USA 40 AIRFOIL (usa40-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: USA 40 AIRFOIL (usa40-il) Reynolds number: 100,000 Max Cl/Cd: 50.08 at α=8.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-usa40-il-100000.txt Download as CSV file: xf-usa40-il-100000.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 = 0.100 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.000 -0.2872 0.09644 0.09162 -0.0319 1.0000 0.1427
-7.750 -0.3340 0.09707 0.09247 -0.0298 1.0000 0.1456
-7.500 -0.3872 0.09761 0.09321 -0.0284 1.0000 0.1463
-7.250 -0.3827 0.09409 0.08975 -0.0247 1.0000 0.1476
-7.000 -0.3624 0.09156 0.08723 -0.0203 1.0000 0.1498
-6.750 -0.3683 0.09035 0.08609 -0.0170 1.0000 0.1520
-6.500 -0.3810 0.08927 0.08510 -0.0145 1.0000 0.1548
-6.250 -0.3876 0.08364 0.07951 -0.0300 0.9931 0.1639
-6.000 -0.3472 0.07901 0.07475 -0.0435 0.9828 0.1787
-5.750 -0.3208 0.07500 0.07081 -0.0420 0.9757 0.1817
-5.500 -0.2831 0.07220 0.06800 -0.0432 0.9701 0.1875
-5.250 -0.2324 0.04064 0.03432 -0.0884 0.9574 0.1045
-5.000 -0.1908 0.03682 0.02998 -0.0930 0.9483 0.1017
-4.750 -0.1412 0.03340 0.02600 -0.0985 0.9422 0.1019
-4.500 -0.1011 0.03088 0.02305 -0.1016 0.9325 0.1020
-4.250 -0.0505 0.02866 0.02033 -0.1059 0.9263 0.1030
-4.000 -0.0123 0.02674 0.01835 -0.1082 0.9165 0.1061
-3.750 0.0371 0.02536 0.01687 -0.1119 0.9101 0.1124
-3.500 0.0778 0.02404 0.01534 -0.1139 0.9004 0.1178
-3.250 0.1226 0.02273 0.01412 -0.1165 0.8931 0.1272
-3.000 0.1581 0.02179 0.01327 -0.1174 0.8820 0.1404
-2.750 0.2009 0.02070 0.01224 -0.1194 0.8742 0.1634
-2.500 0.2322 0.02000 0.01169 -0.1195 0.8611 0.1972
-2.250 0.2672 0.01879 0.01091 -0.1204 0.8508 0.2709
-2.000 0.2995 0.01785 0.01048 -0.1206 0.8390 0.3757
-1.750 0.3264 0.01728 0.01028 -0.1198 0.8248 0.4646
-1.500 0.3525 0.01685 0.01022 -0.1182 0.8112 0.5624
-1.250 0.3778 0.01664 0.01018 -0.1161 0.7983 0.6492
-1.000 0.4010 0.01657 0.01014 -0.1135 0.7841 0.7144
-0.750 0.4201 0.01656 0.01013 -0.1102 0.7682 0.7684
-0.500 0.4365 0.01641 0.00998 -0.1062 0.7524 0.8158
-0.250 0.4526 0.01612 0.00966 -0.1024 0.7367 0.8598
0.000 0.4706 0.01574 0.00924 -0.0990 0.7210 0.9043
0.250 0.5085 0.01537 0.00877 -0.0998 0.7034 0.9625
0.500 0.5456 0.01538 0.00854 -0.1018 0.6851 1.0000
0.750 0.5769 0.01553 0.00846 -0.1026 0.6675 1.0000
1.000 0.6077 0.01573 0.00844 -0.1033 0.6501 1.0000
1.250 0.6379 0.01597 0.00847 -0.1038 0.6333 1.0000
1.500 0.6678 0.01625 0.00854 -0.1042 0.6173 1.0000
1.750 0.6976 0.01653 0.00862 -0.1045 0.6023 1.0000
2.000 0.7265 0.01684 0.00876 -0.1047 0.5875 1.0000
2.250 0.7543 0.01719 0.00900 -0.1047 0.5727 1.0000
2.500 0.7821 0.01755 0.00925 -0.1047 0.5592 1.0000
2.750 0.8107 0.01790 0.00944 -0.1048 0.5474 1.0000
3.000 0.8387 0.01826 0.00968 -0.1048 0.5354 1.0000
3.250 0.8655 0.01868 0.01006 -0.1046 0.5237 1.0000
3.500 0.8941 0.01906 0.01027 -0.1047 0.5139 1.0000
3.750 0.9205 0.01949 0.01070 -0.1045 0.5029 1.0000
4.000 0.9478 0.01995 0.01108 -0.1044 0.4935 1.0000
4.250 0.9750 0.02036 0.01142 -0.1042 0.4837 1.0000
4.500 1.0010 0.02086 0.01191 -0.1039 0.4740 1.0000
4.750 1.0287 0.02128 0.01221 -0.1038 0.4652 1.0000
5.000 1.0535 0.02182 0.01280 -0.1034 0.4555 1.0000
5.250 1.0818 0.02218 0.01299 -0.1033 0.4463 1.0000
5.500 1.1046 0.02263 0.01352 -0.1025 0.4350 1.0000
5.750 1.1305 0.02301 0.01380 -0.1021 0.4248 1.0000
6.000 1.1551 0.02336 0.01414 -0.1016 0.4144 1.0000
6.250 1.1792 0.02385 0.01465 -0.1010 0.4050 1.0000
6.500 1.2048 0.02419 0.01490 -0.1005 0.3956 1.0000
6.750 1.2272 0.02470 0.01550 -0.0997 0.3864 1.0000
7.000 1.2528 0.02505 0.01578 -0.0993 0.3780 1.0000
7.250 1.2750 0.02562 0.01647 -0.0985 0.3700 1.0000
7.500 1.2991 0.02604 0.01689 -0.0980 0.3623 1.0000
7.750 1.3217 0.02653 0.01743 -0.0972 0.3543 1.0000
8.000 1.3438 0.02688 0.01783 -0.0963 0.3457 1.0000
8.250 1.3655 0.02731 0.01830 -0.0954 0.3376 1.0000
8.500 1.3863 0.02768 0.01873 -0.0944 0.3290 1.0000
8.750 1.4075 0.02812 0.01920 -0.0934 0.3215 1.0000
9.000 1.4266 0.02859 0.01981 -0.0922 0.3137 1.0000
9.250 1.4477 0.02895 0.02014 -0.0912 0.3063 1.0000
9.500 1.4635 0.02948 0.02086 -0.0895 0.2979 1.0000
9.750 1.4848 0.02983 0.02115 -0.0886 0.2914 1.0000
10.000 1.4978 0.03050 0.02207 -0.0866 0.2835 1.0000
10.250 1.5152 0.03090 0.02248 -0.0852 0.2766 1.0000
10.500 1.5274 0.03161 0.02339 -0.0831 0.2692 1.0000
10.750 1.5410 0.03219 0.02406 -0.0813 0.2626 1.0000
11.000 1.5510 0.03288 0.02486 -0.0790 0.2557 1.0000
11.250 1.5583 0.03366 0.02579 -0.0764 0.2487 1.0000
11.500 1.5670 0.03443 0.02661 -0.0741 0.2425 1.0000
11.750 1.5713 0.03555 0.02794 -0.0716 0.2352 1.0000
12.000 1.5771 0.03654 0.02895 -0.0694 0.2287 1.0000
12.250 1.5790 0.03804 0.03069 -0.0672 0.2208 1.0000
12.500 1.5807 0.03950 0.03220 -0.0652 0.2134 1.0000
12.750 1.5791 0.04149 0.03440 -0.0635 0.2038 1.0000
13.000 1.5766 0.04371 0.03673 -0.0620 0.1948 1.0000
13.250 1.5709 0.04636 0.03945 -0.0608 0.1852 1.0000
13.500 1.5635 0.04955 0.04279 -0.0601 0.1744 1.0000
13.750 1.5524 0.05334 0.04670 -0.0597 0.1633 1.0000
14.000 1.5376 0.05783 0.05127 -0.0598 0.1523 1.0000
14.250 1.5189 0.06308 0.05658 -0.0605 0.1422 1.0000
14.500 1.4970 0.06906 0.06263 -0.0616 0.1328 1.0000
14.750 1.4739 0.07557 0.06922 -0.0632 0.1243 1.0000
15.000 1.4491 0.08258 0.07628 -0.0652 0.1178 1.0000
15.250 1.4252 0.08969 0.08343 -0.0673 0.1112 1.0000
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Polar data table (+)
Polar graphs
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