XFOIL Version 6.96 Calculated polar for: FX S 02/1-158 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -13.500 -0.6796 0.06995 0.06752 -0.0487 1.0000 0.0047 -13.250 -0.7176 0.05940 0.05674 -0.0560 1.0000 0.0046 -13.000 -0.7284 0.05437 0.05160 -0.0593 1.0000 0.0046 -12.750 -0.7438 0.04904 0.04610 -0.0626 1.0000 0.0046 -12.500 -0.7533 0.04481 0.04173 -0.0651 1.0000 0.0045 -12.250 -0.7775 0.03896 0.03562 -0.0681 1.0000 0.0046 -12.000 -0.7800 0.03602 0.03255 -0.0693 1.0000 0.0046 -11.750 -0.7925 0.03239 0.02872 -0.0702 1.0000 0.0047 -11.500 -0.8008 0.02965 0.02583 -0.0704 1.0000 0.0047 -11.250 -0.7975 0.02738 0.02338 -0.0708 1.0000 0.0048 -11.000 -0.7909 0.02549 0.02133 -0.0701 1.0000 0.0051 -10.750 -0.7593 0.02361 0.01928 -0.0735 0.9735 0.0055 -10.500 -0.6931 0.02152 0.01692 -0.0837 0.9357 0.0060 -10.250 -0.6348 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1.2727 0.02389 0.01664 -0.0645 0.0065 1.0000 9.750 1.2748 0.02560 0.01841 -0.0629 0.0059 1.0000 10.000 1.2769 0.02757 0.02046 -0.0619 0.0055 1.0000 10.250 1.2762 0.03005 0.02303 -0.0611 0.0049 1.0000 10.500 1.2772 0.03255 0.02561 -0.0607 0.0047 1.0000 10.750 1.2780 0.03514 0.02828 -0.0604 0.0046 1.0000 11.000 1.2777 0.03789 0.03112 -0.0601 0.0045 1.0000 11.250 1.2752 0.04091 0.03424 -0.0598 0.0042 1.0000 11.500 1.2754 0.04366 0.03709 -0.0596 0.0043 1.0000 11.750 1.2724 0.04679 0.04031 -0.0595 0.0041 1.0000 12.000 1.2696 0.04991 0.04352 -0.0594 0.0040 1.0000 12.250 1.2659 0.05318 0.04688 -0.0593 0.0039 1.0000 12.500 1.2641 0.05627 0.05004 -0.0594 0.0039 1.0000 12.750 1.2619 0.05955 0.05341 -0.0595 0.0038 1.0000 13.000 1.2603 0.06284 0.05679 -0.0598 0.0038 1.0000 13.250 1.2582 0.06628 0.06031 -0.0601 0.0037 1.0000 13.500 1.2584 0.06950 0.06361 -0.0605 0.0037 1.0000 13.750 1.2564 0.07307 0.06726 -0.0610 0.0037 1.0000 14.000 1.2557 0.07652 0.07079 -0.0615 0.0037 1.0000 14.250 1.2543 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