USA 40 AIRFOIL (usa40-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: USA 40 AIRFOIL (usa40-il) Reynolds number: 200,000 Max Cl/Cd: 69.22 at α=6° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-usa40-il-200000-n5.txt Download as CSV file: xf-usa40-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: USA 40 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
------ -------- --------- --------- -------- -------- --------
-9.750 -0.3314 0.09286 0.08897 -0.0431 1.0000 0.0371
-9.500 -0.3476 0.08668 0.08284 -0.0453 1.0000 0.0366
-9.250 -0.5171 0.04032 0.03555 -0.0905 0.9808 0.0352
-9.000 -0.4990 0.03414 0.02863 -0.0977 0.9721 0.0357
-8.750 -0.4710 0.03223 0.02660 -0.1001 0.9668 0.0361
-8.500 -0.4417 0.03062 0.02485 -0.1022 0.9611 0.0366
-8.250 -0.4090 0.02882 0.02285 -0.1050 0.9571 0.0371
-8.000 -0.3823 0.02708 0.02087 -0.1063 0.9494 0.0376
-7.750 -0.3510 0.02538 0.01892 -0.1082 0.9440 0.0380
-7.500 -0.3206 0.02389 0.01720 -0.1097 0.9379 0.0385
-7.250 -0.2908 0.02258 0.01567 -0.1107 0.9303 0.0391
-7.000 -0.2570 0.02133 0.01420 -0.1123 0.9251 0.0397
-6.750 -0.2288 0.02034 0.01302 -0.1127 0.9150 0.0403
-6.500 -0.1953 0.01936 0.01187 -0.1138 0.9061 0.0410
-6.250 -0.1658 0.01861 0.01111 -0.1143 0.8930 0.0418
-6.000 -0.1363 0.01801 0.01047 -0.1146 0.8801 0.0428
-5.750 -0.1062 0.01742 0.00979 -0.1149 0.8689 0.0439
-5.500 -0.0778 0.01683 0.00911 -0.1150 0.8562 0.0449
-5.250 -0.0496 0.01627 0.00845 -0.1149 0.8432 0.0459
-5.000 -0.0212 0.01575 0.00784 -0.1149 0.8301 0.0469
-4.750 0.0069 0.01529 0.00735 -0.1148 0.8164 0.0481
-4.500 0.0346 0.01492 0.00693 -0.1147 0.8018 0.0497
-4.250 0.0625 0.01458 0.00651 -0.1145 0.7870 0.0517
-4.000 0.0903 0.01427 0.00613 -0.1144 0.7723 0.0543
-3.750 0.1182 0.01401 0.00581 -0.1143 0.7573 0.0575
-3.500 0.1461 0.01375 0.00546 -0.1141 0.7421 0.0615
-3.250 0.1738 0.01351 0.00516 -0.1139 0.7265 0.0665
-3.000 0.2016 0.01328 0.00487 -0.1138 0.7109 0.0731
-2.750 0.2294 0.01309 0.00461 -0.1136 0.6956 0.0820
-2.500 0.2571 0.01290 0.00438 -0.1135 0.6808 0.0937
-2.250 0.2847 0.01270 0.00419 -0.1133 0.6663 0.1107
-1.750 0.3398 0.01236 0.00402 -0.1132 0.6371 0.1924
-1.500 0.3674 0.01224 0.00395 -0.1131 0.6233 0.2306
-1.250 0.3949 0.01212 0.00390 -0.1130 0.6092 0.2701
-1.000 0.4222 0.01202 0.00388 -0.1128 0.5949 0.3150
-0.750 0.4496 0.01192 0.00388 -0.1127 0.5801 0.3686
-0.500 0.4768 0.01181 0.00391 -0.1125 0.5655 0.4297
-0.250 0.5034 0.01170 0.00399 -0.1121 0.5502 0.5036
0.000 0.5294 0.01167 0.00411 -0.1115 0.5345 0.5756
0.250 0.5551 0.01173 0.00423 -0.1108 0.5191 0.6341
0.500 0.5809 0.01185 0.00434 -0.1101 0.5038 0.6751
0.750 0.6066 0.01198 0.00444 -0.1094 0.4886 0.7048
1.000 0.6327 0.01213 0.00453 -0.1088 0.4741 0.7269
1.250 0.6586 0.01230 0.00463 -0.1082 0.4603 0.7455
1.500 0.6843 0.01248 0.00473 -0.1076 0.4473 0.7626
1.750 0.7097 0.01266 0.00484 -0.1069 0.4345 0.7798
2.000 0.7353 0.01281 0.00495 -0.1063 0.4225 0.7973
2.250 0.7603 0.01297 0.00507 -0.1056 0.4121 0.8146
2.500 0.7854 0.01310 0.00519 -0.1049 0.4023 0.8328
2.750 0.8103 0.01324 0.00530 -0.1041 0.3931 0.8537
3.000 0.8344 0.01333 0.00540 -0.1032 0.3826 0.8808
3.250 0.8620 0.01335 0.00546 -0.1029 0.3714 1.0000
3.500 0.8887 0.01365 0.00563 -0.1028 0.3605 1.0000
3.750 0.9160 0.01389 0.00582 -0.1028 0.3496 1.0000
4.000 0.9425 0.01417 0.00602 -0.1026 0.3400 1.0000
4.250 0.9690 0.01444 0.00623 -0.1024 0.3296 1.0000
4.500 0.9953 0.01472 0.00645 -0.1022 0.3202 1.0000
4.750 1.0212 0.01502 0.00669 -0.1019 0.3111 1.0000
5.000 1.0472 0.01530 0.00694 -0.1017 0.3024 1.0000
5.250 1.0724 0.01562 0.00720 -0.1013 0.2932 1.0000
5.500 1.0980 0.01591 0.00748 -0.1010 0.2839 1.0000
5.750 1.1225 0.01627 0.00777 -0.1005 0.2742 1.0000
6.000 1.1476 0.01658 0.00808 -0.1001 0.2645 1.0000
6.250 1.1718 0.01694 0.00840 -0.0996 0.2560 1.0000
6.500 1.1956 0.01732 0.00874 -0.0990 0.2450 1.0000
6.750 1.2192 0.01770 0.00911 -0.0984 0.2346 1.0000
7.000 1.2418 0.01814 0.00950 -0.0977 0.2230 1.0000
7.250 1.2634 0.01865 0.00993 -0.0968 0.2081 1.0000
7.500 1.2838 0.01923 0.01042 -0.0959 0.1906 1.0000
7.750 1.3028 0.01989 0.01099 -0.0947 0.1715 1.0000
8.000 1.3196 0.02071 0.01166 -0.0933 0.1478 1.0000
8.250 1.3251 0.02235 0.01292 -0.0906 0.0993 1.0000
8.500 1.3382 0.02323 0.01374 -0.0886 0.0902 1.0000
8.750 1.3522 0.02404 0.01455 -0.0868 0.0849 1.0000
9.000 1.3674 0.02477 0.01532 -0.0852 0.0807 1.0000
9.250 1.3808 0.02563 0.01620 -0.0834 0.0767 1.0000
9.500 1.3960 0.02636 0.01699 -0.0819 0.0723 1.0000
9.750 1.4101 0.02719 0.01787 -0.0804 0.0644 1.0000
10.000 1.4094 0.02910 0.01963 -0.0777 0.0401 1.0000
10.250 1.4159 0.03056 0.02114 -0.0758 0.0383 1.0000
10.500 1.4230 0.03204 0.02269 -0.0742 0.0372 1.0000
10.750 1.4307 0.03353 0.02428 -0.0727 0.0365 1.0000
11.000 1.4371 0.03519 0.02605 -0.0714 0.0359 1.0000
11.250 1.4422 0.03705 0.02802 -0.0702 0.0354 1.0000
11.500 1.4455 0.03917 0.03026 -0.0691 0.0349 1.0000
11.750 1.4466 0.04161 0.03282 -0.0682 0.0343 1.0000
12.000 1.4459 0.04437 0.03570 -0.0676 0.0339 1.0000
12.250 1.4428 0.04751 0.03897 -0.0671 0.0335 1.0000
12.500 1.4416 0.05057 0.04217 -0.0669 0.0332 1.0000
12.750 1.4393 0.05387 0.04561 -0.0669 0.0329 1.0000
13.000 1.4351 0.05750 0.04937 -0.0671 0.0326 1.0000
13.250 1.4295 0.06142 0.05344 -0.0675 0.0324 1.0000
13.500 1.4224 0.06567 0.05782 -0.0681 0.0321 1.0000
13.750 1.4141 0.07019 0.06248 -0.0689 0.0320 1.0000
14.000 1.4048 0.07499 0.06742 -0.0700 0.0318 1.0000
14.250 1.3948 0.08000 0.07256 -0.0711 0.0316 1.0000
14.500 1.3844 0.08517 0.07785 -0.0725 0.0314 1.0000
14.750 1.3738 0.09044 0.08325 -0.0739 0.0312 1.0000
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