XFOIL Version 6.94 Calculated polar for: NACA4412c 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.100 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0574 0.02760 0.01900 -0.0600 0.9878 0.1325 -2.750 0.0029 0.02601 0.01723 -0.0659 0.9795 0.1475 -2.500 0.0562 0.02461 0.01580 -0.0705 0.9680 0.1653 -2.250 0.1110 0.02302 0.01437 -0.0754 0.9561 0.1882 -2.000 0.1695 0.02136 0.01288 -0.0805 0.9396 0.2199 -1.750 0.2202 0.01971 0.01144 -0.0835 0.9078 0.2592 -1.500 0.2560 0.01849 0.01043 -0.0832 0.8486 0.3088 -1.250 0.2816 0.01754 0.00946 -0.0802 0.6868 0.4080 -1.000 0.2905 0.01640 0.00864 -0.0737 0.4452 0.9999 -0.750 0.3174 0.01722 0.00865 -0.0734 0.3960 0.9999 -0.500 0.3454 0.01779 0.00873 -0.0733 0.3771 0.9999 -0.250 0.3741 0.01826 0.00886 -0.0734 0.3652 0.9999 0.000 0.4027 0.01881 0.00906 -0.0734 0.3569 0.9999 0.250 0.4321 0.01924 0.00928 -0.0736 0.3499 0.9999 0.500 0.4614 0.01976 0.00954 -0.0737 0.3439 0.9999 0.750 0.4907 0.02046 0.00993 -0.0740 0.3390 0.9999 1.000 0.5205 0.02090 0.01026 -0.0743 0.3347 0.9999 1.250 0.5502 0.02138 0.01062 -0.0746 0.3300 0.9999 1.500 0.5797 0.02194 0.01101 -0.0749 0.3254 0.9999 1.750 0.6093 0.02263 0.01150 -0.0753 0.3215 0.9999 2.000 0.6391 0.02338 0.01211 -0.0757 0.3185 0.9999 2.250 0.6688 0.02390 0.01261 -0.0760 0.3159 0.9999 2.500 0.6984 0.02449 0.01318 -0.0763 0.3131 0.9999 2.750 0.7279 0.02513 0.01378 -0.0767 0.3104 0.9999 3.000 0.7573 0.02582 0.01442 -0.0770 0.3079 0.9999 3.250 0.7864 0.02657 0.01509 -0.0774 0.3051 0.9999 3.500 0.8156 0.02757 0.01597 -0.0778 0.3023 0.9999 3.750 0.8442 0.02834 0.01679 -0.0781 0.2999 0.9999 4.000 0.8727 0.02905 0.01760 -0.0784 0.2976 0.9999 4.250 0.9012 0.02987 0.01851 -0.0786 0.2955 0.9999 4.500 0.9294 0.03075 0.01946 -0.0789 0.2934 0.9999 4.750 0.9575 0.03169 0.02044 -0.0791 0.2915 0.9999 5.000 0.9855 0.03267 0.02147 -0.0794 0.2896 0.9999