USA 50 AIRFOIL (usa50-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
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Airfoil: USA 50 AIRFOIL (usa50-il) Reynolds number: 1,000,000 Max Cl/Cd: 92.27 at α=2.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-usa50-il-1000000.txt Download as CSV file: xf-usa50-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: USA 50 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 1.000 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.000 -0.5155 0.08110 0.07957 -0.0167 1.0000 0.0074
-7.750 -0.5264 0.07810 0.07661 -0.0169 1.0000 0.0075
-7.500 -0.5297 0.07375 0.07226 -0.0197 1.0000 0.0075
-7.250 -0.5303 0.06916 0.06766 -0.0225 1.0000 0.0075
-7.000 -0.5285 0.06460 0.06307 -0.0248 1.0000 0.0077
-6.750 -0.5253 0.05987 0.05828 -0.0265 1.0000 0.0079
-6.500 -0.5203 0.05535 0.05369 -0.0273 1.0000 0.0081
-6.250 -0.5142 0.05083 0.04906 -0.0274 1.0000 0.0083
-6.000 -0.4956 0.04557 0.04362 -0.0293 0.9991 0.0090
-5.750 -0.4625 0.04215 0.04006 -0.0314 0.9975 0.0099
-5.500 -0.4577 0.02014 0.01658 -0.0319 0.9930 0.0062
-5.250 -0.4332 0.01650 0.01241 -0.0315 0.9904 0.0063
-5.000 -0.4030 0.01551 0.01128 -0.0324 0.9883 0.0070
-4.750 -0.3714 0.01461 0.01024 -0.0335 0.9866 0.0078
-4.500 -0.3393 0.01342 0.00886 -0.0344 0.9851 0.0085
-4.250 -0.3069 0.01226 0.00750 -0.0354 0.9840 0.0089
-4.000 -0.2736 0.01144 0.00656 -0.0365 0.9830 0.0094
-3.750 -0.2489 0.01061 0.00562 -0.0357 0.9788 0.0098
-3.500 -0.2192 0.00985 0.00476 -0.0360 0.9760 0.0100
-3.250 -0.1907 0.00852 0.00321 -0.0360 0.9735 0.0118
-3.000 -0.1576 0.00807 0.00271 -0.0371 0.9718 0.0140
-2.750 -0.1231 0.00775 0.00232 -0.0386 0.9704 0.0163
-2.500 -0.0989 0.00738 0.00195 -0.0376 0.9640 0.0249
-2.250 -0.0673 0.00706 0.00177 -0.0384 0.9605 0.0583
-2.000 -0.0330 0.00686 0.00159 -0.0398 0.9578 0.0690
-1.750 -0.0071 0.00668 0.00141 -0.0393 0.9489 0.0786
-1.500 0.0278 0.00638 0.00119 -0.0409 0.9409 0.1015
-1.250 0.0661 0.00541 0.00099 -0.0438 0.9312 0.3475
-1.000 0.1071 0.00476 0.00085 -0.0471 0.9143 0.5233
-0.750 0.1398 0.00430 0.00076 -0.0484 0.8881 0.6766
-0.500 0.1599 0.00393 0.00074 -0.0464 0.8552 0.8136
-0.250 0.2304 0.00385 0.00086 -0.0558 0.8139 0.9549
0.000 0.2675 0.00421 0.00096 -0.0577 0.7547 0.9755
0.250 0.2988 0.00449 0.00102 -0.0584 0.7073 0.9841
0.500 0.3386 0.00473 0.00108 -0.0612 0.6647 0.9901
0.750 0.3745 0.00497 0.00113 -0.0631 0.6219 0.9948
1.000 0.4148 0.00519 0.00115 -0.0661 0.5748 0.9986
1.250 0.4459 0.00533 0.00116 -0.0669 0.5493 1.0000
1.500 0.4689 0.00542 0.00119 -0.0659 0.5332 1.0000
1.750 0.4911 0.00557 0.00123 -0.0648 0.5082 1.0000
2.000 0.5144 0.00567 0.00127 -0.0638 0.4866 1.0000
2.250 0.5370 0.00582 0.00132 -0.0627 0.4537 1.0000
2.500 0.5578 0.00612 0.00137 -0.0613 0.3846 1.0000
2.750 0.5760 0.00668 0.00153 -0.0595 0.2940 1.0000
3.000 0.5954 0.00718 0.00172 -0.0579 0.2198 1.0000
3.250 0.6092 0.00824 0.00212 -0.0554 0.0697 1.0000
3.500 0.6290 0.00879 0.00245 -0.0538 0.0175 1.0000
3.750 0.6518 0.00906 0.00277 -0.0526 0.0140 1.0000
4.000 0.6731 0.00951 0.00336 -0.0511 0.0112 1.0000
4.250 0.6948 0.00991 0.00383 -0.0497 0.0108 1.0000
4.500 0.7169 0.01026 0.00422 -0.0484 0.0102 1.0000
4.750 0.7385 0.01065 0.00466 -0.0471 0.0094 1.0000
5.000 0.7590 0.01118 0.00528 -0.0455 0.0088 1.0000
5.250 0.7791 0.01174 0.00591 -0.0439 0.0082 1.0000
5.500 0.7989 0.01234 0.00656 -0.0422 0.0075 1.0000
5.750 0.8152 0.01333 0.00759 -0.0400 0.0066 1.0000
6.000 0.8296 0.01484 0.00920 -0.0372 0.0063 1.0000
6.250 0.8484 0.01602 0.01048 -0.0353 0.0064 1.0000
6.500 0.8689 0.01711 0.01169 -0.0336 0.0067 1.0000
7.250 0.9216 0.02460 0.02008 -0.0262 0.0085 1.0000
7.500 0.9377 0.02612 0.02178 -0.0242 0.0079 1.0000
7.750 0.9524 0.02766 0.02347 -0.0222 0.0074 1.0000
8.000 0.9655 0.02947 0.02545 -0.0200 0.0071 1.0000
8.250 0.9786 0.03104 0.02707 -0.0183 0.0067 1.0000
8.500 0.9846 0.03393 0.03018 -0.0155 0.0064 1.0000
8.750 0.9508 0.04354 0.04039 -0.0086 0.0060 1.0000
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