XFOIL Version 6.94 Calculated polar for: NACA 64A410 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.450 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.3501 0.00690 0.00129 -0.0817 0.7288 0.5868 0.500 0.3998 0.00682 0.00138 -0.0805 0.6952 0.6829 1.000 0.4496 0.00695 0.00145 -0.0793 0.6472 0.7274 1.500 0.4953 0.00740 0.00154 -0.0774 0.5558 0.7652 2.500 0.5681 0.01030 0.00269 -0.0713 0.1180 0.8874 3.000 0.6366 0.01054 0.00301 -0.0745 0.1100 1.0000 3.500 0.6882 0.01089 0.00341 -0.0742 0.1029 1.0000 4.000 0.7403 0.01117 0.00372 -0.0740 0.0935 1.0000 4.500 0.7827 0.01256 0.00489 -0.0719 0.0223 1.0000 5.000 0.8292 0.01354 0.00606 -0.0703 0.0205 1.0000 5.500 0.8712 0.01500 0.00769 -0.0680 0.0200 1.0000 6.000 0.9095 0.01696 0.00984 -0.0651 0.0200 1.0000 6.500 0.9532 0.01818 0.01138 -0.0635 0.0131 1.0000 7.000 0.9891 0.02152 0.01495 -0.0604 0.0098 1.0000 7.500 1.0340 0.02217 0.01562 -0.0594 0.0086 1.0000 8.000 1.0761 0.02340 0.01691 -0.0580 0.0077 1.0000 8.500 1.1154 0.02474 0.01829 -0.0563 0.0070 1.0000 9.000 1.1409 0.03095 0.02501 -0.0524 0.0060 1.0000 10.500 1.1640 0.04647 0.04189 -0.0360 0.0061 1.0000 11.000 1.1565 0.05133 0.04715 -0.0307 0.0065 1.0000 11.500 -0.0206 0.00000 -0.00023 -0.0006 0.0472 0.2893 12.000 -0.0176 0.00000 -0.00021 -0.0004 0.0470 0.2859 12.500 -0.0146 0.00000 -0.00020 -0.0003 0.0468 0.2850 13.000 -0.0115 0.00000 -0.00018 -0.0001 0.0464 0.2884 13.500 -0.0084 0.00000 -0.00015 0.0000 0.0457 0.3108 14.000 -0.0052 0.00000 -0.00013 0.0002 0.0437 0.3412 15.500 -0.0001 0.00000 -0.00003 0.0001 0.0105 0.2411