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 -0.0592 0.02892 0.01881 -0.0109 0.9989 1.0011 -2.750 -0.0207 0.02859 0.01760 -0.0163 0.9989 1.0011 -2.500 0.0088 0.02873 0.01739 -0.0172 0.9989 1.0011 -2.250 0.0364 0.02894 0.01742 -0.0175 0.9989 1.0011 -2.000 0.0631 0.02922 0.01761 -0.0175 0.9989 1.0011 -1.750 0.0892 0.02957 0.01794 -0.0174 0.9989 1.0011 -1.500 0.1148 0.02998 0.01840 -0.0173 0.9989 1.0011 -1.250 0.1401 0.03047 0.01900 -0.0171 0.9989 1.0011 -1.000 0.1653 0.03103 0.01976 -0.0169 0.9989 1.0011 -0.750 0.1903 0.03173 0.02079 -0.0168 0.9989 1.0011 -0.500 0.2129 0.03316 0.02290 -0.0176 0.9989 1.0011 -0.250 0.3416 0.03411 0.02182 -0.0278 0.5487 1.0011 0.000 0.3701 0.03583 0.02301 -0.0268 0.5002 1.0011 0.250 0.4001 0.03750 0.02436 -0.0265 0.4652 1.0011 0.500 0.4311 0.03924 0.02593 -0.0266 0.4384 1.0011 0.750 0.4629 0.04109 0.02773 -0.0270 0.4173 1.0011 1.000 0.4942 0.04306 0.02963 -0.0274 0.4003 1.0011 1.250 0.5268 0.04515 0.03192 -0.0284 0.3859 1.0011 1.500 0.5576 0.04759 0.03439 -0.0290 0.3751 1.0011 1.750 0.5901 0.05031 0.03749 -0.0310 0.3659 1.0011 2.000 0.6208 0.05314 0.04046 -0.0322 0.3578 1.0011 2.250 0.6496 0.05640 0.04368 -0.0328 0.3518 1.0011 2.500 0.6793 0.06034 0.04810 -0.0359 0.3484 1.0011 2.750 0.7064 0.06482 0.05295 -0.0390 0.3474 1.0011 3.000 0.7301 0.06975 0.05819 -0.0423 0.3474 1.0011 3.250 0.7506 0.07505 0.06372 -0.0453 0.3489 1.0011 3.500 0.7683 0.08081 0.06974 -0.0488 0.3537 1.0011 3.750 0.7633 0.09012 0.07943 -0.0573 0.3738 1.0011