XFOIL Version 6.94 Calculated polar for: EPPLER 1098 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 0.4764 0.01748 0.01095 -0.1120 0.6789 0.6229 0.500 0.5296 0.01799 0.01146 -0.1123 0.6761 0.6267 1.000 0.5749 0.01842 0.01200 -0.1115 0.6724 0.6303 1.500 0.6226 0.01886 0.01249 -0.1110 0.6679 0.6342 2.000 0.6763 0.01912 0.01275 -0.1115 0.6639 0.6383 2.500 0.7341 0.01928 0.01288 -0.1125 0.6604 0.6424 3.000 0.7944 0.01930 0.01294 -0.1139 0.6572 0.6469 3.500 0.8300 0.02008 0.01389 -0.1113 0.6504 0.6512 4.000 0.8814 0.02014 0.01405 -0.1110 0.6440 0.6562 4.500 0.9509 0.01954 0.01342 -0.1136 0.6390 0.6617 5.000 1.0113 0.01938 0.01333 -0.1147 0.6332 0.6671 5.500 1.0505 0.01956 0.01373 -0.1123 0.6245 0.6732 6.000 1.1167 0.01899 0.01323 -0.1142 0.6188 0.6813 6.500 1.1780 0.01872 0.01306 -0.1155 0.6124 0.6890 7.000 1.2177 0.01851 0.01310 -0.1128 0.6019 0.6973 7.500 1.2863 0.01746 0.01210 -0.1148 0.5920 0.7078 8.000 1.3239 0.01690 0.01179 -0.1113 0.5783 0.7188 8.500 1.3481 0.01668 0.01179 -0.1056 0.5631 0.7326 9.000 1.3563 0.01664 0.01197 -0.0970 0.5446 0.7496 9.500 1.3588 0.01707 0.01259 -0.0883 0.5182 0.7723 10.000 1.3523 0.01823 0.01380 -0.0791 0.4770 0.8047 10.500 1.3304 0.02050 0.01610 -0.0687 0.4285 0.8989 11.500 1.2942 0.02866 0.02372 -0.0559 0.3330 1.0000 12.000 1.2798 0.03329 0.02814 -0.0511 0.2881 1.0000 12.500 1.2735 0.03777 0.03243 -0.0476 0.2491 1.0000 13.000 1.2703 0.04239 0.03687 -0.0448 0.2105 1.0000 13.500 1.2714 0.04695 0.04131 -0.0428 0.1769 1.0000 14.000 1.2729 0.05175 0.04597 -0.0412 0.1430 1.0000 14.500 1.2763 0.05663 0.05074 -0.0401 0.1152 1.0000 15.000 1.2814 0.06158 0.05562 -0.0393 0.0915 1.0000 15.500 1.2844 0.06702 0.06100 -0.0390 0.0704 1.0000 16.000 1.2866 0.07281 0.06677 -0.0390 0.0505 1.0000 16.500 1.2834 0.07957 0.07348 -0.0395 0.0324 1.0000 17.000 1.2722 0.08777 0.08168 -0.0405 0.0202 1.0000 17.500 1.2610 0.09647 0.09053 -0.0424 0.0149 1.0000 18.000 1.2483 0.10580 0.10005 -0.0450 0.0123 1.0000 18.500 1.2378 0.11507 0.10954 -0.0482 0.0107 1.0000 19.000 1.2259 0.12468 0.11930 -0.0520 0.0099 1.0000 19.500 1.2219 0.13304 0.12789 -0.0556 0.0093 1.0000 20.000 1.2196 0.14105 0.13606 -0.0594 0.0088 1.0000 20.500 1.2200 0.14843 0.14355 -0.0632 0.0084 1.0000 21.000 1.2249 0.15457 0.14973 -0.0662 0.0080 1.0000 21.500 1.2300 0.16099 0.15636 -0.0698 0.0078 1.0000 22.000 1.2339 0.16755 0.16313 -0.0736 0.0076 1.0000 22.500 1.2345 0.17476 0.17057 -0.0780 0.0074 1.0000 23.000 1.2315 0.18278 0.17885 -0.0831 0.0074 1.0000 23.500 1.2225 0.19223 0.18858 -0.0894 0.0074 1.0000 24.000 1.2087 0.20298 0.19963 -0.0967 0.0074 1.0000 24.500 1.1883 0.21566 0.21261 -0.1055 0.0076 1.0000 25.000 1.1610 0.23083 0.22806 -0.1156 0.0078 1.0000