EPPLER 540 AIRFOIL (e540-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: EPPLER 540 AIRFOIL (e540-il) Reynolds number: 100,000 Max Cl/Cd: 38.33 at α=9° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-e540-il-100000.txt Download as CSV file: xf-e540-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 540 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
------ -------- --------- --------- -------- -------- --------
-11.250 -0.4759 0.10657 0.10216 -0.0639 1.0000 0.1214
-11.000 -0.5118 0.10347 0.09917 -0.0629 1.0000 0.1215
-10.750 -0.5432 0.10175 0.09754 -0.0594 1.0000 0.1214
-10.500 -0.5727 0.10013 0.09600 -0.0556 1.0000 0.1214
-10.250 -0.6047 0.09848 0.09440 -0.0521 1.0000 0.1213
-10.000 -0.6371 0.09682 0.09277 -0.0484 1.0000 0.1213
-9.750 -0.6694 0.09526 0.09122 -0.0444 1.0000 0.1212
-9.500 -0.6995 0.09274 0.08861 -0.0433 0.9953 0.1213
-7.750 -0.7103 0.05680 0.05045 -0.0386 0.9572 0.0649
-7.500 -0.7121 0.05303 0.04598 -0.0333 0.9514 0.0533
-7.250 -0.6980 0.04909 0.04176 -0.0323 0.9475 0.0519
-7.000 -0.6762 0.04553 0.03783 -0.0320 0.9446 0.0489
-6.750 -0.6589 0.04248 0.03425 -0.0300 0.9415 0.0461
-6.500 -0.6432 0.04033 0.03161 -0.0275 0.9377 0.0446
-6.250 -0.6145 0.03817 0.02916 -0.0273 0.9346 0.0448
-6.000 -0.5784 0.03630 0.02712 -0.0284 0.9320 0.0464
-5.750 -0.5396 0.03497 0.02564 -0.0301 0.9298 0.0510
-5.500 -0.5167 0.03387 0.02448 -0.0291 0.9269 0.0559
-5.250 -0.5033 0.03290 0.02363 -0.0267 0.9232 0.0620
-5.000 -0.4871 0.03188 0.02269 -0.0248 0.9195 0.0726
-4.750 -0.4712 0.03073 0.02170 -0.0231 0.9163 0.0972
-4.500 -0.4751 0.02922 0.02090 -0.0182 0.9130 0.1703
-3.750 0.1341 0.03966 0.03235 -0.0791 0.9332 0.9856
-3.500 0.1978 0.03776 0.03027 -0.0888 0.9323 1.0000
-3.250 0.2310 0.03705 0.02943 -0.0921 0.9296 1.0000
-3.000 0.2318 0.03735 0.02969 -0.0889 0.9222 1.0000
-2.750 0.2560 0.03698 0.02925 -0.0902 0.9181 1.0000
-2.500 0.2889 0.03643 0.02861 -0.0931 0.9155 1.0000
-2.250 0.2830 0.03707 0.02923 -0.0885 0.9078 1.0000
-2.000 0.3064 0.03686 0.02897 -0.0895 0.9038 1.0000
-1.750 0.3384 0.03648 0.02853 -0.0920 0.9012 1.0000
-1.500 0.3247 0.03741 0.02948 -0.0859 0.8929 1.0000
-1.250 0.3489 0.03730 0.02931 -0.0868 0.8890 1.0000
-1.000 0.3836 0.03694 0.02892 -0.0897 0.8866 1.0000
-0.750 0.3620 0.03814 0.03014 -0.0821 0.8775 1.0000
-0.500 0.3902 0.03799 0.02997 -0.0836 0.8739 1.0000
-0.250 0.3864 0.03875 0.03074 -0.0793 0.8670 1.0000
0.000 0.4013 0.03897 0.03097 -0.0783 0.8615 1.0000
0.250 0.4371 0.03870 0.03068 -0.0810 0.8584 1.0000
0.500 0.4174 0.03987 0.03188 -0.0739 0.8491 1.0000
0.750 0.4494 0.03970 0.03172 -0.0758 0.8451 1.0000
1.000 0.3990 0.04139 0.03342 -0.0624 0.8347 0.9924
1.250 0.4246 0.04131 0.03335 -0.0628 0.8303 0.9897
1.500 0.3629 0.04285 0.03488 -0.0476 0.8193 0.9786
1.750 0.4023 0.04258 0.03463 -0.0504 0.8154 0.9783
2.000 0.3652 0.04356 0.03562 -0.0400 0.8045 0.9736
2.250 0.3970 0.04342 0.03549 -0.0413 0.7996 0.9729
2.500 0.3642 0.04420 0.03628 -0.0315 0.7891 0.9698
2.750 0.3620 0.04457 0.03667 -0.0275 0.7797 0.9684
3.000 0.3947 0.04436 0.03649 -0.0288 0.7739 0.9686
3.250 0.3864 0.04479 0.03695 -0.0236 0.7630 0.9679
3.500 0.4318 0.04431 0.03654 -0.0267 0.7585 0.9679
3.750 0.4196 0.04474 0.03699 -0.0209 0.7466 0.9671
4.000 0.4197 0.04500 0.03730 -0.0170 0.7362 0.9672
4.250 0.4609 0.04445 0.03683 -0.0191 0.7308 0.9674
4.500 0.4524 0.04477 0.03718 -0.0139 0.7189 0.9668
4.750 0.5046 0.04390 0.03642 -0.0174 0.7149 0.9673
5.000 0.5002 0.04410 0.03666 -0.0126 0.7027 0.9680
5.250 0.4963 0.04428 0.03688 -0.0080 0.6906 0.9683
5.500 0.5442 0.04322 0.03597 -0.0105 0.6861 0.9687
5.750 0.5462 0.04312 0.03592 -0.0064 0.6744 0.9691
6.000 0.5485 0.04310 0.03597 -0.0027 0.6622 0.9701
6.250 0.6038 0.04168 0.03472 -0.0061 0.6570 0.9708
6.500 0.6342 0.04089 0.03409 -0.0062 0.6466 0.9719
6.750 0.6518 0.04038 0.03370 -0.0045 0.6347 0.9729
7.000 0.7287 0.03722 0.03080 -0.0095 0.6323 0.9732
7.250 0.8710 0.03177 0.02581 -0.0239 0.6309 0.9733
7.500 0.8973 0.03027 0.02447 -0.0224 0.6174 0.9740
7.750 0.9395 0.02824 0.02263 -0.0232 0.6014 0.9748
8.000 0.9226 0.02804 0.02249 -0.0154 0.5853 0.9762
8.250 0.9241 0.02742 0.02200 -0.0105 0.5655 0.9783
8.500 0.9389 0.02646 0.02113 -0.0076 0.5409 0.9805
8.750 0.9494 0.02578 0.02045 -0.0041 0.5098 0.9826
9.000 0.9689 0.02528 0.01982 -0.0023 0.4653 0.9843
9.250 0.9769 0.02551 0.01979 0.0006 0.4179 0.9865
9.500 0.9748 0.02631 0.02035 0.0044 0.3748 0.9894
9.750 0.9700 0.02740 0.02123 0.0081 0.3344 0.9929
10.000 0.9669 0.02880 0.02241 0.0109 0.2933 0.9971
10.250 0.9585 0.03015 0.02353 0.0146 0.2610 1.0000
10.500 0.9407 0.03099 0.02429 0.0203 0.2397 1.0000
10.750 0.9233 0.03181 0.02496 0.0259 0.2208 1.0000
11.000 0.9091 0.03273 0.02576 0.0307 0.2004 1.0000
11.250 0.9021 0.03415 0.02696 0.0338 0.1774 1.0000
11.500 0.9007 0.03588 0.02857 0.0356 0.1525 1.0000
11.750 0.9051 0.03772 0.03030 0.0368 0.1302 1.0000
12.000 0.9120 0.03957 0.03201 0.0376 0.1121 1.0000
12.250 0.9247 0.04140 0.03363 0.0380 0.0961 1.0000
12.500 0.9465 0.04312 0.03528 0.0381 0.0832 1.0000
12.750 0.9571 0.04501 0.03736 0.0383 0.0744 1.0000
13.000 0.9792 0.04710 0.03956 0.0382 0.0659 1.0000
13.250 1.0011 0.04941 0.04180 0.0378 0.0579 1.0000
13.500 1.0097 0.05194 0.04469 0.0382 0.0532 1.0000
13.750 1.0332 0.05516 0.04782 0.0374 0.0471 1.0000
14.000 1.0282 0.05795 0.05101 0.0381 0.0453 1.0000
14.250 1.0253 0.06144 0.05492 0.0386 0.0433 1.0000
14.500 1.0211 0.06529 0.05910 0.0388 0.0419 1.0000
14.750 1.0172 0.06892 0.06294 0.0386 0.0401 1.0000
15.000 1.0100 0.07281 0.06704 0.0382 0.0391 1.0000
15.250 1.0079 0.07688 0.07117 0.0375 0.0373 1.0000
15.500 0.9931 0.08234 0.07687 0.0366 0.0368 1.0000
15.750 0.9757 0.08743 0.08221 0.0352 0.0369 1.0000
16.000 0.9545 0.09342 0.08844 0.0331 0.0368 1.0000
16.250 0.9329 0.09967 0.09492 0.0303 0.0367 1.0000
16.500 0.9093 0.10635 0.10185 0.0264 0.0372 1.0000
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