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 = 0.022 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1220 0.04002 0.02919 -0.0055 0.9988 0.5959 -2.750 -0.1106 0.03992 0.02941 -0.0012 0.9988 0.6456 -2.500 -0.0929 0.03974 0.02938 0.0010 0.9988 0.6753 -2.250 -0.0665 0.03960 0.02918 0.0002 0.9988 0.6822 -2.000 -0.0389 0.03952 0.02905 -0.0009 0.9988 0.6916 -1.750 -0.0105 0.03948 0.02902 -0.0023 0.9988 0.7057 -1.500 0.0155 0.03933 0.02907 -0.0027 0.9988 0.7275 -1.250 0.0402 0.03909 0.02920 -0.0027 0.9988 0.7661 -1.000 0.0631 0.03832 0.02907 -0.0025 0.9988 0.9290 -0.750 0.0995 0.03880 0.02911 -0.0078 0.9988 1.0012 -0.500 0.1336 0.03978 0.02968 -0.0111 0.9988 1.0012 0.000 0.3456 0.04889 0.03800 -0.0227 0.6528 1.0012 0.250 0.3627 0.05311 0.04200 -0.0208 0.5831 1.0012 0.500 0.3916 0.05516 0.04393 -0.0226 0.5370 1.0012 0.750 0.4322 0.05602 0.04480 -0.0303 0.4965 1.0012 1.000 0.4603 0.05652 0.04510 -0.0290 0.4774 1.0012 1.250 0.4869 0.05676 0.04512 -0.0266 0.4602 1.0012 1.500 0.5121 0.05653 0.04468 -0.0231 0.4442 1.0012 1.750 0.5384 0.05735 0.04533 -0.0215 0.4317 1.0012 2.000 0.5665 0.06034 0.04827 -0.0237 0.4226 1.0012 2.250 0.5927 0.06352 0.05141 -0.0258 0.4133 1.0012 2.500 0.6169 0.06748 0.05536 -0.0289 0.4052 1.0012 2.750 0.6364 0.07386 0.06180 -0.0354 0.4030 1.0012 3.000 0.6483 0.08119 0.06919 -0.0426 0.4071 1.0012 3.250 0.6495 0.08888 0.07692 -0.0498 0.4144 1.0012 3.500 0.6527 0.09497 0.08300 -0.0540 0.4204 1.0012 3.750 0.6450 0.10138 0.08939 -0.0583 0.4309 1.0012 4.000 0.6181 0.10788 0.09585 -0.0616 0.4497 1.0012 4.250 0.6040 0.11390 0.10182 -0.0642 0.4699 1.0012 4.750 0.5750 0.12957 0.11751 -0.0759 0.5805 1.0012 5.000 0.5677 0.13198 0.11987 -0.0767 0.5939 1.0012