XFOIL Version 6.94 Calculated polar for: EPPLER 431 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 0.5085 0.01335 0.00796 -0.1070 0.7298 0.8430 0.500 0.5664 0.01330 0.00777 -0.1076 0.7227 0.8499 1.000 0.6203 0.01338 0.00780 -0.1081 0.7141 0.8567 1.500 0.6719 0.01323 0.00762 -0.1075 0.7051 0.8615 2.000 0.7320 0.01323 0.00751 -0.1087 0.6973 0.8660 2.500 0.7796 0.01320 0.00755 -0.1077 0.6868 0.8723 3.000 0.8423 0.01309 0.00735 -0.1096 0.6776 0.8769 3.500 0.8871 0.01305 0.00740 -0.1077 0.6664 0.8819 4.000 0.9426 0.01291 0.00723 -0.1079 0.6555 0.8876 4.500 0.9916 0.01294 0.00736 -0.1072 0.6426 0.8942 5.000 1.0465 0.01282 0.00720 -0.1073 0.6300 0.8994 5.500 1.0844 0.01275 0.00728 -0.1040 0.6142 0.9071 6.000 1.1279 0.01277 0.00737 -0.1021 0.5970 0.9154 6.500 1.1635 0.01275 0.00743 -0.0983 0.5786 0.9253 7.000 1.1938 0.01281 0.00755 -0.0936 0.5575 0.9374 7.500 1.2147 0.01296 0.00777 -0.0872 0.5336 0.9574 8.000 1.2526 0.01342 0.00827 -0.0850 0.5029 1.0000 8.500 1.2842 0.01432 0.00912 -0.0822 0.4681 1.0000 9.000 1.3058 0.01562 0.01030 -0.0779 0.4293 1.0000 9.500 1.3203 0.01728 0.01186 -0.0728 0.3891 1.0000 10.000 1.3303 0.01932 0.01380 -0.0675 0.3502 1.0000 10.500 1.3363 0.02179 0.01615 -0.0622 0.3129 1.0000 11.000 1.3412 0.02459 0.01885 -0.0574 0.2763 1.0000 11.500 1.3478 0.02759 0.02179 -0.0533 0.2421 1.0000 12.000 1.3508 0.03109 0.02518 -0.0494 0.2098 1.0000 12.500 1.3575 0.03462 0.02868 -0.0464 0.1791 1.0000 13.000 1.3613 0.03864 0.03265 -0.0437 0.1524 1.0000 13.500 1.3662 0.04287 0.03686 -0.0415 0.1276 1.0000 14.000 1.3691 0.04756 0.04155 -0.0398 0.1060 1.0000 14.500 1.3692 0.05286 0.04683 -0.0385 0.0874 1.0000 15.000 1.3706 0.05836 0.05240 -0.0377 0.0717 1.0000 15.500 1.3662 0.06486 0.05892 -0.0375 0.0600 1.0000 16.000 1.3656 0.07130 0.06550 -0.0378 0.0499 1.0000 18.000 1.3452 0.10264 0.09744 -0.0453 0.0275 1.0000 20.000 1.3144 0.13918 0.13476 -0.0621 0.0155 1.0000 20.500 1.3049 0.14885 0.14459 -0.0676 0.0137 1.0000 21.000 1.2961 0.15843 0.15441 -0.0732 0.0120 1.0000 21.500 1.2871 0.16792 0.16402 -0.0791 0.0106 1.0000 22.000 1.2786 0.17756 0.17390 -0.0854 0.0093 1.0000 22.500 1.2714 0.18672 0.18316 -0.0915 0.0083 1.0000 23.000 1.2615 0.19690 0.19359 -0.0984 0.0074 1.0000 23.500 1.2571 0.20544 0.20218 -0.1045 0.0066 1.0000 24.000 1.2453 0.21625 0.21326 -0.1120 0.0061 1.0000 24.500 1.2323 0.22763 0.22489 -0.1199 0.0059 1.0000 25.000 1.2088 0.24261 0.24014 -0.1298 0.0058 1.0000 27.000 1.1460 0.31841 0.31580 -0.1631 0.0131 1.0000 27.500 1.1559 0.32467 0.32206 -0.1676 0.0120 1.0000 28.000 1.1669 0.32933 0.32673 -0.1715 0.0113 1.0000 28.500 1.1801 0.33301 0.33045 -0.1745 0.0109 1.0000 29.000 1.1865 0.34069 0.33809 -0.1800 0.0105 1.0000 29.500 1.1964 0.34613 0.34353 -0.1843 0.0096 1.0000 30.000 1.2062 0.35058 0.34799 -0.1884 0.0090 1.0000 30.500 1.2162 0.35432 0.35176 -0.1921 0.0087 1.0000 31.000 1.2257 0.35896 0.35640 -0.1958 0.0085 1.0000 31.500 1.2336 0.36525 0.36268 -0.2009 0.0080 1.0000 32.000 1.2424 0.36979 0.36723 -0.2051 0.0074 1.0000 32.500 1.2507 0.37346 0.37092 -0.2091 0.0068 1.0000 33.500 1.2666 0.38255 0.38003 -0.2175 0.0061 1.0000 34.000 1.2742 0.38673 0.38422 -0.2216 0.0056 1.0000 34.500 1.2813 0.39029 0.38779 -0.2257 0.0051 1.0000 35.000 1.2876 0.39320 0.39073 -0.2297 0.0048 1.0000 35.500 1.2940 0.39678 0.39432 -0.2336 0.0047 1.0000 36.000 1.3003 0.40131 0.39885 -0.2381 0.0046 1.0000 36.500 1.3073 0.40585 0.40340 -0.2423 0.0040 1.0000 37.000 1.3127 0.40897 0.40654 -0.2465 0.0037 1.0000 37.500 1.3176 0.41170 0.40929 -0.2506 0.0035 1.0000 38.000 1.3219 0.41403 0.41165 -0.2548 0.0033 1.0000 38.500 1.3254 0.41574 0.41339 -0.2587 0.0032 1.0000