XFOIL Version 6.94 Calculated polar for: GOE 439 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 ------- -------- --------- --------- -------- ------- ------- 0.500 0.6051 0.00864 0.00174 -0.1079 0.6456 0.1671 1.000 0.6649 0.00696 0.00192 -0.1093 0.6146 1.0000 2.000 0.7663 0.00756 0.00208 -0.1072 0.5530 1.0000 2.500 0.8168 0.00791 0.00223 -0.1062 0.5239 1.0000 3.000 0.8671 0.00828 0.00244 -0.1053 0.4971 1.0000 3.500 0.9126 0.00888 0.00264 -0.1036 0.4295 1.0000 4.000 0.9520 0.01007 0.00309 -0.1009 0.3070 1.0000 4.500 0.9809 0.01247 0.00433 -0.0969 0.1077 1.0000 5.000 1.0191 0.01408 0.00578 -0.0939 0.0069 1.0000 5.500 1.0665 0.01471 0.00653 -0.0923 0.0075 1.0000 6.000 1.1127 0.01545 0.00743 -0.0903 0.0096 1.0000 6.500 1.1528 0.01671 0.00897 -0.0871 0.0111 1.0000 7.000 1.1795 0.01881 0.01129 -0.0818 0.0113 1.0000 7.500 1.1914 0.02122 0.01388 -0.0743 0.0111 1.0000 8.000 1.1948 0.02402 0.01683 -0.0658 0.0114 1.0000 8.500 1.2142 0.02735 0.02026 -0.0603 0.0111 1.0000 9.500 1.2646 0.03558 0.02890 -0.0542 0.0040 1.0000 11.000 1.2531 0.05159 0.04633 -0.0409 0.0012 1.0000 11.500 1.2434 0.05775 0.05293 -0.0386 0.0011 1.0000 12.000 1.2268 0.06555 0.06114 -0.0381 0.0010 1.0000 12.500 1.2053 0.07482 0.07079 -0.0394 0.0010 1.0000 13.000 1.1802 0.08557 0.08190 -0.0429 0.0010 1.0000 13.500 1.1532 0.09806 0.09471 -0.0485 0.0010 1.0000 14.000 1.1258 0.11202 0.10897 -0.0561 0.0010 1.0000 14.500 1.0994 0.12739 0.12459 -0.0652 0.0010 1.0000 15.000 1.0752 0.14385 0.14127 -0.0754 0.0011 1.0000 15.500 1.0545 0.16121 0.15879 -0.0861 0.0011 1.0000 16.000 1.0420 0.17769 0.17535 -0.0955 0.0012 1.0000