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 = 0.035 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 0.1582 0.05501 0.05061 -0.0214 0.8994 0.9160 -2.750 0.1113 0.05535 0.05105 -0.0125 0.8821 0.8723 -2.500 0.0659 0.05513 0.05094 -0.0048 0.8639 0.8309 -2.000 0.3804 0.04110 0.03397 -0.1032 0.6785 0.2743 -1.750 0.4278 0.03918 0.03098 -0.1049 0.5743 0.2537 -1.500 0.4660 0.03818 0.02880 -0.1057 0.5132 0.2409 -1.250 0.4984 0.03780 0.02746 -0.1058 0.4817 0.2331 -1.000 0.5298 0.03728 0.02630 -0.1059 0.4617 0.2266 -0.750 0.5637 0.03699 0.02543 -0.1063 0.4479 0.2210 -0.500 0.5988 0.03692 0.02474 -0.1067 0.4377 0.2180 -0.250 0.6331 0.03677 0.02423 -0.1071 0.4291 0.2202 0.000 0.6687 0.03677 0.02382 -0.1076 0.4219 0.2268 0.250 0.7018 0.03689 0.02380 -0.1078 0.4150 0.2340 0.500 0.7346 0.03694 0.02374 -0.1079 0.4082 0.2429 0.750 0.7692 0.03722 0.02382 -0.1085 0.4020 0.2611 1.000 0.7992 0.03735 0.02433 -0.1087 0.3968 0.2984 1.250 0.8196 0.03588 0.02474 -0.1064 0.3933 1.0001 1.500 0.8527 0.03719 0.02519 -0.1058 0.3903 1.0001 1.750 0.8826 0.03861 0.02614 -0.1059 0.3876 1.0001 2.000 0.9113 0.04016 0.02742 -0.1062 0.3855 1.0001 2.250 0.9391 0.04189 0.02895 -0.1064 0.3838 1.0001 2.500 0.9627 0.04377 0.03080 -0.1063 0.3823 1.0001 2.750 0.9823 0.04582 0.03296 -0.1058 0.3809 1.0001 3.000 0.9992 0.04810 0.03537 -0.1051 0.3794 1.0001 3.250 1.0134 0.05064 0.03805 -0.1042 0.3779 1.0001 3.500 1.0246 0.05351 0.04107 -0.1033 0.3771 1.0001 3.750 1.0320 0.05684 0.04456 -0.1022 0.3777 1.0001 4.000 1.0342 0.06076 0.04866 -0.1011 0.3795 1.0001 4.250 1.0366 0.06493 0.05295 -0.1002 0.3816 1.0001