XFOIL Version 6.94 Calculated polar for: GOE 504 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 ------- -------- --------- --------- -------- ------- ------- 0.500 0.5570 0.01039 0.00361 -0.0934 0.5638 0.4752 1.500 0.7106 0.00992 0.00443 -0.1025 0.5396 0.9451 2.000 0.7859 0.01034 0.00478 -0.1069 0.5274 0.9732 2.500 0.8698 0.01082 0.00511 -0.1133 0.5145 0.9889 4.500 0.9623 0.01215 0.00581 -0.0897 0.4108 1.0000 5.000 0.9832 0.01274 0.00629 -0.0837 0.3886 1.0000 5.500 0.9873 0.01402 0.00723 -0.0752 0.3353 1.0000 6.000 1.0054 0.01517 0.00820 -0.0695 0.3055 1.0000 6.500 1.0060 0.01729 0.00995 -0.0619 0.2457 1.0000 7.000 1.0131 0.01948 0.01188 -0.0560 0.2003 1.0000 8.000 0.9705 0.02855 0.02017 -0.0411 0.0046 1.0000 8.500 0.9957 0.03062 0.02231 -0.0391 0.0046 1.0000 9.000 1.0189 0.03295 0.02473 -0.0372 0.0046 1.0000 9.500 1.0397 0.03559 0.02748 -0.0354 0.0047 1.0000 10.000 1.0582 0.03855 0.03057 -0.0336 0.0050 1.0000 10.500 1.0748 0.04178 0.03393 -0.0321 0.0052 1.0000 11.000 1.0895 0.04534 0.03763 -0.0306 0.0055 1.0000 11.500 1.0997 0.04948 0.04194 -0.0293 0.0059 1.0000 12.000 1.1036 0.05443 0.04708 -0.0281 0.0063 1.0000 12.500 1.1026 0.06012 0.05295 -0.0271 0.0066 1.0000 13.000 1.1094 0.06517 0.05816 -0.0265 0.0073 1.0000 13.500 1.1033 0.07190 0.06508 -0.0262 0.0078 1.0000 14.000 1.0866 0.08023 0.07361 -0.0263 0.0082 1.0000 14.500 1.0883 0.08642 0.07997 -0.0266 0.0089 1.0000 15.000 1.0764 0.09460 0.08835 -0.0273 0.0096 1.0000 15.500 1.0579 0.10376 0.09767 -0.0284 0.0102 1.0000 16.000 1.0597 0.11030 0.10437 -0.0292 0.0120 1.0000 16.500 1.0642 0.11613 0.11033 -0.0297 0.0141 1.0000