XFOIL Version 6.94 Calculated polar for: agora06 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.026 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0949 0.03883 0.02760 -0.0150 0.9988 0.4791 -2.750 -0.0766 0.03899 0.02795 -0.0133 0.9988 0.5135 -2.500 -0.0613 0.03899 0.02836 -0.0105 0.9988 0.5604 -2.250 -0.0435 0.03895 0.02858 -0.0084 0.9988 0.5967 -2.000 -0.0168 0.03903 0.02857 -0.0091 0.9988 0.6058 -1.750 0.0115 0.03917 0.02860 -0.0103 0.9988 0.6126 -1.500 0.0378 0.03920 0.02876 -0.0107 0.9988 0.6219 -1.250 0.0648 0.03932 0.02901 -0.0114 0.9988 0.6364 -1.000 0.0909 0.03936 0.02937 -0.0118 0.9988 0.6570 -0.750 0.1157 0.03931 0.02985 -0.0118 0.9988 0.6948 -0.500 0.3102 0.03802 0.02874 -0.0283 0.7772 1.0012 -0.250 0.3061 0.04204 0.03245 -0.0171 0.6966 1.0012 0.000 0.3125 0.04618 0.03634 -0.0105 0.6041 1.0012 0.250 0.3509 0.04744 0.03749 -0.0149 0.5189 1.0012 0.500 0.3879 0.04575 0.03554 -0.0155 0.4726 1.0012 0.750 0.4098 0.04448 0.03389 -0.0101 0.4540 1.0012 1.000 0.4312 0.04343 0.03243 -0.0050 0.4362 1.0012 1.250 0.4502 0.04236 0.03084 0.0008 0.4201 1.0012 1.500 0.4739 0.04266 0.03071 0.0037 0.4037 1.0012 1.750 0.4996 0.04358 0.03130 0.0055 0.3920 1.0012 2.000 0.5291 0.04532 0.03289 0.0053 0.3797 1.0012 2.250 0.5572 0.04718 0.03459 0.0054 0.3678 1.0012 2.500 0.5862 0.04952 0.03687 0.0047 0.3586 1.0012 2.750 0.6216 0.05345 0.04110 -0.0001 0.3534 1.0012 3.000 0.6551 0.05838 0.04632 -0.0061 0.3485 1.0012 3.250 0.6835 0.06297 0.05104 -0.0103 0.3434 1.0012 3.750 0.6910 0.09149 0.08043 -0.0472 0.3724 1.0012 4.000 0.6721 0.10138 0.09036 -0.0564 0.3967 1.0012 4.250 0.6448 0.10955 0.09852 -0.0617 0.4283 1.0012