XFOIL Version 6.94 Calculated polar for: manu4/30 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.035 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1241 0.06237 0.05423 0.0093 0.4014 0.7777 -2.750 -0.1455 0.06011 0.05200 0.0152 0.4004 0.7798 -2.000 -0.1511 0.05252 0.04444 0.0210 0.3917 0.7742 -1.750 -0.1433 0.05090 0.04267 0.0186 0.3881 0.7499 -1.500 0.1045 0.05570 0.04475 -0.0500 0.3731 0.3319 -1.250 0.1643 0.05362 0.04209 -0.0576 0.3705 0.2971 -1.000 0.2170 0.05186 0.04001 -0.0631 0.3685 0.2920 -0.750 0.2742 0.05041 0.03815 -0.0693 0.3666 0.2876 -0.500 0.3321 0.04921 0.03652 -0.0755 0.3653 0.2853 -0.250 0.3913 0.04830 0.03535 -0.0819 0.3645 0.2895 0.000 0.4575 0.04792 0.03450 -0.0898 0.3642 0.3011 0.250 0.5235 0.04764 0.03404 -0.0979 0.3645 0.3274 0.500 0.6042 0.04723 0.03356 -0.1091 0.3656 0.3811 0.750 0.6824 0.04664 0.03344 -0.1198 0.3672 0.5159 1.000 0.7228 0.04706 0.03431 -0.1226 0.3691 0.5982 1.250 0.8290 0.04827 0.03660 -0.1406 0.3706 1.0001 1.500 0.8611 0.05006 0.03823 -0.1420 0.3718 1.0001 1.750 0.8897 0.05186 0.03989 -0.1426 0.3724 1.0001 2.000 0.9159 0.05381 0.04170 -0.1428 0.3734 1.0001 2.250 0.9403 0.05593 0.04368 -0.1426 0.3744 1.0001 2.500 0.9512 0.05650 0.04460 -0.1396 0.3796 1.0001 2.750 0.9618 0.05838 0.04667 -0.1372 0.3864 1.0001 3.000 0.9745 0.06068 0.04902 -0.1353 0.3922 1.0001 3.250 0.9927 0.06335 0.05161 -0.1345 0.3973 1.0001 3.500 0.9966 0.06518 0.05371 -0.1314 0.4059 1.0001 3.750 0.9836 0.06832 0.05715 -0.1264 0.4177 1.0001 4.000 1.0088 0.07182 0.06049 -0.1273 0.4251 1.0001 4.250 0.9444 0.07632 0.06561 -0.1169 0.4464 1.0001 4.500 0.8746 0.08387 0.07349 -0.1089 0.4753 1.0001