XFOIL Version 6.94 Calculated polar for: mta3 300.2.14.15.2.c13 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 -0.2163 0.04483 0.03494 -0.0283 1.0000 0.2801 -2.750 -0.1785 0.04239 0.03198 -0.0314 1.0000 0.2636 -2.500 -0.1463 0.03996 0.02929 -0.0323 1.0000 0.2530 -2.250 -0.1104 0.03821 0.02699 -0.0335 1.0000 0.2421 -2.000 -0.0790 0.03633 0.02489 -0.0336 1.0000 0.2361 -1.750 -0.0466 0.03477 0.02308 -0.0334 1.0000 0.2311 -1.500 -0.0149 0.03342 0.02156 -0.0329 1.0000 0.2294 -1.250 0.0159 0.03214 0.02035 -0.0321 1.0000 0.2320 -1.000 0.0467 0.03098 0.01942 -0.0310 1.0000 0.2378 -0.750 0.0790 0.02986 0.01874 -0.0298 1.0000 0.2458 -0.500 0.1143 0.02853 0.01820 -0.0289 1.0000 0.2590 -0.250 0.2167 0.02816 0.01660 -0.0378 0.4210 0.3648 0.000 0.2451 0.02768 0.01663 -0.0349 0.3924 1.0000 0.250 0.2772 0.02899 0.01687 -0.0335 0.3722 1.0000 0.500 0.3103 0.03017 0.01729 -0.0328 0.3591 1.0000 0.750 0.3438 0.03156 0.01799 -0.0325 0.3502 1.0000 1.000 0.3782 0.03284 0.01893 -0.0327 0.3445 1.0000 1.250 0.4126 0.03418 0.02012 -0.0331 0.3406 1.0000 1.500 0.4466 0.03566 0.02153 -0.0337 0.3378 1.0000 1.750 0.4793 0.03728 0.02314 -0.0342 0.3346 1.0000 2.000 0.5107 0.03907 0.02495 -0.0346 0.3314 1.0000 2.250 0.5413 0.04102 0.02699 -0.0352 0.3297 1.0000 2.500 0.5712 0.04299 0.02922 -0.0360 0.3320 1.0000 2.750 0.6009 0.04526 0.03180 -0.0373 0.3375 1.0000 3.000 0.6294 0.04790 0.03465 -0.0386 0.3438 1.0000 3.250 0.6572 0.05068 0.03765 -0.0400 0.3509 1.0000 3.500 0.6823 0.05386 0.04125 -0.0429 0.3635 1.0000 3.750 0.7057 0.05727 0.04493 -0.0456 0.3751 1.0000 4.000 0.7257 0.06126 0.04914 -0.0486 0.3884 1.0000 4.250 0.7388 0.06573 0.05391 -0.0536 0.4075 1.0000 4.500 0.7481 0.07071 0.05906 -0.0590 0.4304 1.0000 4.750 0.7376 0.07663 0.06515 -0.0673 0.4666 1.0000 5.000 0.7241 0.08267 0.07118 -0.0751 0.5104 1.0000