XFOIL Version 6.94 Calculated polar for: manu04 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.005 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -2.750 -0.0812 0.07218 0.05036 -0.0189 0.9999 1.0001 -2.500 -0.0854 0.07017 0.04862 -0.0174 0.9999 1.0001 -2.250 -0.0909 0.06815 0.04687 -0.0158 0.9999 1.0001 -2.000 -0.0967 0.06612 0.04505 -0.0142 0.9999 1.0001 -1.750 -0.0993 0.06422 0.04319 -0.0133 0.9999 1.0001 -1.500 -0.0939 0.06267 0.04142 -0.0139 0.9999 1.0001 -1.250 -0.0758 0.06165 0.03985 -0.0167 0.9999 1.0001 -1.000 -0.0441 0.06121 0.03856 -0.0216 0.9999 1.0001 -0.750 -0.0031 0.06133 0.03751 -0.0276 0.9999 1.0001 -0.500 0.0401 0.06179 0.03678 -0.0330 0.9999 1.0001 -0.250 0.0798 0.06238 0.03626 -0.0368 0.9999 1.0001 0.000 0.1146 0.06297 0.03594 -0.0391 0.9999 1.0001 0.250 0.1456 0.06356 0.03578 -0.0405 0.9999 1.0001 0.500 0.1740 0.06415 0.03586 -0.0414 0.9999 1.0001 0.750 0.2007 0.06479 0.03615 -0.0419 0.9999 1.0001 1.000 0.2257 0.06551 0.03666 -0.0423 0.9999 1.0001 1.250 0.2490 0.06635 0.03742 -0.0427 0.9999 1.0001 1.500 0.2699 0.06741 0.03852 -0.0431 0.9999 1.0001 1.750 0.2866 0.06887 0.04015 -0.0436 0.9999 1.0001 2.000 0.2959 0.07104 0.04254 -0.0442 0.9999 1.0001 2.250 0.2948 0.07421 0.04580 -0.0448 0.9999 1.0001 2.500 0.2890 0.07790 0.04936 -0.0455 0.9999 1.0001 2.750 0.2863 0.08139 0.05261 -0.0463 0.9999 1.0001 3.000 0.2874 0.08461 0.05556 -0.0472 0.9999 1.0001 3.250 0.2913 0.08760 0.05827 -0.0480 0.9999 1.0001 3.500 0.2967 0.09048 0.06087 -0.0488 0.9999 1.0001 3.750 0.3033 0.09326 0.06337 -0.0496 0.9999 1.0001 4.000 0.3108 0.09599 0.06584 -0.0504 0.9999 1.0001 4.250 0.3190 0.09865 0.06824 -0.0512 0.9999 1.0001 4.500 0.3276 0.10131 0.07065 -0.0520 0.9999 1.0001 4.750 0.3368 0.10390 0.07301 -0.0527 0.9999 1.0001 5.000 0.3463 0.10650 0.07539 -0.0534 0.9999 1.0001