EPPLER 67 AIRFOIL (e67-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: EPPLER 67 AIRFOIL (e67-il) Reynolds number: 50,000 Max Cl/Cd: 36.94 at α=7.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e67-il-50000-n5.txt Download as CSV file: xf-e67-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 67 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.500 -0.3932 0.10343 0.09650 -0.0462 1.0000 0.0493
-9.250 -0.4001 0.10008 0.09323 -0.0460 1.0000 0.0489
-9.000 -0.4090 0.09667 0.08993 -0.0457 1.0000 0.0487
-8.750 -0.4199 0.09329 0.08666 -0.0453 1.0000 0.0483
-8.500 -0.4334 0.08990 0.08339 -0.0448 1.0000 0.0480
-8.250 -0.4500 0.08656 0.08017 -0.0440 1.0000 0.0476
-8.000 -0.4703 0.08322 0.07696 -0.0431 1.0000 0.0471
-7.750 -0.4905 0.07865 0.07253 -0.0441 1.0000 0.0466
-7.500 -0.5095 0.07283 0.06673 -0.0469 1.0000 0.0459
-7.250 -0.5254 0.06735 0.06118 -0.0488 1.0000 0.0454
-7.000 -0.5359 0.06174 0.05538 -0.0509 1.0000 0.0451
-6.750 -0.5387 0.05632 0.04964 -0.0528 1.0000 0.0451
-6.500 -0.5337 0.05129 0.04418 -0.0546 1.0000 0.0456
-6.250 -0.5221 0.04692 0.03931 -0.0560 1.0000 0.0467
-6.000 -0.4999 0.04277 0.03444 -0.0584 0.9982 0.0495
-5.750 -0.4689 0.03928 0.03032 -0.0615 0.9943 0.0535
-5.500 -0.4387 0.03712 0.02787 -0.0634 0.9900 0.0575
-5.250 -0.4056 0.03478 0.02486 -0.0652 0.9860 0.0639
-5.000 -0.3745 0.03337 0.02339 -0.0669 0.9818 0.0728
-4.750 -0.3440 0.03197 0.02180 -0.0682 0.9770 0.0841
-4.500 -0.3110 0.03080 0.02042 -0.0698 0.9727 0.0999
-4.250 -0.2803 0.02980 0.01927 -0.0708 0.9677 0.1198
-4.000 -0.2491 0.02899 0.01842 -0.0722 0.9625 0.1456
-3.750 -0.2148 0.02830 0.01776 -0.0741 0.9582 0.1785
-3.500 -0.1870 0.02768 0.01717 -0.0748 0.9520 0.2144
-3.250 -0.1529 0.02721 0.01675 -0.0766 0.9470 0.2629
-3.000 -0.1230 0.02681 0.01649 -0.0776 0.9412 0.3146
-2.750 -0.0921 0.02652 0.01634 -0.0787 0.9353 0.3740
-2.500 -0.0593 0.02633 0.01626 -0.0798 0.9302 0.4399
-2.250 -0.0339 0.02619 0.01626 -0.0795 0.9229 0.5005
-2.000 -0.0021 0.02614 0.01634 -0.0801 0.9179 0.5659
-1.750 0.0192 0.02610 0.01640 -0.0787 0.9099 0.6223
-1.500 0.0474 0.02609 0.01645 -0.0782 0.9043 0.6825
-1.250 0.0657 0.02608 0.01651 -0.0759 0.8961 0.7368
-1.000 0.0908 0.02603 0.01650 -0.0746 0.8901 0.7983
-0.750 0.1090 0.02595 0.01646 -0.0722 0.8816 0.8653
-0.500 0.1569 0.02575 0.01614 -0.0761 0.8759 1.0000
-0.250 0.1849 0.02598 0.01612 -0.0772 0.8672 1.0000
0.000 0.2230 0.02618 0.01606 -0.0797 0.8615 1.0000
0.250 0.2488 0.02647 0.01617 -0.0801 0.8523 1.0000
0.750 0.3110 0.02697 0.01636 -0.0823 0.8370 1.0000
1.000 0.3499 0.02710 0.01635 -0.0845 0.8319 1.0000
1.250 0.3724 0.02747 0.01663 -0.0841 0.8215 1.0000
1.500 0.4028 0.02771 0.01679 -0.0849 0.8137 1.0000
1.750 0.4338 0.02792 0.01694 -0.0857 0.8058 1.0000
2.000 0.4596 0.02825 0.01722 -0.0857 0.7963 1.0000
2.250 0.4959 0.02829 0.01724 -0.0871 0.7900 1.0000
2.500 0.5190 0.02867 0.01760 -0.0866 0.7793 1.0000
2.750 0.5512 0.02876 0.01769 -0.0873 0.7716 1.0000
3.000 0.5806 0.02892 0.01786 -0.0876 0.7626 1.0000
3.250 0.6052 0.02922 0.01821 -0.0872 0.7520 1.0000
3.500 0.6449 0.02898 0.01801 -0.0887 0.7461 1.0000
3.750 0.6671 0.02933 0.01841 -0.0879 0.7342 1.0000
4.000 0.6922 0.02957 0.01875 -0.0874 0.7231 1.0000
4.250 0.7290 0.02934 0.01860 -0.0884 0.7153 1.0000
4.500 0.7572 0.02941 0.01876 -0.0881 0.7043 1.0000
4.750 0.7819 0.02959 0.01904 -0.0875 0.6919 1.0000
5.000 0.8092 0.02965 0.01924 -0.0870 0.6798 1.0000
5.250 0.8390 0.02957 0.01928 -0.0868 0.6678 1.0000
5.500 0.8710 0.02937 0.01921 -0.0868 0.6556 1.0000
5.750 0.9018 0.02918 0.01917 -0.0866 0.6420 1.0000
6.000 0.9287 0.02914 0.01930 -0.0858 0.6267 1.0000
6.250 0.9542 0.02912 0.01943 -0.0849 0.6100 1.0000
6.500 0.9810 0.02903 0.01948 -0.0840 0.5923 1.0000
6.750 1.0116 0.02877 0.01938 -0.0835 0.5735 1.0000
7.000 1.0309 0.02897 0.01971 -0.0816 0.5518 1.0000
7.250 1.0579 0.02886 0.01969 -0.0806 0.5290 1.0000
7.500 1.0741 0.02922 0.02015 -0.0783 0.5036 1.0000
7.750 1.0918 0.02956 0.02054 -0.0762 0.4766 1.0000
8.000 1.1079 0.03001 0.02105 -0.0740 0.4477 1.0000
8.250 1.1216 0.03061 0.02165 -0.0715 0.4176 1.0000
8.500 1.1329 0.03138 0.02238 -0.0689 0.3863 1.0000
8.750 1.1420 0.03234 0.02328 -0.0662 0.3551 1.0000
9.000 1.1483 0.03354 0.02441 -0.0635 0.3238 1.0000
9.250 1.1533 0.03491 0.02572 -0.0608 0.2936 1.0000
9.500 1.1571 0.03647 0.02721 -0.0582 0.2648 1.0000
9.750 1.1598 0.03822 0.02886 -0.0557 0.2379 1.0000
10.000 1.1620 0.04012 0.03074 -0.0535 0.2134 1.0000
10.250 1.1640 0.04217 0.03277 -0.0515 0.1899 1.0000
10.500 1.1651 0.04439 0.03491 -0.0496 0.1704 1.0000
10.750 1.1667 0.04673 0.03725 -0.0479 0.1506 1.0000
11.000 1.1680 0.04919 0.03970 -0.0464 0.1341 1.0000
11.250 1.1686 0.05179 0.04230 -0.0451 0.1193 1.0000
11.500 1.1700 0.05447 0.04501 -0.0440 0.1064 1.0000
11.750 1.1712 0.05725 0.04785 -0.0430 0.0951 1.0000
12.000 1.1730 0.06010 0.05074 -0.0421 0.0858 1.0000
12.250 1.1740 0.06302 0.05369 -0.0414 0.0780 1.0000
12.500 1.1780 0.06596 0.05685 -0.0406 0.0704 1.0000
12.750 1.1802 0.06891 0.05982 -0.0401 0.0652 1.0000
13.000 1.1843 0.07212 0.06336 -0.0396 0.0597 1.0000
13.250 1.1859 0.07535 0.06672 -0.0394 0.0557 1.0000
13.500 1.1928 0.07833 0.06973 -0.0389 0.0526 1.0000
13.750 1.1926 0.08250 0.07433 -0.0388 0.0502 1.0000
14.000 1.1877 0.08704 0.07918 -0.0394 0.0481 1.0000
14.250 1.1810 0.09176 0.08414 -0.0404 0.0464 1.0000
14.500 1.1739 0.09651 0.08906 -0.0418 0.0448 1.0000
14.750 1.1704 0.10074 0.09336 -0.0430 0.0432 1.0000
15.000 1.1615 0.10627 0.09904 -0.0451 0.0422 1.0000
15.250 1.1449 0.11345 0.10651 -0.0486 0.0420 1.0000
15.500 1.1269 0.12133 0.11463 -0.0529 0.0419 1.0000
15.750 1.1075 0.13011 0.12363 -0.0582 0.0420 1.0000
16.000 1.0879 0.13969 0.13338 -0.0642 0.0422 1.0000
16.250 1.0686 0.15005 0.14386 -0.0708 0.0424 1.0000
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