XFOIL Version 6.94 Calculated polar for: SA7036 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 0.3618 0.00695 0.00248 -0.0749 0.8185 1.0000 0.500 0.4099 0.00701 0.00231 -0.0729 0.7860 1.0000 1.000 0.4599 0.00714 0.00221 -0.0713 0.7503 1.0000 1.500 0.5108 0.00736 0.00220 -0.0700 0.7130 1.0000 2.000 0.5622 0.00762 0.00227 -0.0688 0.6732 1.0000 2.500 0.6134 0.00796 0.00239 -0.0677 0.6308 1.0000 3.000 0.6646 0.00834 0.00260 -0.0666 0.5880 1.0000 3.500 0.7158 0.00876 0.00286 -0.0656 0.5446 1.0000 4.000 0.7666 0.00923 0.00321 -0.0646 0.5008 1.0000 4.500 0.8167 0.00976 0.00359 -0.0636 0.4547 1.0000 5.000 0.8664 0.01035 0.00405 -0.0625 0.4096 1.0000 5.500 0.9132 0.01112 0.00458 -0.0611 0.3417 1.0000 6.000 0.9588 0.01205 0.00517 -0.0597 0.2711 1.0000 6.500 1.0042 0.01307 0.00589 -0.0583 0.2057 1.0000 7.000 1.0497 0.01410 0.00677 -0.0569 0.1581 1.0000 7.500 1.0917 0.01548 0.00794 -0.0550 0.1015 1.0000 8.000 1.1281 0.01737 0.00947 -0.0526 0.0396 1.0000 8.500 1.1611 0.01954 0.01153 -0.0493 0.0100 1.0000 9.000 1.1953 0.02141 0.01370 -0.0461 0.0083 1.0000 9.500 1.2181 0.02389 0.01656 -0.0415 0.0077 1.0000 10.000 1.2235 0.02693 0.01994 -0.0347 0.0076 1.0000 10.500 1.2268 0.03046 0.02380 -0.0289 0.0076 1.0000 11.500 1.2328 0.03947 0.03347 -0.0208 0.0076 1.0000 12.000 1.2338 0.04480 0.03916 -0.0184 0.0077 1.0000 12.500 1.2315 0.05075 0.04546 -0.0173 0.0078 1.0000 13.000 1.2231 0.05792 0.05301 -0.0177 0.0080 1.0000 13.500 1.2103 0.06633 0.06179 -0.0198 0.0081 1.0000 14.000 1.1912 0.07661 0.07245 -0.0240 0.0083 1.0000 14.500 1.1662 0.08915 0.08538 -0.0306 0.0085 1.0000 15.000 1.1359 0.10431 0.10091 -0.0395 0.0087 1.0000 15.500 1.0994 0.12275 0.11971 -0.0510 0.0089 1.0000 16.000 1.0566 0.14519 0.14247 -0.0650 0.0091 1.0000 16.500 0.9734 0.18793 0.18526 -0.0870 0.0110 1.0000 17.000 0.9689 0.20335 0.20062 -0.0945 0.0126 1.0000