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 = 0.050 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -6.500 -0.6776 0.05318 0.04526 -0.0137 1.0000 0.1538 -6.250 -0.6581 0.04768 0.03931 -0.0137 1.0000 0.1334 -6.000 -0.6385 0.04335 0.03430 -0.0129 1.0000 0.1207 -5.750 -0.6169 0.03981 0.02991 -0.0117 1.0000 0.1124 -5.500 -0.5944 0.03658 0.02617 -0.0107 1.0000 0.1114 -5.250 -0.5707 0.03411 0.02306 -0.0094 1.0000 0.1138 -5.000 -0.5447 0.03118 0.01984 -0.0085 1.0000 0.1154 -4.750 -0.5167 0.02860 0.01707 -0.0078 1.0000 0.1177 -4.500 -0.4873 0.02645 0.01478 -0.0069 1.0000 0.1233 -4.250 -0.4593 0.02459 0.01290 -0.0060 1.0000 0.1384 -4.000 -0.4333 0.02272 0.01108 -0.0048 1.0000 0.1595 -3.750 -0.2491 0.02188 0.01257 -0.0098 1.0000 0.9931 -3.500 -0.2151 0.02087 0.01115 -0.0131 1.0000 1.0000 -3.250 -0.1968 0.02014 0.01020 -0.0134 1.0000 1.0000 -3.000 -0.1783 0.01948 0.00935 -0.0135 1.0000 1.0000 -2.750 -0.1598 0.01888 0.00859 -0.0135 1.0000 1.0000 -2.500 -0.1413 0.01834 0.00791 -0.0134 1.0000 1.0000 -2.250 -0.1231 0.01785 0.00731 -0.0132 1.0000 1.0000 -2.000 -0.1050 0.01742 0.00674 -0.0128 1.0000 1.0000 -1.750 -0.0874 0.01703 0.00628 -0.0123 1.0000 1.0000 -1.500 -0.0704 0.01670 0.00589 -0.0115 1.0000 1.0000 -1.250 -0.0543 0.01641 0.00557 -0.0105 1.0000 1.0000 -1.000 -0.0395 0.01616 0.00531 -0.0093 1.0000 1.0000 -0.750 -0.0264 0.01597 0.00509 -0.0076 1.0000 1.0000 -0.500 -0.0155 0.01582 0.00495 -0.0055 1.0000 1.0000 -0.250 -0.0072 0.01573 0.00488 -0.0028 1.0000 1.0000 0.000 0.0000 0.01570 0.00485 0.0000 1.0000 1.0000 0.250 0.0072 0.01573 0.00488 0.0028 1.0000 1.0000 0.500 0.0156 0.01582 0.00495 0.0054 1.0000 1.0000 0.750 0.0264 0.01596 0.00509 0.0076 1.0000 1.0000 1.000 0.0395 0.01616 0.00530 0.0093 1.0000 1.0000 1.250 0.0544 0.01640 0.00556 0.0105 1.0000 1.0000 1.500 0.0705 0.01669 0.00589 0.0115 1.0000 1.0000 1.750 0.0875 0.01703 0.00628 0.0123 1.0000 1.0000 2.000 0.1051 0.01741 0.00674 0.0128 1.0000 1.0000 2.250 0.1231 0.01785 0.00730 0.0132 1.0000 1.0000 2.500 0.1414 0.01833 0.00790 0.0134 1.0000 1.0000 2.750 0.1599 0.01887 0.00858 0.0135 1.0000 1.0000 3.000 0.1785 0.01947 0.00934 0.0135 1.0000 1.0000 3.250 0.1970 0.02013 0.01019 0.0133 1.0000 1.0000 3.500 0.2153 0.02086 0.01114 0.0131 1.0000 1.0000 3.750 0.2495 0.02187 0.01256 0.0097 0.9930 1.0000 4.000 0.4333 0.02273 0.01108 0.0048 0.1595 1.0000 4.250 0.4592 0.02459 0.01290 0.0060 0.1384 1.0000 4.500 0.4872 0.02645 0.01478 0.0069 0.1233 1.0000 4.750 0.5166 0.02860 0.01707 0.0078 0.1177 1.0000 5.000 0.5447 0.03118 0.01983 0.0085 0.1153 1.0000 5.250 0.5707 0.03411 0.02306 0.0094 0.1138 1.0000 5.500 0.5944 0.03658 0.02616 0.0107 0.1114 1.0000 5.750 0.6169 0.03982 0.02992 0.0117 0.1124 1.0000 6.000 0.6385 0.04335 0.03430 0.0129 0.1207 1.0000 6.250 0.6582 0.04769 0.03932 0.0136 0.1334 1.0000 6.500 0.6778 0.05320 0.04528 0.0137 0.1539 1.0000 8.250 0.6249 0.10021 0.09354 -0.0334 0.3816 1.0000 8.500 0.6488 0.10514 0.09850 -0.0316 0.3606 1.0000