USA 40 AIRFOIL (usa40-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: USA 40 AIRFOIL (usa40-il) Reynolds number: 50,000 Max Cl/Cd: 23.34 at α=4° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-usa40-il-50000.txt Download as CSV file: xf-usa40-il-50000.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 = 0.050 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.500 -0.2618 0.10912 0.10207 -0.0260 1.0000 0.2561
-8.250 -0.2560 0.10670 0.09972 -0.0250 1.0000 0.2640
-8.000 -0.2988 0.10920 0.10247 -0.0240 1.0000 0.2698
-7.750 -0.2579 0.10230 0.09553 -0.0229 1.0000 0.2767
-7.500 -0.2722 0.10182 0.09520 -0.0210 1.0000 0.2860
-7.250 -0.2778 0.09963 0.09314 -0.0191 1.0000 0.2913
-7.000 -0.2720 0.09742 0.09099 -0.0168 1.0000 0.3014
-6.750 -0.3213 0.09964 0.09347 -0.0120 1.0000 0.3060
-6.500 -0.2876 0.09443 0.08825 -0.0111 1.0000 0.3163
-6.250 -0.3270 0.09567 0.08969 -0.0060 1.0000 0.3232
-6.000 -0.3257 0.09300 0.08712 -0.0035 1.0000 0.3296
-5.750 -0.3387 0.09223 0.08646 0.0002 1.0000 0.3388
-5.500 -0.3656 0.09168 0.08608 0.0029 1.0000 0.3457
-5.250 -0.3599 0.08951 0.08396 0.0060 1.0000 0.3551
-5.000 -0.3820 0.08862 0.08321 0.0079 1.0000 0.3649
-4.750 -0.3830 0.08697 0.08163 0.0107 1.0000 0.3760
-4.500 -0.3903 0.08514 0.07989 0.0127 1.0000 0.3863
-4.250 -0.4024 0.08386 0.07871 0.0138 1.0000 0.4019
-4.000 -0.3228 0.05494 0.04826 -0.0505 1.0000 0.1817
-3.750 -0.2939 0.05086 0.04384 -0.0553 1.0000 0.1749
-3.500 -0.2641 0.04736 0.03997 -0.0597 1.0000 0.1725
-3.250 -0.1926 0.04276 0.03459 -0.0713 0.9889 0.1701
-3.000 -0.1281 0.04001 0.03118 -0.0799 0.9767 0.1744
-2.750 -0.0692 0.03810 0.02883 -0.0867 0.9634 0.1836
-2.500 -0.0125 0.03674 0.02702 -0.0925 0.9494 0.1973
-2.250 0.0390 0.03558 0.02586 -0.0972 0.9349 0.2183
-2.000 0.0903 0.03456 0.02484 -0.1015 0.9203 0.2535
-1.750 0.1424 0.03345 0.02411 -0.1055 0.9062 0.3167
-1.500 0.1898 0.03217 0.02353 -0.1082 0.8914 0.4529
-1.250 0.2199 0.03177 0.02392 -0.1068 0.8751 0.5908
-1.000 0.2394 0.03188 0.02440 -0.1027 0.8586 0.7037
-0.750 0.2495 0.03170 0.02441 -0.0967 0.8426 0.8065
-0.500 0.2622 0.03066 0.02359 -0.0906 0.8274 0.9505
-0.250 0.3246 0.03057 0.02301 -0.0984 0.8096 1.0000
0.000 0.3741 0.03059 0.02268 -0.1024 0.7936 1.0000
0.250 0.4206 0.03052 0.02231 -0.1052 0.7791 1.0000
0.500 0.4681 0.03021 0.02174 -0.1076 0.7659 1.0000
0.750 0.5012 0.03045 0.02179 -0.1083 0.7494 1.0000
1.000 0.5322 0.03081 0.02197 -0.1087 0.7335 1.0000
1.250 0.5625 0.03119 0.02220 -0.1090 0.7181 1.0000
1.500 0.5937 0.03152 0.02238 -0.1092 0.7037 1.0000
1.750 0.6382 0.03096 0.02163 -0.1101 0.6941 1.0000
2.000 0.6618 0.03175 0.02232 -0.1097 0.6787 1.0000
2.250 0.6847 0.03264 0.02313 -0.1092 0.6640 1.0000
2.500 0.7108 0.03337 0.02377 -0.1090 0.6513 1.0000
2.750 0.7487 0.03317 0.02344 -0.1092 0.6416 1.0000
3.000 0.7653 0.03464 0.02489 -0.1084 0.6274 1.0000
3.250 0.7876 0.03576 0.02596 -0.1080 0.6158 1.0000
3.500 0.8214 0.03588 0.02598 -0.1079 0.6064 1.0000
3.750 0.8319 0.03802 0.02816 -0.1071 0.5935 1.0000
4.000 0.8743 0.03746 0.02748 -0.1073 0.5866 1.0000
4.250 0.8746 0.04055 0.03067 -0.1061 0.5735 1.0000
4.500 0.8974 0.04174 0.03185 -0.1056 0.5647 1.0000
4.750 0.9110 0.04373 0.03387 -0.1049 0.5546 1.0000
5.000 0.9153 0.04676 0.03696 -0.1041 0.5451 1.0000
5.250 0.9333 0.04843 0.03865 -0.1036 0.5366 1.0000
5.500 0.9124 0.05407 0.04437 -0.1026 0.5265 1.0000
5.750 0.9236 0.05656 0.04690 -0.1021 0.5190 1.0000
6.000 0.8779 0.06530 0.05567 -0.1024 0.5111 1.0000
6.250 0.8500 0.07208 0.06247 -0.1030 0.5049 1.0000
6.500 0.9084 0.06993 0.06036 -0.1019 0.4987 1.0000
6.750 0.8640 0.07887 0.06932 -0.1032 0.4961 1.0000
7.000 0.8417 0.08523 0.07571 -0.1042 0.4952 1.0000
7.250 0.8311 0.09047 0.08098 -0.1051 0.4951 1.0000
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