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 = 0.300 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.7164 0.00987 0.00369 -0.1679 0.7756 0.3950 0.500 0.7706 0.00975 0.00371 -0.1673 0.7589 0.4677 1.000 0.8248 0.00964 0.00379 -0.1667 0.7425 0.5632 1.500 0.8732 0.00914 0.00385 -0.1647 0.7256 0.7800 2.000 0.9319 0.00914 0.00384 -0.1650 0.7064 1.0000 4.000 1.1078 0.01086 0.00447 -0.1544 0.5054 1.0000 4.500 1.1409 0.01190 0.00507 -0.1501 0.4308 1.0000 5.000 1.1740 0.01298 0.00578 -0.1460 0.3617 1.0000 5.500 1.1900 0.01486 0.00695 -0.1392 0.2342 1.0000 6.000 1.1737 0.01862 0.00947 -0.1278 0.0357 1.0000 6.500 1.2061 0.01972 0.01055 -0.1241 0.0262 1.0000 7.000 1.2349 0.02107 0.01186 -0.1200 0.0059 1.0000 7.500 1.2645 0.02243 0.01331 -0.1162 0.0048 1.0000 8.000 1.2925 0.02395 0.01498 -0.1124 0.0047 1.0000 8.500 1.3177 0.02572 0.01693 -0.1084 0.0047 1.0000 9.000 1.3405 0.02776 0.01916 -0.1045 0.0048 1.0000 9.500 1.3615 0.03006 0.02164 -0.1006 0.0050 1.0000 10.000 1.3800 0.03270 0.02447 -0.0969 0.0052 1.0000 10.500 1.3948 0.03584 0.02784 -0.0933 0.0054 1.0000 11.000 1.4044 0.03967 0.03190 -0.0898 0.0057 1.0000 11.500 1.4088 0.04432 0.03678 -0.0867 0.0060 1.0000 12.000 1.4074 0.04999 0.04265 -0.0841 0.0062 1.0000 12.500 1.4164 0.05488 0.04776 -0.0824 0.0065 1.0000 13.000 1.4215 0.06059 0.05374 -0.0809 0.0070 1.0000 13.500 1.4255 0.06682 0.06021 -0.0796 0.0076 1.0000 14.000 1.4397 0.07226 0.06579 -0.0778 0.0081 1.0000 14.500 1.4542 0.07894 0.07317 -0.0760 0.0096 1.0000 15.000 1.4396 0.09211 0.08736 -0.0756 0.0117 1.0000 15.500 1.2643 0.11320 0.10926 -0.0859 0.0098 1.0000 16.000 1.2276 0.12671 0.12341 -0.0922 0.0107 1.0000 16.500 1.1893 0.14136 0.13848 -0.1010 0.0107 1.0000 17.000 1.1500 0.15774 0.15524 -0.1115 0.0106 1.0000 17.500 1.1215 0.17271 0.17039 -0.1211 0.0102 1.0000