XFOIL Version 6.94 Calculated polar for: GOE 406c 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.040 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 0.2511 0.04760 0.04291 -0.0391 0.7849 0.9999 -2.750 0.2157 0.04717 0.04234 -0.0287 0.7337 0.9742 -2.500 0.1581 0.04729 0.04265 -0.0157 0.7406 0.9322 -2.250 0.1033 0.04688 0.04242 -0.0048 0.7400 0.8905 -2.000 0.0514 0.04579 0.04147 0.0045 0.7321 0.8505 -1.750 0.3025 0.04414 0.03538 -0.0920 0.4850 0.3204 -1.500 0.3586 0.04341 0.03350 -0.0980 0.4588 0.2615 -1.250 0.3997 0.04237 0.03183 -0.1004 0.4419 0.2361 -1.000 0.4402 0.04183 0.03054 -0.1022 0.4285 0.2173 -0.750 0.4760 0.04086 0.02921 -0.1033 0.4169 0.2079 -0.500 0.5133 0.04045 0.02827 -0.1045 0.4078 0.2023 -0.250 0.5492 0.04014 0.02762 -0.1053 0.4008 0.2007 0.000 0.5844 0.03994 0.02702 -0.1058 0.3952 0.2001 0.250 0.6190 0.03988 0.02663 -0.1061 0.3907 0.1997 0.500 0.6517 0.03977 0.02643 -0.1063 0.3866 0.2017 0.750 0.6832 0.03983 0.02643 -0.1062 0.3823 0.2070 1.000 0.7143 0.04016 0.02660 -0.1062 0.3778 0.2180 1.250 0.7453 0.04057 0.02691 -0.1064 0.3736 0.2347 1.500 0.7768 0.04096 0.02758 -0.1069 0.3698 0.2609 1.750 0.7981 0.03923 0.02797 -0.1053 0.3670 0.9999 2.000 0.8307 0.04067 0.02866 -0.1048 0.3649 0.9999 2.250 0.8602 0.04225 0.02992 -0.1051 0.3634 0.9999 2.500 0.8887 0.04398 0.03153 -0.1056 0.3623 0.9999 2.750 0.9161 0.04588 0.03340 -0.1061 0.3616 0.9999 3.000 0.9422 0.04796 0.03551 -0.1066 0.3613 0.9999 3.250 0.9669 0.05019 0.03779 -0.1070 0.3606 0.9999 3.500 0.9903 0.05257 0.04023 -0.1074 0.3597 0.9999 3.750 1.0120 0.05513 0.04287 -0.1076 0.3588 0.9999 4.000 1.0308 0.05810 0.04599 -0.1079 0.3588 0.9999 4.250 1.0411 0.06237 0.05061 -0.1083 0.3617 0.9999 4.500 1.0447 0.06763 0.05620 -0.1087 0.3670 0.9999 4.750 1.0552 0.07212 0.06078 -0.1091 0.3712 0.9999