XFOIL Version 6.94 Calculated polar for: Agora0 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 -0.1375 0.04274 0.03204 -0.0192 0.9992 0.4576 -2.750 -0.1121 0.04189 0.03123 -0.0193 0.9992 0.4719 -2.500 -0.0836 0.04139 0.03060 -0.0202 0.9992 0.4893 -2.250 -0.0579 0.04078 0.03016 -0.0201 0.9992 0.5072 -2.000 -0.0288 0.04040 0.02979 -0.0210 0.9992 0.5247 -1.750 -0.0005 0.04005 0.02959 -0.0215 0.9992 0.5445 -1.500 0.0277 0.03970 0.02956 -0.0218 0.9992 0.5717 -1.250 0.0550 0.03927 0.02969 -0.0219 0.9992 0.6159 -1.000 0.0748 0.03814 0.02986 -0.0193 0.9992 0.7425 -0.750 0.2765 0.03764 0.02823 -0.0368 0.7840 1.0008 -0.500 0.2717 0.04137 0.03177 -0.0254 0.7037 1.0008 -0.250 0.2748 0.04548 0.03568 -0.0174 0.6175 1.0008 0.000 0.3010 0.04731 0.03741 -0.0170 0.5402 1.0008 0.250 0.3501 0.04564 0.03571 -0.0228 0.4753 1.0008 0.500 0.3715 0.04438 0.03412 -0.0172 0.4568 1.0008 0.750 0.3920 0.04315 0.03249 -0.0115 0.4394 1.0008 1.000 0.4117 0.04220 0.03107 -0.0061 0.4224 1.0008 1.250 0.4347 0.04231 0.03075 -0.0026 0.4068 1.0008 1.500 0.4615 0.04327 0.03142 -0.0011 0.3946 1.0008 1.750 0.4891 0.04465 0.03255 -0.0001 0.3825 1.0008 2.000 0.5191 0.04661 0.03442 -0.0005 0.3701 1.0008 2.250 0.5488 0.04886 0.03662 -0.0012 0.3610 1.0008 2.500 0.5835 0.05236 0.04035 -0.0050 0.3557 1.0008 2.750 0.6163 0.05638 0.04455 -0.0091 0.3501 1.0008 3.000 0.6467 0.06083 0.04917 -0.0134 0.3446 1.0008 3.500 0.6714 0.08772 0.07699 -0.0494 0.3687 1.0008 3.750 0.6584 0.09803 0.08736 -0.0596 0.3910 1.0008 4.000 0.6392 0.10673 0.09608 -0.0666 0.4196 1.0008 4.500 0.5374 0.12555 0.11494 -0.0810 0.5694 1.0008 4.750 0.5520 0.13084 0.12017 -0.0834 0.5804 1.0008 5.000 0.5440 0.13255 0.12184 -0.0835 0.5867 1.0008