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 = 0.022 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0551 0.03062 0.01912 -0.0134 0.9989 1.0011 -2.750 -0.0223 0.03054 0.01842 -0.0160 0.9989 1.0011 -2.500 0.0064 0.03066 0.01822 -0.0167 0.9989 1.0011 -2.250 0.0337 0.03087 0.01824 -0.0170 0.9989 1.0011 -2.000 0.0601 0.03113 0.01841 -0.0170 0.9989 1.0011 -1.750 0.0859 0.03147 0.01872 -0.0169 0.9989 1.0011 -1.500 0.1112 0.03187 0.01916 -0.0167 0.9989 1.0011 -1.250 0.1362 0.03234 0.01975 -0.0165 0.9989 1.0011 -1.000 0.1610 0.03290 0.02049 -0.0163 0.9989 1.0011 -0.750 0.1855 0.03358 0.02146 -0.0162 0.9989 1.0011 -0.500 0.2090 0.03458 0.02294 -0.0165 0.9989 1.0011 -0.250 0.3601 0.03621 0.02396 -0.0348 0.6301 1.0011 0.000 0.3860 0.03806 0.02517 -0.0326 0.5702 1.0011 0.250 0.4149 0.03994 0.02668 -0.0318 0.5290 1.0011 0.500 0.4450 0.04186 0.02840 -0.0317 0.4969 1.0011 0.750 0.4758 0.04390 0.03033 -0.0319 0.4720 1.0011 1.000 0.5065 0.04608 0.03243 -0.0322 0.4521 1.0011 1.250 0.5378 0.04839 0.03481 -0.0331 0.4348 1.0011 1.500 0.5697 0.05116 0.03780 -0.0348 0.4221 1.0011 1.750 0.6004 0.05403 0.04085 -0.0364 0.4111 1.0011 2.000 0.6294 0.05715 0.04397 -0.0372 0.4027 1.0011 2.250 0.6595 0.06126 0.04853 -0.0411 0.3974 1.0011 2.500 0.6861 0.06579 0.05338 -0.0447 0.3946 1.0011 2.750 0.7084 0.07077 0.05862 -0.0484 0.3934 1.0011 3.000 0.7261 0.07630 0.06437 -0.0523 0.3954 1.0011 3.250 0.7408 0.08204 0.07025 -0.0556 0.3995 1.0011 3.500 0.7512 0.08810 0.07646 -0.0593 0.4058 1.0011 3.750 0.7360 0.09669 0.08520 -0.0662 0.4264 1.0011 4.000 0.6868 0.10608 0.09465 -0.0740 0.4675 1.0011 4.250 0.6416 0.11444 0.10296 -0.0795 0.5286 1.0011 4.500 0.3491 0.11449 0.10276 -0.0626 0.9989 1.0011 4.750 0.3608 0.11807 0.10630 -0.0641 0.9989 1.0011 5.000 0.3781 0.12196 0.11017 -0.0670 0.9949 1.0011