XFOIL Version 6.94 Calculated polar for: manu4/33 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.060 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0998 0.05839 0.04979 -0.0376 0.3076 0.3417 -2.750 -0.0744 0.05637 0.04765 -0.0385 0.3052 0.3594 -2.500 -0.0447 0.05420 0.04537 -0.0397 0.3028 0.3813 -2.250 -0.0106 0.05169 0.04279 -0.0412 0.3004 0.4082 -1.750 0.1369 0.04138 0.02986 -0.0656 0.2959 0.1889 -1.500 0.1917 0.03950 0.02753 -0.0711 0.2939 0.1870 -1.250 0.2432 0.03812 0.02562 -0.0758 0.2922 0.1887 -1.000 0.2910 0.03742 0.02462 -0.0797 0.2912 0.1941 -0.750 0.3337 0.03705 0.02404 -0.0827 0.2904 0.1989 -0.500 0.3760 0.03667 0.02328 -0.0854 0.2899 0.2051 -0.250 0.4173 0.03659 0.02319 -0.0881 0.2900 0.2139 0.000 0.4734 0.03671 0.02321 -0.0941 0.2902 0.2312 0.250 0.5319 0.03706 0.02364 -0.1007 0.2906 0.2691 0.500 0.5942 0.03727 0.02402 -0.1084 0.2902 0.3563 0.750 0.6353 0.03792 0.02484 -0.1114 0.2893 0.4048 1.000 0.6707 0.03872 0.02577 -0.1132 0.2887 0.4384 1.250 0.7039 0.03952 0.02672 -0.1144 0.2891 0.4667 1.500 0.7362 0.04017 0.02782 -0.1155 0.2912 0.5030 1.750 0.7671 0.04096 0.02918 -0.1164 0.2945 0.5535 2.250 0.8974 0.04495 0.03419 -0.1340 0.3032 1.0001 2.500 0.9221 0.04692 0.03604 -0.1336 0.3059 1.0001 2.750 0.9480 0.04926 0.03822 -0.1336 0.3087 1.0001 3.000 0.9601 0.04998 0.03961 -0.1304 0.3239 1.0001 3.250 0.9821 0.05256 0.04210 -0.1298 0.3291 1.0001 3.500 0.9963 0.05465 0.04467 -0.1279 0.3506 1.0001 3.750 1.0208 0.05758 0.04787 -0.1286 0.3802 1.0001 4.250 0.4171 0.09763 0.08933 -0.0857 0.7347 0.4409 4.500 0.4650 0.09985 0.09174 -0.0899 0.7105 0.4861 5.000 0.5897 0.10485 0.09761 -0.1041 0.6426 1.0001