EPPLER E852 AIRFOIL (e852-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: EPPLER E852 AIRFOIL (e852-il) Reynolds number: 500,000 Max Cl/Cd: 119.56 at α=3.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-e852-il-500000.txt Download as CSV file: xf-e852-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER E852 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 -0.4430 0.08559 0.08350 -0.0427 1.0000 0.0110
-8.750 -0.4508 0.08192 0.07989 -0.0430 1.0000 0.0111
-8.500 -0.4611 0.07902 0.07703 -0.0423 1.0000 0.0114
-8.250 -0.4668 0.07480 0.07286 -0.0446 0.9989 0.0117
-8.000 -0.4612 0.06086 0.05886 -0.0654 0.9915 0.0111
-7.750 -0.4455 0.05206 0.04982 -0.0787 0.9840 0.0112
-7.500 -0.4287 0.04647 0.04400 -0.0852 0.9784 0.0116
-7.250 -0.4094 0.04177 0.03906 -0.0898 0.9722 0.0124
-7.000 -0.3836 0.03662 0.03357 -0.0945 0.9692 0.0131
-6.750 -0.3639 0.03255 0.02912 -0.0959 0.9614 0.0140
-6.500 -0.3330 0.02569 0.02146 -0.0972 0.9581 0.0098
-6.250 -0.2992 0.02362 0.01913 -0.0993 0.9564 0.0096
-6.000 -0.2792 0.01864 0.01361 -0.0997 0.9492 0.0111
-5.750 -0.2460 0.01661 0.01132 -0.1011 0.9466 0.0115
-5.500 -0.2112 0.01487 0.00934 -0.1025 0.9444 0.0105
-5.250 -0.1809 0.01356 0.00790 -0.1031 0.9394 0.0102
-5.000 -0.1499 0.01214 0.00631 -0.1040 0.9344 0.0103
-4.750 -0.1156 0.01078 0.00476 -0.1057 0.9309 0.0114
-4.500 -0.0861 0.00972 0.00359 -0.1060 0.9237 0.0294
-4.250 -0.0536 0.00906 0.00312 -0.1075 0.9187 0.0804
-4.000 -0.0257 0.00818 0.00267 -0.1084 0.9118 0.2034
-3.750 0.0027 0.00726 0.00237 -0.1094 0.9056 0.3927
-3.500 0.0312 0.00709 0.00229 -0.1097 0.8990 0.4516
-3.250 0.0606 0.00702 0.00222 -0.1100 0.8930 0.4871
-3.000 0.0893 0.00700 0.00215 -0.1103 0.8870 0.5087
-2.750 0.1176 0.00698 0.00210 -0.1104 0.8808 0.5281
-2.500 0.1466 0.00698 0.00204 -0.1107 0.8757 0.5431
-2.250 0.1741 0.00697 0.00200 -0.1106 0.8697 0.5560
-2.000 0.2030 0.00698 0.00199 -0.1109 0.8649 0.5722
-1.750 0.2303 0.00700 0.00200 -0.1108 0.8594 0.5879
-1.500 0.2583 0.00700 0.00197 -0.1109 0.8543 0.5994
-1.250 0.2870 0.00701 0.00196 -0.1111 0.8500 0.6091
-1.000 0.3144 0.00702 0.00196 -0.1111 0.8447 0.6183
-0.750 0.3425 0.00701 0.00196 -0.1112 0.8402 0.6266
-0.500 0.3708 0.00703 0.00198 -0.1114 0.8360 0.6355
-0.250 0.3982 0.00704 0.00201 -0.1114 0.8310 0.6452
0.000 0.4263 0.00704 0.00204 -0.1115 0.8267 0.6547
0.250 0.4543 0.00706 0.00208 -0.1116 0.8223 0.6647
0.500 0.4814 0.00706 0.00213 -0.1115 0.8168 0.6757
0.750 0.5095 0.00708 0.00218 -0.1116 0.8120 0.6876
1.000 0.5363 0.00708 0.00225 -0.1114 0.8059 0.7003
1.250 0.5636 0.00708 0.00228 -0.1112 0.7993 0.7139
1.500 0.5899 0.00706 0.00234 -0.1109 0.7915 0.7286
1.750 0.6169 0.00705 0.00237 -0.1106 0.7837 0.7448
2.000 0.6423 0.00701 0.00244 -0.1100 0.7735 0.7631
2.250 0.6673 0.00697 0.00248 -0.1093 0.7617 0.7834
2.500 0.6919 0.00692 0.00252 -0.1084 0.7491 0.8070
2.750 0.7163 0.00689 0.00260 -0.1075 0.7382 0.8338
3.000 0.7393 0.00683 0.00268 -0.1063 0.7270 0.8667
3.250 0.7589 0.00674 0.00272 -0.1042 0.7116 0.9144
3.500 0.7890 0.00669 0.00276 -0.1045 0.6919 1.0000
3.750 0.8130 0.00680 0.00283 -0.1037 0.6634 1.0000
4.000 0.8323 0.00710 0.00291 -0.1018 0.5993 1.0000
4.250 0.8356 0.00841 0.00334 -0.0971 0.4254 1.0000
4.500 0.8391 0.01016 0.00409 -0.0931 0.2341 1.0000
4.750 0.8540 0.01120 0.00463 -0.0911 0.1376 1.0000
5.000 0.8707 0.01211 0.00515 -0.0895 0.0699 1.0000
5.250 0.8901 0.01281 0.00564 -0.0881 0.0334 1.0000
5.500 0.9063 0.01390 0.00658 -0.0859 0.0058 1.0000
5.750 0.9254 0.01473 0.00757 -0.0841 0.0048 1.0000
6.000 0.9427 0.01572 0.00878 -0.0821 0.0046 1.0000
6.250 0.9570 0.01708 0.01029 -0.0796 0.0044 1.0000
6.500 0.9721 0.01860 0.01195 -0.0772 0.0044 1.0000
6.750 0.9910 0.02000 0.01347 -0.0757 0.0044 1.0000
7.000 1.0122 0.02141 0.01500 -0.0745 0.0045 1.0000
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Polar data table (+)
Polar graphs
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