XFOIL Version 6.94 Calculated polar for: GOA 3 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 -0.1494 0.02343 0.01245 -0.0317 0.9999 0.0916 -2.750 -0.1226 0.02234 0.01135 -0.0307 0.9999 0.1028 -2.500 -0.0951 0.02123 0.01036 -0.0299 0.9999 0.1289 -2.250 -0.0664 0.01954 0.00953 -0.0298 0.9999 0.2475 -2.000 -0.0491 0.01681 0.00931 -0.0259 0.9999 1.0001 -1.750 -0.0234 0.01690 0.00923 -0.0247 0.9999 1.0001 -1.500 0.0027 0.01700 0.00934 -0.0234 0.9999 1.0001 -1.250 0.1071 0.01942 0.00847 -0.0338 0.2346 1.0001 -1.000 0.1310 0.02023 0.00875 -0.0326 0.2134 1.0001 -0.750 0.1552 0.02103 0.00907 -0.0316 0.1951 1.0001 -0.500 0.1812 0.02177 0.00951 -0.0307 0.1856 1.0001 -0.250 0.2088 0.02259 0.01006 -0.0299 0.1809 1.0001 0.000 0.2387 0.02359 0.01078 -0.0295 0.1778 1.0001 0.250 0.2714 0.02458 0.01164 -0.0296 0.1765 1.0001 0.500 0.3042 0.02565 0.01265 -0.0296 0.1764 1.0001 0.750 0.3356 0.02691 0.01384 -0.0295 0.1771 1.0001 1.000 0.3658 0.02792 0.01504 -0.0291 0.1796 1.0001 1.250 0.3962 0.02922 0.01659 -0.0288 0.1839 1.0001 1.500 0.4260 0.03085 0.01836 -0.0287 0.1887 1.0001 1.750 0.4546 0.03292 0.02042 -0.0286 0.1929 1.0001 2.000 0.4865 0.03431 0.02242 -0.0285 0.2044 1.0001 2.250 0.5176 0.03626 0.02480 -0.0287 0.2172 1.0001 2.500 0.5494 0.03861 0.02763 -0.0293 0.2361 1.0001 2.750 0.5804 0.04171 0.03099 -0.0300 0.2589 1.0001 3.250 0.6322 0.05782 0.05031 -0.0886 0.6857 1.0001 3.500 0.5554 0.06289 0.05519 -0.0888 0.7844 1.0001 4.000 0.3757 0.06067 0.05265 -0.0606 0.9999 1.0001 4.250 0.3879 0.06352 0.05543 -0.0619 0.9999 1.0001 4.500 0.3999 0.06644 0.05829 -0.0632 0.9999 1.0001 4.750 0.4341 0.07110 0.06294 -0.0698 0.9831 1.0001 5.000 0.5258 0.07954 0.07145 -0.0854 0.8749 1.0001