XFOIL Version 6.94 Calculated polar for: mta1 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.022 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.2919 0.05035 0.04050 -0.0210 1.0000 0.2985 -2.750 -0.2516 0.04727 0.03701 -0.0250 1.0000 0.2752 -2.500 -0.2116 0.04475 0.03395 -0.0282 1.0000 0.2581 -2.250 -0.1757 0.04238 0.03123 -0.0298 1.0000 0.2482 -2.000 -0.1380 0.04069 0.02896 -0.0312 1.0000 0.2404 -1.750 -0.1041 0.03883 0.02686 -0.0316 1.0000 0.2373 -1.500 -0.0699 0.03725 0.02505 -0.0317 1.0000 0.2346 -1.250 -0.0359 0.03588 0.02354 -0.0315 1.0000 0.2325 -1.000 -0.0025 0.03467 0.02236 -0.0309 1.0000 0.2327 -0.750 0.0310 0.03357 0.02150 -0.0300 1.0000 0.2373 -0.500 0.0670 0.03233 0.02094 -0.0293 1.0000 0.2468 -0.250 0.1090 0.03007 0.01979 -0.0296 0.7047 0.2629 0.000 0.2317 0.02936 0.01851 -0.0422 0.3962 1.0000 0.250 0.2742 0.03072 0.01859 -0.0423 0.3734 1.0000 0.500 0.3178 0.03214 0.01899 -0.0433 0.3587 1.0000 0.750 0.3615 0.03340 0.01969 -0.0448 0.3499 1.0000 1.000 0.4050 0.03484 0.02072 -0.0467 0.3436 1.0000 1.250 0.4475 0.03647 0.02208 -0.0487 0.3398 1.0000 1.500 0.4870 0.03816 0.02367 -0.0504 0.3366 1.0000 1.750 0.5228 0.03977 0.02540 -0.0515 0.3338 1.0000 2.000 0.5564 0.04153 0.02732 -0.0525 0.3319 1.0000 2.250 0.5892 0.04353 0.02949 -0.0535 0.3331 1.0000 2.500 0.6214 0.04586 0.03195 -0.0545 0.3367 1.0000 2.750 0.6498 0.04789 0.03442 -0.0555 0.3442 1.0000 3.000 0.6769 0.05049 0.03729 -0.0565 0.3526 1.0000 3.250 0.7022 0.05314 0.04022 -0.0576 0.3626 1.0000 3.500 0.7249 0.05620 0.04354 -0.0590 0.3744 1.0000 3.750 0.7435 0.05924 0.04693 -0.0609 0.3884 1.0000 4.000 0.7626 0.06271 0.05057 -0.0625 0.4015 1.0000 4.250 0.7789 0.06670 0.05469 -0.0647 0.4185 1.0000 4.500 0.7739 0.07066 0.05901 -0.0691 0.4475 1.0000 4.750 0.7805 0.07552 0.06396 -0.0735 0.4787 1.0000 5.000 0.7518 0.08053 0.06908 -0.0792 0.5239 1.0000