EPPLER E854 AIRFOIL (e854-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER E854 AIRFOIL (e854-il) Reynolds number: 100,000 Max Cl/Cd: 57.37 at α=8.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-e854-il-100000.txt Download as CSV file: xf-e854-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER E854 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 -0.3733 0.09816 0.09363 -0.0426 1.0000 0.1241
-8.750 -0.4009 0.09532 0.09095 -0.0448 1.0000 0.1305
-8.500 -0.4470 0.09265 0.08850 -0.0468 1.0000 0.1314
-8.250 -0.4036 0.08909 0.08487 -0.0416 1.0000 0.1357
-8.000 -0.4105 0.08678 0.08265 -0.0397 1.0000 0.1391
-7.750 -0.4374 0.08490 0.08092 -0.0380 1.0000 0.1427
-7.250 -0.5402 0.07932 0.07559 -0.0414 1.0000 0.1454
-6.750 -0.5644 0.05043 0.04476 -0.0558 0.9954 0.0503
-6.500 -0.5337 0.04480 0.03875 -0.0599 0.9891 0.0493
-6.250 -0.4965 0.04013 0.03336 -0.0645 0.9833 0.0496
-6.000 -0.4631 0.03595 0.02863 -0.0671 0.9761 0.0491
-5.750 -0.4207 0.03279 0.02480 -0.0704 0.9710 0.0500
-5.500 -0.3889 0.02976 0.02169 -0.0722 0.9635 0.0550
-5.250 -0.3458 0.02773 0.01933 -0.0750 0.9586 0.0640
-5.000 -0.3132 0.02598 0.01761 -0.0764 0.9499 0.0785
-4.750 -0.2742 0.02444 0.01608 -0.0787 0.9436 0.1030
-4.500 -0.2397 0.02275 0.01471 -0.0806 0.9358 0.1447
-4.250 -0.2038 0.02005 0.01353 -0.0841 0.9301 0.3558
-4.000 -0.1694 0.02059 0.01424 -0.0845 0.9208 0.4986
-3.750 -0.1389 0.02110 0.01463 -0.0843 0.9112 0.5405
-3.500 -0.0993 0.02143 0.01486 -0.0855 0.9053 0.5744
-3.250 -0.0723 0.02168 0.01503 -0.0848 0.8951 0.5976
-3.000 -0.0304 0.02170 0.01492 -0.0866 0.8906 0.6224
-2.750 -0.0058 0.02184 0.01500 -0.0855 0.8804 0.6403
-2.500 0.0348 0.02175 0.01485 -0.0867 0.8765 0.6616
-2.250 0.0590 0.02180 0.01483 -0.0858 0.8666 0.6776
-2.000 0.1032 0.02147 0.01436 -0.0884 0.8628 0.6923
-1.750 0.1295 0.02140 0.01421 -0.0882 0.8533 0.7028
-1.500 0.1731 0.02100 0.01369 -0.0911 0.8490 0.7140
-1.250 0.2008 0.02095 0.01354 -0.0914 0.8401 0.7243
-1.000 0.2400 0.02061 0.01313 -0.0932 0.8351 0.7335
-0.750 0.2671 0.02059 0.01305 -0.0934 0.8264 0.7432
-0.500 0.3055 0.02032 0.01269 -0.0954 0.8209 0.7538
-0.250 0.3296 0.02037 0.01273 -0.0948 0.8126 0.7627
0.000 0.3649 0.02016 0.01247 -0.0962 0.8067 0.7731
0.250 0.3901 0.02026 0.01255 -0.0960 0.7986 0.7842
0.500 0.4211 0.02016 0.01243 -0.0965 0.7925 0.7950
0.750 0.4438 0.02029 0.01258 -0.0957 0.7847 0.8064
1.000 0.4727 0.02024 0.01255 -0.0959 0.7783 0.8191
1.250 0.4955 0.02040 0.01273 -0.0951 0.7712 0.8329
1.500 0.5204 0.02045 0.01281 -0.0946 0.7644 0.8483
1.750 0.5451 0.02052 0.01291 -0.0940 0.7582 0.8659
2.000 0.5635 0.02072 0.01320 -0.0924 0.7508 0.8870
2.250 0.5976 0.02047 0.01299 -0.0931 0.7467 0.9123
2.500 0.6222 0.02092 0.01358 -0.0934 0.7376 0.9497
2.750 0.6825 0.02075 0.01345 -0.1002 0.7333 1.0000
3.000 0.7041 0.02151 0.01420 -0.1012 0.7246 1.0000
3.250 0.7436 0.02163 0.01428 -0.1039 0.7195 1.0000
3.500 0.7645 0.02243 0.01509 -0.1039 0.7113 1.0000
3.750 0.7991 0.02262 0.01530 -0.1054 0.7057 1.0000
4.000 0.8194 0.02337 0.01609 -0.1048 0.6978 1.0000
4.250 0.8514 0.02359 0.01634 -0.1057 0.6916 1.0000
4.500 0.8719 0.02428 0.01711 -0.1049 0.6835 1.0000
4.750 0.9049 0.02437 0.01724 -0.1056 0.6768 1.0000
5.000 0.9236 0.02505 0.01802 -0.1045 0.6678 1.0000
5.250 0.9625 0.02479 0.01780 -0.1058 0.6612 1.0000
5.500 0.9805 0.02536 0.01851 -0.1043 0.6506 1.0000
5.750 1.0108 0.02526 0.01848 -0.1041 0.6407 1.0000
6.000 1.0522 0.02447 0.01773 -0.1052 0.6302 1.0000
6.250 1.0817 0.02413 0.01749 -0.1047 0.6174 1.0000
6.500 1.1079 0.02393 0.01738 -0.1037 0.6038 1.0000
6.750 1.1342 0.02370 0.01725 -0.1028 0.5900 1.0000
7.000 1.1616 0.02336 0.01700 -0.1019 0.5753 1.0000
7.250 1.1902 0.02288 0.01659 -0.1010 0.5592 1.0000
7.500 1.2081 0.02272 0.01662 -0.0987 0.5406 1.0000
7.750 1.2281 0.02232 0.01633 -0.0964 0.5192 1.0000
8.000 1.2430 0.02206 0.01618 -0.0934 0.4944 1.0000
8.250 1.2537 0.02187 0.01608 -0.0897 0.4634 1.0000
8.500 1.2586 0.02194 0.01616 -0.0852 0.4224 1.0000
8.750 1.2541 0.02245 0.01644 -0.0793 0.3616 1.0000
9.000 1.2392 0.02382 0.01729 -0.0724 0.2922 1.0000
9.250 1.2218 0.02591 0.01888 -0.0662 0.2372 1.0000
9.500 1.2074 0.02824 0.02084 -0.0611 0.1973 1.0000
9.750 1.1967 0.03064 0.02297 -0.0570 0.1675 1.0000
10.000 1.1903 0.03304 0.02515 -0.0537 0.1447 1.0000
10.250 1.1881 0.03536 0.02732 -0.0510 0.1253 1.0000
10.500 1.1884 0.03770 0.02951 -0.0488 0.1091 1.0000
10.750 1.1940 0.03995 0.03164 -0.0469 0.0944 1.0000
11.000 1.2031 0.04214 0.03375 -0.0453 0.0816 1.0000
11.250 1.2122 0.04424 0.03589 -0.0439 0.0715 1.0000
11.500 1.2238 0.04651 0.03828 -0.0425 0.0622 1.0000
11.750 1.2393 0.04899 0.04079 -0.0415 0.0542 1.0000
12.000 1.2523 0.05135 0.04319 -0.0405 0.0483 1.0000
12.250 1.2673 0.05453 0.04663 -0.0394 0.0433 1.0000
12.500 1.2710 0.05719 0.04951 -0.0381 0.0399 1.0000
12.750 1.2835 0.06035 0.05279 -0.0374 0.0375 1.0000
13.000 1.2965 0.06618 0.05892 -0.0369 0.0357 1.0000
13.250 1.2847 0.06969 0.06280 -0.0354 0.0353 1.0000
13.500 1.2710 0.07367 0.06712 -0.0344 0.0349 1.0000
13.750 1.2552 0.07801 0.07179 -0.0340 0.0345 1.0000
14.000 1.2384 0.08286 0.07694 -0.0341 0.0343 1.0000
14.250 1.2205 0.08810 0.08246 -0.0349 0.0343 1.0000
14.500 1.2009 0.09378 0.08840 -0.0363 0.0343 1.0000
14.750 1.1806 0.09987 0.09473 -0.0383 0.0344 1.0000
15.000 1.1598 0.10646 0.10154 -0.0410 0.0346 1.0000
15.250 1.1401 0.11344 0.10870 -0.0443 0.0350 1.0000
15.500 1.1193 0.12098 0.11641 -0.0483 0.0353 1.0000
15.750 1.1014 0.12881 0.12432 -0.0525 0.0357 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER E854 AIRFOIL (e854-il)