XFOIL Version 6.94 Calculated polar for: agora10 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 -0.0016 0.03238 0.02219 -0.0238 0.9316 0.3726 -2.750 0.0238 0.03315 0.02308 -0.0226 0.9038 0.4038 -2.500 0.0406 0.03402 0.02422 -0.0193 0.8758 0.4407 -2.250 0.0558 0.03494 0.02533 -0.0156 0.8497 0.4753 -2.000 0.0747 0.03596 0.02621 -0.0126 0.8218 0.4832 -1.750 0.0955 0.03698 0.02709 -0.0101 0.7957 0.4852 -1.500 0.1169 0.03806 0.02808 -0.0077 0.7678 0.4881 -1.250 0.1378 0.03930 0.02931 -0.0053 0.7314 0.4919 -1.000 0.1534 0.04129 0.03139 -0.0013 0.6854 0.4972 -0.750 0.1653 0.04523 0.03539 0.0033 0.6159 0.5040 -0.500 0.1834 0.05085 0.04103 0.0046 0.5274 0.5140 -0.250 0.2067 0.05459 0.04490 0.0043 0.4625 0.5288 0.000 0.2288 0.05656 0.04705 0.0053 0.4215 0.5541 0.250 0.2629 0.05631 0.04771 0.0025 0.3941 0.6338 0.500 0.3021 0.05558 0.04746 -0.0001 0.3750 1.0012 0.750 0.3357 0.05577 0.04721 -0.0007 0.3583 1.0012 1.000 0.3669 0.05655 0.04767 -0.0009 0.3457 1.0012 1.250 0.3963 0.05760 0.04846 -0.0009 0.3364 1.0012 1.500 0.4272 0.05913 0.04981 -0.0023 0.3257 1.0012 1.750 0.4563 0.06063 0.05116 -0.0031 0.3159 1.0012 2.000 0.4892 0.06456 0.05505 -0.0078 0.3107 1.0012 2.250 0.5194 0.06858 0.05906 -0.0122 0.3066 1.0012 2.500 0.5488 0.07422 0.06476 -0.0197 0.3030 1.0012 2.750 0.5730 0.08018 0.07076 -0.0269 0.3032 1.0012 3.000 0.5869 0.08924 0.07997 -0.0391 0.3172 1.0012 3.250 0.5870 0.09844 0.08927 -0.0494 0.3438 1.0012 4.000 0.5661 0.11786 0.10869 -0.0634 0.4120 1.0012 4.250 0.5642 0.12098 0.11174 -0.0644 0.4140 1.0012 4.500 0.5679 0.12451 0.11521 -0.0654 0.4154 1.0012 4.750 0.5736 0.12818 0.11881 -0.0664 0.4166 1.0012 5.000 0.5796 0.13176 0.12234 -0.0674 0.4168 1.0012