XFOIL Version 6.94 Calculated polar for: manu4/30 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.050 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0941 0.06700 0.05821 -0.0328 0.3224 0.3884 -2.750 -0.0676 0.06302 0.05426 -0.0330 0.3206 0.4011 -1.750 0.0669 0.05217 0.04322 -0.0383 0.3140 0.5799 -1.500 0.1102 0.04967 0.04061 -0.0424 0.3130 0.5996 -1.250 0.2855 0.04625 0.03414 -0.0829 0.3109 0.2116 -1.000 0.3298 0.04507 0.03269 -0.0862 0.3107 0.2125 -0.750 0.3702 0.04461 0.03221 -0.0888 0.3106 0.2185 -0.500 0.4087 0.04413 0.03147 -0.0908 0.3111 0.2233 -0.250 0.4479 0.04372 0.03073 -0.0929 0.3118 0.2277 0.000 0.4867 0.04371 0.03077 -0.0952 0.3129 0.2348 0.250 0.5393 0.04390 0.03078 -0.1004 0.3136 0.2489 0.500 0.5949 0.04435 0.03120 -0.1065 0.3138 0.2753 0.750 0.6660 0.04453 0.03151 -0.1161 0.3136 0.3631 1.000 0.7100 0.04522 0.03249 -0.1198 0.3137 0.4255 1.250 0.7443 0.04634 0.03373 -0.1214 0.3150 0.4613 1.500 0.7799 0.04763 0.03528 -0.1235 0.3177 0.4972 1.750 0.8123 0.04912 0.03708 -0.1250 0.3201 0.5433 2.000 0.8437 0.05123 0.03946 -0.1265 0.3225 0.5997 2.250 0.8626 0.05045 0.03959 -0.1246 0.3323 0.6471 2.500 0.9527 0.05442 0.04420 -0.1398 0.3436 1.0001 2.750 0.9821 0.05794 0.04743 -0.1410 0.3482 1.0001 3.000 0.9859 0.05858 0.04872 -0.1371 0.3687 1.0001 3.250 0.9943 0.06091 0.05142 -0.1350 0.3926 1.0001 3.500 1.0059 0.06430 0.05508 -0.1345 0.4227 1.0001 4.000 0.2888 0.10726 0.09803 -0.0721 0.8045 0.2652 4.250 0.3431 0.11066 0.10145 -0.0804 0.7836 0.3235 4.750 0.4022 0.11375 0.10494 -0.0866 0.7313 0.4304 5.000 0.4403 0.11609 0.10740 -0.0898 0.7076 0.4734