XFOIL Version 6.96 Calculated polar for: NACA 0008-34 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 1.000 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -8.750 -0.6414 0.07735 0.07582 -0.0176 1.0000 0.0052 -8.500 -0.6508 0.07219 0.07059 -0.0209 1.0000 0.0052 -8.250 -0.6588 0.06787 0.06620 -0.0215 1.0000 0.0052 -8.000 -0.6612 0.06339 0.06162 -0.0219 1.0000 0.0052 -7.750 -0.6612 0.05906 0.05717 -0.0217 1.0000 0.0052 -7.500 -0.6589 0.05489 0.05286 -0.0209 1.0000 0.0053 -7.250 -0.6549 0.05091 0.04872 -0.0196 1.0000 0.0053 -7.000 -0.6497 0.04713 0.04477 -0.0178 1.0000 0.0053 -6.750 -0.6442 0.04361 0.04108 -0.0154 1.0000 0.0053 -6.500 -0.6381 0.04030 0.03758 -0.0126 1.0000 0.0053 -5.750 -0.5868 0.02918 0.02565 -0.0096 0.9953 0.0053 -5.500 -0.5705 0.02353 0.01952 -0.0086 0.9920 0.0042 -5.250 -0.5444 0.02018 0.01580 -0.0080 0.9896 0.0037 -5.000 -0.5160 0.01706 0.01231 -0.0076 0.9879 0.0033 -4.750 -0.4868 0.01405 0.00892 -0.0071 0.9864 0.0028 -4.500 -0.4615 0.01181 0.00647 -0.0061 0.9830 0.0025 -4.250 -0.4357 0.01039 0.00486 -0.0055 0.9793 0.0024 -4.000 -0.4075 0.00952 0.00385 -0.0057 0.9765 0.0022 -3.750 -0.3814 0.00897 0.00319 -0.0054 0.9711 0.0022 -3.500 -0.3530 0.00855 0.00267 -0.0056 0.9669 0.0022 -3.250 -0.3255 0.00824 0.00228 -0.0056 0.9617 0.0023 -3.000 -0.2979 0.00801 0.00197 -0.0055 0.9556 0.0026 -2.750 -0.2704 0.00782 0.00178 -0.0055 0.9492 0.0047 -2.500 -0.2435 0.00767 0.00164 -0.0054 0.9415 0.0101 -2.250 -0.2164 0.00758 0.00144 -0.0054 0.9332 0.0048 -2.000 -0.1899 0.00743 0.00132 -0.0052 0.9238 0.0147 -1.750 -0.1706 0.00619 0.00099 -0.0041 0.9123 0.2852 -1.500 -0.1487 0.00546 0.00082 -0.0033 0.9003 0.4495 -1.250 -0.1266 0.00485 0.00070 -0.0024 0.8885 0.5927 -1.000 -0.1034 0.00447 0.00065 -0.0016 0.8768 0.6926 -0.750 -0.0795 0.00423 0.00063 -0.0008 0.8647 0.7638 -0.500 -0.0532 0.00417 0.00061 -0.0005 0.8522 0.7894 -0.250 -0.0267 0.00414 0.00060 -0.0002 0.8389 0.8084 0.000 0.0000 0.00413 0.00059 0.0000 0.8245 0.8245 0.250 0.0266 0.00414 0.00060 0.0002 0.8084 0.8390 0.500 0.0532 0.00417 0.00061 0.0005 0.7894 0.8522 0.750 0.0794 0.00423 0.00063 0.0008 0.7637 0.8647 1.000 0.1033 0.00447 0.00065 0.0016 0.6907 0.8768 1.250 0.1266 0.00484 0.00070 0.0024 0.5947 0.8886 1.500 0.1485 0.00547 0.00082 0.0033 0.4466 0.9004 1.750 0.1705 0.00620 0.00099 0.0041 0.2833 0.9123 2.000 0.1898 0.00743 0.00132 0.0052 0.0142 0.9239 2.250 0.2163 0.00758 0.00144 0.0054 0.0048 0.9333 2.500 0.2434 0.00767 0.00164 0.0055 0.0106 0.9417 2.750 0.2703 0.00782 0.00178 0.0055 0.0048 0.9494 3.000 0.2978 0.00800 0.00196 0.0056 0.0026 0.9557 3.250 0.3254 0.00824 0.00228 0.0056 0.0023 0.9618 3.500 0.3531 0.00855 0.00268 0.0056 0.0022 0.9670 3.750 0.3815 0.00897 0.00319 0.0054 0.0022 0.9713 4.000 0.4076 0.00953 0.00386 0.0057 0.0023 0.9765 4.250 0.4358 0.01039 0.00487 0.0055 0.0024 0.9793 4.500 0.4616 0.01182 0.00647 0.0061 0.0025 0.9830 4.750 0.4868 0.01404 0.00891 0.0071 0.0028 0.9864 5.000 0.5160 0.01705 0.01231 0.0076 0.0033 0.9879 5.250 0.5445 0.02018 0.01580 0.0080 0.0037 0.9896 5.500 0.5705 0.02355 0.01954 0.0086 0.0043 0.9920 5.750 0.5869 0.02917 0.02564 0.0095 0.0052 0.9953 6.000 0.6091 0.03309 0.02987 0.0097 0.0053 0.9974 6.250 0.6301 0.03707 0.03414 0.0098 0.0053 0.9997 6.500 0.6381 0.04031 0.03759 0.0125 0.0053 1.0000 6.750 0.6442 0.04362 0.04108 0.0154 0.0053 1.0000 7.000 0.6497 0.04713 0.04477 0.0178 0.0053 1.0000 7.250 0.6549 0.05091 0.04872 0.0196 0.0053 1.0000 7.500 0.6589 0.05490 0.05286 0.0209 0.0053 1.0000 7.750 0.6612 0.05906 0.05717 0.0217 0.0052 1.0000 8.000 0.6613 0.06339 0.06162 0.0219 0.0052 1.0000 8.250 0.6589 0.06788 0.06621 0.0215 0.0052 1.0000 8.500 0.6509 0.07222 0.07062 0.0208 0.0052 1.0000 8.750 0.6418 0.07738 0.07585 0.0175 0.0052 1.0000