XFOIL Version 6.94 Calculated polar for: FX 63-100 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 0.6793 0.00753 0.00249 -0.1570 0.7517 0.6560 0.500 0.7346 0.00763 0.00261 -0.1566 0.7337 0.6843 1.000 0.7892 0.00775 0.00270 -0.1559 0.7115 0.7079 1.500 0.8431 0.00789 0.00278 -0.1552 0.6843 0.7279 2.000 0.8973 0.00806 0.00292 -0.1546 0.6607 0.7484 2.500 0.9501 0.00823 0.00308 -0.1537 0.6291 0.7717 4.500 1.1081 0.01264 0.00536 -0.1415 0.1542 1.0000 5.000 1.1463 0.01483 0.00673 -0.1391 0.0327 1.0000 5.500 1.1959 0.01562 0.00749 -0.1382 0.0255 1.0000 6.000 1.2412 0.01681 0.00854 -0.1366 0.0029 1.0000 6.500 1.2886 0.01769 0.00948 -0.1352 0.0025 1.0000 7.000 1.3345 0.01862 0.01051 -0.1335 0.0026 1.0000 7.500 1.3781 0.01967 0.01167 -0.1315 0.0028 1.0000 8.000 1.4185 0.02090 0.01307 -0.1290 0.0029 1.0000 8.500 1.4555 0.02228 0.01461 -0.1260 0.0032 1.0000 9.000 1.4878 0.02362 0.01612 -0.1221 0.0036 1.0000 9.500 1.5137 0.02527 0.01796 -0.1175 0.0040 1.0000 10.000 1.5349 0.02737 0.02029 -0.1125 0.0043 1.0000 10.500 1.5583 0.02939 0.02251 -0.1083 0.0049 1.0000 11.000 1.5709 0.03246 0.02582 -0.1034 0.0054 1.0000 11.500 1.5764 0.03641 0.03000 -0.0989 0.0058 1.0000 12.000 1.5839 0.04062 0.03449 -0.0954 0.0066 1.0000 12.500 1.5768 0.04688 0.04103 -0.0923 0.0070 1.0000 13.000 1.5647 0.05458 0.04898 -0.0904 0.0073 1.0000 13.500 1.5666 0.06122 0.05600 -0.0895 0.0080 1.0000 14.000 1.5550 0.07095 0.06618 -0.0889 0.0089 1.0000 15.000 1.4913 0.10032 0.09694 -0.0951 0.0111 1.0000 16.000 1.4079 0.13717 0.13478 -0.1160 0.0121 1.0000 16.500 1.3740 0.15627 0.15421 -0.1295 0.0120 1.0000 17.000 1.3509 0.17380 0.17192 -0.1421 0.0118 1.0000