XFOIL Version 6.94 Calculated polar for: GOE 500 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 0.7239 0.00828 0.00237 -0.1680 0.7462 0.3439 0.500 0.7782 0.00826 0.00244 -0.1676 0.7290 0.4039 1.000 0.8320 0.00827 0.00258 -0.1671 0.7115 0.4795 1.500 0.8852 0.00826 0.00278 -0.1664 0.6943 0.5825 2.000 0.9358 0.00767 0.00291 -0.1650 0.6707 1.0000 2.500 0.9874 0.00795 0.00305 -0.1640 0.6437 1.0000 3.000 1.0387 0.00826 0.00324 -0.1629 0.6180 1.0000 3.500 1.0751 0.00908 0.00349 -0.1589 0.5185 1.0000 4.000 1.1079 0.01041 0.00415 -0.1547 0.4132 1.0000 4.500 1.1464 0.01151 0.00479 -0.1517 0.3300 1.0000 5.000 1.1396 0.01585 0.00731 -0.1418 0.0240 1.0000 5.500 1.1741 0.01692 0.00820 -0.1381 0.0042 1.0000 6.000 1.2155 0.01751 0.00886 -0.1356 0.0031 1.0000 6.500 1.2523 0.01841 0.00988 -0.1324 0.0030 1.0000 7.000 1.2874 0.01942 0.01098 -0.1291 0.0030 1.0000 7.500 1.3209 0.02055 0.01224 -0.1257 0.0030 1.0000 8.000 1.3518 0.02190 0.01374 -0.1220 0.0031 1.0000 8.500 1.3800 0.02347 0.01546 -0.1183 0.0033 1.0000 9.000 1.4059 0.02529 0.01743 -0.1144 0.0035 1.0000 9.500 1.4289 0.02741 0.01976 -0.1105 0.0037 1.0000 10.000 1.4477 0.03000 0.02255 -0.1064 0.0040 1.0000 10.500 1.4617 0.03315 0.02594 -0.1023 0.0042 1.0000 11.000 1.4697 0.03705 0.03011 -0.0982 0.0045 1.0000 11.500 1.4710 0.04194 0.03527 -0.0943 0.0048 1.0000 12.000 1.4659 0.04795 0.04159 -0.0908 0.0050 1.0000 12.500 1.4666 0.05377 0.04767 -0.0882 0.0053 1.0000 13.000 1.4759 0.05884 0.05299 -0.0866 0.0057 1.0000 13.500 1.4837 0.06438 0.05880 -0.0851 0.0063 1.0000 14.000 1.4874 0.07140 0.06633 -0.0821 0.0077 1.0000 14.500 1.4729 0.08226 0.07784 -0.0804 0.0092 1.0000 15.000 1.4379 0.09594 0.09208 -0.0822 0.0098 1.0000