XFOIL Version 6.94 Calculated polar for: EPPLER E854 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.450 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.4856 0.00870 0.00254 -0.1166 0.6584 0.5935 0.500 0.5407 0.00880 0.00260 -0.1166 0.6455 0.6061 1.000 0.5954 0.00889 0.00272 -0.1164 0.6331 0.6207 1.500 0.6506 0.00902 0.00285 -0.1164 0.6217 0.6356 2.000 0.7050 0.00911 0.00303 -0.1162 0.6108 0.6535 2.500 0.7602 0.00931 0.00325 -0.1162 0.5999 0.6734 3.000 0.8133 0.00936 0.00349 -0.1157 0.5894 0.6978 3.500 0.8665 0.00953 0.00376 -0.1153 0.5769 0.7282 4.000 0.9163 0.00956 0.00398 -0.1140 0.5620 0.7661 4.500 0.9627 0.00956 0.00422 -0.1119 0.5441 0.8223 5.000 1.0065 0.00940 0.00442 -0.1090 0.5276 0.9767 5.500 1.0575 0.00966 0.00470 -0.1082 0.5094 1.0000 6.000 1.1053 0.00993 0.00502 -0.1067 0.4865 1.0000 6.500 1.1483 0.01032 0.00537 -0.1043 0.4539 1.0000 7.000 1.1862 0.01088 0.00586 -0.1009 0.4015 1.0000 7.500 1.2012 0.01231 0.00678 -0.0935 0.2959 1.0000 8.000 1.2044 0.01437 0.00826 -0.0845 0.1916 1.0000 8.500 1.2140 0.01626 0.00978 -0.0771 0.1233 1.0000 9.000 1.2254 0.01810 0.01140 -0.0704 0.0811 1.0000 9.500 1.2373 0.02000 0.01318 -0.0643 0.0542 1.0000 10.000 1.2498 0.02203 0.01518 -0.0588 0.0366 1.0000 10.500 1.2608 0.02437 0.01754 -0.0539 0.0254 1.0000 11.000 1.2683 0.02723 0.02043 -0.0493 0.0168 1.0000 11.500 1.2716 0.03074 0.02407 -0.0450 0.0109 1.0000 12.000 1.2737 0.03477 0.02826 -0.0416 0.0080 1.0000 12.500 1.2721 0.03952 0.03320 -0.0388 0.0067 1.0000 13.000 1.2718 0.04455 0.03841 -0.0370 0.0058 1.0000 13.500 1.2646 0.05073 0.04479 -0.0357 0.0054 1.0000 14.000 1.2639 0.05666 0.05096 -0.0352 0.0051 1.0000 14.500 1.2630 0.06301 0.05755 -0.0352 0.0049 1.0000 15.000 1.2616 0.06983 0.06462 -0.0359 0.0047 1.0000 15.500 1.2588 0.07726 0.07232 -0.0372 0.0046 1.0000 16.000 1.2535 0.08553 0.08088 -0.0393 0.0045 1.0000 16.500 1.2450 0.09476 0.09043 -0.0423 0.0044 1.0000 17.000 1.2315 0.10538 0.10140 -0.0466 0.0044 1.0000 17.500 1.2123 0.11761 0.11401 -0.0524 0.0045 1.0000 18.000 1.1862 0.13199 0.12879 -0.0603 0.0045 1.0000 18.500 1.1557 0.14824 0.14543 -0.0701 0.0047 1.0000 19.000 1.1227 0.16633 0.16385 -0.0817 0.0048 1.0000 19.500 1.0848 0.18744 0.18524 -0.0949 0.0051 1.0000 21.500 0.8818 0.28756 0.28605 -0.1472 0.0118 1.0000 22.000 0.8831 0.30131 0.29979 -0.1528 0.0110 1.0000 22.500 0.8906 0.31088 0.30937 -0.1565 0.0102 1.0000 23.000 0.9026 0.31691 0.31543 -0.1584 0.0094 1.0000 23.500 0.9045 0.33081 0.32932 -0.1638 0.0088 1.0000