XFOIL Version 6.94 Calculated polar for: HSNLF(1)-0213 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 ------- -------- --------- --------- -------- ------- ------- 0.000 0.0710 0.01628 0.01131 -0.0016 0.7707 0.8714 0.500 0.1263 0.01620 0.01120 -0.0023 0.7643 0.8746 1.000 0.1818 0.01597 0.01091 -0.0025 0.7589 0.8771 1.500 0.2375 0.01593 0.01087 -0.0032 0.7522 0.8792 2.000 0.2948 0.01573 0.01069 -0.0043 0.7427 0.8810 2.500 0.3527 0.01512 0.01003 -0.0042 0.7354 0.8818 3.000 0.4107 0.01453 0.00950 -0.0048 0.7210 0.8823 3.500 0.4694 0.01366 0.00862 -0.0047 0.7076 0.8835 4.000 0.5294 0.01263 0.00755 -0.0048 0.6920 0.8839 4.500 0.5884 0.01187 0.00691 -0.0053 0.6693 0.8842 5.000 0.6467 0.01122 0.00636 -0.0058 0.6258 0.8845 5.500 0.6934 0.01160 0.00595 -0.0044 0.4581 0.8852 6.000 0.7260 0.01381 0.00716 -0.0027 0.2581 0.8868 6.500 0.7617 0.01584 0.00847 -0.0017 0.1303 0.8874 7.000 0.8009 0.01746 0.00980 -0.0008 0.0843 0.8881 7.500 0.8400 0.01896 0.01125 0.0002 0.0681 0.8888 9.500 0.9882 0.02571 0.01817 0.0051 0.0467 0.8935 10.000 1.0319 0.02774 0.02023 0.0061 0.0442 0.8945 10.500 1.0751 0.02973 0.02244 0.0070 0.0418 0.8959 11.000 1.1251 0.03225 0.02497 0.0071 0.0394 0.8972 11.500 1.1572 0.03511 0.02833 0.0087 0.0379 0.8993 12.000 1.1872 0.03852 0.03214 0.0103 0.0366 0.9019 12.500 1.2181 0.04167 0.03545 0.0114 0.0352 0.9050 13.000 1.2221 0.04672 0.04102 0.0144 0.0342 0.9107 13.500 1.2028 0.05277 0.04774 0.0179 0.0337 0.9219 14.000 1.1712 0.06011 0.05572 0.0205 0.0335 0.9996 14.500 1.1305 0.06990 0.06601 0.0202 0.0335 1.0000 15.000 1.0800 0.08209 0.07864 0.0167 0.0337 1.0000 15.500 1.0258 0.09751 0.09443 0.0090 0.0341 1.0000 16.000 0.9789 0.11517 0.11231 -0.0014 0.0345 1.0000