XFOIL Version 6.94 Calculated polar for: EPPLER 67 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 0.00896 0.00401 -0.1258 0.8316 0.8059 0.500 0.5542 0.00886 0.00401 -0.1229 0.8147 0.8580 1.000 0.5966 0.00876 0.00396 -0.1192 0.7974 0.9111 1.500 0.6558 0.00863 0.00383 -0.1192 0.7801 0.9696 2.000 0.7259 0.00862 0.00374 -0.1224 0.7607 1.0000 2.500 0.7792 0.00874 0.00377 -0.1220 0.7389 1.0000 3.000 0.8322 0.00889 0.00384 -0.1214 0.7150 1.0000 3.500 0.8850 0.00911 0.00396 -0.1206 0.6893 1.0000 4.000 0.9331 0.00936 0.00419 -0.1190 0.6595 1.0000 4.500 0.9813 0.00970 0.00444 -0.1173 0.6273 1.0000 5.000 1.0254 0.01007 0.00477 -0.1148 0.5902 1.0000 5.500 1.0669 0.01057 0.00518 -0.1119 0.5481 1.0000 6.000 1.1041 0.01119 0.00567 -0.1082 0.4992 1.0000 6.500 1.1353 0.01199 0.00630 -0.1034 0.4425 1.0000 7.000 1.1588 0.01295 0.00704 -0.0973 0.3810 1.0000 7.500 1.1758 0.01425 0.00802 -0.0904 0.3068 1.0000 8.000 1.1911 0.01589 0.00927 -0.0837 0.2298 1.0000 8.500 1.2108 0.01755 0.01066 -0.0782 0.1727 1.0000 9.000 1.2299 0.01936 0.01227 -0.0729 0.1256 1.0000 9.500 1.2474 0.02140 0.01412 -0.0678 0.0845 1.0000 10.000 1.2575 0.02407 0.01658 -0.0622 0.0492 1.0000 10.500 1.2636 0.02719 0.01967 -0.0565 0.0345 1.0000 11.500 1.2810 0.03366 0.02641 -0.0477 0.0259 1.0000 12.000 1.2891 0.03736 0.03022 -0.0442 0.0236 1.0000 12.500 1.2977 0.04140 0.03447 -0.0413 0.0220 1.0000 13.000 1.3087 0.04554 0.03872 -0.0390 0.0205 1.0000 13.500 1.3206 0.04996 0.04338 -0.0372 0.0189 1.0000 14.000 1.3321 0.05454 0.04803 -0.0361 0.0170 1.0000 14.500 1.3343 0.06012 0.05403 -0.0354 0.0155 1.0000 15.000 1.3380 0.06604 0.06012 -0.0354 0.0142 1.0000 15.500 1.3323 0.07351 0.06801 -0.0363 0.0131 1.0000 16.000 1.3297 0.08096 0.07569 -0.0379 0.0123 1.0000 17.500 1.2623 0.11706 0.11317 -0.0533 0.0111 1.0000 18.000 1.2312 0.13244 0.12897 -0.0625 0.0110 1.0000 18.500 1.1965 0.14979 0.14670 -0.0738 0.0111 1.0000 19.000 1.1592 0.16927 0.16651 -0.0869 0.0114 1.0000 19.500 1.1246 0.18942 0.18685 -0.1000 0.0118 1.0000