EPPLER 625 AIRFOIL (e625-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: EPPLER 625 AIRFOIL (e625-il) Reynolds number: 100,000 Max Cl/Cd: 44.54 at α=7° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e625-il-100000-n5.txt Download as CSV file: xf-e625-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 625 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.500 -0.4574 0.08161 0.07675 -0.0321 1.0000 0.0346
-9.250 -0.4616 0.07802 0.07318 -0.0330 1.0000 0.0343
-9.000 -0.4708 0.07425 0.06942 -0.0335 1.0000 0.0339
-8.750 -0.4836 0.07060 0.06575 -0.0328 1.0000 0.0335
-8.500 -0.4979 0.06730 0.06242 -0.0308 1.0000 0.0331
-8.250 -0.5103 0.06378 0.05884 -0.0285 1.0000 0.0326
-8.000 -0.5212 0.06004 0.05498 -0.0260 1.0000 0.0321
-7.750 -0.5299 0.05643 0.05123 -0.0230 1.0000 0.0316
-7.500 -0.5392 0.05274 0.04736 -0.0193 1.0000 0.0311
-7.250 -0.5499 0.04927 0.04367 -0.0147 1.0000 0.0307
-7.000 -0.5381 0.04383 0.03767 -0.0145 0.9740 0.0302
-6.750 -0.5160 0.03997 0.03335 -0.0152 0.9524 0.0303
-6.500 -0.4894 0.03737 0.03041 -0.0161 0.9326 0.0313
-6.250 -0.4632 0.03474 0.02734 -0.0164 0.9137 0.0324
-6.000 -0.4373 0.03212 0.02422 -0.0162 0.8952 0.0330
-5.750 -0.4112 0.02983 0.02147 -0.0157 0.8773 0.0333
-5.500 -0.3845 0.02792 0.01914 -0.0151 0.8598 0.0337
-5.250 -0.3573 0.02626 0.01711 -0.0145 0.8431 0.0342
-5.000 -0.3296 0.02491 0.01544 -0.0140 0.8266 0.0348
-4.750 -0.3028 0.02372 0.01416 -0.0136 0.8105 0.0357
-4.500 -0.2772 0.02290 0.01326 -0.0129 0.7946 0.0375
-4.250 -0.2512 0.02213 0.01236 -0.0122 0.7791 0.0396
-4.000 -0.2250 0.02135 0.01141 -0.0115 0.7640 0.0417
-3.750 -0.2005 0.02055 0.01062 -0.0107 0.7495 0.0437
-3.500 -0.1766 0.01992 0.00991 -0.0097 0.7352 0.0462
-3.250 -0.1530 0.01936 0.00922 -0.0085 0.7215 0.0495
-3.000 -0.1300 0.01885 0.00869 -0.0073 0.7083 0.0560
-2.750 -0.1070 0.01833 0.00815 -0.0061 0.6950 0.0668
-2.500 -0.0840 0.01777 0.00768 -0.0049 0.6816 0.0905
-2.250 -0.0604 0.01725 0.00734 -0.0039 0.6686 0.1420
-2.000 0.0216 0.01590 0.00881 -0.0116 0.6529 0.8484
-1.750 0.0601 0.01678 0.00941 -0.0119 0.6388 0.8870
-1.500 0.1062 0.01745 0.00979 -0.0141 0.6241 0.9117
-1.250 0.1616 0.01792 0.00996 -0.0185 0.6085 0.9318
-1.000 0.2185 0.01824 0.01002 -0.0234 0.5926 0.9545
-0.750 0.2711 0.01817 0.00970 -0.0281 0.5774 0.9682
-0.500 0.3025 0.01804 0.00941 -0.0289 0.5639 0.9722
-0.250 0.3321 0.01792 0.00914 -0.0295 0.5507 0.9751
0.000 0.3603 0.01785 0.00893 -0.0297 0.5381 0.9782
0.250 0.3869 0.01784 0.00877 -0.0297 0.5264 0.9813
0.500 0.4149 0.01779 0.00857 -0.0299 0.5143 0.9840
0.750 0.4444 0.01770 0.00839 -0.0305 0.5018 0.9864
1.000 0.4727 0.01766 0.00825 -0.0308 0.4900 0.9888
1.250 0.5006 0.01766 0.00813 -0.0310 0.4789 0.9915
1.500 0.5280 0.01769 0.00807 -0.0312 0.4676 0.9938
1.750 0.5566 0.01767 0.00798 -0.0316 0.4562 0.9959
2.000 0.5852 0.01768 0.00789 -0.0321 0.4456 0.9983
2.250 0.6125 0.01773 0.00787 -0.0322 0.4351 1.0000
2.500 0.6343 0.01786 0.00796 -0.0313 0.4253 1.0000
3.000 0.6775 0.01815 0.00816 -0.0293 0.4068 1.0000
3.250 0.6990 0.01833 0.00826 -0.0283 0.3987 1.0000
3.500 0.7205 0.01850 0.00844 -0.0273 0.3895 1.0000
3.750 0.7418 0.01871 0.00859 -0.0263 0.3816 1.0000
4.000 0.7630 0.01891 0.00879 -0.0253 0.3733 1.0000
4.250 0.7841 0.01914 0.00898 -0.0242 0.3658 1.0000
4.500 0.8050 0.01937 0.00923 -0.0232 0.3577 1.0000
4.750 0.8257 0.01962 0.00944 -0.0221 0.3506 1.0000
5.000 0.8463 0.01988 0.00973 -0.0210 0.3430 1.0000
5.250 0.8666 0.02015 0.00997 -0.0198 0.3364 1.0000
5.500 0.8867 0.02044 0.01032 -0.0186 0.3289 1.0000
5.750 0.9066 0.02073 0.01057 -0.0174 0.3228 1.0000
6.000 0.9261 0.02106 0.01097 -0.0162 0.3158 1.0000
6.250 0.9454 0.02139 0.01132 -0.0149 0.3094 1.0000
6.500 0.9644 0.02174 0.01169 -0.0135 0.3035 1.0000
6.750 0.9829 0.02211 0.01213 -0.0121 0.2970 1.0000
7.000 1.0013 0.02248 0.01246 -0.0107 0.2915 1.0000
7.250 1.0188 0.02290 0.01301 -0.0092 0.2852 1.0000
7.500 1.0360 0.02330 0.01345 -0.0076 0.2794 1.0000
7.750 1.0532 0.02373 0.01387 -0.0060 0.2744 1.0000
8.000 1.0689 0.02421 0.01450 -0.0043 0.2682 1.0000
8.250 1.0848 0.02465 0.01496 -0.0025 0.2629 1.0000
8.500 1.1001 0.02515 0.01550 -0.0007 0.2579 1.0000
8.750 1.1139 0.02568 0.01616 0.0014 0.2523 1.0000
9.000 1.1280 0.02616 0.01667 0.0034 0.2475 1.0000
9.250 1.1411 0.02671 0.01729 0.0055 0.2427 1.0000
9.500 1.1522 0.02731 0.01802 0.0079 0.2373 1.0000
9.750 1.1640 0.02784 0.01858 0.0102 0.2328 1.0000
10.000 1.1749 0.02844 0.01924 0.0125 0.2286 1.0000
10.250 1.1821 0.02913 0.02008 0.0154 0.2233 1.0000
10.500 1.1905 0.02973 0.02074 0.0181 0.2190 1.0000
10.750 1.2003 0.03031 0.02130 0.0205 0.2153 1.0000
11.000 1.2028 0.03118 0.02236 0.0239 0.2109 1.0000
11.250 1.2051 0.03193 0.02321 0.0272 0.2066 1.0000
11.500 1.2078 0.03257 0.02387 0.0306 0.2032 1.0000
11.750 1.2092 0.03337 0.02471 0.0339 0.1997 1.0000
12.000 1.2051 0.03448 0.02599 0.0373 0.1957 1.0000
12.250 1.2058 0.03557 0.02717 0.0399 0.1920 1.0000
12.500 1.2102 0.03653 0.02814 0.0419 0.1885 1.0000
12.750 1.2107 0.03789 0.02960 0.0439 0.1850 1.0000
13.000 1.2055 0.03973 0.03163 0.0457 0.1811 1.0000
13.250 1.2050 0.04137 0.03336 0.0471 0.1775 1.0000
13.500 1.2102 0.04265 0.03466 0.0482 0.1741 1.0000
13.750 1.2060 0.04486 0.03700 0.0492 0.1708 1.0000
14.000 1.1956 0.04778 0.04011 0.0497 0.1672 1.0000
14.250 1.1911 0.05029 0.04273 0.0500 0.1638 1.0000
14.500 1.1958 0.05191 0.04436 0.0505 0.1606 1.0000
14.750 1.1866 0.05515 0.04775 0.0503 0.1575 1.0000
15.000 1.1636 0.06022 0.05303 0.0491 0.1544 1.0000
15.250 1.1462 0.06487 0.05782 0.0479 0.1510 1.0000
15.500 1.1487 0.06704 0.06002 0.0477 0.1479 1.0000
15.750 1.1447 0.07018 0.06323 0.0470 0.1450 1.0000
16.000 0.9233 0.10978 0.10313 0.0281 0.1358 1.0000
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Polar data table (+)
Polar graphs
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