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 = 0.018 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0544 0.03317 0.02006 -0.0139 0.9989 1.0011 -2.750 -0.0243 0.03313 0.01951 -0.0155 0.9989 1.0011 -2.500 0.0036 0.03323 0.01931 -0.0162 0.9989 1.0011 -2.250 0.0304 0.03342 0.01932 -0.0164 0.9989 1.0011 -2.000 0.0565 0.03367 0.01947 -0.0164 0.9989 1.0011 -1.750 0.0820 0.03399 0.01976 -0.0163 0.9989 1.0011 -1.500 0.1071 0.03438 0.02019 -0.0161 0.9989 1.0011 -1.250 0.1317 0.03484 0.02075 -0.0159 0.9989 1.0011 -1.000 0.1561 0.03539 0.02147 -0.0157 0.9989 1.0011 -0.750 0.1801 0.03606 0.02239 -0.0156 0.9989 1.0011 -0.500 0.2033 0.03693 0.02365 -0.0156 0.9989 1.0011 -0.250 0.2220 0.03866 0.02601 -0.0164 0.9989 1.0011 0.000 0.4149 0.04177 0.02882 -0.0457 0.6651 1.0011 0.250 0.4417 0.04376 0.03036 -0.0435 0.6159 1.0011 0.500 0.4704 0.04596 0.03237 -0.0430 0.5780 1.0011 0.750 0.4993 0.04826 0.03451 -0.0429 0.5484 1.0011 1.000 0.5289 0.05082 0.03702 -0.0435 0.5249 1.0011 1.250 0.5578 0.05349 0.03963 -0.0439 0.5054 1.0011 1.500 0.5873 0.05652 0.04277 -0.0455 0.4893 1.0011 1.750 0.6153 0.06000 0.04637 -0.0475 0.4777 1.0011 2.000 0.6411 0.06403 0.05064 -0.0508 0.4688 1.0011 2.250 0.6654 0.06805 0.05477 -0.0529 0.4620 1.0011 2.500 0.6901 0.07224 0.05897 -0.0545 0.4564 1.0011 2.750 0.7072 0.07758 0.06452 -0.0584 0.4559 1.0011 3.250 0.6996 0.09114 0.07844 -0.0688 0.4734 1.0011 3.500 0.7055 0.09690 0.08420 -0.0713 0.4802 1.0011 3.750 0.6756 0.10399 0.09129 -0.0753 0.5021 1.0011 4.000 0.6479 0.11024 0.09748 -0.0775 0.5276 1.0011 4.250 0.6335 0.11648 0.10367 -0.0807 0.5600 1.0011 4.750 0.3573 0.11811 0.10495 -0.0617 0.9989 1.0011 5.000 0.3690 0.12166 0.10846 -0.0633 0.9989 1.0011