EPPLER 502 AIRFOIL (e502-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: EPPLER 502 AIRFOIL (e502-il) Reynolds number: 100,000 Max Cl/Cd: 45.6 at α=8° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-e502-il-100000.txt Download as CSV file: xf-e502-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 502 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
------ -------- --------- --------- -------- -------- --------
-11.000 -0.4776 0.09898 0.09473 -0.0594 1.0000 0.1243
-10.750 -0.4359 0.09721 0.09295 -0.0529 1.0000 0.1289
-10.500 -0.4425 0.09319 0.08902 -0.0544 1.0000 0.1345
-10.250 -0.4899 0.08805 0.08406 -0.0589 1.0000 0.1372
-10.000 -0.5204 0.08857 0.08473 -0.0521 1.0000 0.1364
-9.750 -0.5547 0.08557 0.08174 -0.0532 0.9955 0.1369
-8.500 -0.6439 0.05378 0.04758 -0.0532 0.9502 0.0556
-8.250 -0.6345 0.04913 0.04270 -0.0522 0.9450 0.0527
-8.000 -0.6195 0.04476 0.03781 -0.0514 0.9412 0.0500
-7.750 -0.6190 0.04218 0.03478 -0.0472 0.9346 0.0488
-7.500 -0.6028 0.04005 0.03230 -0.0460 0.9303 0.0512
-7.250 -0.5767 0.03759 0.02927 -0.0459 0.9274 0.0547
-7.000 -0.5651 0.03603 0.02727 -0.0428 0.9221 0.0562
-6.750 -0.5362 0.03324 0.02441 -0.0435 0.9188 0.0618
-6.500 -0.5045 0.03185 0.02265 -0.0439 0.9158 0.0697
-6.250 -0.4540 0.02924 0.02009 -0.0482 0.9149 0.0847
-6.000 -0.4096 0.02745 0.01847 -0.0515 0.9135 0.1018
-5.750 -0.3873 0.02653 0.01766 -0.0509 0.9102 0.1186
-5.500 -0.3935 0.02639 0.01758 -0.0451 0.9038 0.1290
-5.250 -0.3821 0.02545 0.01689 -0.0426 0.9002 0.1607
-5.000 -0.3921 0.02443 0.01659 -0.0366 0.8962 0.2429
-4.750 -0.4216 0.02431 0.01684 -0.0269 0.8897 0.2995
-4.500 -0.1273 0.03321 0.02671 -0.0529 0.9033 0.8039
-4.250 -0.0368 0.03369 0.02686 -0.0619 0.9041 0.8326
-4.000 0.0108 0.03362 0.02657 -0.0654 0.9019 0.8518
-3.750 0.0524 0.03344 0.02621 -0.0683 0.8994 0.8692
-3.500 0.1042 0.03289 0.02548 -0.0732 0.8977 0.8840
-3.250 0.1258 0.03305 0.02556 -0.0730 0.8932 0.8977
-3.000 0.1648 0.03273 0.02513 -0.0760 0.8900 0.9121
-2.750 0.2057 0.03233 0.02462 -0.0793 0.8872 0.9259
-2.500 0.2558 0.03154 0.02373 -0.0845 0.8850 0.9387
-2.250 0.3020 0.03084 0.02295 -0.0892 0.8830 0.9510
-2.000 0.3241 0.03086 0.02294 -0.0898 0.8781 0.9604
-1.750 0.3457 0.03094 0.02298 -0.0901 0.8739 0.9665
-1.500 0.3892 0.03026 0.02226 -0.0945 0.8714 0.9754
-1.250 0.4186 0.03020 0.02216 -0.0961 0.8687 0.9813
-1.000 0.4121 0.03118 0.02317 -0.0916 0.8619 0.9804
-0.750 0.4426 0.03096 0.02294 -0.0936 0.8584 0.9847
-0.500 0.4792 0.03060 0.02257 -0.0964 0.8556 0.9890
-0.250 0.4584 0.03215 0.02416 -0.0893 0.8481 0.9861
0.000 0.4815 0.03227 0.02429 -0.0897 0.8439 0.9881
0.250 0.5140 0.03212 0.02415 -0.0915 0.8410 0.9902
0.500 0.4824 0.03397 0.02603 -0.0826 0.8319 0.9882
0.750 0.5144 0.03387 0.02595 -0.0843 0.8282 0.9899
1.000 0.4925 0.03536 0.02747 -0.0770 0.8197 0.9895
1.250 0.5173 0.03549 0.02765 -0.0774 0.8149 0.9906
1.500 0.5534 0.03537 0.02757 -0.0796 0.8112 0.9923
1.750 0.5254 0.03693 0.02915 -0.0714 0.8012 0.9917
2.000 0.5779 0.03637 0.02866 -0.0760 0.7982 0.9936
2.250 0.5434 0.03804 0.03035 -0.0668 0.7868 0.9935
2.500 0.5860 0.03778 0.03019 -0.0699 0.7819 0.9946
2.750 0.5853 0.03854 0.03099 -0.0661 0.7722 0.9958
3.000 0.6481 0.03760 0.03016 -0.0719 0.7690 0.9970
3.250 0.6317 0.03882 0.03142 -0.0657 0.7571 0.9974
3.500 0.6252 0.03972 0.03238 -0.0610 0.7457 0.9988
3.750 0.6931 0.03848 0.03131 -0.0671 0.7419 0.9997
4.000 0.6873 0.03933 0.03222 -0.0623 0.7297 1.0000
4.250 0.7448 0.03795 0.03101 -0.0660 0.7257 1.0000
4.500 0.7470 0.03839 0.03153 -0.0621 0.7138 1.0000
4.750 0.7484 0.03884 0.03207 -0.0581 0.7016 1.0000
5.000 0.7988 0.03729 0.03068 -0.0602 0.6961 1.0000
5.250 0.8187 0.03676 0.03028 -0.0583 0.6855 1.0000
5.500 0.8310 0.03647 0.03014 -0.0553 0.6734 1.0000
5.750 0.8867 0.03391 0.02777 -0.0572 0.6675 1.0000
6.000 0.9155 0.03247 0.02649 -0.0559 0.6568 1.0000
6.250 0.9360 0.03137 0.02555 -0.0534 0.6443 1.0000
6.500 0.9611 0.02988 0.02426 -0.0514 0.6317 1.0000
6.750 0.9894 0.02803 0.02258 -0.0496 0.6178 1.0000
7.000 1.0196 0.02590 0.02062 -0.0479 0.6010 1.0000
7.250 1.0293 0.02498 0.01984 -0.0439 0.5798 1.0000
7.500 1.0382 0.02407 0.01905 -0.0397 0.5533 1.0000
7.750 1.0416 0.02351 0.01853 -0.0347 0.5192 1.0000
8.000 1.0460 0.02294 0.01779 -0.0298 0.4687 1.0000
8.250 1.0380 0.02309 0.01759 -0.0233 0.4140 1.0000
8.500 1.0155 0.02357 0.01787 -0.0150 0.3822 1.0000
8.750 0.9873 0.02402 0.01816 -0.0061 0.3612 1.0000
9.000 0.9576 0.02443 0.01843 0.0026 0.3436 1.0000
9.250 0.9268 0.02474 0.01861 0.0112 0.3290 1.0000
9.500 0.9063 0.02554 0.01916 0.0176 0.3010 1.0000
9.750 0.8946 0.02703 0.02032 0.0218 0.2606 1.0000
10.000 0.8864 0.02882 0.02177 0.0251 0.2201 1.0000
10.250 0.8807 0.03077 0.02337 0.0279 0.1833 1.0000
10.500 0.8775 0.03277 0.02507 0.0303 0.1513 1.0000
10.750 0.8782 0.03470 0.02676 0.0324 0.1255 1.0000
11.000 0.8816 0.03662 0.02843 0.0342 0.1049 1.0000
11.250 0.8900 0.03838 0.03014 0.0357 0.0882 1.0000
11.500 0.9009 0.04019 0.03188 0.0370 0.0744 1.0000
11.750 0.9150 0.04206 0.03372 0.0381 0.0630 1.0000
12.000 0.9406 0.04411 0.03568 0.0387 0.0542 1.0000
12.250 0.9481 0.04600 0.03779 0.0399 0.0484 1.0000
12.500 0.9814 0.04925 0.04113 0.0398 0.0431 1.0000
12.750 0.9860 0.05182 0.04406 0.0413 0.0411 1.0000
13.000 0.9895 0.05467 0.04721 0.0425 0.0393 1.0000
13.250 0.9909 0.05723 0.04994 0.0435 0.0373 1.0000
13.500 0.9902 0.05996 0.05280 0.0443 0.0356 1.0000
13.750 0.9897 0.06375 0.05675 0.0449 0.0346 1.0000
14.000 0.9801 0.06853 0.06179 0.0457 0.0340 1.0000
14.250 0.9667 0.07247 0.06600 0.0462 0.0341 1.0000
14.500 0.9505 0.07582 0.06964 0.0464 0.0348 1.0000
14.750 0.9308 0.08069 0.07477 0.0460 0.0348 1.0000
15.000 0.9075 0.08604 0.08037 0.0449 0.0350 1.0000
15.250 0.8772 0.09264 0.08727 0.0425 0.0357 1.0000
15.500 0.8279 0.10302 0.09801 0.0370 0.0374 1.0000
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