USA 41 AIRFOIL (usa41-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: USA 41 AIRFOIL (usa41-il) Reynolds number: 50,000 Max Cl/Cd: 41.15 at α=7.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-usa41-il-50000.txt Download as CSV file: xf-usa41-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: USA 41 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
------ -------- --------- --------- -------- -------- --------
-7.750 -0.3416 0.10157 0.09480 -0.0209 1.0000 0.1614
-7.500 -0.3436 0.09979 0.09312 -0.0207 1.0000 0.1675
-7.250 -0.3584 0.10033 0.09385 -0.0224 1.0000 0.1701
-7.000 -0.3418 0.09451 0.08803 -0.0193 1.0000 0.1769
-6.750 -0.3473 0.09347 0.08712 -0.0208 1.0000 0.1833
-6.500 -0.3439 0.09021 0.08394 -0.0205 1.0000 0.1871
-6.250 -0.3388 0.08730 0.08109 -0.0195 1.0000 0.1945
-6.000 -0.3418 0.08589 0.07980 -0.0228 1.0000 0.1994
-5.750 -0.3339 0.08203 0.07599 -0.0194 1.0000 0.2059
-5.500 -0.3335 0.08031 0.07435 -0.0229 1.0000 0.2133
-5.250 -0.3276 0.07689 0.07098 -0.0197 1.0000 0.2206
-5.000 -0.3234 0.07432 0.06846 -0.0212 1.0000 0.2289
-4.750 -0.3165 0.07194 0.06610 -0.0232 1.0000 0.2412
-4.500 -0.3096 0.06920 0.06339 -0.0230 1.0000 0.2553
-4.000 -0.2969 0.06350 0.05777 -0.0198 1.0000 0.2873
-3.750 -0.2888 0.06091 0.05520 -0.0187 1.0000 0.3070
-3.500 -0.2804 0.05832 0.05265 -0.0179 1.0000 0.3316
-3.250 -0.2734 0.05589 0.05027 -0.0159 1.0000 0.3612
-3.000 -0.2683 0.05363 0.04807 -0.0127 1.0000 0.3955
-2.250 -0.2754 0.04766 0.04240 0.0057 1.0000 0.5474
-2.000 -0.0567 0.03745 0.02970 -0.0552 1.0000 0.1899
-1.750 -0.0201 0.03457 0.02619 -0.0579 1.0000 0.1649
-1.500 0.0071 0.03283 0.02407 -0.0588 1.0000 0.1630
-1.250 0.0329 0.03137 0.02223 -0.0594 1.0000 0.1623
-1.000 0.0585 0.03007 0.02048 -0.0597 1.0000 0.1598
-0.750 0.0817 0.02913 0.01924 -0.0597 1.0000 0.1603
-0.500 0.1026 0.02852 0.01839 -0.0595 1.0000 0.1633
-0.250 0.1229 0.02814 0.01773 -0.0592 1.0000 0.1683
0.000 0.1514 0.02794 0.01738 -0.0606 0.9966 0.1815
0.250 0.2152 0.02723 0.01640 -0.0675 0.9827 0.2311
0.500 0.2774 0.02374 0.01512 -0.0738 0.9710 1.0000
0.750 0.3304 0.02460 0.01536 -0.0792 0.9505 1.0000
1.000 0.3849 0.02536 0.01572 -0.0847 0.9309 1.0000
1.250 0.4350 0.02599 0.01609 -0.0892 0.9117 1.0000
1.500 0.4787 0.02652 0.01646 -0.0924 0.8914 1.0000
1.750 0.5270 0.02690 0.01672 -0.0962 0.8732 1.0000
2.000 0.5777 0.02710 0.01684 -0.0999 0.8562 1.0000
2.250 0.6131 0.02750 0.01722 -0.1012 0.8367 1.0000
2.500 0.6515 0.02775 0.01745 -0.1026 0.8186 1.0000
2.750 0.6906 0.02789 0.01760 -0.1039 0.8017 1.0000
3.000 0.7290 0.02795 0.01772 -0.1049 0.7855 1.0000
3.250 0.7656 0.02800 0.01781 -0.1054 0.7696 1.0000
3.500 0.7927 0.02844 0.01830 -0.1048 0.7518 1.0000
3.750 0.8202 0.02888 0.01880 -0.1041 0.7345 1.0000
4.000 0.8489 0.02925 0.01929 -0.1035 0.7181 1.0000
4.250 0.8773 0.02963 0.01976 -0.1029 0.7021 1.0000
4.500 0.9053 0.03005 0.02028 -0.1021 0.6863 1.0000
4.750 0.9325 0.03052 0.02086 -0.1012 0.6707 1.0000
5.000 0.9592 0.03103 0.02154 -0.1002 0.6552 1.0000
5.250 0.9853 0.03160 0.02224 -0.0992 0.6399 1.0000
5.500 1.0107 0.03224 0.02304 -0.0981 0.6246 1.0000
5.750 1.0359 0.03292 0.02389 -0.0969 0.6094 1.0000
6.000 1.0559 0.03407 0.02529 -0.0956 0.5936 1.0000
6.250 1.0759 0.03524 0.02669 -0.0942 0.5777 1.0000
6.500 1.0958 0.03644 0.02814 -0.0928 0.5618 1.0000
6.750 1.1170 0.03730 0.02925 -0.0910 0.5437 1.0000
7.000 1.1565 0.03326 0.02518 -0.0858 0.4977 1.0000
7.250 1.1775 0.03016 0.02188 -0.0801 0.4411 1.0000
7.500 1.1873 0.02885 0.02061 -0.0748 0.3873 1.0000
7.750 1.1757 0.02884 0.02019 -0.0673 0.2736 1.0000
8.000 1.1634 0.03217 0.02223 -0.0619 0.1689 1.0000
8.250 1.1662 0.03518 0.02474 -0.0586 0.1366 1.0000
8.500 1.1771 0.03766 0.02712 -0.0562 0.1176 1.0000
8.750 1.1987 0.04023 0.02969 -0.0546 0.1059 1.0000
9.000 1.2335 0.04341 0.03275 -0.0545 0.0973 1.0000
9.250 1.2582 0.04633 0.03595 -0.0536 0.0904 1.0000
9.500 1.2891 0.05070 0.04031 -0.0538 0.0859 1.0000
9.750 1.3045 0.05444 0.04459 -0.0521 0.0851 1.0000
10.000 1.3155 0.05855 0.04920 -0.0502 0.0849 1.0000
10.250 1.3206 0.06287 0.05398 -0.0481 0.0849 1.0000
10.500 1.3196 0.06719 0.05878 -0.0459 0.0849 1.0000
10.750 1.3132 0.07149 0.06348 -0.0436 0.0850 1.0000
11.000 1.3030 0.07589 0.06821 -0.0415 0.0852 1.0000
11.250 1.2892 0.08024 0.07283 -0.0394 0.0855 1.0000
11.500 1.2756 0.08485 0.07762 -0.0378 0.0859 1.0000
11.750 1.0747 0.08693 0.08061 -0.0323 0.0917 1.0000
12.000 1.0295 0.09578 0.08963 -0.0366 0.0937 1.0000
|
Polar data table (+)
Polar graphs
<< Back to USA 41 AIRFOIL (usa41-il)