XFOIL Version 6.96 Calculated polar for: FX 61-184 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.100 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -13.500 -0.4496 0.09525 0.09013 -0.0732 1.0000 0.0277 -13.250 -0.4828 0.08569 0.08048 -0.0787 1.0000 0.0271 -13.000 -0.5122 0.07851 0.07319 -0.0823 1.0000 0.0266 -12.750 -0.5406 0.07272 0.06729 -0.0845 1.0000 0.0265 -12.500 -0.5666 0.06829 0.06277 -0.0851 1.0000 0.0264 -12.250 -0.5776 0.06255 0.05677 -0.0897 0.9739 0.0269 -12.000 -0.5841 0.05745 0.05133 -0.0942 0.9539 0.0272 -11.750 -0.5845 0.05272 0.04616 -0.0989 0.9396 0.0279 -11.500 -0.5783 0.04837 0.04129 -0.1032 0.9281 0.0285 -11.250 -0.5643 0.04482 0.03730 -0.1063 0.9172 0.0291 -11.000 -0.5394 0.04233 0.03463 -0.1087 0.9091 0.0298 -10.750 -0.5146 0.04032 0.03245 -0.1109 0.9011 0.0309 -10.500 -0.4920 0.03857 0.03049 -0.1124 0.8930 0.0324 -10.250 -0.4686 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0.7032 4.500 0.8388 0.02574 0.01714 -0.0880 0.6035 0.7053 4.750 0.8669 0.02566 0.01713 -0.0882 0.5983 0.7078 5.000 0.9039 0.02531 0.01680 -0.0897 0.5949 0.7106 5.250 0.9134 0.02596 0.01756 -0.0877 0.5857 0.7136 5.500 0.9499 0.02562 0.01726 -0.0892 0.5808 0.7163 5.750 0.9621 0.02601 0.01778 -0.0871 0.5722 0.7182 6.000 0.9923 0.02578 0.01761 -0.0874 0.5661 0.7202 6.250 1.0083 0.02606 0.01799 -0.0859 0.5577 0.7226 6.500 1.0385 0.02583 0.01783 -0.0863 0.5504 0.7254 6.750 1.0518 0.02617 0.01826 -0.0844 0.5408 0.7286 7.000 1.0884 0.02586 0.01797 -0.0859 0.5335 0.7318 7.250 1.0937 0.02646 0.01873 -0.0829 0.5237 0.7341 7.500 1.1270 0.02618 0.01850 -0.0836 0.5167 0.7364 7.750 1.1335 0.02689 0.01935 -0.0809 0.5060 0.7391 8.000 1.1525 0.02719 0.01974 -0.0800 0.4962 0.7421 8.250 1.1796 0.02717 0.01977 -0.0801 0.4862 0.7454 8.500 1.1877 0.02797 0.02069 -0.0780 0.4739 0.7487 8.750 1.1990 0.02858 0.02140 -0.0761 0.4619 0.7516 9.000 1.2134 0.02911 0.02199 -0.0747 0.4494 0.7549 9.250 1.2279 0.02969 0.02264 -0.0735 0.4362 0.7583 9.500 1.2415 0.03039 0.02338 -0.0723 0.4223 0.7618 9.750 1.2519 0.03127 0.02432 -0.0709 0.4078 0.7649 10.000 1.2586 0.03234 0.02545 -0.0691 0.3922 0.7681 10.250 1.2635 0.03363 0.02677 -0.0673 0.3752 0.7718 10.500 1.2680 0.03507 0.02824 -0.0658 0.3573 0.7760 10.750 1.2723 0.03660 0.02979 -0.0643 0.3387 0.7802 11.000 1.2746 0.03822 0.03142 -0.0626 0.3199 0.7840 11.250 1.2767 0.03997 0.03314 -0.0612 0.3012 0.7882 11.500 1.2782 0.04193 0.03506 -0.0599 0.2823 0.7926 11.750 1.2780 0.04409 0.03721 -0.0587 0.2635 0.7968 12.000 1.2773 0.04631 0.03943 -0.0575 0.2460 0.8010 12.250 1.2771 0.04864 0.04176 -0.0565 0.2297 0.8060 12.500 1.2771 0.05110 0.04424 -0.0558 0.2143 0.8114 12.750 1.2762 0.05360 0.04678 -0.0550 0.2000 0.8169 13.000 1.2753 0.05629 0.04950 -0.0544 0.1860 0.8235 13.250 1.2740 0.05905 0.05230 -0.0539 0.1731 0.8304 13.500 1.2721 0.06197 0.05526 -0.0535 0.1605 0.8386 13.750 1.2693 0.06501 0.05836 -0.0531 0.1482 0.8479 14.000 1.2665 0.06811 0.06152 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