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 = 0.026 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0083 0.03217 0.01890 -0.0091 0.9989 0.5021 -2.750 0.0166 0.03193 0.01897 -0.0082 0.9989 0.5394 -2.500 0.0410 0.03164 0.01901 -0.0071 0.9989 0.5759 -2.250 0.0617 0.03113 0.01913 -0.0050 0.9989 0.6386 -1.500 0.1326 0.02932 0.01865 -0.0015 0.9989 1.0011 -1.250 0.1607 0.02981 0.01894 -0.0023 0.9989 1.0011 -1.000 0.1873 0.03037 0.01941 -0.0027 0.9989 1.0011 -0.750 0.2130 0.03101 0.02005 -0.0029 0.9989 1.0011 -0.500 0.2379 0.03175 0.02091 -0.0031 0.9989 1.0011 -0.250 0.2619 0.03276 0.02224 -0.0035 0.9989 1.0011 0.000 0.3938 0.03326 0.02128 -0.0164 0.5811 1.0011 0.250 0.4209 0.03510 0.02239 -0.0152 0.5211 1.0011 0.500 0.4501 0.03683 0.02366 -0.0149 0.4810 1.0011 0.750 0.4795 0.03854 0.02503 -0.0148 0.4489 1.0011 1.000 0.5109 0.04039 0.02676 -0.0154 0.4257 1.0011 1.250 0.5417 0.04227 0.02857 -0.0159 0.4047 1.0011 1.500 0.5725 0.04428 0.03060 -0.0167 0.3868 1.0011 1.750 0.6038 0.04665 0.03301 -0.0176 0.3764 1.0011 2.000 0.6361 0.04943 0.03616 -0.0196 0.3687 1.0011 2.250 0.6660 0.05218 0.03892 -0.0205 0.3605 1.0011 2.500 0.6958 0.05553 0.04271 -0.0231 0.3512 1.0011 2.750 0.7235 0.05874 0.04603 -0.0243 0.3437 1.0011 3.000 0.7514 0.06304 0.05067 -0.0271 0.3432 1.0011 3.250 0.7752 0.06895 0.05715 -0.0325 0.3482 1.0011 3.500 0.7930 0.07500 0.06356 -0.0370 0.3525 1.0011 3.750 0.8059 0.08098 0.06976 -0.0406 0.3550 1.0011 4.000 0.8180 0.08687 0.07578 -0.0434 0.3584 1.0011 5.000 0.5156 0.12582 0.11504 -0.0759 0.8909 1.0011