XFOIL Version 6.94 Calculated polar for: EPPLER 335 BL AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.026 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 0.0128 0.02797 0.01846 -0.0151 0.9994 1.0006 -2.750 0.0263 0.02904 0.01926 -0.0146 0.9994 1.0006 -2.500 0.0405 0.03025 0.02023 -0.0143 0.9994 1.0006 -2.250 0.0551 0.03160 0.02140 -0.0145 0.9994 1.0006 -2.000 0.2133 0.02972 0.01889 -0.0347 0.8694 1.0006 -1.750 0.2290 0.03109 0.01976 -0.0275 0.7615 1.0006 -1.500 0.2386 0.03262 0.02062 -0.0202 0.6628 1.0006 -1.250 0.2596 0.03396 0.02138 -0.0172 0.5981 1.0006 -1.000 0.2871 0.03528 0.02228 -0.0167 0.5599 1.0006 -0.750 0.3164 0.03658 0.02327 -0.0170 0.5345 1.0006 -0.500 0.3461 0.03790 0.02432 -0.0174 0.5167 1.0006 -0.250 0.3768 0.03932 0.02556 -0.0185 0.5020 1.0006 0.000 0.4091 0.04106 0.02721 -0.0205 0.4898 1.0006 0.250 0.4389 0.04254 0.02855 -0.0215 0.4791 1.0006 0.500 0.4699 0.04445 0.03042 -0.0235 0.4697 1.0006 0.750 0.5008 0.04662 0.03260 -0.0261 0.4612 1.0006 1.000 0.5302 0.04868 0.03459 -0.0277 0.4559 1.0006 1.250 0.5586 0.05084 0.03666 -0.0290 0.4523 1.0006 1.500 0.5887 0.05440 0.04036 -0.0336 0.4505 1.0006 1.750 0.6158 0.05830 0.04440 -0.0383 0.4492 1.0006 2.000 0.6390 0.06247 0.04869 -0.0428 0.4482 1.0006 2.250 0.6581 0.06687 0.05318 -0.0469 0.4477 1.0006 2.500 0.6740 0.07143 0.05780 -0.0506 0.4490 1.0006 2.750 0.6860 0.07661 0.06306 -0.0548 0.4540 1.0006 3.000 0.6721 0.08488 0.07146 -0.0622 0.4706 1.0006 3.500 0.6425 0.09859 0.08523 -0.0709 0.5200 1.0006