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 = 0.005 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0557 0.05418 0.03205 0.0003 0.9988 1.0012 -2.750 -0.0337 0.05431 0.03166 -0.0014 0.9988 1.0012 -2.500 -0.0130 0.05463 0.03152 -0.0028 0.9988 1.0012 -2.250 0.0058 0.05513 0.03159 -0.0037 0.9988 1.0012 -2.000 0.0226 0.05579 0.03187 -0.0042 0.9988 1.0012 -1.750 0.0380 0.05661 0.03233 -0.0044 0.9988 1.0012 -1.500 0.0526 0.05755 0.03294 -0.0046 0.9988 1.0012 -1.250 0.0673 0.05860 0.03368 -0.0048 0.9988 1.0012 -1.000 0.0818 0.05975 0.03456 -0.0050 0.9988 1.0012 -0.750 0.0963 0.06101 0.03559 -0.0054 0.9988 1.0012 -0.500 0.1108 0.06240 0.03678 -0.0059 0.9988 1.0012 -0.250 0.1250 0.06394 0.03815 -0.0066 0.9988 1.0012 0.000 0.1388 0.06563 0.03972 -0.0074 0.9988 1.0012 0.250 0.1519 0.06751 0.04149 -0.0085 0.9988 1.0012 0.500 0.1640 0.06958 0.04347 -0.0097 0.9988 1.0012 0.750 0.1752 0.07187 0.04568 -0.0112 0.9988 1.0012 1.000 0.1855 0.07435 0.04807 -0.0129 0.9988 1.0012 1.250 0.1952 0.07699 0.05060 -0.0147 0.9988 1.0012 1.500 0.2047 0.07973 0.05322 -0.0166 0.9988 1.0012 1.750 0.2142 0.08252 0.05589 -0.0185 0.9988 1.0012 2.000 0.2238 0.08535 0.05859 -0.0203 0.9988 1.0012 2.250 0.2336 0.08819 0.06130 -0.0222 0.9988 1.0012 2.500 0.2436 0.09104 0.06403 -0.0240 0.9988 1.0012 2.750 0.2539 0.09392 0.06679 -0.0258 0.9988 1.0012 3.000 0.2645 0.09683 0.06957 -0.0276 0.9988 1.0012 3.250 0.2752 0.09974 0.07237 -0.0293 0.9988 1.0012 3.500 0.2861 0.10266 0.07516 -0.0311 0.9988 1.0012 3.750 0.2971 0.10559 0.07798 -0.0328 0.9988 1.0012 4.000 0.3081 0.10853 0.08081 -0.0345 0.9988 1.0012 4.250 0.3193 0.11148 0.08365 -0.0362 0.9988 1.0012 4.500 0.3306 0.11443 0.08650 -0.0379 0.9988 1.0012 4.750 0.3420 0.11739 0.08937 -0.0396 0.9988 1.0012 5.000 0.3535 0.12036 0.09225 -0.0413 0.9988 1.0012