XFOIL Version 6.94 Calculated polar for: E174 (Dicke 8.92%) 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 0.4051 0.00802 0.00266 -0.0827 0.6859 1.0000 0.500 0.4603 0.00823 0.00267 -0.0825 0.6732 1.0000 1.000 0.5156 0.00846 0.00274 -0.0824 0.6615 1.0000 1.500 0.5710 0.00870 0.00284 -0.0822 0.6499 1.0000 2.000 0.6263 0.00891 0.00301 -0.0821 0.6373 1.0000 2.500 0.6814 0.00916 0.00323 -0.0819 0.6248 1.0000 3.000 0.7363 0.00940 0.00342 -0.0817 0.6112 1.0000 3.500 0.7909 0.00961 0.00363 -0.0814 0.5959 1.0000 4.000 0.8452 0.00980 0.00385 -0.0810 0.5795 1.0000 4.500 0.8991 0.00999 0.00413 -0.0805 0.5613 1.0000 5.000 0.9523 0.01014 0.00434 -0.0799 0.5380 1.0000 5.500 1.0036 0.01030 0.00453 -0.0789 0.5017 1.0000 6.000 1.0534 0.01066 0.00486 -0.0777 0.4500 1.0000 6.500 1.0960 0.01174 0.00554 -0.0758 0.3394 1.0000 7.000 1.1265 0.01426 0.00718 -0.0727 0.1772 1.0000 7.500 1.1487 0.01757 0.00956 -0.0687 0.0362 1.0000 8.000 1.1819 0.01959 0.01171 -0.0654 0.0232 1.0000 8.500 1.2068 0.02196 0.01427 -0.0613 0.0201 1.0000 9.000 1.2243 0.02431 0.01682 -0.0562 0.0186 1.0000 9.500 1.2364 0.02705 0.01977 -0.0509 0.0178 1.0000 10.000 1.2509 0.03029 0.02321 -0.0468 0.0173 1.0000 10.500 1.2712 0.03388 0.02706 -0.0434 0.0167 1.0000 11.000 1.2979 0.03811 0.03162 -0.0407 0.0166 1.0000 11.500 1.3223 0.04391 0.03795 -0.0380 0.0170 1.0000 12.000 1.3245 0.05116 0.04580 -0.0350 0.0178 1.0000 12.500 1.3093 0.05875 0.05391 -0.0329 0.0179 1.0000 13.000 1.2833 0.06785 0.06349 -0.0327 0.0182 1.0000 13.500 1.2513 0.07822 0.07428 -0.0349 0.0182 1.0000 14.000 1.2121 0.09099 0.08744 -0.0398 0.0186 1.0000 14.500 1.1687 0.10636 0.10315 -0.0477 0.0190 1.0000