XFOIL Version 6.96 Calculated polar for: FX 60-100 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -9.500 -0.2879 0.08431 0.08223 -0.0408 1.0000 0.0195 -9.250 -0.2935 0.08055 0.07851 -0.0409 1.0000 0.0195 -9.000 -0.3021 0.07732 0.07532 -0.0398 1.0000 0.0195 -8.750 -0.3151 0.07473 0.07279 -0.0377 1.0000 0.0195 -8.500 -0.3277 0.06558 0.06367 -0.0423 0.9983 0.0200 -8.250 -0.3177 0.06163 0.05973 -0.0451 0.9960 0.0207 -8.000 -0.3017 0.05900 0.05709 -0.0477 0.9937 0.0219 -7.750 -0.2974 0.05280 0.05090 -0.0532 0.9893 0.0224 -7.500 -0.3612 0.05081 0.04841 -0.0817 0.9849 0.0195 -7.250 -0.3481 0.04027 0.03763 -0.0942 0.9782 0.0209 -7.000 -0.3180 0.03782 0.03510 -0.0987 0.9750 0.0222 -6.750 -0.2828 0.03391 0.03088 -0.1049 0.9725 0.0245 -6.500 -0.2437 0.02840 0.02463 -0.1126 0.9705 0.0286 -6.250 -0.2162 0.02634 0.02257 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3.750 0.9093 0.00847 0.00309 -0.1131 0.4613 0.8362 4.000 0.9300 0.00847 0.00319 -0.1114 0.4364 1.0000 4.250 0.9561 0.00878 0.00339 -0.1112 0.4094 1.0000 4.500 0.9817 0.00915 0.00362 -0.1108 0.3777 1.0000 4.750 1.0065 0.00958 0.00388 -0.1104 0.3412 1.0000 5.000 1.0310 0.01007 0.00419 -0.1099 0.3009 1.0000 5.250 1.0552 0.01058 0.00453 -0.1095 0.2649 1.0000 5.500 1.0792 0.01112 0.00489 -0.1090 0.2284 1.0000 5.750 1.1028 0.01170 0.00529 -0.1084 0.1923 1.0000 6.000 1.1263 0.01229 0.00574 -0.1078 0.1632 1.0000 6.250 1.1497 0.01287 0.00618 -0.1073 0.1371 1.0000 6.500 1.1737 0.01337 0.00663 -0.1067 0.1213 1.0000 6.750 1.1975 0.01387 0.00707 -0.1062 0.1046 1.0000 7.000 1.2206 0.01444 0.00757 -0.1055 0.0915 1.0000 7.250 1.2441 0.01494 0.00808 -0.1049 0.0802 1.0000 7.500 1.2671 0.01549 0.00857 -0.1042 0.0680 1.0000 7.750 1.2893 0.01610 0.00913 -0.1035 0.0587 1.0000 8.000 1.3096 0.01690 0.00989 -0.1024 0.0514 1.0000 8.250 1.3329 0.01734 0.01041 -0.1017 0.0482 1.0000 8.500 1.3552 0.01786 0.01100 -0.1009 0.0442 1.0000 8.750 1.3747 0.01865 0.01180 -0.0998 0.0398 1.0000 9.000 1.3976 0.01903 0.01225 -0.0991 0.0378 1.0000 9.250 1.4197 0.01947 0.01269 -0.0984 0.0344 1.0000 9.500 1.4395 0.02011 0.01331 -0.0974 0.0299 1.0000 9.750 1.4613 0.02051 0.01374 -0.0966 0.0256 1.0000 10.000 1.4801 0.02118 0.01440 -0.0954 0.0190 1.0000 10.250 1.4927 0.02239 0.01550 -0.0934 0.0092 1.0000 10.500 1.5044 0.02354 0.01670 -0.0911 0.0070 1.0000 10.750 1.5152 0.02454 0.01783 -0.0887 0.0063 1.0000 11.000 1.5239 0.02566 0.01906 -0.0861 0.0059 1.0000 11.250 1.5309 0.02693 0.02045 -0.0834 0.0055 1.0000 11.500 1.5352 0.02846 0.02211 -0.0806 0.0051 1.0000 11.750 1.5359 0.03031 0.02413 -0.0777 0.0049 1.0000 12.000 1.5390 0.03205 0.02600 -0.0753 0.0048 1.0000 12.250 1.5417 0.03390 0.02800 -0.0732 0.0047 1.0000 12.500 1.5427 0.03599 0.03023 -0.0713 0.0046 1.0000 12.750 1.5424 0.03833 0.03272 -0.0695 0.0046 1.0000 13.000 1.5404 0.04097 0.03551 -0.0681 0.0045 1.0000 13.250 1.5370 0.04391 0.03862 -0.0669 0.0045 1.0000 13.500 1.5325 0.04717 0.04205 -0.0662 0.0044 1.0000 13.750 1.5265 0.05081 0.04585 -0.0658 0.0044 1.0000 14.000 1.5193 0.05485 0.05005 -0.0660 0.0043 1.0000 14.250 1.5107 0.05930 0.05467 -0.0666 0.0043 1.0000 14.500 1.5013 0.06418 0.05971 -0.0677 0.0043 1.0000 14.750 1.4907 0.06954 0.06524 -0.0693 0.0043 1.0000 15.000 1.4791 0.07538 0.07125 -0.0714 0.0043 1.0000 15.250 1.4667 0.08165 0.07769 -0.0740 0.0043 1.0000 15.500 1.4535 0.08836 0.08457 -0.0770 0.0043 1.0000 15.750 1.4397 0.09546 0.09183 -0.0804 0.0043 1.0000 16.000 1.4255 0.10291 0.09945 -0.0842 0.0043 1.0000 16.250 1.4102 0.11080 0.10750 -0.0883 0.0043 1.0000 16.500 1.3954 0.11884 0.11569 -0.0927 0.0043 1.0000