XFOIL Version 6.94 Calculated polar for: manu4/38 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 -0.3985 0.05922 0.05334 0.0595 0.9999 0.8217 -2.750 -0.4500 0.05667 0.05093 0.0690 0.9999 0.8067 -2.500 -0.4831 0.05397 0.04831 0.0740 0.9999 0.7902 -2.250 -0.4751 0.05149 0.04585 0.0704 0.9999 0.7639 -2.000 -0.3812 0.05292 0.04683 0.0395 0.9999 0.6115 -1.750 -0.0823 0.04441 0.03606 -0.0265 0.7367 0.2853 -1.500 0.0291 0.04122 0.02998 -0.0407 0.4586 0.2839 -1.250 0.0611 0.04024 0.02839 -0.0416 0.4421 0.2821 -1.000 0.0989 0.03931 0.02684 -0.0434 0.4290 0.2835 -0.750 0.1515 0.03862 0.02537 -0.0482 0.4147 0.2898 -0.500 0.2089 0.03788 0.02412 -0.0540 0.4034 0.3051 -0.250 0.2826 0.03720 0.02287 -0.0632 0.3944 0.3402 0.000 0.3916 0.03506 0.02100 -0.0796 0.3876 0.4840 0.250 0.5350 0.03399 0.02113 -0.1045 0.3812 1.0001 0.500 0.5868 0.03482 0.02152 -0.1094 0.3793 1.0001 0.750 0.6320 0.03566 0.02207 -0.1129 0.3780 1.0001 1.000 0.6716 0.03654 0.02274 -0.1152 0.3776 1.0001 1.250 0.7071 0.03747 0.02356 -0.1168 0.3777 1.0001 1.500 0.7396 0.03846 0.02452 -0.1177 0.3783 1.0001 1.750 0.7696 0.03953 0.02557 -0.1182 0.3789 1.0001 2.000 0.7976 0.04066 0.02671 -0.1184 0.3792 1.0001 2.250 0.8240 0.04186 0.02795 -0.1183 0.3792 1.0001 2.500 0.8487 0.04313 0.02930 -0.1179 0.3796 1.0001 2.750 0.8721 0.04450 0.03073 -0.1173 0.3798 1.0001 3.000 0.8944 0.04596 0.03228 -0.1165 0.3806 1.0001 3.250 0.9157 0.04751 0.03397 -0.1158 0.3828 1.0001 3.500 0.9359 0.04924 0.03586 -0.1149 0.3860 1.0001 3.750 0.9559 0.05114 0.03784 -0.1140 0.3891 1.0001 4.000 0.9770 0.05326 0.04000 -0.1134 0.3919 1.0001 4.250 0.9854 0.05480 0.04199 -0.1110 0.3986 1.0001 4.500 0.9941 0.05715 0.04461 -0.1090 0.4053 1.0001 4.750 1.0095 0.05970 0.04726 -0.1080 0.4107 1.0001 5.000 1.0095 0.06228 0.05018 -0.1054 0.4197 1.0001