XFOIL Version 6.94 Calculated polar for: GOE 278 (DAIMLER IX) AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.300 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 1.500 0.6339 0.01024 0.00276 -0.0877 0.6623 0.0584 2.000 0.6839 0.01043 0.00268 -0.0863 0.5988 0.0674 2.500 0.7333 0.00889 0.00280 -0.0852 0.5296 1.0000 3.000 0.7814 0.00959 0.00299 -0.0838 0.4595 1.0000 3.500 0.8306 0.01027 0.00332 -0.0827 0.4126 1.0000 4.000 0.8814 0.01084 0.00379 -0.0819 0.3835 1.0000 4.500 0.9287 0.01160 0.00416 -0.0806 0.3171 1.0000 5.000 0.9747 0.01260 0.00461 -0.0794 0.2137 1.0000 6.000 1.0564 0.01618 0.00711 -0.0755 0.0297 1.0000 6.500 1.1044 0.01695 0.00792 -0.0744 0.0233 1.0000 7.000 1.1522 0.01772 0.00880 -0.0733 0.0195 1.0000 7.500 1.1959 0.01889 0.00996 -0.0715 0.0048 1.0000 8.000 1.2386 0.02010 0.01139 -0.0696 0.0047 1.0000 8.500 1.2788 0.02143 0.01295 -0.0672 0.0049 1.0000 9.000 1.3158 0.02293 0.01468 -0.0645 0.0052 1.0000 9.500 1.3469 0.02475 0.01676 -0.0609 0.0056 1.0000 10.000 1.3686 0.02686 0.01914 -0.0560 0.0060 1.0000 10.500 1.3867 0.02917 0.02174 -0.0511 0.0065 1.0000 11.000 1.3919 0.03264 0.02552 -0.0457 0.0071 1.0000 11.500 1.3872 0.03741 0.03056 -0.0409 0.0075 1.0000 12.000 1.3878 0.04231 0.03579 -0.0377 0.0080 1.0000 12.500 1.3793 0.04888 0.04277 -0.0344 0.0088 1.0000 13.000 1.3701 0.05626 0.05049 -0.0315 0.0095 1.0000 14.000 1.3037 0.08058 0.07627 -0.0318 0.0112 1.0000 14.500 1.2677 0.09388 0.08999 -0.0373 0.0114 1.0000 15.000 1.2308 0.10899 0.10548 -0.0457 0.0114 1.0000 15.500 1.1934 0.12608 0.12291 -0.0563 0.0112 1.0000 16.000 1.1608 0.14383 0.14093 -0.0679 0.0108 1.0000 16.500 1.1380 0.16067 0.15794 -0.0789 0.0103 1.0000 17.000 1.1191 0.17810 0.17545 -0.0896 0.0100 1.0000