XFOIL Version 6.94 Calculated polar for: EPPLER 604 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.300 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.4760 0.01584 0.00987 -0.1043 0.6978 0.7161 0.500 0.5321 0.01575 0.00972 -0.1051 0.6912 0.7210 1.000 0.5962 0.01562 0.00945 -0.1078 0.6856 0.7258 1.500 0.6463 0.01565 0.00951 -0.1075 0.6785 0.7301 2.000 0.6973 0.01564 0.00956 -0.1071 0.6711 0.7335 2.500 0.7583 0.01559 0.00946 -0.1087 0.6648 0.7373 3.000 0.8096 0.01572 0.00964 -0.1086 0.6571 0.7418 3.500 0.8637 0.01569 0.00963 -0.1091 0.6484 0.7471 4.000 0.9275 0.01557 0.00946 -0.1112 0.6406 0.7510 4.500 0.9678 0.01561 0.00968 -0.1087 0.6304 0.7550 5.000 1.0269 0.01544 0.00951 -0.1097 0.6209 0.7595 5.500 1.0714 0.01551 0.00970 -0.1082 0.6096 0.7649 6.000 1.1280 0.01535 0.00953 -0.1088 0.5984 0.7702 6.500 1.1625 0.01537 0.00974 -0.1051 0.5849 0.7751 7.000 1.2073 0.01532 0.00975 -0.1034 0.5713 0.7807 7.500 1.2394 0.01530 0.00979 -0.0993 0.5548 0.7871 8.000 1.2557 0.01548 0.01011 -0.0923 0.5363 0.7931 8.500 1.2748 0.01590 0.01062 -0.0861 0.5152 0.8003 9.000 1.2931 0.01659 0.01132 -0.0803 0.4908 0.8087 9.500 1.3027 0.01757 0.01234 -0.0735 0.4633 0.8162 10.000 1.3073 0.01908 0.01387 -0.0666 0.4312 0.8253 10.500 1.3066 0.02110 0.01584 -0.0598 0.3963 0.8346 11.000 1.3013 0.02373 0.01839 -0.0531 0.3591 0.8459 11.500 1.2959 0.02670 0.02134 -0.0472 0.3223 0.8585 12.000 1.2894 0.03017 0.02475 -0.0419 0.2875 0.8733 12.500 1.2844 0.03382 0.02838 -0.0374 0.2544 0.8927 13.000 1.2832 0.03750 0.03214 -0.0339 0.2244 0.9431 14.500 1.2964 0.05142 0.04585 -0.0309 0.1444 1.0000 15.000 1.2979 0.05675 0.05112 -0.0302 0.1243 1.0000 15.500 1.3024 0.06200 0.05640 -0.0300 0.1067 1.0000 16.000 1.3052 0.06769 0.06212 -0.0302 0.0915 1.0000 16.500 1.3053 0.07393 0.06839 -0.0308 0.0787 1.0000 17.000 1.3034 0.08069 0.07518 -0.0318 0.0681 1.0000 17.500 1.3026 0.08758 0.08216 -0.0332 0.0588 1.0000 18.000 1.3013 0.09476 0.08945 -0.0351 0.0510 1.0000 18.500 1.2985 0.10233 0.09713 -0.0374 0.0446 1.0000 19.000 1.2951 0.11013 0.10504 -0.0402 0.0393 1.0000 19.500 1.2910 0.11808 0.11309 -0.0434 0.0351 1.0000 20.000 1.2883 0.12593 0.12110 -0.0469 0.0313 1.0000 20.500 1.2859 0.13375 0.12909 -0.0507 0.0280 1.0000 21.000 1.2826 0.14166 0.13714 -0.0548 0.0251 1.0000 21.500 1.2799 0.14934 0.14488 -0.0591 0.0227 1.0000 22.000 1.2763 0.15744 0.15320 -0.0639 0.0204 1.0000 22.500 1.2733 0.16515 0.16104 -0.0687 0.0185 1.0000 23.000 1.2702 0.17294 0.16894 -0.0738 0.0167 1.0000 24.000 1.2612 0.18902 0.18531 -0.0850 0.0138 1.0000 25.000 1.2463 0.20649 0.20315 -0.0976 0.0116 1.0000 25.500 1.2421 0.21431 0.21107 -0.1035 0.0107 1.0000 26.000 1.2251 0.22546 0.22251 -0.1115 0.0101 1.0000 26.500 1.2057 0.23750 0.23479 -0.1200 0.0096 1.0000