTIONALONG FLEXIBLE WALLS
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
'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.$7S
o.sse
0.525 , / _
o.s_}8
O. 475, " ...
1 i i i i i , i i i
-. BOTH 'dALLS SET _'STRAZS?IT"
FIGURE 2. I0
TS_ITMACH NO. DZSTRZBUTZON.............. ALONGn_T_, F UALLS
RUN NO ALPHA M-_,.
_.a_e,TOP _'ALL I
t ie.s._. T
I
_.._2_ ' Ii
!9..._,
I ":.! I
o', __,,, _T_,L ,koo ._ l i i i i ; i i l i 1
J .T. --._ -4 -::I -2 -I _ I 2 :3 ,4 l;POgIT"ZON DO_ZlqgTR_'_.AM'OF
-_ , FmDEL IQ 0 . ¢ 0_ l T
' BO'}'TOM _ALL i
1
f I" I I t I T T I I" i
_0, H _4ALLS SET "STRAZC_HT:'
I I! i
i i . _ j
FIGURE2.11
TSWT HACH NO. DISTRIBUTIONALONG FLEXIBLE!-,_ALLS
" RUN NO ALPHA HACH NO_2 -2. _ _. 499
:.. o.6_a. TOP WALL T
• 8.57_, I.. _) ,_=__,
_"- T
_. _._as. T
' N " :i
z 1- ] T T I" r l' i i J iI ;.
'° I I oo , : TRo,o.. MODEL I/4 Ct_PD PT.CCHOF,DS)
o' e'6e_"l BOTTOMWALL t..:.:'2!1.52_ ' I
i T I T i _ I I i I 1
BOTH WALLSSET "S'TRA_TGHT=
FIGURE 2.12
TSWT HACH NO. DISTRIBUTION........................ ALONG FLEXIBLE IJALLS .....
RUN NO ALPHA MACH NO• 51 -3.9 0.5_5
0._8,TOP t_ALL
0.576.
0._-25
t_
' _ _.475-, ;2o : , i = , i = z '" 1 z i=: --5 -4 -3 -2 - t 8 t 2 3 4 5I -.._
MODEL t/4c_k_ PT.CQ_RDS)
_.ee_ BOTTOM WALL
_OTH WALL,.?,.SET "3TRAZGHT"
FIGURE 2.13
TSWT MACH NO. DISTRIBUTIONALONG FLEXIBLE WALLS
RUN NO ALPHA MACH NO63 -4.3 0.584
II
t1
F~S!iION pot.n~3T;;~M CF1/4 C;.I,cN) PT 0 {Cr-:O:'DS'
..1I...
I j i I2 3 <4 5
,-
BOTTOM WALL
TOP 'WALL
----~I I I I i I
-04 -3 -2 -1 0 t•
1 HODEL-Jtj
39
•6031
0.575
0.5("-01
0.5'251a0 5001
10.476I
0.6231
2.5751
e.zF.31
C·~-5t~ 0.50.11lJ:en ~ 4-,r::s~...:l
Z ...., -----r-----.,-----y-----~---__T----_,_.----....__---___T----~---___,
(3 -5
~
BOTH t!ALLS SET ltSTRAIGHT·
FIGURE2.14
7SWT HACH NO. DZSTRZ_UTZONA'_,_'_-FLEXZE_LE_JALL$ ........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 -
NACA 0012-64 Section
- 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 -
NACA 0012-64 Section
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
NACA 0012-64 Section
,,, 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
#
II
v,