XFOIL Version 6.94 Calculated polar for: WBL FX 71-089A AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.035 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 0.0027 0.02977 0.01685 -0.0112 0.9989 0.3528 -2.750 0.0302 0.02972 0.01704 -0.0110 0.9989 0.3887 -2.500 0.0571 0.02967 0.01721 -0.0107 0.9989 0.4202 -2.250 0.0838 0.02963 0.01738 -0.0104 0.9989 0.4457 -2.000 0.1094 0.02953 0.01764 -0.0100 0.9989 0.4835 -1.750 0.1344 0.02942 0.01793 -0.0094 0.9989 0.5211 -1.500 0.1582 0.02924 0.01830 -0.0086 0.9989 0.5658 -1.250 0.1787 0.02880 0.01874 -0.0069 0.9989 0.6463 -1.000 0.1897 0.02746 0.01869 -0.0027 0.9989 0.9980 -0.750 0.2182 0.02812 0.01929 -0.0036 0.9989 1.0011 -0.500 0.2451 0.02890 0.02021 -0.0041 0.9989 1.0011 -0.250 0.3479 0.02897 0.01814 -0.0114 0.5155 1.0011 0.000 0.3752 0.03066 0.01889 -0.0107 0.4521 1.0011 0.250 0.4047 0.03209 0.01978 -0.0107 0.4121 1.0011 0.500 0.4352 0.03346 0.02081 -0.0109 0.3833 1.0011 0.750 0.4661 0.03491 0.02198 -0.0112 0.3613 1.0011 1.000 0.4970 0.03643 0.02330 -0.0116 0.3418 1.0011 1.250 0.5289 0.03806 0.02503 -0.0123 0.3287 1.0011 1.500 0.5602 0.04002 0.02698 -0.0130 0.3195 1.0011 1.750 0.5922 0.04205 0.02921 -0.0140 0.3108 1.0011 2.000 0.6218 0.04427 0.03142 -0.0146 0.3010 1.0011 2.250 0.6522 0.04653 0.03394 -0.0157 0.2914 1.0011 2.500 0.6820 0.04924 0.03677 -0.0166 0.2869 1.0011 2.750 0.7115 0.05248 0.04019 -0.0178 0.2845 1.0011 3.000 0.7409 0.05609 0.04429 -0.0198 0.2831 1.0011 3.250 0.7676 0.06004 0.04876 -0.0221 0.2790 1.0011 3.500 0.7922 0.06460 0.05373 -0.0246 0.2780 1.0011 3.750 0.8144 0.06989 0.05941 -0.0273 0.2807 1.0011 4.000 0.8326 0.07546 0.06527 -0.0300 0.2822 1.0011 4.250 0.8501 0.08113 0.07116 -0.0322 0.2856 1.0011 4.750 0.5964 0.12113 0.11208 -0.0796 0.6910 1.0011 5.000 0.6175 0.12541 0.11632 -0.0804 0.6664 1.0011