NASA/LANGLEY RC10-64C AIRFOIL (rc1064c-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: NASA/LANGLEY RC10-64C AIRFOIL (rc1064c-il) Reynolds number: 200,000 Max Cl/Cd: 46.62 at α=3.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-rc1064c-il-200000-n5.txt Download as CSV file: xf-rc1064c-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NASA/LANGLEY RC10-64C AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.200 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.500 -0.6411 0.08136 0.07765 -0.0310 1.0000 0.0206
-10.250 -0.6924 0.06546 0.06168 -0.0431 1.0000 0.0200
-10.000 -0.7547 0.05446 0.05036 -0.0416 1.0000 0.0196
-9.750 -0.7843 0.04832 0.04387 -0.0379 1.0000 0.0196
-9.500 -0.7974 0.04413 0.03935 -0.0343 1.0000 0.0198
-9.250 -0.8030 0.04066 0.03554 -0.0309 1.0000 0.0200
-9.000 -0.8031 0.03771 0.03226 -0.0276 1.0000 0.0204
-8.750 -0.7984 0.03539 0.02964 -0.0247 1.0000 0.0210
-8.500 -0.7921 0.03309 0.02700 -0.0218 1.0000 0.0217
-8.250 -0.7840 0.03084 0.02436 -0.0189 1.0000 0.0224
-8.000 -0.7732 0.02879 0.02193 -0.0162 1.0000 0.0228
-7.750 -0.7597 0.02704 0.01983 -0.0139 1.0000 0.0232
-7.500 -0.7441 0.02561 0.01810 -0.0119 1.0000 0.0236
-7.250 -0.7280 0.02396 0.01627 -0.0101 1.0000 0.0241
-7.000 -0.7104 0.02281 0.01502 -0.0085 1.0000 0.0247
-6.750 -0.6920 0.02189 0.01401 -0.0070 1.0000 0.0253
-6.500 -0.6684 0.02110 0.01311 -0.0065 0.9991 0.0264
-6.250 -0.6373 0.02027 0.01217 -0.0076 0.9964 0.0279
-6.000 -0.6060 0.01938 0.01114 -0.0086 0.9939 0.0292
-5.750 -0.5768 0.01851 0.01015 -0.0091 0.9905 0.0302
-5.500 -0.5478 0.01754 0.00913 -0.0097 0.9869 0.0314
-5.250 -0.5171 0.01682 0.00839 -0.0107 0.9835 0.0333
-5.000 -0.4887 0.01624 0.00776 -0.0111 0.9789 0.0354
-4.750 -0.4577 0.01573 0.00719 -0.0119 0.9748 0.0387
-4.500 -0.4257 0.01513 0.00661 -0.0131 0.9716 0.0440
-4.250 -0.3953 0.01467 0.00611 -0.0138 0.9672 0.0503
-4.000 -0.3623 0.01421 0.00566 -0.0151 0.9636 0.0612
-3.750 -0.3273 0.01374 0.00526 -0.0169 0.9608 0.0768
-3.500 -0.2924 0.01329 0.00491 -0.0186 0.9579 0.0999
-3.250 -0.2582 0.01286 0.00465 -0.0203 0.9547 0.1341
-3.000 -0.2216 0.01272 0.00459 -0.0223 0.9520 0.1812
-2.750 -0.1850 0.01246 0.00424 -0.0243 0.9497 0.2010
-2.250 -0.1222 0.01119 0.00365 -0.0268 0.9427 0.3286
-2.000 -0.0932 0.01038 0.00346 -0.0275 0.9395 0.4882
-1.750 -0.0693 0.00989 0.00350 -0.0266 0.9335 0.6211
-1.500 -0.0436 0.00967 0.00369 -0.0257 0.9272 0.7325
-1.250 -0.0166 0.00971 0.00396 -0.0248 0.9197 0.8060
-1.000 0.0078 0.00984 0.00420 -0.0234 0.9097 0.8517
-0.750 0.0178 0.00987 0.00431 -0.0193 0.8876 0.8777
-0.500 0.0463 0.00989 0.00433 -0.0191 0.8780 0.8958
-0.250 0.0802 0.00988 0.00429 -0.0200 0.8715 0.9063
0.000 0.1124 0.00982 0.00420 -0.0206 0.8649 0.9150
0.250 0.1460 0.00974 0.00409 -0.0215 0.8587 0.9223
0.500 0.1775 0.00967 0.00398 -0.0220 0.8510 0.9295
0.750 0.2129 0.00961 0.00386 -0.0233 0.8434 0.9338
1.000 0.2435 0.00960 0.00380 -0.0236 0.8297 0.9394
1.250 0.2717 0.00963 0.00383 -0.0235 0.8111 0.9450
1.500 0.3035 0.00965 0.00382 -0.0241 0.7909 0.9489
1.750 0.3335 0.00968 0.00379 -0.0243 0.7700 0.9537
2.000 0.3615 0.00973 0.00379 -0.0241 0.7444 0.9591
2.250 0.3944 0.00981 0.00379 -0.0250 0.7159 0.9625
2.500 0.4251 0.00992 0.00381 -0.0255 0.6838 0.9669
2.750 0.4526 0.01009 0.00386 -0.0253 0.6457 0.9721
3.000 0.4821 0.01040 0.00387 -0.0255 0.5769 0.9754
3.250 0.5077 0.01089 0.00396 -0.0252 0.4881 0.9798
3.500 0.5325 0.01147 0.00415 -0.0248 0.4016 0.9844
3.750 0.5589 0.01223 0.00445 -0.0252 0.3005 0.9880
4.000 0.5849 0.01302 0.00481 -0.0255 0.2137 0.9921
4.250 0.6139 0.01365 0.00517 -0.0263 0.1558 0.9955
4.500 0.6442 0.01423 0.00557 -0.0274 0.1171 0.9989
4.750 0.6666 0.01473 0.00594 -0.0267 0.0914 1.0000
5.000 0.6850 0.01517 0.00632 -0.0250 0.0769 1.0000
5.250 0.7034 0.01561 0.00673 -0.0234 0.0675 1.0000
5.500 0.7220 0.01602 0.00717 -0.0217 0.0619 1.0000
5.750 0.7402 0.01648 0.00764 -0.0200 0.0570 1.0000
6.000 0.7572 0.01703 0.00818 -0.0181 0.0530 1.0000
6.250 0.7755 0.01747 0.00869 -0.0163 0.0500 1.0000
6.500 0.7932 0.01798 0.00924 -0.0145 0.0476 1.0000
6.750 0.8105 0.01855 0.00984 -0.0127 0.0455 1.0000
7.000 0.8263 0.01929 0.01060 -0.0107 0.0434 1.0000
7.250 0.8429 0.02003 0.01138 -0.0088 0.0418 1.0000
7.500 0.8610 0.02061 0.01206 -0.0071 0.0401 1.0000
7.750 0.8789 0.02129 0.01281 -0.0054 0.0384 1.0000
8.000 0.8968 0.02204 0.01363 -0.0038 0.0371 1.0000
8.250 0.9147 0.02283 0.01449 -0.0023 0.0359 1.0000
8.500 0.9325 0.02370 0.01541 -0.0008 0.0348 1.0000
8.750 0.9497 0.02493 0.01667 0.0006 0.0335 1.0000
9.000 0.9681 0.02603 0.01790 0.0019 0.0325 1.0000
9.250 0.9866 0.02694 0.01898 0.0033 0.0313 1.0000
9.500 1.0044 0.02806 0.02030 0.0046 0.0303 1.0000
9.750 1.0214 0.02929 0.02170 0.0061 0.0293 1.0000
10.000 1.0375 0.03054 0.02311 0.0075 0.0284 1.0000
10.250 1.0525 0.03165 0.02435 0.0090 0.0275 1.0000
10.500 1.0666 0.03268 0.02546 0.0106 0.0267 1.0000
10.750 1.0787 0.03412 0.02697 0.0122 0.0258 1.0000
11.000 1.0875 0.03596 0.02907 0.0143 0.0251 1.0000
11.250 1.0935 0.03778 0.03122 0.0168 0.0245 1.0000
11.500 1.0949 0.03971 0.03344 0.0198 0.0238 1.0000
11.750 1.0938 0.04172 0.03571 0.0227 0.0231 1.0000
12.000 1.0913 0.04377 0.03799 0.0252 0.0225 1.0000
12.250 1.0884 0.04581 0.04024 0.0274 0.0219 1.0000
12.500 1.0845 0.04800 0.04261 0.0292 0.0214 1.0000
12.750 1.0812 0.05020 0.04496 0.0304 0.0209 1.0000
13.000 1.0758 0.05284 0.04774 0.0311 0.0206 1.0000
13.250 1.0690 0.05584 0.05089 0.0313 0.0203 1.0000
13.500 1.0583 0.05963 0.05484 0.0308 0.0201 1.0000
13.750 1.0436 0.06434 0.05973 0.0293 0.0200 1.0000
14.000 1.0235 0.07042 0.06601 0.0264 0.0199 1.0000
14.250 0.9808 0.08176 0.07770 0.0192 0.0201 1.0000
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Polar data table (+)
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