XFOIL Version 6.94 Calculated polar for: NACA4412c 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.005 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0418 0.06787 0.04624 -0.0281 1.0001 0.9999 -2.750 -0.0535 0.06580 0.04449 -0.0269 1.0001 0.9999 -2.500 -0.0677 0.06367 0.04267 -0.0255 1.0001 0.9999 -2.250 -0.0797 0.06162 0.04069 -0.0249 1.0001 0.9999 -2.000 -0.0731 0.06008 0.03845 -0.0278 1.0001 0.9999 -1.750 -0.0492 0.05954 0.03662 -0.0318 1.0001 0.9999 -1.500 -0.0238 0.05954 0.03539 -0.0340 1.0001 0.9999 -1.250 0.0001 0.05976 0.03462 -0.0351 1.0001 0.9999 -1.000 0.0226 0.06011 0.03414 -0.0356 1.0001 0.9999 -0.750 0.0442 0.06057 0.03388 -0.0359 1.0001 0.9999 -0.500 0.0650 0.06111 0.03380 -0.0361 1.0001 0.9999 -0.250 0.0851 0.06176 0.03393 -0.0361 1.0001 0.9999 0.000 0.1045 0.06250 0.03423 -0.0361 1.0001 0.9999 0.250 0.1231 0.06335 0.03470 -0.0361 1.0001 0.9999 0.500 0.1408 0.06432 0.03534 -0.0361 1.0001 0.9999 0.750 0.1576 0.06542 0.03619 -0.0362 1.0001 0.9999 1.000 0.1732 0.06669 0.03725 -0.0363 1.0001 0.9999 1.250 0.1874 0.06815 0.03855 -0.0365 1.0001 0.9999 1.500 0.1999 0.06984 0.04011 -0.0369 1.0001 0.9999 1.750 0.2107 0.07179 0.04193 -0.0374 1.0001 0.9999 2.000 0.2195 0.07400 0.04403 -0.0381 1.0001 0.9999 2.250 0.2269 0.07645 0.04637 -0.0389 1.0001 0.9999 2.500 0.2335 0.07908 0.04887 -0.0399 1.0001 0.9999 2.750 0.2399 0.08180 0.05143 -0.0410 1.0001 0.9999 3.000 0.2467 0.08454 0.05400 -0.0422 1.0001 0.9999 3.250 0.2542 0.08727 0.05654 -0.0434 1.0001 0.9999 3.500 0.2625 0.08998 0.05907 -0.0446 1.0001 0.9999 3.750 0.2714 0.09268 0.06158 -0.0457 1.0001 0.9999 4.000 0.2807 0.09535 0.06407 -0.0469 1.0001 0.9999 4.250 0.2905 0.09801 0.06654 -0.0481 1.0001 0.9999 4.500 0.3007 0.10065 0.06900 -0.0492 1.0001 0.9999 4.750 0.3112 0.10329 0.07147 -0.0504 1.0001 0.9999 5.000 0.3219 0.10592 0.07394 -0.0515 1.0001 0.9999