XFOIL Version 6.94 Calculated polar for: GOE 328 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 0.4126 0.01472 0.00728 -0.0828 0.7210 0.0465 0.500 0.4693 0.01256 0.00488 -0.0825 0.6987 0.0483 1.000 0.5253 0.01188 0.00415 -0.0822 0.6741 0.0532 1.500 0.5811 0.01144 0.00363 -0.0820 0.6468 0.0587 2.000 0.6371 0.01106 0.00322 -0.0819 0.6157 0.0619 2.500 0.6905 0.00937 0.00310 -0.0820 0.5816 0.7222 3.500 0.7957 0.00955 0.00321 -0.0798 0.5026 1.0000 4.000 0.8463 0.01032 0.00347 -0.0790 0.4285 1.0000 4.500 0.8980 0.01104 0.00387 -0.0786 0.3681 1.0000 5.000 0.9497 0.01175 0.00426 -0.0781 0.3049 1.0000 5.500 0.9979 0.01295 0.00493 -0.0774 0.2203 1.0000 6.500 1.0841 0.01668 0.00760 -0.0747 0.0108 1.0000 7.000 1.1324 0.01766 0.00858 -0.0736 0.0070 1.0000 7.500 1.1796 0.01868 0.00975 -0.0724 0.0064 1.0000 8.000 1.2243 0.01989 0.01119 -0.0708 0.0062 1.0000 8.500 1.2649 0.02140 0.01304 -0.0687 0.0062 1.0000 9.000 1.2989 0.02334 0.01527 -0.0657 0.0062 1.0000 9.500 1.3239 0.02569 0.01789 -0.0617 0.0063 1.0000 10.000 1.3351 0.02844 0.02090 -0.0561 0.0064 1.0000 10.500 1.3418 0.03194 0.02466 -0.0514 0.0066 1.0000 11.000 1.3462 0.03623 0.02922 -0.0479 0.0068 1.0000 11.500 1.3491 0.04118 0.03446 -0.0451 0.0070 1.0000 12.000 1.3508 0.04666 0.04031 -0.0425 0.0073 1.0000 12.500 1.3510 0.05279 0.04682 -0.0397 0.0078 1.0000 13.000 1.3461 0.06010 0.05460 -0.0373 0.0083 1.0000 13.500 1.3297 0.06921 0.06421 -0.0366 0.0088 1.0000 14.000 1.3040 0.07990 0.07535 -0.0383 0.0092 1.0000 14.500 1.2723 0.09227 0.08815 -0.0425 0.0095 1.0000 15.000 1.2389 0.10635 0.10262 -0.0492 0.0096 1.0000 15.500 1.2051 0.12212 0.11873 -0.0582 0.0098 1.0000 16.000 1.1714 0.13968 0.13660 -0.0693 0.0098 1.0000 16.500 1.1380 0.15988 0.15706 -0.0827 0.0099 1.0000 17.000 1.0980 0.18781 0.18508 -0.0996 0.0102 1.0000