USA 45 M AIRFOIL (usa45m-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: USA 45 M AIRFOIL (usa45m-il) Reynolds number: 50,000 Max Cl/Cd: 20.38 at α=2° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-usa45m-il-50000.txt Download as CSV file: xf-usa45m-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: USA 45 M AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 -0.3473 0.10733 0.10051 -0.0116 1.0000 0.2598
-8.750 -0.3586 0.10572 0.09903 -0.0124 1.0000 0.2699
-8.500 -0.3590 0.10365 0.09702 -0.0120 1.0000 0.2836
-8.250 -0.3447 0.09988 0.09328 -0.0109 1.0000 0.2984
-8.000 -0.3406 0.09712 0.09059 -0.0100 1.0000 0.3145
-7.750 -0.3413 0.09482 0.08839 -0.0089 1.0000 0.3313
-7.500 -0.3468 0.09292 0.08660 -0.0076 1.0000 0.3483
-7.250 -0.3143 0.08808 0.08174 -0.0058 1.0000 0.3720
-7.000 -0.3164 0.08614 0.07991 -0.0039 1.0000 0.3943
-6.500 -0.3121 0.08168 0.07562 0.0010 1.0000 0.4441
-6.250 -0.2953 0.07865 0.07265 0.0034 1.0000 0.4751
-6.000 -0.2942 0.07685 0.07096 0.0070 1.0000 0.5084
-5.750 -0.2637 0.07341 0.06751 0.0093 1.0000 0.5550
-4.250 -0.1324 0.05597 0.05041 0.0145 1.0000 0.7725
-4.000 -0.2195 0.05745 0.05230 0.0247 1.0000 0.7033
-3.750 -0.2841 0.05796 0.05314 0.0332 1.0000 0.6722
-3.500 -0.3782 0.04538 0.03840 -0.0128 1.0000 0.2209
-3.250 -0.3263 0.04124 0.03346 -0.0192 0.9875 0.1891
-3.000 -0.2603 0.03776 0.02927 -0.0268 0.9672 0.1708
-2.750 -0.1968 0.03554 0.02606 -0.0330 0.9446 0.1590
-2.500 -0.1324 0.03346 0.02355 -0.0393 0.9231 0.1590
-2.250 -0.0758 0.03168 0.02148 -0.0440 0.8995 0.1599
-2.000 -0.0069 0.02976 0.01938 -0.0502 0.8800 0.1613
-1.750 0.0622 0.02805 0.01760 -0.0559 0.8614 0.1683
-1.500 0.1083 0.02701 0.01653 -0.0581 0.8376 0.1804
-1.250 0.1517 0.02595 0.01547 -0.0595 0.8174 0.1958
-1.000 0.3071 0.02133 0.01337 -0.0772 0.7987 1.0000
-0.750 0.3294 0.02162 0.01323 -0.0757 0.7751 1.0000
-0.500 0.3530 0.02188 0.01312 -0.0744 0.7553 1.0000
-0.250 0.3759 0.02218 0.01311 -0.0730 0.7376 1.0000
0.000 0.3976 0.02255 0.01322 -0.0717 0.7211 1.0000
0.250 0.4183 0.02302 0.01346 -0.0703 0.7058 1.0000
0.500 0.4383 0.02356 0.01381 -0.0689 0.6913 1.0000
0.750 0.4580 0.02416 0.01425 -0.0675 0.6779 1.0000
1.000 0.4782 0.02477 0.01470 -0.0662 0.6656 1.0000
1.250 0.5011 0.02526 0.01499 -0.0650 0.6551 1.0000
1.500 0.5192 0.02606 0.01571 -0.0637 0.6434 1.0000
1.750 0.5368 0.02695 0.01652 -0.0624 0.6325 1.0000
2.000 0.5603 0.02749 0.01691 -0.0613 0.6235 1.0000
2.250 0.5756 0.02859 0.01800 -0.0600 0.6131 1.0000
2.500 0.5935 0.02958 0.01894 -0.0587 0.6039 1.0000
2.750 0.6141 0.03038 0.01967 -0.0575 0.5952 1.0000
3.000 0.6268 0.03183 0.02113 -0.0561 0.5862 1.0000
3.250 0.6503 0.03250 0.02171 -0.0551 0.5784 1.0000
3.500 0.6581 0.03435 0.02364 -0.0536 0.5692 1.0000
3.750 0.6843 0.03491 0.02411 -0.0527 0.5621 1.0000
4.000 0.6845 0.03743 0.02674 -0.0509 0.5531 1.0000
4.250 0.7145 0.03774 0.02698 -0.0502 0.5461 1.0000
4.500 0.7049 0.04111 0.03047 -0.0483 0.5371 1.0000
4.750 0.7393 0.04116 0.03048 -0.0477 0.5303 1.0000
5.000 0.7152 0.04576 0.03521 -0.0455 0.5215 1.0000
5.250 0.7570 0.04530 0.03471 -0.0450 0.5146 1.0000
5.500 0.7039 0.05238 0.04189 -0.0428 0.5072 1.0000
5.750 0.7592 0.05084 0.04034 -0.0423 0.4990 1.0000
6.000 0.6833 0.05977 0.04926 -0.0404 0.4946 1.0000
6.250 0.6522 0.06512 0.05459 -0.0397 0.4915 1.0000
6.500 0.6920 0.06532 0.05483 -0.0390 0.4817 1.0000
6.750 0.6608 0.07097 0.06045 -0.0390 0.4802 1.0000
7.000 0.6463 0.07547 0.06495 -0.0392 0.4795 1.0000
7.250 0.6398 0.07958 0.06908 -0.0397 0.4800 1.0000
7.500 0.6465 0.08343 0.07298 -0.0406 0.4821 1.0000
7.750 0.5582 0.09619 0.08580 -0.0488 0.5927 1.0000
8.000 0.5665 0.09830 0.08793 -0.0481 0.5779 1.0000
8.250 0.5627 0.10000 0.08963 -0.0470 0.5653 1.0000
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Polar data table (+)
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