XFOIL Version 6.94 Calculated polar for: manu01 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.022 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.4651 0.06780 0.06015 0.0621 0.9999 0.7331 -2.750 -0.4584 0.06628 0.05854 0.0547 0.9999 0.6706 -2.500 -0.4150 0.06530 0.05732 0.0404 0.9999 0.5907 -2.250 -0.3648 0.06392 0.05562 0.0276 0.9999 0.5221 -2.000 -0.3213 0.06213 0.05366 0.0196 0.9999 0.4811 -1.750 -0.2772 0.06069 0.05198 0.0124 0.9999 0.4499 -1.500 -0.2370 0.05875 0.04999 0.0077 0.9999 0.4286 -1.250 -0.2000 0.05633 0.04769 0.0049 0.9999 0.4149 -1.000 -0.0212 0.04950 0.04035 -0.0247 0.8238 0.3952 -0.500 0.1987 0.04303 0.03016 -0.0520 0.5483 0.4229 -0.250 0.2735 0.04257 0.02872 -0.0615 0.5233 0.4512 0.000 0.3497 0.04143 0.02726 -0.0709 0.5048 0.5232 0.250 0.5226 0.03942 0.02561 -0.1013 0.4862 1.0001 0.500 0.5755 0.04054 0.02610 -0.1067 0.4817 1.0001 0.750 0.6209 0.04159 0.02673 -0.1105 0.4784 1.0001 1.000 0.6623 0.04266 0.02750 -0.1134 0.4761 1.0001 1.250 0.7003 0.04377 0.02841 -0.1157 0.4746 1.0001 1.500 0.7353 0.04494 0.02946 -0.1175 0.4739 1.0001 1.750 0.7683 0.04620 0.03066 -0.1190 0.4737 1.0001 2.000 0.7987 0.04756 0.03202 -0.1201 0.4742 1.0001 2.250 0.8263 0.04902 0.03353 -0.1207 0.4752 1.0001 2.500 0.8517 0.05061 0.03518 -0.1211 0.4768 1.0001 2.750 0.8751 0.05235 0.03699 -0.1212 0.4786 1.0001 3.000 0.8959 0.05424 0.03897 -0.1210 0.4803 1.0001 3.250 0.9134 0.05632 0.04117 -0.1202 0.4817 1.0001 3.500 0.9285 0.05858 0.04352 -0.1193 0.4830 1.0001 3.750 0.9432 0.06102 0.04603 -0.1184 0.4847 1.0001 4.000 0.9454 0.06371 0.04892 -0.1159 0.4872 1.0001 4.250 0.9180 0.06761 0.05312 -0.1105 0.4937 1.0001 4.500 0.8963 0.07209 0.05775 -0.1065 0.5004 1.0001 4.750 0.8938 0.07616 0.06184 -0.1049 0.5059 1.0001