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 = 0.045 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 0.0057 0.02753 0.01539 -0.0116 0.9989 0.2646 -2.750 0.0345 0.02765 0.01549 -0.0114 0.9989 0.2903 -2.500 0.0624 0.02767 0.01572 -0.0113 0.9989 0.3236 -2.250 0.0897 0.02769 0.01604 -0.0112 0.9989 0.3595 -2.000 0.1167 0.02778 0.01631 -0.0110 0.9989 0.3824 -1.750 0.1429 0.02785 0.01670 -0.0108 0.9989 0.4110 -1.500 0.1686 0.02797 0.01714 -0.0105 0.9989 0.4426 -1.250 0.1941 0.02813 0.01765 -0.0103 0.9989 0.4649 -1.000 0.2187 0.02820 0.01834 -0.0099 0.9989 0.5080 -0.750 0.2435 0.02821 0.01916 -0.0095 0.9989 0.5630 -0.500 0.3107 0.02591 0.01647 -0.0097 0.4858 1.0011 -0.250 0.3390 0.02752 0.01693 -0.0095 0.4068 1.0011 0.000 0.3689 0.02874 0.01751 -0.0098 0.3678 1.0011 0.250 0.3992 0.02995 0.01821 -0.0100 0.3414 1.0011 0.500 0.4302 0.03111 0.01909 -0.0104 0.3197 1.0011 0.750 0.4605 0.03246 0.02014 -0.0107 0.3033 1.0011 1.000 0.4923 0.03378 0.02148 -0.0112 0.2914 1.0011 1.250 0.5231 0.03548 0.02305 -0.0116 0.2824 1.0011 1.500 0.5541 0.03708 0.02473 -0.0122 0.2730 1.0011 1.750 0.5839 0.03896 0.02654 -0.0128 0.2634 1.0011 2.000 0.6145 0.04078 0.02856 -0.0135 0.2552 1.0011 2.250 0.6445 0.04300 0.03081 -0.0141 0.2504 1.0011 2.500 0.6742 0.04577 0.03367 -0.0149 0.2470 1.0011 2.750 0.7047 0.04839 0.03679 -0.0161 0.2440 1.0011 3.000 0.7333 0.05112 0.03993 -0.0173 0.2376 1.0011 3.250 0.7608 0.05420 0.04326 -0.0183 0.2335 1.0011 3.500 0.7859 0.05773 0.04669 -0.0187 0.2281 1.0011 3.750 0.8104 0.06179 0.05145 -0.0207 0.2252 1.0011 4.000 0.8336 0.06666 0.05666 -0.0223 0.2266 1.0011 4.250 0.8523 0.07510 0.06591 -0.0272 0.2453 1.0011 4.750 0.6768 0.11773 0.10984 -0.0738 0.5236 1.0011