EPPLER 559 AIRFOIL (e559-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: EPPLER 559 AIRFOIL (e559-il) Reynolds number: 50,000 Max Cl/Cd: 8.77 at α=4.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-e559-il-50000.txt Download as CSV file: xf-e559-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 559 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.250 -0.2735 0.12349 0.11634 -0.0318 1.0000 0.2695
-10.000 -0.2702 0.12098 0.11387 -0.0311 1.0000 0.2797
-9.750 -0.2949 0.12081 0.11385 -0.0311 1.0000 0.2825
-9.500 -0.2722 0.11592 0.10892 -0.0297 1.0000 0.2969
-9.250 -0.2631 0.11225 0.10528 -0.0285 1.0000 0.3074
-9.000 -0.2834 0.11176 0.10493 -0.0275 1.0000 0.3149
-8.750 -0.2730 0.10831 0.10149 -0.0259 1.0000 0.3301
-8.500 -0.2662 0.10507 0.09828 -0.0245 1.0000 0.3409
-8.250 -0.2928 0.10514 0.09852 -0.0222 1.0000 0.3487
-8.000 -0.3794 0.09590 0.08963 -0.0280 1.0000 0.2505
-7.750 -0.3950 0.09314 0.08696 -0.0258 1.0000 0.2492
-7.500 -0.4156 0.09084 0.08477 -0.0231 1.0000 0.2486
-7.250 -0.4405 0.08868 0.08275 -0.0203 1.0000 0.2479
-7.000 -0.4655 0.08590 0.08008 -0.0187 1.0000 0.2472
-6.750 -0.4903 0.08235 0.07664 -0.0186 1.0000 0.2468
-6.500 -0.5154 0.07731 0.07167 -0.0208 1.0000 0.2470
-6.250 -0.5128 0.07715 0.07153 -0.0171 1.0000 0.2586
-6.000 -0.5313 0.07048 0.06482 -0.0227 1.0000 0.2621
-5.750 -0.5326 0.06706 0.06136 -0.0241 1.0000 0.2712
-5.500 -0.5299 0.06007 0.05410 -0.0320 1.0000 0.2842
-5.250 -0.5220 0.06113 0.05520 -0.0274 1.0000 0.2942
-5.000 -0.5130 0.05936 0.05335 -0.0276 1.0000 0.3059
-4.750 -0.4716 0.05778 0.05152 -0.0332 0.9891 0.3259
-4.500 -0.4283 0.05530 0.04875 -0.0404 0.9771 0.3459
-4.250 -0.3825 0.05227 0.04532 -0.0492 0.9651 0.3669
-4.000 -0.3437 0.05123 0.04408 -0.0531 0.9526 0.3840
-3.750 -0.3127 0.05137 0.04421 -0.0531 0.9401 0.3968
-3.500 -0.2748 0.05041 0.04307 -0.0565 0.9281 0.4126
-3.250 -0.2322 0.04934 0.04180 -0.0608 0.9168 0.4287
-3.000 -0.2006 0.04833 0.04062 -0.0635 0.9040 0.4427
-2.750 -0.1657 0.04735 0.03945 -0.0668 0.8917 0.4575
-2.500 -0.1259 0.04651 0.03839 -0.0708 0.8806 0.4730
-2.250 -0.0830 0.04575 0.03744 -0.0748 0.8694 0.4888
-2.000 -0.0613 0.04564 0.03729 -0.0743 0.8569 0.4996
-1.750 -0.0289 0.04525 0.03678 -0.0763 0.8458 0.5126
-1.500 0.0157 0.04468 0.03606 -0.0799 0.8357 0.5282
-1.250 0.0403 0.04454 0.03577 -0.0813 0.8234 0.5409
-1.000 0.0698 0.04449 0.03565 -0.0823 0.8128 0.5542
-0.750 0.1075 0.04422 0.03531 -0.0842 0.8029 0.5685
-0.500 0.1253 0.04453 0.03557 -0.0840 0.7913 0.5799
-0.250 0.1644 0.04435 0.03530 -0.0862 0.7824 0.5957
0.000 0.1873 0.04467 0.03554 -0.0869 0.7713 0.6095
0.250 0.2076 0.04515 0.03600 -0.0866 0.7612 0.6229
0.500 0.2438 0.04510 0.03591 -0.0881 0.7525 0.6395
0.750 0.2527 0.04605 0.03687 -0.0867 0.7419 0.6515
1.000 0.2971 0.04577 0.03654 -0.0888 0.7344 0.6720
1.250 0.2968 0.04727 0.03805 -0.0869 0.7237 0.6836
1.500 0.3446 0.04688 0.03762 -0.0890 0.7169 0.7077
1.750 0.3361 0.04881 0.03959 -0.0864 0.7063 0.7196
2.000 0.3814 0.04847 0.03925 -0.0880 0.7000 0.7470
2.250 0.3679 0.05081 0.04164 -0.0854 0.6901 0.7618
2.500 0.4041 0.05062 0.04152 -0.0855 0.6837 0.7942
2.750 0.3896 0.05291 0.04391 -0.0827 0.6755 0.8146
3.000 0.4049 0.05338 0.04454 -0.0809 0.6690 0.8586
3.250 0.4140 0.05456 0.04596 -0.0807 0.6618 1.0000
3.500 0.4567 0.05662 0.04779 -0.0880 0.6534 1.0000
3.750 0.4915 0.05858 0.04951 -0.0923 0.6463 1.0000
4.000 0.4955 0.06137 0.05218 -0.0932 0.6400 1.0000
4.250 0.5453 0.06216 0.05277 -0.0961 0.6328 1.0000
4.500 0.5328 0.06571 0.05627 -0.0954 0.6291 1.0000
4.750 0.5308 0.06875 0.05926 -0.0953 0.6259 1.0000
5.000 0.5355 0.07158 0.06204 -0.0956 0.6235 1.0000
5.250 0.5371 0.07478 0.06522 -0.0960 0.6248 1.0000
5.500 0.5418 0.07803 0.06845 -0.0966 0.6278 1.0000
5.750 0.5578 0.08122 0.07161 -0.0980 0.6312 1.0000
6.000 0.4690 0.08958 0.08018 -0.0988 0.7320 1.0000
6.250 0.4858 0.09153 0.08208 -0.0992 0.7191 1.0000
6.500 0.4915 0.09346 0.08401 -0.0987 0.7093 1.0000
6.750 0.5253 0.09690 0.08741 -0.1009 0.7008 1.0000
7.000 0.5245 0.09828 0.08878 -0.0998 0.6889 1.0000
7.250 0.5508 0.10172 0.09221 -0.1014 0.6815 1.0000
7.500 0.5612 0.10356 0.09405 -0.1012 0.6690 1.0000
7.750 0.5672 0.10581 0.09632 -0.1010 0.6592 1.0000
8.000 0.6025 0.10960 0.10011 -0.1030 0.6498 1.0000
8.250 0.5977 0.11092 0.10145 -0.1018 0.6378 1.0000
8.500 0.6235 0.11476 0.10531 -0.1033 0.6312 1.0000
8.750 0.6317 0.11660 0.10719 -0.1031 0.6181 1.0000
9.000 0.6345 0.11886 0.10948 -0.1028 0.6077 1.0000
9.250 0.6682 0.12298 0.11364 -0.1046 0.5994 1.0000
9.500 0.6616 0.12436 0.11506 -0.1037 0.5874 1.0000
9.750 0.6812 0.12803 0.11876 -0.1047 0.5798 1.0000
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Polar data table (+)
Polar graphs
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