XFOIL Version 6.94 Calculated polar for: ARA-D 13% AIRFOIL 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 -0.2324 0.04310 0.03388 -0.0286 1.0000 0.2599 -2.750 -0.2114 0.04160 0.03224 -0.0291 1.0000 0.2679 -2.500 -0.1892 0.04054 0.03099 -0.0297 1.0000 0.2777 -2.250 -0.1688 0.04062 0.03081 -0.0300 1.0000 0.2979 -2.000 -0.1494 0.03920 0.02947 -0.0304 1.0000 0.3178 -1.750 -0.1274 0.03875 0.02863 -0.0310 1.0000 0.3303 -1.500 -0.1037 0.03855 0.02810 -0.0318 1.0000 0.3431 -1.250 -0.0795 0.03812 0.02749 -0.0328 1.0000 0.3601 -1.000 -0.0565 0.03811 0.02748 -0.0335 0.9994 0.3849 -0.750 0.0129 0.03750 0.02714 -0.0417 0.9802 0.4620 -0.500 0.0599 0.03569 0.02719 -0.0438 0.9627 0.6947 -0.250 0.1185 0.03610 0.02675 -0.0500 0.9378 1.0000 0.000 0.1740 0.03722 0.02735 -0.0558 0.9131 1.0000 0.250 0.2296 0.03838 0.02813 -0.0613 0.8878 1.0000 0.500 0.2860 0.03899 0.02848 -0.0662 0.8600 1.0000 0.750 0.3251 0.04013 0.02945 -0.0688 0.8379 1.0000 1.000 0.3698 0.04074 0.02992 -0.0715 0.8122 1.0000 1.250 0.4135 0.04146 0.03051 -0.0736 0.7861 1.0000 1.500 0.4705 0.04104 0.02997 -0.0758 0.7579 1.0000 1.750 0.4894 0.04220 0.03105 -0.0744 0.7312 1.0000 2.000 0.5346 0.04216 0.03092 -0.0750 0.7090 1.0000 2.250 0.5533 0.04371 0.03245 -0.0739 0.6870 1.0000 2.500 0.5798 0.04456 0.03324 -0.0731 0.6642 1.0000 2.750 0.6222 0.04456 0.03307 -0.0722 0.6463 1.0000 3.000 0.6243 0.04681 0.03535 -0.0705 0.6197 1.0000 3.250 0.6817 0.04462 0.03290 -0.0683 0.6013 1.0000 3.500 0.6767 0.04767 0.03601 -0.0668 0.5743 1.0000 3.750 0.7196 0.04653 0.03471 -0.0649 0.5571 1.0000 4.000 0.7313 0.04915 0.03731 -0.0640 0.5412 1.0000 4.250 0.7273 0.05301 0.04126 -0.0635 0.5236 1.0000 4.500 0.7634 0.05291 0.04106 -0.0622 0.5102 1.0000 4.750 0.8037 0.05298 0.04094 -0.0605 0.4991 1.0000 5.000 0.7576 0.06150 0.04974 -0.0613 0.4865 1.0000