XFOIL Version 6.96 Calculated polar for: FX 66-S-161 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 = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -11.250 -0.2963 0.11162 0.10847 -0.0482 1.0000 0.0564 -11.000 -0.3195 0.10517 0.10214 -0.0560 1.0000 0.0585 -10.750 -0.3298 0.09944 0.09650 -0.0603 1.0000 0.0587 -10.500 -0.2979 0.09602 0.09309 -0.0585 0.9608 0.0608 -10.250 -0.2506 0.09069 0.08759 -0.0686 0.9128 0.0639 -10.000 -0.1790 0.06974 0.06617 -0.0841 0.7865 0.0734 -9.750 -0.1656 0.06804 0.06436 -0.0827 0.7724 0.0751 -9.500 -0.1670 0.06400 0.06027 -0.0838 0.7626 0.0768 -9.250 -0.1782 0.05823 0.05448 -0.0867 0.7549 0.0784 -9.000 -0.2123 0.04934 0.04558 -0.0932 0.7501 0.0792 -8.750 -0.3900 0.04165 0.03620 -0.1043 0.7702 0.0374 -8.500 -0.3771 0.03766 0.03205 -0.1039 0.7602 0.0352 -8.250 -0.3696 0.03424 0.02825 -0.1028 0.7498 0.0341 -8.000 -0.3575 0.03122 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1.1358 0.02089 0.01455 -0.0924 0.5134 1.0000 6.750 1.1618 0.02117 0.01489 -0.0923 0.5100 1.0000 7.000 1.1910 0.02134 0.01507 -0.0927 0.5073 1.0000 7.250 1.2184 0.02165 0.01542 -0.0928 0.5043 1.0000 7.500 1.2342 0.02236 0.01635 -0.0914 0.4988 1.0000 7.750 1.2594 0.02264 0.01672 -0.0912 0.4950 1.0000 8.000 1.2900 0.02268 0.01678 -0.0917 0.4921 1.0000 8.250 1.3115 0.02311 0.01735 -0.0909 0.4874 1.0000 8.500 1.3319 0.02335 0.01774 -0.0899 0.4814 1.0000 8.750 1.3689 0.02218 0.01648 -0.0906 0.4740 1.0000 9.000 1.3936 0.02067 0.01493 -0.0892 0.4591 1.0000 9.250 1.4126 0.01984 0.01412 -0.0872 0.4448 1.0000 9.500 1.4296 0.01961 0.01399 -0.0853 0.4331 1.0000 9.750 1.4428 0.01967 0.01421 -0.0830 0.4208 1.0000 10.000 1.4555 0.01973 0.01436 -0.0805 0.4064 1.0000 10.250 1.4639 0.01988 0.01454 -0.0773 0.3886 1.0000 10.500 1.4633 0.02037 0.01505 -0.0729 0.3674 1.0000 10.750 1.4564 0.02128 0.01588 -0.0681 0.3397 1.0000 11.000 1.4410 0.02292 0.01736 -0.0631 0.3083 1.0000 11.250 1.4197 0.02545 0.01973 -0.0587 0.2719 1.0000 11.500 1.3931 0.02900 0.02308 -0.0553 0.2353 1.0000 11.750 1.3665 0.03320 0.02710 -0.0529 0.2031 1.0000 12.000 1.3386 0.03792 0.03165 -0.0511 0.1715 1.0000 12.250 1.3127 0.04274 0.03631 -0.0498 0.1428 1.0000 12.500 1.2861 0.04786 0.04125 -0.0489 0.1132 1.0000 12.750 1.2606 0.05315 0.04634 -0.0483 0.0849 1.0000 13.000 1.2385 0.05839 0.05141 -0.0480 0.0638 1.0000 13.250 1.2214 0.06334 0.05626 -0.0480 0.0515 1.0000 13.500 1.2110 0.06766 0.06059 -0.0481 0.0439 1.0000 13.750 1.2025 0.07191 0.06489 -0.0483 0.0393 1.0000 14.000 1.1978 0.07575 0.06878 -0.0487 0.0354 1.0000 14.250 1.1902 0.07999 0.07306 -0.0491 0.0333 1.0000 14.500 1.1885 0.08354 0.07669 -0.0494 0.0308 1.0000 14.750 1.1891 0.08680 0.08004 -0.0498 0.0285 1.0000 15.000 1.1891 0.09012 0.08339 -0.0502 0.0269 1.0000 15.250 1.1894 0.09298 0.08620 -0.0499 0.0252 1.0000