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 = 0.012 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.3613 0.07139 0.05774 0.0200 1.0001 0.6156 -2.750 -0.3462 0.06884 0.05518 0.0191 1.0001 0.6250 -2.500 -0.3291 0.06632 0.05255 0.0170 1.0001 0.6376 -2.250 -0.3097 0.06408 0.05031 0.0161 1.0001 0.6537 -2.000 -0.2880 0.06189 0.04810 0.0145 1.0001 0.6728 -1.750 -0.2658 0.05990 0.04612 0.0136 1.0001 0.7003 -1.500 -0.2439 0.05805 0.04440 0.0139 1.0001 0.7403 -1.250 -0.2166 0.05618 0.04276 0.0140 1.0001 0.8037 -1.000 -0.0981 0.05099 0.03728 -0.0176 1.0001 0.9999 -0.750 -0.0537 0.05099 0.03506 -0.0272 1.0001 0.9999 -0.500 -0.0302 0.05170 0.03492 -0.0280 1.0001 0.9999 -0.250 -0.0091 0.05249 0.03511 -0.0283 1.0001 0.9999 0.000 0.0108 0.05335 0.03550 -0.0285 1.0001 0.9999 0.250 0.0301 0.05430 0.03605 -0.0286 1.0001 0.9999 0.500 0.0487 0.05533 0.03677 -0.0287 1.0001 0.9999 0.750 0.0667 0.05647 0.03765 -0.0289 1.0001 0.9999 1.000 0.0838 0.05773 0.03870 -0.0291 1.0001 0.9999 1.250 0.1000 0.05913 0.03992 -0.0294 1.0001 0.9999 1.500 0.1151 0.06070 0.04138 -0.0298 1.0001 0.9999 1.750 0.1288 0.06251 0.04310 -0.0304 1.0001 0.9999 2.000 0.1409 0.06459 0.04512 -0.0311 1.0001 0.9999 2.250 0.1510 0.06697 0.04746 -0.0320 1.0001 0.9999 2.500 0.1593 0.06964 0.05012 -0.0331 1.0001 0.9999 2.750 0.1666 0.07254 0.05299 -0.0343 1.0001 0.9999 3.000 0.1736 0.07554 0.05595 -0.0357 1.0001 0.9999 3.250 0.1811 0.07856 0.05892 -0.0371 1.0001 0.9999 3.500 0.1892 0.08155 0.06185 -0.0384 1.0001 0.9999 3.750 0.1980 0.08450 0.06473 -0.0398 1.0001 0.9999 4.000 0.2075 0.08741 0.06757 -0.0411 1.0001 0.9999 4.250 0.2176 0.09029 0.07037 -0.0423 1.0001 0.9999 4.500 0.2280 0.09314 0.07314 -0.0436 1.0001 0.9999 4.750 0.2389 0.09598 0.07590 -0.0448 1.0001 0.9999 5.000 0.2501 0.09880 0.07864 -0.0460 1.0001 0.9999