XFOIL Version 6.94 Calculated polar for: manu04 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.070 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0687 0.04443 0.03482 -0.0537 0.3141 0.2116 -2.750 -0.0537 0.04443 0.03503 -0.0510 0.3088 0.2405 -2.500 -0.0373 0.04412 0.03483 -0.0485 0.3029 0.2779 -2.250 -0.0179 0.04331 0.03401 -0.0468 0.2978 0.3175 -2.000 0.0338 0.03500 0.02362 -0.0519 0.2955 0.1719 -1.750 0.0813 0.03332 0.02131 -0.0556 0.2916 0.1706 -1.500 0.1280 0.03205 0.01964 -0.0591 0.2887 0.1716 -1.250 0.1782 0.03112 0.01823 -0.0633 0.2862 0.1756 -1.000 0.2268 0.03066 0.01746 -0.0674 0.2838 0.1836 -0.750 0.2763 0.03019 0.01670 -0.0718 0.2817 0.1917 -0.500 0.3241 0.02977 0.01617 -0.0759 0.2801 0.2029 -0.250 0.3870 0.02969 0.01616 -0.0837 0.2785 0.2250 0.000 0.4433 0.02985 0.01645 -0.0899 0.2774 0.2795 0.250 0.4881 0.03016 0.01687 -0.0936 0.2770 0.3496 0.500 0.5287 0.03056 0.01743 -0.0965 0.2770 0.3914 0.750 0.5658 0.03107 0.01807 -0.0985 0.2771 0.4215 1.000 0.6006 0.03161 0.01876 -0.1000 0.2771 0.4488 1.250 0.6337 0.03212 0.01951 -0.1012 0.2770 0.4863 1.750 0.7785 0.03428 0.02268 -0.1219 0.2751 1.0001 2.000 0.8042 0.03512 0.02357 -0.1212 0.2761 1.0001 2.250 0.8288 0.03606 0.02465 -0.1203 0.2784 1.0001 2.500 0.8529 0.03725 0.02598 -0.1195 0.2811 1.0001 2.750 0.8767 0.03858 0.02741 -0.1187 0.2838 1.0001 3.000 0.8997 0.04004 0.02893 -0.1178 0.2866 1.0001 3.250 0.9225 0.04159 0.03055 -0.1170 0.2891 1.0001 3.500 0.9456 0.04344 0.03237 -0.1163 0.2913 1.0001 3.750 0.9629 0.04464 0.03414 -0.1142 0.3014 1.0001 4.000 0.9823 0.04680 0.03643 -0.1130 0.3074 1.0001 4.250 1.0051 0.04862 0.03834 -0.1124 0.3141 1.0001 4.500 1.0208 0.05202 0.04209 -0.1112 0.3304 1.0001