XFOIL Version 6.94 Calculated polar for: GOE 182 (MVA H.27) AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.450 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.5300 0.01032 0.00235 -0.0978 0.4528 0.0702 0.500 0.5844 0.01025 0.00240 -0.0975 0.4064 0.1514 1.000 0.6384 0.01035 0.00256 -0.0972 0.3634 0.2550 1.500 0.6915 0.01017 0.00288 -0.0969 0.3341 0.5160 2.000 0.7465 0.00957 0.00308 -0.0965 0.3145 1.0000 2.500 0.8007 0.01004 0.00334 -0.0960 0.3003 1.0000 3.000 0.8551 0.01044 0.00361 -0.0955 0.2896 1.0000 3.500 0.9086 0.01095 0.00403 -0.0950 0.2811 1.0000 4.000 0.9631 0.01119 0.00410 -0.0947 0.2581 1.0000 4.500 1.0166 0.01158 0.00430 -0.0943 0.2338 1.0000 5.000 1.0696 0.01204 0.00464 -0.0938 0.2166 1.0000 5.500 1.1216 0.01260 0.00505 -0.0932 0.1932 1.0000 6.000 1.1573 0.01545 0.00698 -0.0908 0.0268 1.0000 6.500 1.2052 0.01648 0.00797 -0.0896 0.0037 1.0000 7.000 1.2542 0.01727 0.00884 -0.0885 0.0036 1.0000 7.500 1.3019 0.01817 0.00985 -0.0872 0.0038 1.0000 8.000 1.3474 0.01923 0.01105 -0.0857 0.0040 1.0000 8.500 1.3895 0.02056 0.01257 -0.0837 0.0043 1.0000 9.000 1.4301 0.02187 0.01404 -0.0816 0.0047 1.0000 9.500 1.4655 0.02348 0.01584 -0.0788 0.0053 1.0000 10.000 1.4864 0.02590 0.01849 -0.0744 0.0058 1.0000 10.500 1.5058 0.02797 0.02075 -0.0700 0.0064 1.0000 11.000 1.5076 0.03171 0.02475 -0.0654 0.0071 1.0000 11.500 1.4935 0.03827 0.03159 -0.0634 0.0074 1.0000 12.000 1.4907 0.04504 0.03863 -0.0641 0.0080 1.0000 12.500 1.4730 0.05442 0.04832 -0.0660 0.0086 1.0000 13.000 1.4459 0.06533 0.05949 -0.0685 0.0090 1.0000 13.500 1.4185 0.07589 0.07024 -0.0704 0.0093 1.0000 14.000 1.4093 0.08404 0.07857 -0.0714 0.0100 1.0000 14.500 1.4207 0.08666 0.08135 -0.0664 0.0127 1.0000 15.000 1.1348 0.11221 0.10755 -0.0610 0.0076 1.0000 15.500 1.1335 0.11769 0.11324 -0.0614 0.0084 1.0000 16.000 1.1462 0.11912 0.11472 -0.0589 0.0099 1.0000