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 = 0.080 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0565 0.03999 0.03027 -0.0568 0.2908 0.1707 -2.750 -0.0370 0.03602 0.02557 -0.0551 0.2884 0.1595 -2.500 -0.0119 0.03427 0.02347 -0.0545 0.2852 0.1587 -2.250 0.0194 0.03264 0.02142 -0.0549 0.2819 0.1588 -2.000 0.0625 0.03102 0.01928 -0.0578 0.2784 0.1581 -1.750 0.1173 0.02964 0.01730 -0.0632 0.2750 0.1586 -1.500 0.1692 0.02870 0.01590 -0.0679 0.2726 0.1606 -1.250 0.2133 0.02836 0.01544 -0.0712 0.2711 0.1649 -1.000 0.2514 0.02838 0.01538 -0.0731 0.2698 0.1722 -0.750 0.2910 0.02830 0.01519 -0.0754 0.2685 0.1795 -0.500 0.3308 0.02835 0.01515 -0.0777 0.2676 0.1896 -0.250 0.3805 0.02847 0.01526 -0.0823 0.2663 0.2095 0.000 0.4385 0.02860 0.01557 -0.0890 0.2645 0.2632 0.250 0.4785 0.02899 0.01600 -0.0916 0.2627 0.3236 0.500 0.5173 0.02941 0.01648 -0.0942 0.2609 0.3718 0.750 0.5548 0.02989 0.01709 -0.0964 0.2600 0.4024 1.000 0.5898 0.03043 0.01775 -0.0979 0.2601 0.4286 1.250 0.6232 0.03098 0.01849 -0.0992 0.2606 0.4574 1.500 0.6558 0.03142 0.01932 -0.1003 0.2613 0.5042 1.750 0.7781 0.03287 0.02185 -0.1220 0.2633 1.0001 2.000 0.8045 0.03392 0.02290 -0.1215 0.2649 1.0001 2.250 0.8300 0.03505 0.02406 -0.1210 0.2666 1.0001 2.500 0.8550 0.03627 0.02528 -0.1204 0.2687 1.0001 2.750 0.8794 0.03762 0.02663 -0.1197 0.2705 1.0001 3.000 0.9035 0.03915 0.02815 -0.1191 0.2723 1.0001 3.250 0.9281 0.04111 0.03005 -0.1188 0.2740 1.0001 3.500 0.9451 0.04194 0.03159 -0.1162 0.2852 1.0001 3.750 0.9669 0.04391 0.03360 -0.1153 0.2895 1.0001 4.000 0.9908 0.04580 0.03548 -0.1148 0.2934 1.0001 4.250 1.0060 0.04850 0.03866 -0.1129 0.3095 1.0001 5.000 0.7493 0.07949 0.07252 -0.0995 0.5553 1.0001