XFOIL Version 6.96 Calculated polar for: NACA 64-008A AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.100 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -8.500 -0.6995 0.08631 0.08172 -0.0046 1.0000 0.1120 -8.250 -0.5878 0.07394 0.06960 -0.0071 1.0000 0.1349 -8.000 -0.7019 0.07714 0.07254 -0.0076 1.0000 0.1206 -7.750 -0.7297 0.07381 0.06869 -0.0138 1.0000 0.1276 -7.500 -0.7029 0.06836 0.06362 -0.0108 1.0000 0.1352 -7.250 -0.7020 0.06404 0.05917 -0.0120 1.0000 0.1455 -7.000 -0.6965 0.06023 0.05526 -0.0122 1.0000 0.1587 -6.750 -0.6867 0.05681 0.05179 -0.0117 1.0000 0.1741 -6.500 -0.6806 0.05364 0.04847 -0.0113 1.0000 0.1983 -5.750 -0.6102 0.03502 0.02693 -0.0101 1.0000 0.0725 -5.500 -0.5886 0.03161 0.02314 -0.0090 1.0000 0.0710 -5.250 -0.5648 0.02830 0.01940 -0.0077 1.0000 0.0664 -5.000 -0.5390 0.02571 0.01623 -0.0062 1.0000 0.0633 -4.750 -0.5129 0.02356 0.01384 -0.0052 1.0000 0.0634 -4.500 -0.4873 0.02173 0.01192 -0.0044 1.0000 0.0659 -4.250 -0.4624 0.02054 0.01058 -0.0034 1.0000 0.0728 -4.000 -0.4389 0.01873 0.00888 -0.0023 1.0000 0.0773 -3.750 -0.4172 0.01748 0.00763 -0.0009 1.0000 0.0853 -3.500 -0.3977 0.01620 0.00644 0.0007 1.0000 0.1023 -3.250 -0.3865 0.01327 0.00499 0.0030 1.0000 0.3228 -3.000 -0.3875 0.01194 0.00559 0.0116 1.0000 0.7642 -2.750 -0.3727 0.01215 0.00585 0.0167 1.0000 0.8366 -2.500 -0.3452 0.01294 0.00658 0.0212 1.0000 0.9076 -2.250 -0.2475 0.01397 0.00706 0.0109 1.0000 0.9550 -2.000 -0.1923 0.01375 0.00659 0.0049 1.0000 0.9681 -1.750 -0.1395 0.01341 0.00605 -0.0008 1.0000 0.9794 -1.500 -0.0894 0.01303 0.00554 -0.0063 1.0000 0.9902 -1.250 -0.0418 0.01264 0.00505 -0.0114 1.0000 1.0000 -1.000 -0.0245 0.01233 0.00473 -0.0108 1.0000 1.0000 -0.750 -0.0092 0.01206 0.00447 -0.0097 1.0000 1.0000 -0.500 0.0023 0.01186 0.00428 -0.0080 1.0000 1.0000 -0.250 0.0060 0.01171 0.00418 -0.0048 1.0000 1.0000 0.000 0.0000 0.01166 0.00415 0.0000 1.0000 1.0000 0.250 -0.0059 0.01171 0.00418 0.0048 1.0000 1.0000 0.500 -0.0023 0.01185 0.00428 0.0080 1.0000 1.0000 0.750 0.0092 0.01206 0.00447 0.0097 1.0000 1.0000 1.000 0.0246 0.01232 0.00473 0.0108 1.0000 1.0000 1.250 0.0418 0.01264 0.00505 0.0114 1.0000 1.0000 1.500 0.0894 0.01303 0.00553 0.0063 0.9902 1.0000 1.750 0.1395 0.01341 0.00605 0.0009 0.9794 1.0000 2.000 0.1923 0.01374 0.00659 -0.0049 0.9681 1.0000 2.250 0.2474 0.01397 0.00706 -0.0109 0.9550 1.0000 2.500 0.3452 0.01294 0.00658 -0.0212 0.9077 1.0000 2.750 0.3726 0.01215 0.00585 -0.0167 0.8366 1.0000 3.000 0.3875 0.01194 0.00558 -0.0115 0.7642 1.0000 3.250 0.3864 0.01327 0.00499 -0.0030 0.3233 1.0000 3.500 0.3977 0.01620 0.00644 -0.0007 0.1023 1.0000 3.750 0.4172 0.01748 0.00763 0.0009 0.0853 1.0000 4.000 0.4388 0.01873 0.00888 0.0023 0.0773 1.0000 4.250 0.4623 0.02054 0.01058 0.0035 0.0728 1.0000 4.500 0.4872 0.02173 0.01191 0.0044 0.0659 1.0000 4.750 0.5128 0.02356 0.01384 0.0052 0.0634 1.0000 5.000 0.5389 0.02571 0.01623 0.0062 0.0633 1.0000 5.250 0.5647 0.02830 0.01939 0.0077 0.0664 1.0000 5.500 0.5886 0.03161 0.02314 0.0090 0.0710 1.0000 5.750 0.6102 0.03502 0.02693 0.0101 0.0725 1.0000 7.500 0.7033 0.06840 0.06365 0.0107 0.1350 1.0000 7.750 0.7296 0.07379 0.06868 0.0137 0.1276 1.0000 8.000 0.7025 0.07717 0.07257 0.0074 0.1205 1.0000 8.250 0.7272 0.08167 0.07687 0.0116 0.1137 1.0000 8.500 0.6996 0.08639 0.08181 0.0043 0.1119 1.0000 8.750 0.6830 0.09248 0.08785 -0.0039 0.1091 1.0000