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 = 0.008 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0287 0.05739 0.04096 -0.0252 0.9999 1.0001 -2.750 -0.0452 0.05505 0.03894 -0.0233 0.9999 1.0001 -2.500 -0.0624 0.05263 0.03675 -0.0219 0.9999 1.0001 -2.250 -0.0443 0.05068 0.03387 -0.0297 0.9999 1.0001 -2.000 -0.0013 0.05072 0.03200 -0.0370 0.9999 1.0001 -1.750 0.0281 0.05117 0.03125 -0.0386 0.9999 1.0001 -1.500 0.0530 0.05170 0.03092 -0.0390 0.9999 1.0001 -1.250 0.0761 0.05230 0.03084 -0.0391 0.9999 1.0001 -1.000 0.0981 0.05297 0.03097 -0.0390 0.9999 1.0001 -0.750 0.1192 0.05373 0.03129 -0.0389 0.9999 1.0001 -0.500 0.1395 0.05458 0.03179 -0.0387 0.9999 1.0001 -0.250 0.1590 0.05554 0.03245 -0.0386 0.9999 1.0001 0.000 0.1775 0.05664 0.03333 -0.0385 0.9999 1.0001 0.250 0.1947 0.05791 0.03445 -0.0385 0.9999 1.0001 0.500 0.2103 0.05941 0.03585 -0.0387 0.9999 1.0001 0.750 0.2235 0.06125 0.03764 -0.0389 0.9999 1.0001 1.000 0.2334 0.06353 0.03994 -0.0395 0.9999 1.0001 1.250 0.2389 0.06641 0.04284 -0.0404 0.9999 1.0001 1.500 0.2410 0.06981 0.04624 -0.0417 0.9999 1.0001 1.750 0.2426 0.07337 0.04974 -0.0431 0.9999 1.0001 2.000 0.2462 0.07681 0.05309 -0.0447 0.9999 1.0001 2.250 0.2518 0.08008 0.05622 -0.0462 0.9999 1.0001 2.500 0.2590 0.08319 0.05919 -0.0476 0.9999 1.0001 2.750 0.2673 0.08620 0.06205 -0.0489 0.9999 1.0001 3.000 0.2765 0.08913 0.06482 -0.0502 0.9999 1.0001 3.250 0.2863 0.09200 0.06755 -0.0515 0.9999 1.0001 3.500 0.2966 0.09485 0.07025 -0.0527 0.9999 1.0001 3.750 0.3073 0.09767 0.07292 -0.0539 0.9999 1.0001 4.000 0.3184 0.10048 0.07558 -0.0550 0.9999 1.0001 4.250 0.3297 0.10327 0.07824 -0.0562 0.9999 1.0001 4.500 0.3412 0.10606 0.08090 -0.0573 0.9999 1.0001 4.750 0.3528 0.10884 0.08357 -0.0585 0.9999 1.0001 5.000 0.3646 0.11163 0.08624 -0.0596 0.9999 1.0001