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 = 0.045 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 0.3843 0.04205 0.03619 -0.0568 0.5266 0.9298 -2.750 0.3295 0.04291 0.03734 -0.0454 0.5362 0.8771 -2.250 0.2224 0.04335 0.03830 -0.0255 0.5472 0.8019 -2.000 0.1695 0.04281 0.03801 -0.0161 0.5506 0.7751 -1.750 0.1204 0.04182 0.03716 -0.0073 0.5434 0.7501 -1.250 0.1205 0.03949 0.03336 -0.0273 0.4808 0.4977 -1.000 0.1918 0.03926 0.03111 -0.0409 0.4434 0.3196 -0.750 0.2297 0.03889 0.02964 -0.0422 0.4260 0.2555 -0.500 0.2597 0.03789 0.02803 -0.0420 0.4129 0.2300 -0.250 0.2937 0.03769 0.02701 -0.0419 0.4014 0.2108 0.000 0.3274 0.03638 0.02546 -0.0426 0.3922 0.2035 0.250 0.3701 0.03625 0.02458 -0.0439 0.3848 0.1901 0.500 0.4208 0.03541 0.02334 -0.0474 0.3783 0.1845 0.750 0.4668 0.03493 0.02262 -0.0500 0.3731 0.1804 1.000 0.5302 0.03482 0.02207 -0.0559 0.3667 0.1800 1.250 0.5928 0.03512 0.02187 -0.0617 0.3603 0.1879 1.500 0.6395 0.03517 0.02195 -0.0649 0.3554 0.1959 1.750 0.6745 0.03546 0.02228 -0.0659 0.3512 0.2066 2.000 0.7086 0.03580 0.02273 -0.0669 0.3482 0.2281 2.250 0.8293 0.03554 0.02372 -0.0844 0.3434 1.0000 2.500 0.8525 0.03684 0.02477 -0.0835 0.3421 1.0000 2.750 0.8742 0.03824 0.02605 -0.0825 0.3411 1.0000 3.000 0.8947 0.03976 0.02749 -0.0814 0.3402 1.0000 3.250 0.9143 0.04146 0.02911 -0.0802 0.3391 1.0000 3.500 0.9314 0.04333 0.03094 -0.0788 0.3380 1.0000 3.750 0.9428 0.04503 0.03275 -0.0765 0.3373 1.0000 4.000 0.9514 0.04689 0.03475 -0.0739 0.3366 1.0000 4.250 0.9587 0.04896 0.03695 -0.0714 0.3362 1.0000 4.500 0.9669 0.05126 0.03933 -0.0691 0.3366 1.0000 4.750 0.9759 0.05380 0.04193 -0.0672 0.3375 1.0000 5.000 0.9465 0.05717 0.04581 -0.0611 0.3413 1.0000