XFOIL Version 6.94 Calculated polar for: manu4/25 2 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 ------- -------- --------- --------- -------- ------- ------- -3.000 0.0098 0.10934 0.09271 -0.0161 0.9999 1.0001 -2.750 0.0123 0.10724 0.09098 -0.0156 0.9999 1.0001 -2.500 0.0142 0.10518 0.08931 -0.0150 0.9999 1.0001 -2.250 0.0152 0.10314 0.08767 -0.0143 0.9999 1.0001 -2.000 0.0150 0.10115 0.08610 -0.0135 0.9999 1.0001 -1.750 0.0136 0.09922 0.08461 -0.0126 0.9999 1.0001 -1.500 0.0104 0.09740 0.08327 -0.0115 0.9999 1.0001 -1.250 0.0046 0.09569 0.08205 -0.0103 0.9999 1.0001 -1.000 -0.0042 0.09421 0.08106 -0.0087 0.9999 1.0001 -0.750 -0.0167 0.09314 0.08051 -0.0069 0.9999 1.0001 -0.500 -0.0338 0.09277 0.08063 -0.0049 0.9999 1.0001 -0.250 -0.0558 0.09352 0.08172 -0.0029 0.9999 1.0001 0.000 -0.0791 0.09495 0.08326 -0.0012 0.9999 1.0001 0.250 -0.0981 0.09634 0.08459 -0.0002 0.9999 1.0001 0.500 -0.1104 0.09750 0.08556 -0.0001 0.9999 1.0001 0.750 -0.1158 0.09864 0.08641 -0.0009 0.9999 1.0001 1.000 -0.1148 0.09995 0.08735 -0.0028 0.9999 1.0001 1.250 -0.1077 0.10152 0.08846 -0.0056 0.9999 1.0001 1.500 -0.0957 0.10341 0.08979 -0.0092 0.9999 1.0001 1.750 -0.0789 0.10558 0.09136 -0.0134 0.9999 1.0001 2.000 -0.0586 0.10804 0.09316 -0.0182 0.9999 1.0001 2.250 -0.0352 0.11075 0.09515 -0.0234 0.9999 1.0001 2.500 -0.0094 0.11370 0.09732 -0.0288 0.9999 1.0001 2.750 0.0176 0.11685 0.09962 -0.0342 0.9999 1.0001 3.000 0.0456 0.12014 0.10202 -0.0395 0.9999 1.0001 3.250 0.0733 0.12351 0.10446 -0.0445 0.9999 1.0001 3.500 0.0993 0.12684 0.10685 -0.0488 0.9999 1.0001 3.750 0.1229 0.13006 0.10915 -0.0523 0.9999 1.0001 4.000 0.1445 0.13316 0.11137 -0.0551 0.9999 1.0001 4.250 0.1640 0.13613 0.11348 -0.0574 0.9999 1.0001 4.500 0.1819 0.13900 0.11558 -0.0593 0.9999 1.0001 4.750 0.1989 0.14179 0.11764 -0.0609 0.9999 1.0001 5.000 0.2149 0.14453 0.11971 -0.0623 0.9999 1.0001