USA 40 AIRFOIL (usa40-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: USA 40 AIRFOIL (usa40-il) Reynolds number: 500,000 Max Cl/Cd: 95.44 at α=5.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-usa40-il-500000.txt Download as CSV file: xf-usa40-il-500000.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 = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.750 -0.3106 0.09682 0.09440 -0.0401 1.0000 0.0427
-9.500 -0.3092 0.09407 0.09168 -0.0405 1.0000 0.0436
-9.250 -0.5152 0.03184 0.02795 -0.1014 0.9828 0.0359
-9.000 -0.4903 0.02820 0.02391 -0.1046 0.9771 0.0359
-8.750 -0.4585 0.02596 0.02149 -0.1074 0.9739 0.0363
-8.500 -0.4232 0.02471 0.02019 -0.1100 0.9719 0.0368
-8.250 -0.3870 0.02316 0.01850 -0.1128 0.9702 0.0373
-8.000 -0.3587 0.02139 0.01649 -0.1139 0.9637 0.0375
-7.750 -0.3252 0.01985 0.01475 -0.1157 0.9590 0.0378
-7.500 -0.2934 0.01855 0.01328 -0.1169 0.9524 0.0382
-7.250 -0.2633 0.01746 0.01204 -0.1175 0.9435 0.0386
-7.000 -0.2344 0.01653 0.01097 -0.1178 0.9343 0.0391
-6.750 -0.2046 0.01570 0.01001 -0.1181 0.9257 0.0396
-6.500 -0.1773 0.01505 0.00924 -0.1180 0.9146 0.0401
-6.250 -0.1486 0.01460 0.00866 -0.1179 0.9043 0.0406
-6.000 -0.1221 0.01359 0.00760 -0.1177 0.8926 0.0415
-5.750 -0.0950 0.01306 0.00703 -0.1175 0.8803 0.0422
-5.500 -0.0675 0.01263 0.00656 -0.1173 0.8682 0.0429
-5.250 -0.0399 0.01226 0.00611 -0.1170 0.8554 0.0438
-5.000 -0.0125 0.01190 0.00568 -0.1168 0.8417 0.0447
-4.750 0.0152 0.01159 0.00529 -0.1165 0.8281 0.0457
-4.500 0.0427 0.01123 0.00485 -0.1163 0.8139 0.0468
-4.250 0.0702 0.01086 0.00444 -0.1161 0.7993 0.0485
-4.000 0.0980 0.01067 0.00419 -0.1159 0.7842 0.0505
-3.750 0.1258 0.01052 0.00395 -0.1157 0.7683 0.0527
-3.500 0.1535 0.01023 0.00363 -0.1155 0.7521 0.0559
-3.250 0.1814 0.01010 0.00343 -0.1153 0.7361 0.0598
-3.000 0.2093 0.00989 0.00319 -0.1152 0.7205 0.0664
-2.750 0.2372 0.00971 0.00298 -0.1150 0.7047 0.0767
-2.500 0.2650 0.00956 0.00279 -0.1149 0.6893 0.0898
-2.250 0.2929 0.00939 0.00263 -0.1148 0.6748 0.1086
-2.000 0.3207 0.00916 0.00254 -0.1148 0.6604 0.1585
-1.750 0.3487 0.00905 0.00251 -0.1147 0.6461 0.2025
-1.500 0.3767 0.00900 0.00247 -0.1146 0.6321 0.2300
-1.250 0.4046 0.00893 0.00243 -0.1146 0.6186 0.2595
-1.000 0.4323 0.00884 0.00241 -0.1145 0.6046 0.3052
-0.750 0.4603 0.00867 0.00243 -0.1145 0.5897 0.3824
-0.500 0.4880 0.00850 0.00246 -0.1145 0.5742 0.4609
-0.250 0.5153 0.00838 0.00252 -0.1143 0.5575 0.5438
0.000 0.5422 0.00837 0.00262 -0.1140 0.5388 0.6150
0.250 0.5689 0.00846 0.00271 -0.1135 0.5181 0.6585
0.500 0.5955 0.00861 0.00279 -0.1130 0.4971 0.6874
0.750 0.6222 0.00877 0.00290 -0.1126 0.4778 0.7151
1.000 0.6490 0.00895 0.00300 -0.1122 0.4611 0.7347
1.500 0.7025 0.00927 0.00323 -0.1114 0.4336 0.7692
1.750 0.7292 0.00941 0.00333 -0.1110 0.4215 0.7823
2.000 0.7557 0.00959 0.00345 -0.1105 0.4094 0.7942
2.250 0.7824 0.00973 0.00355 -0.1102 0.3973 0.8078
2.500 0.8091 0.00986 0.00366 -0.1098 0.3866 0.8220
3.000 0.8616 0.01008 0.00387 -0.1089 0.3672 0.8491
3.250 0.8872 0.01022 0.00398 -0.1083 0.3581 0.8645
3.500 0.9127 0.01024 0.00407 -0.1077 0.3499 0.8869
3.750 0.9391 0.01025 0.00414 -0.1072 0.3407 1.0000
4.000 0.9674 0.01043 0.00426 -0.1073 0.3307 1.0000
4.250 0.9949 0.01065 0.00442 -0.1073 0.3218 1.0000
4.500 1.0222 0.01087 0.00458 -0.1073 0.3117 1.0000
4.750 1.0494 0.01109 0.00474 -0.1072 0.3018 1.0000
5.000 1.0758 0.01136 0.00493 -0.1069 0.2915 1.0000
5.250 1.1028 0.01157 0.00511 -0.1068 0.2813 1.0000
5.500 1.1290 0.01183 0.00532 -0.1066 0.2710 1.0000
5.750 1.1546 0.01214 0.00554 -0.1062 0.2578 1.0000
6.000 1.1798 0.01247 0.00579 -0.1058 0.2437 1.0000
6.250 1.2045 0.01284 0.00606 -0.1054 0.2269 1.0000
6.500 1.2281 0.01330 0.00638 -0.1048 0.2057 1.0000
6.750 1.2502 0.01388 0.00680 -0.1040 0.1810 1.0000
7.000 1.2702 0.01464 0.00734 -0.1029 0.1487 1.0000
7.250 1.2820 0.01610 0.00839 -0.1008 0.0915 1.0000
7.500 1.3030 0.01669 0.00892 -0.0998 0.0820 1.0000
7.750 1.3245 0.01720 0.00941 -0.0989 0.0761 1.0000
8.000 1.3458 0.01770 0.00988 -0.0979 0.0690 1.0000
8.250 1.3619 0.01859 0.01061 -0.0963 0.0413 1.0000
8.500 1.3797 0.01928 0.01129 -0.0949 0.0377 1.0000
8.750 1.3984 0.01985 0.01189 -0.0935 0.0365 1.0000
9.000 1.4148 0.02044 0.01254 -0.0918 0.0357 1.0000
9.250 1.4302 0.02111 0.01326 -0.0900 0.0350 1.0000
9.500 1.4442 0.02185 0.01407 -0.0881 0.0343 1.0000
9.750 1.4566 0.02271 0.01499 -0.0861 0.0336 1.0000
10.000 1.4672 0.02370 0.01605 -0.0839 0.0329 1.0000
10.250 1.4791 0.02462 0.01704 -0.0821 0.0326 1.0000
10.500 1.4897 0.02567 0.01816 -0.0802 0.0322 1.0000
10.750 1.4985 0.02688 0.01945 -0.0784 0.0319 1.0000
11.000 1.5060 0.02826 0.02090 -0.0766 0.0315 1.0000
11.250 1.5117 0.02985 0.02257 -0.0749 0.0312 1.0000
11.500 1.5158 0.03167 0.02448 -0.0733 0.0309 1.0000
11.750 1.5179 0.03377 0.02667 -0.0719 0.0306 1.0000
12.000 1.5178 0.03622 0.02921 -0.0707 0.0302 1.0000
12.250 1.5153 0.03905 0.03213 -0.0697 0.0299 1.0000
12.500 1.5100 0.04236 0.03554 -0.0690 0.0296 1.0000
12.750 1.5024 0.04608 0.03937 -0.0687 0.0294 1.0000
13.000 1.4922 0.05030 0.04369 -0.0686 0.0291 1.0000
13.250 1.4808 0.05474 0.04824 -0.0687 0.0290 1.0000
13.500 1.4693 0.05930 0.05289 -0.0690 0.0288 1.0000
13.750 1.4670 0.06285 0.05655 -0.0693 0.0287 1.0000
14.000 1.4637 0.06658 0.06036 -0.0697 0.0285 1.0000
14.250 1.4599 0.07042 0.06430 -0.0702 0.0284 1.0000
14.500 1.4560 0.07432 0.06828 -0.0707 0.0282 1.0000
14.750 1.4522 0.07822 0.07226 -0.0713 0.0281 1.0000
15.000 1.4487 0.08208 0.07620 -0.0718 0.0279 1.0000
15.250 1.4460 0.08585 0.08003 -0.0724 0.0276 1.0000
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