XFOIL Version 6.94 Calculated polar for: USA 22 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.055 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 1.000 0.5537 0.02257 0.01287 -0.0942 0.7559 1.0000 2.000 0.6561 0.02434 0.01413 -0.0909 0.6984 1.0000 3.000 0.7556 0.02633 0.01586 -0.0871 0.6444 1.0000 4.000 0.8515 0.02867 0.01818 -0.0833 0.5900 1.0000 5.000 0.9431 0.03170 0.02144 -0.0795 0.5357 1.0000 5.000 0.9431 0.03170 0.02144 -0.0795 0.5357 1.0000 5.100 0.9547 0.03175 0.02147 -0.0791 0.5313 1.0000 5.200 0.9632 0.03218 0.02193 -0.0787 0.5260 1.0000 5.300 0.9705 0.03272 0.02254 -0.0784 0.5200 1.0000 5.400 0.9827 0.03272 0.02252 -0.0780 0.5160 1.0000 5.500 0.9892 0.03339 0.02330 -0.0776 0.5102 1.0000 5.600 0.9975 0.03383 0.02379 -0.0773 0.5046 1.0000 5.700 1.0103 0.03377 0.02370 -0.0769 0.5008 1.0000 5.800 1.0150 0.03465 0.02470 -0.0765 0.4946 1.0000 5.900 1.0238 0.03503 0.02512 -0.0762 0.4895 1.0000 6.000 1.0370 0.03496 0.02502 -0.0758 0.4859 1.0000 6.100 1.0404 0.03602 0.02622 -0.0755 0.4799 1.0000 6.200 1.0485 0.03654 0.02680 -0.0751 0.4751 1.0000 6.300 1.0612 0.03658 0.02683 -0.0748 0.4717 1.0000 6.400 1.0660 0.03754 0.02790 -0.0745 0.4668 1.0000 6.500 1.0710 0.03843 0.02890 -0.0742 0.4617 1.0000 6.600 1.0827 0.03856 0.02908 -0.0738 0.4580 1.0000 6.700 1.0952 0.03866 0.02917 -0.0735 0.4548 1.0000 6.800 1.0929 0.04039 0.03112 -0.0732 0.4488 1.0000 6.900 1.1019 0.04082 0.03159 -0.0728 0.4448 1.0000 7.000 1.1161 0.04070 0.03147 -0.0725 0.4418 1.0000 7.100 1.1170 0.04213 0.03304 -0.0721 0.4373 1.0000 7.200 1.1184 0.04349 0.03454 -0.0718 0.4328 1.0000 7.300 1.1276 0.04397 0.03508 -0.0715 0.4296 1.0000 7.400 1.1411 0.04401 0.03514 -0.0712 0.4271 1.0000 7.500 1.1415 0.04559 0.03686 -0.0710 0.4237 1.0000 7.600 1.1341 0.04804 0.03948 -0.0708 0.4195 1.0000 7.700 1.1388 0.04907 0.04059 -0.0705 0.4163 1.0000 7.800 1.1510 0.04920 0.04078 -0.0701 0.4137 1.0000 7.900 1.1633 0.04906 0.04068 -0.0695 0.4095 1.0000 8.000 1.1439 0.05316 0.04494 -0.0697 0.4069 1.0000 8.100 1.1288 0.05658 0.04845 -0.0698 0.4037 1.0000 8.200 1.2209 0.04168 0.03318 -0.0651 0.3725 1.0000 8.300 1.2378 0.04031 0.03174 -0.0644 0.3649 1.0000 8.400 1.2430 0.04048 0.03204 -0.0636 0.3580 1.0000 8.500 1.2589 0.03919 0.03068 -0.0629 0.3506 1.0000 8.600 1.2684 0.03822 0.02974 -0.0619 0.3404 1.0000 8.700 1.2750 0.03795 0.02957 -0.0610 0.3325 1.0000 8.800 1.2882 0.03645 0.02796 -0.0600 0.3225 1.0000 8.900 1.2924 0.03598 0.02760 -0.0587 0.3119 1.0000 9.000 1.2985 0.03579 0.02751 -0.0578 0.3045 1.0000 9.100 1.3061 0.03515 0.02688 -0.0567 0.2956 1.0000 9.200 1.3085 0.03512 0.02700 -0.0555 0.2856 1.0000 9.300 1.3132 0.03493 0.02688 -0.0544 0.2765 1.0000 9.400 1.3152 0.03477 0.02674 -0.0530 0.2639 1.0000 9.500 1.3153 0.03488 0.02696 -0.0517 0.2494 1.0000 9.600 1.3147 0.03521 0.02730 -0.0504 0.2333 1.0000 9.700 1.3139 0.03570 0.02779 -0.0492 0.2166 1.0000 9.800 1.3120 0.03638 0.02841 -0.0480 0.1989 1.0000 9.900 1.3091 0.03724 0.02920 -0.0468 0.1800 1.0000 10.000 1.3053 0.03820 0.03011 -0.0455 0.1598 1.0000