EPPLER 593 AIRFOIL (e593-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 593 AIRFOIL (e593-il) Reynolds number: 500,000 Max Cl/Cd: 113.45 at α=8.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e593-il-500000-n5.txt Download as CSV file: xf-e593-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 593 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.750 -0.3850 0.09432 0.09034 0.0035 0.4744 0.0069
-8.500 -0.3843 0.09010 0.08615 0.0015 0.4731 0.0067
-8.000 -0.5593 0.02863 0.02353 -0.0376 0.4800 0.0052
-7.750 -0.5403 0.02549 0.01997 -0.0379 0.4784 0.0052
-7.500 -0.5186 0.02328 0.01742 -0.0379 0.4765 0.0053
-7.250 -0.4954 0.02157 0.01540 -0.0377 0.4745 0.0053
-7.000 -0.4710 0.02019 0.01376 -0.0374 0.4723 0.0054
-6.750 -0.4459 0.01907 0.01240 -0.0371 0.4701 0.0054
-6.500 -0.4202 0.01810 0.01123 -0.0368 0.4682 0.0055
-6.250 -0.3940 0.01727 0.01020 -0.0365 0.4664 0.0055
-6.000 -0.3675 0.01655 0.00932 -0.0363 0.4648 0.0056
-5.750 -0.3409 0.01575 0.00839 -0.0360 0.4632 0.0058
-5.500 -0.3140 0.01506 0.00759 -0.0357 0.4616 0.0060
-5.250 -0.2866 0.01454 0.00696 -0.0355 0.4599 0.0063
-5.000 -0.2591 0.01408 0.00640 -0.0352 0.4581 0.0066
-4.750 -0.2314 0.01366 0.00588 -0.0350 0.4563 0.0071
-4.500 -0.2036 0.01329 0.00541 -0.0348 0.4544 0.0076
-4.250 -0.1759 0.01289 0.00493 -0.0345 0.4527 0.0085
-4.000 -0.1481 0.01259 0.00454 -0.0343 0.4511 0.0098
-3.750 -0.1202 0.01228 0.00419 -0.0341 0.4495 0.0126
-3.500 -0.0922 0.01199 0.00389 -0.0339 0.4483 0.0170
-3.250 -0.0642 0.01174 0.00365 -0.0337 0.4469 0.0232
-3.000 -0.0362 0.01151 0.00346 -0.0336 0.4454 0.0316
-2.750 -0.0080 0.01133 0.00329 -0.0335 0.4439 0.0414
-2.500 0.0201 0.01116 0.00317 -0.0334 0.4422 0.0521
-2.250 0.0484 0.01102 0.00303 -0.0332 0.4404 0.0603
-2.000 0.0767 0.01091 0.00290 -0.0331 0.4388 0.0680
-1.750 0.1049 0.01080 0.00278 -0.0330 0.4374 0.0763
-1.500 0.1333 0.01072 0.00268 -0.0329 0.4360 0.0847
-1.250 0.1615 0.01067 0.00261 -0.0328 0.4346 0.0942
-1.000 0.1899 0.01059 0.00255 -0.0327 0.4334 0.1050
-0.750 0.2183 0.01051 0.00250 -0.0326 0.4322 0.1155
-0.500 0.2468 0.01045 0.00245 -0.0325 0.4307 0.1266
-0.250 0.2752 0.01038 0.00241 -0.0325 0.4291 0.1403
0.000 0.3036 0.01030 0.00238 -0.0324 0.4274 0.1572
0.250 0.3319 0.01023 0.00236 -0.0324 0.4259 0.1781
0.500 0.3602 0.01015 0.00235 -0.0323 0.4245 0.2073
0.750 0.3883 0.01006 0.00235 -0.0323 0.4232 0.2451
1.000 0.4163 0.00994 0.00237 -0.0322 0.4219 0.3011
1.250 0.4433 0.00963 0.00239 -0.0322 0.4206 0.4226
1.750 0.5190 0.00829 0.00250 -0.0360 0.4175 1.0000
2.000 0.5464 0.00837 0.00255 -0.0358 0.4160 1.0000
2.250 0.5738 0.00845 0.00260 -0.0355 0.4145 1.0000
2.500 0.6012 0.00853 0.00266 -0.0353 0.4131 1.0000
2.750 0.6287 0.00861 0.00272 -0.0351 0.4115 1.0000
3.000 0.6562 0.00869 0.00278 -0.0349 0.4098 1.0000
3.250 0.6838 0.00877 0.00284 -0.0347 0.4082 1.0000
3.500 0.7113 0.00887 0.00291 -0.0345 0.4068 1.0000
3.750 0.7388 0.00899 0.00299 -0.0344 0.4052 1.0000
4.000 0.7663 0.00910 0.00310 -0.0342 0.4036 1.0000
4.250 0.7940 0.00918 0.00321 -0.0341 0.4018 1.0000
4.500 0.8217 0.00927 0.00332 -0.0341 0.3996 1.0000
4.750 0.8494 0.00935 0.00342 -0.0340 0.3971 1.0000
5.000 0.8771 0.00942 0.00351 -0.0339 0.3944 1.0000
5.250 0.9047 0.00951 0.00358 -0.0338 0.3918 1.0000
5.500 0.9321 0.00962 0.00368 -0.0337 0.3893 1.0000
5.750 0.9599 0.00971 0.00384 -0.0336 0.3864 1.0000
6.000 0.9875 0.00979 0.00397 -0.0336 0.3831 1.0000
6.250 1.0151 0.00988 0.00408 -0.0336 0.3796 1.0000
6.500 1.0424 0.00998 0.00418 -0.0335 0.3761 1.0000
6.750 1.0700 0.01009 0.00435 -0.0335 0.3721 1.0000
7.000 1.0974 0.01019 0.00450 -0.0334 0.3678 1.0000
7.250 1.1247 0.01031 0.00464 -0.0334 0.3634 1.0000
7.500 1.1519 0.01044 0.00482 -0.0333 0.3587 1.0000
8.000 1.2060 0.01072 0.00518 -0.0333 0.3448 1.0000
8.250 1.2328 0.01089 0.00538 -0.0332 0.3358 1.0000
8.500 1.2593 0.01110 0.00562 -0.0332 0.3256 1.0000
8.750 1.2851 0.01138 0.00590 -0.0331 0.3117 1.0000
9.000 1.3101 0.01177 0.00625 -0.0329 0.2938 1.0000
9.250 1.3336 0.01234 0.00675 -0.0327 0.2704 1.0000
9.500 1.3552 0.01314 0.00744 -0.0324 0.2435 1.0000
9.750 1.3758 0.01403 0.00824 -0.0321 0.2189 1.0000
10.000 1.3959 0.01494 0.00908 -0.0317 0.1970 1.0000
10.250 1.4149 0.01589 0.00998 -0.0313 0.1774 1.0000
10.500 1.4328 0.01690 0.01094 -0.0307 0.1583 1.0000
10.750 1.4487 0.01801 0.01200 -0.0301 0.1414 1.0000
11.000 1.4631 0.01917 0.01313 -0.0294 0.1272 1.0000
11.250 1.4749 0.02044 0.01439 -0.0285 0.1146 1.0000
11.500 1.4828 0.02192 0.01588 -0.0276 0.1036 1.0000
11.750 1.4795 0.02400 0.01803 -0.0262 0.0959 1.0000
12.000 1.4667 0.02820 0.02231 -0.0275 0.0907 1.0000
12.250 1.4584 0.03188 0.02605 -0.0281 0.0854 1.0000
12.500 1.4497 0.03539 0.02961 -0.0283 0.0807 1.0000
12.750 1.4398 0.03895 0.03320 -0.0284 0.0765 1.0000
13.000 1.4326 0.04221 0.03650 -0.0285 0.0719 1.0000
13.250 1.4236 0.04567 0.03999 -0.0286 0.0676 1.0000
13.500 1.4179 0.04884 0.04320 -0.0288 0.0638 1.0000
13.750 1.4125 0.05201 0.04640 -0.0290 0.0596 1.0000
14.000 1.4065 0.05533 0.04975 -0.0294 0.0560 1.0000
14.250 1.4060 0.05821 0.05270 -0.0297 0.0532 1.0000
14.500 1.4024 0.06155 0.05606 -0.0303 0.0491 1.0000
14.750 1.4015 0.06463 0.05918 -0.0309 0.0459 1.0000
15.000 1.4006 0.06780 0.06240 -0.0315 0.0430 1.0000
15.250 1.3983 0.07119 0.06581 -0.0323 0.0397 1.0000
15.500 1.3981 0.07439 0.06906 -0.0330 0.0368 1.0000
15.750 1.3965 0.07780 0.07250 -0.0338 0.0339 1.0000
16.000 1.3954 0.08122 0.07598 -0.0347 0.0319 1.0000
16.250 1.3952 0.08456 0.07937 -0.0357 0.0293 1.0000
16.500 1.3925 0.08836 0.08321 -0.0368 0.0273 1.0000
16.750 1.3925 0.09177 0.08669 -0.0379 0.0255 1.0000
17.000 1.3908 0.09547 0.09046 -0.0391 0.0237 1.0000
17.250 1.3880 0.09943 0.09446 -0.0404 0.0218 1.0000
17.500 1.3867 0.10321 0.09830 -0.0418 0.0198 1.0000
17.750 1.3834 0.10735 0.10248 -0.0433 0.0182 1.0000
18.000 1.3815 0.11132 0.10652 -0.0448 0.0169 1.0000
18.250 1.3796 0.11534 0.11062 -0.0465 0.0158 1.0000
18.500 1.3754 0.11982 0.11514 -0.0484 0.0144 1.0000
18.750 1.3740 0.12388 0.11929 -0.0501 0.0138 1.0000
19.000 1.3710 0.12820 0.12368 -0.0521 0.0125 1.0000
19.250 1.3681 0.13258 0.12814 -0.0542 0.0119 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 593 AIRFOIL (e593-il)