XFOIL Version 6.94 Calculated polar for: mta2 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.018 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.2926 0.05113 0.04072 -0.0211 1.0000 0.3149 -2.750 -0.2517 0.04820 0.03732 -0.0253 1.0000 0.2980 -2.500 -0.2116 0.04570 0.03432 -0.0285 1.0000 0.2853 -2.250 -0.1736 0.04346 0.03165 -0.0304 1.0000 0.2768 -2.000 -0.1359 0.04164 0.02942 -0.0318 1.0000 0.2728 -1.750 -0.1005 0.03990 0.02752 -0.0322 1.0000 0.2734 -1.500 -0.0659 0.03827 0.02596 -0.0322 1.0000 0.2769 -1.250 -0.0304 0.03682 0.02470 -0.0319 1.0000 0.2824 -1.000 0.0071 0.03549 0.02369 -0.0315 1.0000 0.2901 -0.750 0.0474 0.03389 0.02286 -0.0317 1.0000 0.3067 -0.500 0.1644 0.03153 0.02058 -0.0442 0.4504 0.4668 -0.250 0.2049 0.03164 0.02007 -0.0428 0.4181 1.0000 0.000 0.2445 0.03306 0.02004 -0.0427 0.3959 1.0000 0.250 0.2845 0.03435 0.02029 -0.0433 0.3801 1.0000 0.500 0.3258 0.03566 0.02085 -0.0446 0.3709 1.0000 0.750 0.3670 0.03701 0.02170 -0.0463 0.3651 1.0000 1.000 0.4075 0.03849 0.02282 -0.0482 0.3614 1.0000 1.250 0.4469 0.04006 0.02419 -0.0500 0.3597 1.0000 1.500 0.4845 0.04163 0.02573 -0.0516 0.3596 1.0000 1.750 0.5200 0.04323 0.02743 -0.0532 0.3613 1.0000 2.000 0.5535 0.04493 0.02931 -0.0545 0.3644 1.0000 2.250 0.5849 0.04682 0.03138 -0.0558 0.3687 1.0000 2.500 0.6148 0.04895 0.03364 -0.0569 0.3734 1.0000 2.750 0.6423 0.05106 0.03600 -0.0581 0.3795 1.0000 3.000 0.6675 0.05341 0.03863 -0.0595 0.3879 1.0000 3.250 0.6934 0.05606 0.04144 -0.0609 0.3959 1.0000 3.500 0.7131 0.05881 0.04449 -0.0627 0.4075 1.0000 3.750 0.7304 0.06178 0.04772 -0.0648 0.4199 1.0000 4.000 0.7481 0.06509 0.05118 -0.0669 0.4325 1.0000 4.250 0.7624 0.06875 0.05498 -0.0692 0.4471 1.0000 4.500 0.7685 0.07262 0.05899 -0.0721 0.4649 1.0000 4.750 0.7559 0.07679 0.06332 -0.0755 0.4884 1.0000 5.000 0.7450 0.08131 0.06787 -0.0789 0.5143 1.0000