XFOIL Version 6.94 Calculated polar for: tilt 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.008 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1463 0.05325 0.03974 -0.0066 0.9998 0.7314 -2.750 -0.0758 0.04981 0.03697 -0.0169 0.9998 0.9376 -2.500 -0.0218 0.04886 0.03430 -0.0292 0.9998 1.0002 -2.250 0.0207 0.04930 0.03299 -0.0333 0.9998 1.0002 -2.000 0.0509 0.04978 0.03227 -0.0338 0.9998 1.0002 -1.750 0.0768 0.05027 0.03185 -0.0334 0.9998 1.0002 -1.500 0.1008 0.05078 0.03162 -0.0327 0.9998 1.0002 -1.250 0.1237 0.05132 0.03155 -0.0319 0.9998 1.0002 -1.000 0.1459 0.05190 0.03162 -0.0312 0.9998 1.0002 -0.750 0.1675 0.05253 0.03182 -0.0305 0.9998 1.0002 -0.500 0.1887 0.05320 0.03204 -0.0299 0.9998 1.0002 -0.250 0.2094 0.05394 0.03249 -0.0294 0.9998 1.0002 0.000 0.2297 0.05476 0.03307 -0.0289 0.9998 1.0002 0.250 0.2493 0.05565 0.03380 -0.0285 0.9998 1.0002 0.500 0.2681 0.05667 0.03472 -0.0282 0.9998 1.0002 0.750 0.2858 0.05784 0.03587 -0.0281 0.9998 1.0002 1.000 0.3016 0.05927 0.03730 -0.0282 0.9998 1.0002 1.250 0.3140 0.06114 0.03931 -0.0285 0.9998 1.0002 1.500 0.3196 0.06389 0.04228 -0.0294 0.9998 1.0002 1.750 0.3148 0.06797 0.04654 -0.0311 0.9998 1.0002 2.000 0.3092 0.07236 0.05095 -0.0331 0.9998 1.0002 2.250 0.3086 0.07624 0.05475 -0.0349 0.9998 1.0002 2.500 0.3116 0.07973 0.05812 -0.0365 0.9998 1.0002 2.750 0.3167 0.08299 0.06124 -0.0379 0.9998 1.0002 3.000 0.3231 0.08611 0.06421 -0.0393 0.9998 1.0002 3.250 0.3306 0.08913 0.06708 -0.0405 0.9998 1.0002 3.500 0.3387 0.09209 0.06989 -0.0418 0.9998 1.0002 3.750 0.3474 0.09502 0.07266 -0.0430 0.9998 1.0002 4.000 0.3565 0.09792 0.07537 -0.0441 0.9998 1.0002 4.250 0.3660 0.10080 0.07810 -0.0453 0.9998 1.0002 4.500 0.3757 0.10366 0.08081 -0.0465 0.9998 1.0002 4.750 0.3856 0.10651 0.08352 -0.0476 0.9998 1.0002 5.000 0.3957 0.10936 0.08623 -0.0488 0.9998 1.0002