XFOIL Version 6.96 Calculated polar for: HAM-STD HS1-404 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.200 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -7.750 -0.4454 0.09557 0.09204 -0.0097 1.0000 0.0135 -7.500 -0.4404 0.09242 0.08893 -0.0112 1.0000 0.0138 -7.250 -0.4353 0.08927 0.08582 -0.0129 1.0000 0.0141 -7.000 -0.4268 0.08580 0.08238 -0.0157 1.0000 0.0143 -6.750 -0.4167 0.08221 0.07883 -0.0189 1.0000 0.0146 -6.500 -0.4054 0.07851 0.07515 -0.0224 1.0000 0.0148 -5.000 -0.3007 0.05267 0.04910 -0.0457 1.0000 0.0074 -4.750 -0.2737 0.04754 0.04385 -0.0506 0.9993 0.0071 -4.500 -0.2297 0.04110 0.03718 -0.0586 0.9965 0.0070 -4.250 -0.1845 0.03449 0.03025 -0.0655 0.9943 0.0069 -4.000 -0.1415 0.02812 0.02343 -0.0705 0.9918 0.0073 -3.750 -0.1019 0.02306 0.01773 -0.0750 0.9893 0.0107 -3.500 -0.0647 0.02013 0.01428 -0.0775 0.9868 0.0123 -3.250 -0.0271 0.01716 0.01063 -0.0793 0.9848 0.0117 -3.000 0.0088 0.01519 0.00821 -0.0807 0.9831 0.0114 -2.750 0.0425 0.01378 0.00653 -0.0817 0.9803 0.0112 -2.500 0.0758 0.01275 0.00531 -0.0827 0.9768 0.0113 -2.250 0.1103 0.01201 0.00433 -0.0840 0.9738 0.0120 -2.000 0.1451 0.01147 0.00359 -0.0854 0.9712 0.0136 -1.750 0.1792 0.01079 0.00302 -0.0867 0.9682 0.0612 -1.500 0.2103 0.01033 0.00308 -0.0878 0.9632 0.1947 -1.250 0.2432 0.01019 0.00292 -0.0890 0.9591 0.2220 -1.000 0.2763 0.01002 0.00276 -0.0902 0.9554 0.2492 -0.750 0.3070 0.00984 0.00258 -0.0909 0.9485 0.2790 -0.250 0.3632 0.00793 0.00238 -0.0906 0.9325 1.0000 0.000 0.3933 0.00792 0.00228 -0.0908 0.9216 1.0000 0.250 0.4226 0.00791 0.00219 -0.0908 0.9094 1.0000 0.500 0.4508 0.00792 0.00217 -0.0906 0.8969 1.0000 0.750 0.4786 0.00796 0.00216 -0.0903 0.8862 1.0000 1.250 0.5334 0.00807 0.00225 -0.0898 0.8668 1.0000 1.500 0.5603 0.00814 0.00233 -0.0894 0.8558 1.0000 1.750 0.5871 0.00822 0.00243 -0.0890 0.8438 1.0000 2.000 0.6137 0.00830 0.00254 -0.0885 0.8300 1.0000 2.250 0.6393 0.00837 0.00271 -0.0876 0.8078 1.0000 2.500 0.6631 0.00847 0.00271 -0.0861 0.7623 1.0000 2.750 0.6854 0.00872 0.00267 -0.0842 0.6691 1.0000 3.000 0.7027 0.00979 0.00281 -0.0818 0.4883 1.0000 3.250 0.7227 0.01096 0.00328 -0.0806 0.3418 1.0000 3.500 0.7454 0.01187 0.00377 -0.0800 0.2372 1.0000 3.750 0.7656 0.01336 0.00448 -0.0793 0.0866 1.0000 4.000 0.7894 0.01423 0.00531 -0.0788 0.0506 1.0000 4.250 0.8143 0.01490 0.00606 -0.0782 0.0394 1.0000 4.500 0.8381 0.01581 0.00708 -0.0774 0.0325 1.0000 4.750 0.8625 0.01651 0.00790 -0.0768 0.0257 1.0000 5.000 0.8847 0.01780 0.00931 -0.0758 0.0214 1.0000 5.250 0.9096 0.01843 0.01006 -0.0752 0.0141 1.0000 5.750 0.9550 0.02093 0.01279 -0.0734 0.0050 1.0000 6.000 0.9768 0.02304 0.01518 -0.0720 0.0041 1.0000 6.250 0.9984 0.02599 0.01867 -0.0704 0.0035 1.0000 6.500 1.0186 0.02969 0.02293 -0.0685 0.0032 1.0000 6.750 1.0357 0.03420 0.02807 -0.0662 0.0031 1.0000 7.000 1.0485 0.03941 0.03391 -0.0636 0.0031 1.0000 7.250 1.0573 0.04501 0.04005 -0.0610 0.0032 1.0000 7.500 1.0624 0.05065 0.04615 -0.0586 0.0032 1.0000 7.750 1.0641 0.05623 0.05211 -0.0565 0.0033 1.0000 8.000 1.0620 0.06174 0.05794 -0.0547 0.0033 1.0000 8.250 1.0562 0.06711 0.06356 -0.0533 0.0034 1.0000 8.500 1.0465 0.07234 0.06899 -0.0524 0.0034 1.0000 8.750 1.0313 0.07708 0.07389 -0.0515 0.0035 1.0000 9.000 1.0149 0.08237 0.07929 -0.0527 0.0035 1.0000