EPPLER E850 AIRFOIL (e850-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
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Airfoil: EPPLER E850 AIRFOIL (e850-il) Reynolds number: 1,000,000 Max Cl/Cd: 81.66 at α=0.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e850-il-1000000-n5.txt Download as CSV file: xf-e850-il-1000000-n5.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER E850 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 1.000 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-6.500 -0.3747 0.01487 0.01190 -0.0855 0.9693 0.0051
-6.250 -0.3556 0.01305 0.00990 -0.0858 0.9663 0.0058
-6.000 -0.3357 0.01134 0.00799 -0.0856 0.9625 0.0067
-5.750 -0.3117 0.00953 0.00589 -0.0858 0.9601 0.0078
-5.500 -0.2856 0.00797 0.00409 -0.0859 0.9584 0.0089
-4.750 -0.2153 0.01248 0.00761 -0.0866 0.9563 0.0043
-4.500 -0.1917 0.01036 0.00532 -0.0857 0.9532 0.0029
-4.250 -0.1666 0.00922 0.00401 -0.0851 0.9499 0.0025
-4.000 -0.1384 0.00842 0.00304 -0.0852 0.9475 0.0022
-3.750 -0.1084 0.00794 0.00242 -0.0856 0.9458 0.0020
-3.500 -0.0778 0.00764 0.00203 -0.0863 0.9443 0.0019
-3.250 -0.0469 0.00744 0.00172 -0.0871 0.9430 0.0018
-3.000 -0.0200 0.00730 0.00153 -0.0869 0.9400 0.0019
-2.750 0.0080 0.00718 0.00137 -0.0870 0.9372 0.0021
-2.500 0.0369 0.00706 0.00124 -0.0874 0.9348 0.0028
-2.250 0.0666 0.00696 0.00117 -0.0879 0.9328 0.0037
-2.000 0.0920 0.00523 0.00082 -0.0889 0.9303 0.4199
-1.750 0.1184 0.00498 0.00083 -0.0888 0.9270 0.5036
-1.500 0.1462 0.00492 0.00081 -0.0889 0.9233 0.5216
-1.250 0.1755 0.00487 0.00078 -0.0893 0.9195 0.5355
-1.000 0.2042 0.00483 0.00075 -0.0895 0.9143 0.5453
-0.750 0.2320 0.00480 0.00072 -0.0895 0.9074 0.5524
-0.500 0.2602 0.00478 0.00070 -0.0896 0.9006 0.5595
-0.250 0.2882 0.00476 0.00069 -0.0897 0.8922 0.5664
0.000 0.3154 0.00475 0.00068 -0.0895 0.8823 0.5734
0.250 0.3426 0.00474 0.00068 -0.0894 0.8698 0.5809
0.500 0.3694 0.00476 0.00067 -0.0891 0.8525 0.5881
0.750 0.3944 0.00483 0.00067 -0.0884 0.8173 0.5960
1.000 0.4102 0.00533 0.00071 -0.0857 0.7001 0.6033
1.250 0.4223 0.00623 0.00095 -0.0825 0.5337 0.6110
1.500 0.4470 0.00644 0.00113 -0.0819 0.4941 0.6196
1.750 0.4659 0.00713 0.00137 -0.0804 0.3659 0.6282
2.000 0.4854 0.00786 0.00162 -0.0790 0.2331 0.6377
2.250 0.5086 0.00829 0.00183 -0.0783 0.1615 0.6479
2.500 0.5321 0.00870 0.00206 -0.0776 0.0991 0.6586
2.750 0.5560 0.00908 0.00230 -0.0770 0.0526 0.6702
3.000 0.5809 0.00936 0.00253 -0.0766 0.0286 0.6828
3.250 0.6056 0.00968 0.00285 -0.0761 0.0097 0.6960
3.500 0.6303 0.01004 0.00327 -0.0754 0.0034 0.7103
3.750 0.6543 0.01049 0.00387 -0.0745 0.0019 0.7263
4.000 0.6791 0.01080 0.00427 -0.0740 0.0018 0.7435
4.250 0.7042 0.01103 0.00459 -0.0736 0.0016 0.7629
4.500 0.7288 0.01131 0.00499 -0.0730 0.0014 0.7848
4.750 0.7526 0.01169 0.00551 -0.0722 0.0012 0.8094
5.000 0.7757 0.01208 0.00605 -0.0713 0.0009 0.8391
5.250 0.7968 0.01254 0.00671 -0.0699 0.0008 0.8807
5.500 0.8190 0.01299 0.00741 -0.0686 0.0007 0.9946
5.750 0.8418 0.01385 0.00838 -0.0676 0.0006 1.0000
6.000 0.8644 0.01496 0.00964 -0.0666 0.0006 1.0000
6.250 0.8867 0.01639 0.01134 -0.0654 0.0005 1.0000
6.500 0.9071 0.01893 0.01426 -0.0637 0.0005 1.0000
6.750 0.8963 0.03457 0.03133 -0.0528 0.0008 1.0000
7.000 0.8992 0.04123 0.03842 -0.0481 0.0009 1.0000
7.250 0.9017 0.04690 0.04440 -0.0444 0.0010 1.0000
7.500 0.9019 0.05227 0.05002 -0.0411 0.0011 1.0000
7.750 0.8999 0.05739 0.05535 -0.0383 0.0012 1.0000
8.000 0.8939 0.06248 0.06062 -0.0358 0.0013 1.0000
8.250 0.8871 0.06706 0.06533 -0.0338 0.0013 1.0000
8.500 0.8739 0.07139 0.06977 -0.0315 0.0014 1.0000
8.750 0.8558 0.07512 0.07357 -0.0291 0.0014 1.0000
9.000 0.8390 0.07932 0.07784 -0.0286 0.0014 1.0000
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Polar data table (+)
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