XFOIL Version 6.94 Calculated polar for: nacc632 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.012 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0237 0.03765 0.01720 -0.0228 0.9994 1.0006 -2.750 -0.0029 0.03790 0.01715 -0.0224 0.9994 1.0006 -2.500 0.0177 0.03821 0.01718 -0.0220 0.9994 1.0006 -2.250 0.0382 0.03860 0.01733 -0.0216 0.9994 1.0006 -2.000 0.0585 0.03904 0.01757 -0.0211 0.9994 1.0006 -1.750 0.0787 0.03955 0.01789 -0.0207 0.9994 1.0006 -1.500 0.0988 0.04012 0.01829 -0.0202 0.9994 1.0006 -1.250 0.1187 0.04075 0.01879 -0.0198 0.9994 1.0006 -1.000 0.1386 0.04144 0.01937 -0.0195 0.9994 1.0006 -0.750 0.1584 0.04220 0.02004 -0.0192 0.9994 1.0006 -0.500 0.1781 0.04301 0.02079 -0.0189 0.9994 1.0006 -0.250 0.1976 0.04390 0.02165 -0.0188 0.9994 1.0006 0.000 0.2169 0.04485 0.02260 -0.0186 0.9994 1.0006 0.250 0.2358 0.04588 0.02364 -0.0186 0.9994 1.0006 0.500 0.2543 0.04699 0.02481 -0.0186 0.9994 1.0006 0.750 0.2723 0.04822 0.02612 -0.0187 0.9994 1.0006 1.000 0.2897 0.04957 0.02758 -0.0189 0.9994 1.0006 1.250 0.3061 0.05108 0.02924 -0.0192 0.9994 1.0006 1.500 0.3212 0.05281 0.03114 -0.0197 0.9994 1.0006 1.750 0.3344 0.05485 0.03338 -0.0205 0.9994 1.0006 2.000 0.3448 0.05733 0.03608 -0.0216 0.9994 1.0006 2.250 0.3514 0.06041 0.03937 -0.0233 0.9994 1.0006 2.500 0.3549 0.06405 0.04319 -0.0254 0.9994 1.0006 2.750 0.3580 0.06795 0.04721 -0.0278 0.9994 1.0006 3.000 0.3620 0.07180 0.05117 -0.0302 0.9994 1.0006 3.250 0.3672 0.07554 0.05500 -0.0325 0.9994 1.0006 3.500 0.3734 0.07920 0.05873 -0.0346 0.9994 1.0006 3.750 0.3803 0.08278 0.06239 -0.0366 0.9994 1.0006 4.000 0.3878 0.08632 0.06600 -0.0386 0.9994 1.0006 4.250 0.3959 0.08984 0.06958 -0.0405 0.9994 1.0006 4.500 0.4042 0.09333 0.07313 -0.0423 0.9994 1.0006 4.750 0.4446 0.09861 0.07858 -0.0513 0.9749 1.0006 5.000 0.4715 0.10318 0.08332 -0.0569 0.9566 1.0006