XFOIL Version 6.94 Calculated polar for: manu4/38 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.030 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.4030 0.05995 0.05339 0.0593 0.9999 0.8202 -2.750 -0.4629 0.05746 0.05104 0.0687 0.9999 0.7917 -2.500 -0.4676 0.05584 0.04935 0.0614 0.9999 0.7169 -2.250 -0.3497 0.05765 0.05020 0.0210 0.9999 0.4908 -2.000 -0.2828 0.05556 0.04749 0.0062 0.9999 0.3813 -1.750 -0.2412 0.05328 0.04492 0.0013 0.9999 0.3424 -1.250 -0.0405 0.04328 0.03416 -0.0251 0.7393 0.3144 -1.000 0.0980 0.03981 0.02756 -0.0435 0.4891 0.3169 -0.750 0.1480 0.03930 0.02624 -0.0479 0.4695 0.3279 -0.500 0.2077 0.03881 0.02497 -0.0542 0.4531 0.3502 -0.250 0.2787 0.03797 0.02364 -0.0628 0.4375 0.3853 0.000 0.3696 0.03577 0.02179 -0.0752 0.4250 0.5365 0.250 0.5134 0.03524 0.02168 -0.0999 0.4148 1.0001 0.500 0.5636 0.03611 0.02207 -0.1044 0.4121 1.0001 0.750 0.6080 0.03692 0.02262 -0.1078 0.4102 1.0001 1.000 0.6497 0.03778 0.02328 -0.1106 0.4089 1.0001 1.250 0.6888 0.03869 0.02409 -0.1129 0.4082 1.0001 1.500 0.7251 0.03968 0.02503 -0.1147 0.4081 1.0001 1.750 0.7587 0.04075 0.02608 -0.1160 0.4085 1.0001 2.000 0.7899 0.04192 0.02724 -0.1169 0.4093 1.0001 2.250 0.8192 0.04319 0.02853 -0.1175 0.4103 1.0001 2.500 0.8465 0.04454 0.02994 -0.1177 0.4112 1.0001 2.750 0.8714 0.04598 0.03144 -0.1175 0.4118 1.0001 3.000 0.8946 0.04751 0.03306 -0.1170 0.4123 1.0001 3.250 0.9168 0.04913 0.03476 -0.1165 0.4129 1.0001 3.500 0.9381 0.05089 0.03664 -0.1158 0.4137 1.0001 3.750 0.9538 0.05241 0.03843 -0.1143 0.4159 1.0001 4.000 0.9656 0.05421 0.04052 -0.1125 0.4196 1.0001 4.250 0.9755 0.05643 0.04299 -0.1107 0.4249 1.0001 4.500 0.9879 0.05886 0.04558 -0.1093 0.4298 1.0001 4.750 1.0060 0.06147 0.04824 -0.1088 0.4341 1.0001 5.000 0.9882 0.06457 0.05178 -0.1044 0.4438 1.0001