XFOIL Version 6.94 Calculated polar for: GOE 431 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.450 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.8891 0.01000 0.00357 -0.2058 0.7439 0.1928 0.500 0.9400 0.00975 0.00335 -0.2044 0.7251 0.1974 1.000 0.9876 0.00963 0.00322 -0.2024 0.6942 0.2024 1.500 0.9917 0.01088 0.00352 -0.1915 0.5141 0.2061 2.000 1.0131 0.01192 0.00418 -0.1848 0.4459 0.2130 2.500 1.0440 0.01251 0.00466 -0.1799 0.4169 0.2256 3.000 1.0818 0.01295 0.00516 -0.1765 0.4001 0.2634 3.500 1.1137 0.01354 0.00577 -0.1723 0.3684 0.3518 4.000 1.1445 0.01421 0.00653 -0.1679 0.3262 0.4663 4.500 1.1801 0.01486 0.00719 -0.1644 0.3016 0.5401 5.000 1.2159 0.01552 0.00785 -0.1610 0.2788 0.5903 5.500 1.2387 0.01679 0.00893 -0.1559 0.1990 0.6786 7.000 1.3057 0.02256 0.01407 -0.1415 0.0046 1.0000 7.500 1.3408 0.02378 0.01537 -0.1387 0.0052 1.0000 8.000 1.3740 0.02513 0.01682 -0.1358 0.0058 1.0000 8.500 1.4056 0.02668 0.01849 -0.1327 0.0066 1.0000 9.000 1.4346 0.02845 0.02041 -0.1294 0.0074 1.0000 9.500 1.4572 0.03082 0.02296 -0.1256 0.0076 1.0000 10.000 1.4719 0.03393 0.02628 -0.1214 0.0078 1.0000 10.500 1.4806 0.03771 0.03025 -0.1170 0.0081 1.0000 11.000 1.4888 0.04175 0.03449 -0.1131 0.0084 1.0000 11.500 1.4895 0.04683 0.03980 -0.1093 0.0089 1.0000 12.000 1.4813 0.05332 0.04652 -0.1059 0.0095 1.0000 12.500 1.4685 0.06099 0.05437 -0.1034 0.0099 1.0000 13.000 1.4749 0.06690 0.06053 -0.1017 0.0111 1.0000 13.500 1.5049 0.07054 0.06408 -0.0988 0.0125 1.0000 14.500 1.5966 0.07762 0.07161 -0.0956 0.0122 1.0000 15.000 1.5819 0.08571 0.08019 -0.0947 0.0117 1.0000