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 = 0.040 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 0.2885 0.04722 0.04272 -0.0421 0.7830 0.8907 -2.750 0.2410 0.04744 0.04302 -0.0341 0.7716 0.8318 -2.500 0.1894 0.04777 0.04348 -0.0252 0.7596 0.7962 -2.250 0.1407 0.04726 0.04308 -0.0174 0.7471 0.7612 -1.750 0.0419 0.04588 0.04190 -0.0015 0.7183 0.7140 -1.250 0.1400 0.04177 0.03552 -0.0383 0.5779 0.3656 -1.000 0.1986 0.03955 0.03183 -0.0431 0.5039 0.2885 -0.750 0.2332 0.03904 0.03017 -0.0434 0.4721 0.2545 -0.500 0.2631 0.03882 0.02903 -0.0427 0.4523 0.2344 -0.250 0.2917 0.03741 0.02721 -0.0426 0.4369 0.2262 0.000 0.3269 0.03744 0.02641 -0.0426 0.4237 0.2111 0.250 0.3640 0.03642 0.02506 -0.0439 0.4129 0.2048 0.500 0.4083 0.03593 0.02406 -0.0460 0.4039 0.1981 0.750 0.4706 0.03578 0.02325 -0.0514 0.3963 0.1944 1.000 0.5308 0.03564 0.02286 -0.0569 0.3896 0.1998 1.250 0.5898 0.03575 0.02263 -0.0619 0.3834 0.2076 1.500 0.6417 0.03595 0.02267 -0.0659 0.3778 0.2167 1.750 0.6746 0.03629 0.02311 -0.0666 0.3733 0.2327 2.000 0.7083 0.03651 0.02362 -0.0677 0.3688 0.2684 2.250 0.8215 0.03658 0.02452 -0.0832 0.3625 1.0000 2.500 0.8457 0.03800 0.02566 -0.0825 0.3609 1.0000 2.750 0.8682 0.03957 0.02707 -0.0817 0.3597 1.0000 3.000 0.8875 0.04123 0.02866 -0.0806 0.3590 1.0000 3.250 0.9031 0.04293 0.03039 -0.0789 0.3587 1.0000 3.500 0.9160 0.04477 0.03230 -0.0769 0.3584 1.0000 3.750 0.9262 0.04677 0.03439 -0.0747 0.3581 1.0000 4.000 0.9325 0.04897 0.03671 -0.0721 0.3576 1.0000 4.250 0.9356 0.05140 0.03927 -0.0694 0.3573 1.0000 4.500 0.9358 0.05412 0.04211 -0.0665 0.3574 1.0000 4.750 0.9332 0.05717 0.04527 -0.0636 0.3580 1.0000