XFOIL Version 6.94 Calculated polar for: GOE 406b 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.030 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 0.0336 0.06897 0.06435 0.0035 0.9874 0.9579 -2.750 -0.0094 0.06879 0.06431 0.0113 0.9707 0.9103 -2.500 -0.0438 0.06783 0.06346 0.0154 0.9517 0.8552 -2.250 -0.0850 0.06697 0.06270 0.0203 0.9346 0.8120 -2.000 -0.1331 0.06595 0.06180 0.0267 0.9189 0.7823 -1.750 -0.1816 0.06469 0.06064 0.0333 0.9029 0.7586 -1.500 -0.1994 0.06207 0.05801 0.0292 0.8787 0.6963 -1.250 0.0170 0.05417 0.04829 -0.0305 0.7710 0.3893 -1.000 0.1351 0.04722 0.04043 -0.0413 0.6842 0.3165 -0.750 0.1962 0.04404 0.03639 -0.0441 0.6173 0.2856 -0.500 0.2479 0.04195 0.03342 -0.0460 0.5708 0.2664 -0.250 0.2977 0.04064 0.03109 -0.0478 0.5408 0.2528 0.250 0.3904 0.03923 0.02811 -0.0522 0.4986 0.2443 0.500 0.4369 0.03898 0.02727 -0.0548 0.4819 0.2459 0.750 0.4835 0.03896 0.02680 -0.0576 0.4677 0.2481 1.000 0.5363 0.03897 0.02631 -0.0613 0.4561 0.2528 1.250 0.5795 0.03923 0.02655 -0.0640 0.4484 0.2645 1.500 0.6254 0.03935 0.02664 -0.0669 0.4419 0.2881 1.750 0.6608 0.03974 0.02726 -0.0683 0.4375 0.3243 2.000 0.7516 0.03994 0.02869 -0.0799 0.4296 1.0000 2.250 0.7732 0.04152 0.02972 -0.0785 0.4254 1.0000 2.500 0.7994 0.04297 0.03068 -0.0781 0.4209 1.0000 2.750 0.8119 0.04501 0.03259 -0.0763 0.4180 1.0000 3.000 0.8148 0.04753 0.03517 -0.0737 0.4164 1.0000 3.250 0.8154 0.05040 0.03810 -0.0712 0.4160 1.0000 3.500 0.8115 0.05373 0.04149 -0.0686 0.4166 1.0000 3.750 0.8020 0.05757 0.04541 -0.0658 0.4180 1.0000 4.500 0.5244 0.09188 0.08225 -0.0635 0.4735 1.0000