USA 50 AIRFOIL (usa50-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: USA 50 AIRFOIL (usa50-il) Reynolds number: 50,000 Max Cl/Cd: 36.45 at α=5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-usa50-il-50000.txt Download as CSV file: xf-usa50-il-50000.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 = 0.050 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.000 -0.4981 0.09927 0.09258 -0.0080 1.0000 0.2257
-7.750 -0.4989 0.09641 0.08980 -0.0073 1.0000 0.2397
-7.500 -0.5050 0.09397 0.08747 -0.0065 1.0000 0.2542
-7.250 -0.4905 0.08994 0.08338 -0.0041 1.0000 0.2772
-7.000 -0.4983 0.08770 0.08127 -0.0027 1.0000 0.2942
-6.750 -0.4895 0.08418 0.07780 -0.0002 1.0000 0.3177
-6.500 -0.5056 0.08263 0.07640 0.0001 1.0000 0.3373
-6.250 -0.4831 0.07837 0.07213 0.0049 1.0000 0.3709
-6.000 -0.4857 0.07608 0.06994 0.0080 1.0000 0.4047
-5.750 -0.4761 0.07301 0.06692 0.0124 1.0000 0.4458
-4.750 -0.3738 0.05795 0.05170 0.0260 1.0000 0.6787
-4.500 -0.4134 0.04343 0.03556 -0.0285 1.0000 0.1561
-4.250 -0.3884 0.03904 0.03038 -0.0284 1.0000 0.1291
-4.000 -0.3647 0.03569 0.02627 -0.0273 1.0000 0.1172
-3.750 -0.3424 0.03305 0.02305 -0.0259 1.0000 0.1181
-3.500 -0.3205 0.03025 0.02006 -0.0248 1.0000 0.1228
-3.250 -0.2954 0.02787 0.01724 -0.0235 1.0000 0.1237
-3.000 -0.2692 0.02575 0.01469 -0.0222 1.0000 0.1289
-2.750 -0.2433 0.02388 0.01249 -0.0210 1.0000 0.1467
-2.500 -0.2138 0.02167 0.01037 -0.0204 1.0000 0.1779
-2.250 -0.1253 0.01585 0.00783 -0.0288 1.0000 1.0000
-2.000 -0.1050 0.01581 0.00707 -0.0270 1.0000 1.0000
-1.750 -0.0857 0.01580 0.00660 -0.0254 1.0000 1.0000
-1.500 -0.0663 0.01581 0.00618 -0.0240 1.0000 1.0000
-1.250 -0.0468 0.01584 0.00591 -0.0226 1.0000 1.0000
-1.000 -0.0273 0.01591 0.00571 -0.0212 1.0000 1.0000
-0.750 -0.0078 0.01599 0.00558 -0.0199 1.0000 1.0000
-0.500 0.0116 0.01611 0.00551 -0.0187 1.0000 1.0000
-0.250 0.0309 0.01626 0.00550 -0.0174 1.0000 1.0000
0.000 0.0498 0.01644 0.00552 -0.0162 1.0000 1.0000
0.250 0.0684 0.01666 0.00563 -0.0150 1.0000 1.0000
0.500 0.0863 0.01694 0.00582 -0.0139 1.0000 1.0000
0.750 0.1036 0.01727 0.00609 -0.0127 1.0000 1.0000
1.000 0.1199 0.01767 0.00646 -0.0116 1.0000 1.0000
1.250 0.1350 0.01816 0.00693 -0.0105 1.0000 1.0000
1.500 0.1491 0.01876 0.00751 -0.0096 1.0000 1.0000
1.750 0.1621 0.01946 0.00821 -0.0087 1.0000 1.0000
2.000 0.1746 0.02026 0.00901 -0.0080 1.0000 1.0000
2.250 0.1871 0.02113 0.00989 -0.0075 1.0000 1.0000
2.500 0.2431 0.02243 0.01130 -0.0153 0.9813 1.0000
2.750 0.3082 0.02344 0.01253 -0.0240 0.9566 1.0000
3.000 0.3674 0.02409 0.01340 -0.0310 0.9319 1.0000
3.250 0.4229 0.02451 0.01410 -0.0368 0.9067 1.0000
3.500 0.4783 0.02474 0.01473 -0.0420 0.8811 1.0000
3.750 0.5371 0.02466 0.01506 -0.0472 0.8551 1.0000
4.000 0.5972 0.02433 0.01523 -0.0519 0.8274 1.0000
4.250 0.6538 0.02382 0.01531 -0.0551 0.7975 1.0000
4.500 0.7011 0.02315 0.01512 -0.0555 0.7587 1.0000
4.750 0.7359 0.02119 0.01324 -0.0494 0.6799 1.0000
5.000 0.7450 0.02044 0.01219 -0.0412 0.5853 1.0000
5.250 0.7447 0.02057 0.01192 -0.0331 0.4547 1.0000
5.500 0.7337 0.02423 0.01313 -0.0258 0.1577 1.0000
5.750 0.7486 0.02684 0.01535 -0.0235 0.1146 1.0000
6.000 0.7719 0.02915 0.01764 -0.0218 0.1000 1.0000
6.250 0.8031 0.03188 0.02031 -0.0213 0.0917 1.0000
6.500 0.8350 0.03485 0.02360 -0.0207 0.0882 1.0000
6.750 0.8595 0.03762 0.02681 -0.0193 0.0835 1.0000
7.000 0.8822 0.04097 0.03028 -0.0184 0.0789 1.0000
7.250 0.9013 0.04430 0.03412 -0.0165 0.0792 1.0000
7.500 0.9124 0.04787 0.03855 -0.0136 0.0820 1.0000
7.750 0.9210 0.05215 0.04343 -0.0111 0.0853 1.0000
8.000 0.9287 0.05665 0.04830 -0.0091 0.0884 1.0000
8.250 0.9321 0.06099 0.05315 -0.0069 0.0931 1.0000
8.500 0.9190 0.06617 0.05890 -0.0048 0.0986 1.0000
8.750 0.9235 0.07130 0.06422 -0.0039 0.1053 1.0000
9.000 0.8910 0.07643 0.06972 -0.0031 0.1093 1.0000
9.250 0.8874 0.08225 0.07564 -0.0032 0.1182 1.0000
9.500 0.8523 0.08708 0.08052 -0.0042 0.1196 1.0000
9.750 0.8190 0.09461 0.08800 -0.0105 0.1224 1.0000
10.000 0.8000 0.10354 0.09684 -0.0177 0.1348 1.0000
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