EPPLER 541 AIRFOIL (e541-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: EPPLER 541 AIRFOIL (e541-il) Reynolds number: 100,000 Max Cl/Cd: 40.54 at α=8.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-e541-il-100000.txt Download as CSV file: xf-e541-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 541 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
------ -------- --------- --------- -------- -------- --------
-10.750 -0.5100 0.11493 0.11088 -0.0433 1.0000 0.1090
-10.500 -0.5515 0.11007 0.10610 -0.0463 1.0000 0.1099
-10.250 -0.5861 0.10561 0.10163 -0.0485 0.9987 0.1101
-10.000 -0.5377 0.10333 0.09941 -0.0443 0.9977 0.1158
-9.750 -0.5504 0.09742 0.09349 -0.0491 0.9943 0.1199
-9.500 -0.5816 0.09229 0.08831 -0.0520 0.9902 0.1216
-9.250 -0.6238 0.08896 0.08488 -0.0518 0.9856 0.1228
-9.000 -0.6607 0.08705 0.08284 -0.0490 0.9810 0.1233
-8.750 -0.7061 0.08627 0.08185 -0.0433 0.9755 0.1241
-6.750 -0.7392 0.05297 0.04601 -0.0164 0.9595 0.0531
-6.500 -0.7350 0.04924 0.04134 -0.0113 0.9590 0.0451
-6.250 -0.7191 0.04675 0.03849 -0.0096 0.9576 0.0451
-6.000 -0.6967 0.04495 0.03631 -0.0088 0.9559 0.0447
-5.750 -0.6777 0.04280 0.03384 -0.0073 0.9549 0.0448
-5.500 -0.6598 0.04078 0.03153 -0.0054 0.9543 0.0449
-5.250 -0.6397 0.03906 0.02956 -0.0039 0.9530 0.0449
-5.000 -0.6140 0.03705 0.02747 -0.0033 0.9513 0.0461
-4.750 -0.5912 0.03573 0.02616 -0.0023 0.9504 0.0487
-4.500 -0.5725 0.03497 0.02535 -0.0008 0.9498 0.0536
-4.250 -0.5539 0.03386 0.02430 0.0007 0.9483 0.0608
-4.000 0.0593 0.04190 0.03480 -0.0742 0.9492 1.0000
-3.750 0.0739 0.04161 0.03437 -0.0740 0.9460 1.0000
-3.500 0.0922 0.04135 0.03400 -0.0744 0.9434 1.0000
-3.250 0.1156 0.04108 0.03361 -0.0757 0.9410 1.0000
-3.000 0.1101 0.04130 0.03379 -0.0713 0.9368 1.0000
-2.750 0.1195 0.04132 0.03374 -0.0698 0.9333 1.0000
-2.500 0.1370 0.04126 0.03358 -0.0697 0.9302 1.0000
-2.250 0.1615 0.04118 0.03341 -0.0710 0.9277 1.0000
-2.000 0.1596 0.04149 0.03368 -0.0671 0.9237 1.0000
-1.750 0.1653 0.04169 0.03385 -0.0647 0.9198 1.0000
-1.500 0.1832 0.04176 0.03386 -0.0646 0.9163 1.0000
-1.250 0.2114 0.04180 0.03382 -0.0663 0.9134 1.0000
-1.000 0.2045 0.04223 0.03424 -0.0614 0.9084 1.0000
-0.750 0.2158 0.04245 0.03443 -0.0599 0.9040 1.0000
-0.500 0.2451 0.04251 0.03445 -0.0617 0.9000 1.0000
-0.250 0.2452 0.04292 0.03485 -0.0581 0.8946 1.0000
0.000 0.2581 0.04318 0.03508 -0.0568 0.8893 1.0000
0.250 0.2952 0.04324 0.03512 -0.0598 0.8851 1.0000
0.500 0.2859 0.04374 0.03564 -0.0545 0.8777 1.0000
0.750 0.3155 0.04386 0.03574 -0.0560 0.8724 1.0000
1.000 0.3197 0.04426 0.03615 -0.0531 0.8645 1.0000
1.250 0.3478 0.04438 0.03627 -0.0542 0.8585 1.0000
1.500 0.3542 0.04477 0.03668 -0.0516 0.8499 1.0000
1.750 0.3880 0.04480 0.03673 -0.0536 0.8437 1.0000
2.000 0.3922 0.04519 0.03716 -0.0505 0.8338 1.0000
2.250 0.4360 0.04509 0.03709 -0.0540 0.8285 1.0000
2.500 0.4363 0.04549 0.03752 -0.0502 0.8173 1.0000
2.750 0.4512 0.04575 0.03782 -0.0488 0.8077 1.0000
3.000 0.4884 0.04559 0.03774 -0.0509 0.8011 1.0000
3.250 0.4959 0.04588 0.03808 -0.0482 0.7896 1.0000
3.500 0.5143 0.04600 0.03826 -0.0471 0.7798 1.0000
3.750 0.5537 0.04561 0.03797 -0.0493 0.7729 1.0000
4.000 0.5625 0.04584 0.03826 -0.0466 0.7611 1.0000
4.250 0.5801 0.04584 0.03835 -0.0453 0.7504 1.0000
4.500 0.6275 0.04498 0.03761 -0.0482 0.7444 1.0000
4.750 0.6395 0.04496 0.03769 -0.0458 0.7321 1.0000
5.000 0.6553 0.04484 0.03767 -0.0439 0.7206 1.0000
5.250 0.7111 0.04334 0.03634 -0.0475 0.7160 1.0000
5.500 0.7245 0.04305 0.03616 -0.0450 0.7035 1.0000
6.000 0.7817 0.04106 0.03450 -0.0438 0.6832 1.0000
6.250 0.8275 0.03909 0.03273 -0.0451 0.6754 1.0000
6.500 0.9029 0.03547 0.02944 -0.0501 0.6717 1.0000
6.750 0.9410 0.03323 0.02742 -0.0499 0.6605 1.0000
7.000 0.9673 0.03140 0.02578 -0.0480 0.6469 1.0000
7.250 1.0062 0.02894 0.02353 -0.0476 0.6311 1.0000
7.500 0.9757 0.02966 0.02428 -0.0379 0.6113 1.0000
7.750 0.9985 0.02799 0.02274 -0.0354 0.5879 1.0000
8.000 1.0077 0.02718 0.02201 -0.0312 0.5557 1.0000
8.250 1.0394 0.02570 0.02034 -0.0299 0.4978 1.0000
8.500 1.0451 0.02578 0.02008 -0.0256 0.4413 1.0000
8.750 1.0343 0.02655 0.02056 -0.0193 0.3962 1.0000
9.000 1.0209 0.02750 0.02127 -0.0131 0.3580 1.0000
9.250 1.0060 0.02856 0.02211 -0.0069 0.3229 1.0000
9.500 0.9912 0.02969 0.02301 -0.0010 0.2906 1.0000
9.750 0.9768 0.03087 0.02400 0.0046 0.2596 1.0000
10.000 0.9634 0.03207 0.02501 0.0099 0.2316 1.0000
10.250 0.9505 0.03330 0.02605 0.0149 0.2060 1.0000
10.500 0.9388 0.03453 0.02714 0.0197 0.1814 1.0000
10.750 0.9280 0.03576 0.02821 0.0243 0.1603 1.0000
11.000 0.9194 0.03694 0.02927 0.0287 0.1417 1.0000
11.250 0.9124 0.03806 0.03027 0.0328 0.1263 1.0000
11.500 0.9071 0.03911 0.03124 0.0367 0.1130 1.0000
11.750 0.9029 0.04011 0.03221 0.0404 0.1014 1.0000
12.000 0.9045 0.04117 0.03323 0.0436 0.0907 1.0000
12.250 0.9163 0.04236 0.03439 0.0457 0.0809 1.0000
12.500 0.9240 0.04356 0.03557 0.0476 0.0731 1.0000
12.750 0.9399 0.04529 0.03748 0.0487 0.0654 1.0000
13.000 0.9607 0.04738 0.03953 0.0488 0.0578 1.0000
13.250 0.9662 0.04942 0.04188 0.0502 0.0532 1.0000
13.500 0.9908 0.05260 0.04501 0.0496 0.0470 1.0000
13.750 0.9817 0.05486 0.04768 0.0514 0.0449 1.0000
14.000 0.9806 0.05798 0.05115 0.0524 0.0427 1.0000
14.250 0.9785 0.06142 0.05491 0.0532 0.0412 1.0000
14.500 0.9755 0.06478 0.05849 0.0536 0.0397 1.0000
14.750 0.9681 0.06854 0.06250 0.0538 0.0389 1.0000
15.000 0.9658 0.07318 0.06718 0.0533 0.0366 1.0000
15.500 0.9312 0.08299 0.07752 0.0522 0.0364 1.0000
15.750 0.9077 0.08840 0.08317 0.0506 0.0362 1.0000
16.000 0.8899 0.09405 0.08906 0.0485 0.0364 1.0000
16.250 0.8676 0.10033 0.09556 0.0455 0.0365 1.0000
16.500 0.7818 0.11887 0.11474 0.0333 0.0441 1.0000
16.750 0.7570 0.12925 0.12518 0.0269 0.0456 1.0000
17.000 0.7375 0.13953 0.13547 0.0209 0.0466 1.0000
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