XFOIL Version 6.94 Calculated polar for: GOE 484 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 0.6066 0.00859 0.00250 -0.1436 0.8605 0.0437 1.000 0.7107 0.00814 0.00188 -0.1406 0.7991 0.0531 1.500 0.7597 0.00791 0.00186 -0.1387 0.7505 0.1744 2.000 0.8044 0.00685 0.00207 -0.1363 0.6488 1.0000 2.500 0.8446 0.00777 0.00240 -0.1328 0.5529 1.0000 3.000 0.8517 0.01182 0.00391 -0.1244 0.0545 1.0000 3.500 0.9003 0.01244 0.00444 -0.1229 0.0410 1.0000 4.000 0.9488 0.01302 0.00488 -0.1215 0.0295 1.0000 4.500 0.9978 0.01352 0.00537 -0.1201 0.0270 1.0000 5.000 1.0460 0.01404 0.00580 -0.1187 0.0162 1.0000 5.500 1.0904 0.01494 0.00654 -0.1166 0.0043 1.0000 6.000 1.1366 0.01561 0.00730 -0.1147 0.0039 1.0000 6.500 1.1795 0.01655 0.00835 -0.1122 0.0038 1.0000 7.000 1.2212 0.01753 0.00947 -0.1095 0.0038 1.0000 7.500 1.2611 0.01860 0.01078 -0.1066 0.0038 1.0000 8.000 1.2977 0.01985 0.01224 -0.1030 0.0040 1.0000 8.500 1.3299 0.02125 0.01388 -0.0988 0.0041 1.0000 9.000 1.3597 0.02284 0.01569 -0.0943 0.0043 1.0000 9.500 1.3862 0.02474 0.01783 -0.0894 0.0046 1.0000 10.000 1.4077 0.02722 0.02061 -0.0842 0.0049 1.0000 10.500 1.4249 0.03049 0.02432 -0.0788 0.0051 1.0000 11.000 1.4359 0.03545 0.02986 -0.0731 0.0056 1.0000 11.500 1.4322 0.04252 0.03767 -0.0669 0.0059 1.0000 12.000 1.4040 0.05174 0.04764 -0.0607 0.0061 1.0000 12.500 1.3636 0.06218 0.05868 -0.0570 0.0062 1.0000 13.000 1.3168 0.07468 0.07168 -0.0576 0.0063 1.0000 13.500 1.2689 0.08976 0.08717 -0.0633 0.0063 1.0000 14.000 1.2293 0.10657 0.10430 -0.0728 0.0062 1.0000