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 = 0.005 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0242 0.05653 0.02758 0.0018 0.9989 1.0011 -2.750 -0.0013 0.05646 0.02717 0.0017 0.9989 1.0011 -2.500 0.0219 0.05647 0.02687 0.0016 0.9989 1.0011 -2.250 0.0451 0.05654 0.02668 0.0016 0.9989 1.0011 -2.000 0.0683 0.05669 0.02663 0.0015 0.9989 1.0011 -1.750 0.0913 0.05690 0.02669 0.0015 0.9989 1.0011 -1.500 0.1142 0.05719 0.02689 0.0015 0.9989 1.0011 -1.250 0.1367 0.05754 0.02720 0.0015 0.9989 1.0011 -1.000 0.1588 0.05798 0.02766 0.0015 0.9989 1.0011 -0.750 0.1805 0.05850 0.02826 0.0016 0.9989 1.0011 -0.500 0.2017 0.05911 0.02902 0.0016 0.9989 1.0011 -0.250 0.2223 0.05983 0.02994 0.0017 0.9989 1.0011 0.000 0.2422 0.06069 0.03107 0.0016 0.9989 1.0011 0.250 0.2612 0.06171 0.03242 0.0015 0.9989 1.0011 0.500 0.2787 0.06296 0.03405 0.0013 0.9989 1.0011 0.750 0.2942 0.06453 0.03604 0.0009 0.9989 1.0011 1.000 0.3065 0.06659 0.03854 0.0003 0.9989 1.0011 1.250 0.3144 0.06938 0.04169 -0.0008 0.9989 1.0011 1.500 0.3167 0.07299 0.04551 -0.0025 0.9989 1.0011 1.750 0.3158 0.07717 0.04973 -0.0047 0.9989 1.0011 2.000 0.3151 0.08141 0.05394 -0.0071 0.9989 1.0011 2.250 0.3164 0.08547 0.05792 -0.0096 0.9989 1.0011 2.500 0.3195 0.08932 0.06170 -0.0119 0.9989 1.0011 2.750 0.3241 0.09300 0.06529 -0.0142 0.9989 1.0011 3.000 0.3298 0.09656 0.06878 -0.0163 0.9989 1.0011 3.250 0.3366 0.10006 0.07221 -0.0184 0.9989 1.0011 3.500 0.3442 0.10351 0.07559 -0.0205 0.9989 1.0011 3.750 0.3522 0.10690 0.07892 -0.0225 0.9989 1.0011 4.000 0.3608 0.11026 0.08221 -0.0245 0.9989 1.0011 4.250 0.3697 0.11358 0.08548 -0.0265 0.9989 1.0011 4.500 0.3789 0.11689 0.08874 -0.0284 0.9989 1.0011 4.750 0.3884 0.12018 0.09198 -0.0303 0.9989 1.0011 5.000 0.3981 0.12345 0.09522 -0.0322 0.9989 1.0011