XFOIL Version 6.96 Calculated polar for: FX 61-163 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 1.000 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -14.500 -0.7146 0.07490 0.07275 -0.0548 1.0001 0.0034 -14.250 -0.7547 0.06370 0.06129 -0.0616 1.0001 0.0034 -14.000 -0.7750 0.05714 0.05456 -0.0650 1.0001 0.0034 -13.750 -0.7896 0.05216 0.04943 -0.0671 1.0001 0.0034 -13.500 -0.8006 0.04807 0.04520 -0.0683 1.0001 0.0034 -13.250 -0.8083 0.04466 0.04166 -0.0689 1.0001 0.0034 -13.000 -0.8142 0.04167 0.03854 -0.0690 1.0001 0.0034 -12.750 -0.8183 0.03904 0.03579 -0.0687 1.0001 0.0035 -12.500 -0.8152 0.03642 0.03303 -0.0695 0.9990 0.0035 -12.250 -0.8115 0.03406 0.03055 -0.0696 0.9886 0.0035 -12.000 -0.7533 0.03052 0.02672 -0.0821 0.9393 0.0036 -11.750 -0.7278 0.02878 0.02440 -0.0855 0.8253 0.0037 -11.500 -0.7187 0.02747 0.02281 -0.0848 0.7845 0.0038 -11.250 -0.7072 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8.250 1.3480 0.01656 0.00929 -0.1048 0.0434 0.6845 8.500 1.3612 0.01742 0.01004 -0.1027 0.0268 0.6857 8.750 1.3760 0.01806 0.01067 -0.1009 0.0200 0.6870 9.000 1.3888 0.01866 0.01128 -0.0987 0.0160 0.6886 9.250 1.4029 0.01929 0.01194 -0.0968 0.0132 0.6902 9.500 1.4158 0.02000 0.01268 -0.0948 0.0100 0.6916 9.750 1.4245 0.02096 0.01365 -0.0925 0.0049 0.6931 10.000 1.4327 0.02201 0.01473 -0.0902 0.0028 0.6945 10.250 1.4435 0.02297 0.01575 -0.0885 0.0025 0.6959 10.500 1.4537 0.02407 0.01690 -0.0869 0.0022 0.6975 10.750 1.4632 0.02529 0.01818 -0.0855 0.0021 0.6991 11.000 1.4724 0.02665 0.01961 -0.0843 0.0019 0.7005 11.250 1.4805 0.02819 0.02123 -0.0832 0.0018 0.7019 11.500 1.4891 0.02979 0.02292 -0.0824 0.0017 0.7033 11.750 1.4973 0.03150 0.02471 -0.0816 0.0017 0.7048 12.000 1.5045 0.03334 0.02663 -0.0810 0.0017 0.7062 12.250 1.5110 0.03528 0.02866 -0.0804 0.0016 0.7077 12.500 1.5164 0.03739 0.03084 -0.0798 0.0016 0.7092 12.750 1.5208 0.03958 0.03311 -0.0793 0.0015 0.7108 13.000 1.5243 0.04188 0.03550 -0.0787 0.0015 0.7124 13.250 1.5268 0.04430 0.03799 -0.0782 0.0014 0.7140 13.500 1.5282 0.04685 0.04063 -0.0776 0.0014 0.7155 13.750 1.5289 0.04947 0.04333 -0.0772 0.0013 0.7168 14.000 1.5295 0.05214 0.04609 -0.0768 0.0013 0.7183 14.250 1.5291 0.05500 0.04904 -0.0764 0.0013 0.7198 14.500 1.5280 0.05798 0.05212 -0.0762 0.0012 0.7214 14.750 1.5270 0.06100 0.05523 -0.0760 0.0012 0.7230 15.000 1.5251 0.06421 0.05853 -0.0759 0.0012 0.7246 15.250 1.5226 0.06757 0.06198 -0.0760 0.0012 0.7262 15.500 1.5192 0.07107 0.06557 -0.0761 0.0012 0.7278 15.750 1.5152 0.07470 0.06929 -0.0763 0.0011 0.7293 16.000 1.5117 0.07835 0.07303 -0.0766 0.0011 0.7307 16.250 1.5065 0.08231 0.07710 -0.0771 0.0011 0.7324 16.500 1.5019 0.08623 0.08112 -0.0776 0.0011 0.7343 16.750 1.4954 0.09047 0.08547 -0.0783 0.0011 0.7362 17.000 1.4889 0.09478 0.08989 -0.0791 0.0011 0.7382 17.250 1.4805 0.09946 0.09467 -0.0802 0.0011 0.7402 17.500 1.4724 0.10415 0.09947 -0.0813 0.0010 0.7422 17.750 1.4635 0.10899 0.10442 -0.0826 0.0010 0.7440 18.000 1.4550 0.11385 0.10938 -0.0840 0.0010 0.7458 18.250 1.4457 0.11897 0.11461 -0.0857 0.0010 0.7477