XFOIL Version 6.94 Calculated polar for: lb3000 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 -0.2572 0.04656 0.03609 -0.0373 1.0001 0.2339 -2.750 -0.2312 0.04556 0.03482 -0.0379 1.0001 0.2343 -2.500 -0.2052 0.04480 0.03377 -0.0385 1.0001 0.2355 -2.250 -0.1805 0.04405 0.03293 -0.0388 1.0001 0.2392 -2.000 -0.1569 0.04362 0.03252 -0.0390 1.0001 0.2453 -1.750 -0.1327 0.04339 0.03218 -0.0391 1.0001 0.2514 -1.500 -0.1077 0.04328 0.03185 -0.0393 1.0001 0.2566 -1.250 -0.0836 0.04309 0.03169 -0.0394 1.0001 0.2624 -1.000 -0.0595 0.04318 0.03178 -0.0394 1.0001 0.2710 -0.750 -0.0347 0.04340 0.03197 -0.0397 1.0001 0.2839 -0.250 0.0564 0.04303 0.03217 -0.0476 0.9863 0.3558 0.000 0.1104 0.03976 0.03111 -0.0498 0.9680 0.9999 0.250 0.1863 0.03998 0.03069 -0.0588 0.9467 0.9999 0.500 0.2500 0.03989 0.03028 -0.0654 0.9240 0.9999 0.750 0.3110 0.03953 0.02972 -0.0711 0.8991 0.9999 1.000 0.3831 0.03858 0.02861 -0.0780 0.8738 0.9999 1.250 0.4588 0.03706 0.02696 -0.0843 0.8477 0.9999 1.500 0.5115 0.03593 0.02575 -0.0857 0.8179 0.9999 1.750 0.5506 0.03494 0.02469 -0.0841 0.7837 0.9999 2.000 0.5790 0.03400 0.02366 -0.0803 0.7432 0.9999 2.250 0.6002 0.03300 0.02257 -0.0757 0.6906 0.9999 2.500 0.6249 0.03165 0.02102 -0.0719 0.6420 0.9999 2.750 0.6511 0.03064 0.01975 -0.0691 0.6000 0.9999 3.000 0.6788 0.03001 0.01872 -0.0670 0.5686 0.9999 3.250 0.7061 0.03017 0.01852 -0.0659 0.5398 0.9999 3.500 0.7338 0.03063 0.01865 -0.0653 0.5167 0.9999 3.750 0.7618 0.03133 0.01906 -0.0650 0.4977 0.9999 4.000 0.7903 0.03214 0.01959 -0.0648 0.4823 0.9999 4.250 0.8187 0.03315 0.02049 -0.0650 0.4691 0.9999 4.500 0.8471 0.03433 0.02158 -0.0654 0.4586 0.9999 4.750 0.8756 0.03546 0.02261 -0.0657 0.4486 0.9999 5.000 0.9037 0.03675 0.02384 -0.0661 0.4399 0.9999