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 = 0.012 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0557 0.03929 0.02252 -0.0136 0.9989 1.0011 -2.750 -0.0284 0.03924 0.02207 -0.0145 0.9989 1.0011 -2.500 -0.0019 0.03930 0.02185 -0.0150 0.9989 1.0011 -2.250 0.0240 0.03945 0.02182 -0.0151 0.9989 1.0011 -2.000 0.0494 0.03968 0.02193 -0.0152 0.9989 1.0011 -1.750 0.0743 0.03997 0.02219 -0.0151 0.9989 1.0011 -1.500 0.0988 0.04033 0.02258 -0.0149 0.9989 1.0011 -1.250 0.1228 0.04077 0.02311 -0.0147 0.9989 1.0011 -1.000 0.1464 0.04130 0.02380 -0.0145 0.9989 1.0011 -0.750 0.1695 0.04193 0.02466 -0.0144 0.9989 1.0011 -0.500 0.1920 0.04272 0.02575 -0.0143 0.9989 1.0011 -0.250 0.2131 0.04378 0.02723 -0.0144 0.9989 1.0011 0.000 0.2296 0.04556 0.02958 -0.0150 0.9989 1.0011 0.250 0.2247 0.05046 0.03496 -0.0176 0.9989 1.0011 0.500 0.2040 0.05768 0.04218 -0.0224 0.9989 1.0011 1.000 0.4761 0.06690 0.05173 -0.0727 0.7692 1.0011 1.250 0.5271 0.06991 0.05468 -0.0766 0.7289 1.0011 1.500 0.5371 0.07384 0.05857 -0.0771 0.7102 1.0011 1.750 0.5509 0.07791 0.06255 -0.0780 0.6954 1.0011 2.000 0.5584 0.08208 0.06667 -0.0786 0.6841 1.0011 2.250 0.5771 0.08631 0.07083 -0.0800 0.6729 1.0011 2.500 0.5730 0.09084 0.07529 -0.0801 0.6710 1.0011 2.750 0.5683 0.09525 0.07961 -0.0799 0.6708 1.0011 3.000 0.5674 0.09973 0.08401 -0.0802 0.6719 1.0011 3.250 0.5717 0.10437 0.08858 -0.0813 0.6742 1.0011 3.500 0.5471 0.10831 0.09242 -0.0799 0.6880 1.0011 3.750 0.5488 0.11280 0.09684 -0.0810 0.6963 1.0011 4.000 0.5390 0.11675 0.10072 -0.0812 0.7127 1.0011 4.250 0.5462 0.12179 0.10570 -0.0836 0.7292 1.0011 4.500 0.5266 0.12500 0.10886 -0.0831 0.7621 1.0011 4.750 0.5040 0.12787 0.11163 -0.0820 0.8071 1.0011 5.000 0.4657 0.12937 0.11303 -0.0782 0.8817 1.0011