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 -0.0683 0.02787 0.01835 -0.0123 0.9989 0.7351 -2.750 -0.0187 0.02554 0.01638 -0.0166 0.9989 1.0011 -2.500 0.0132 0.02570 0.01610 -0.0181 0.9989 1.0011 -2.250 0.0416 0.02594 0.01615 -0.0185 0.9989 1.0011 -2.000 0.0688 0.02624 0.01636 -0.0185 0.9989 1.0011 -1.750 0.0953 0.02660 0.01672 -0.0184 0.9989 1.0011 -1.500 0.1214 0.02703 0.01721 -0.0182 0.9989 1.0011 -1.250 0.1473 0.02752 0.01784 -0.0181 0.9989 1.0011 -1.000 0.1734 0.02809 0.01866 -0.0179 0.9989 1.0011 -0.500 0.2958 0.02989 0.01830 -0.0233 0.4765 1.0011 -0.250 0.3243 0.03140 0.01917 -0.0227 0.4297 1.0011 0.000 0.3552 0.03272 0.02022 -0.0227 0.3973 1.0011 0.250 0.3863 0.03412 0.02138 -0.0228 0.3731 1.0011 0.500 0.4179 0.03558 0.02269 -0.0230 0.3537 1.0011 0.750 0.4505 0.03707 0.02430 -0.0235 0.3385 1.0011 1.000 0.4807 0.03899 0.02601 -0.0235 0.3263 1.0011 1.250 0.5137 0.04072 0.02809 -0.0244 0.3163 1.0011 1.500 0.5448 0.04271 0.03012 -0.0248 0.3068 1.0011 1.750 0.5751 0.04531 0.03271 -0.0254 0.3002 1.0011 2.000 0.6073 0.04773 0.03558 -0.0267 0.2942 1.0011 2.250 0.6383 0.05057 0.03877 -0.0280 0.2897 1.0011 2.500 0.6679 0.05362 0.04208 -0.0293 0.2853 1.0011 2.750 0.6958 0.05684 0.04545 -0.0302 0.2810 1.0011 3.000 0.7221 0.06063 0.04925 -0.0307 0.2772 1.0011 3.250 0.7476 0.06505 0.05400 -0.0326 0.2758 1.0011 3.500 0.7718 0.06986 0.05927 -0.0352 0.2777 1.0011 3.750 0.7858 0.07886 0.06909 -0.0430 0.2966 1.0011 4.500 0.5515 0.12114 0.11222 -0.0902 0.6905 1.0011 4.750 0.5694 0.12490 0.11594 -0.0900 0.6611 1.0011 5.000 0.5856 0.12869 0.11969 -0.0897 0.6336 1.0011