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 = 0.040 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1312 0.04063 0.03104 -0.0491 1.0001 0.2376 -2.750 -0.1016 0.03888 0.02884 -0.0501 1.0001 0.2462 -2.500 -0.0773 0.03770 0.02764 -0.0498 1.0001 0.2599 -2.250 -0.0507 0.03657 0.02638 -0.0500 1.0001 0.2745 -2.000 -0.0237 0.03569 0.02539 -0.0502 1.0001 0.2939 -1.750 0.0031 0.03502 0.02466 -0.0502 1.0001 0.3181 -1.500 0.0314 0.03449 0.02404 -0.0504 1.0001 0.3492 -1.250 0.0589 0.03396 0.02368 -0.0505 1.0001 0.3903 -1.000 0.0868 0.03336 0.02353 -0.0506 1.0001 0.4531 -0.750 0.1122 0.03170 0.02353 -0.0491 1.0001 0.6546 -0.500 0.1186 0.03123 0.02313 -0.0455 1.0001 0.9999 -0.250 0.2689 0.02887 0.01990 -0.0666 0.9331 0.9999 0.000 0.3959 0.02424 0.01484 -0.0777 0.7771 0.9999 0.250 0.4357 0.02412 0.01348 -0.0760 0.6351 0.9999 0.500 0.4629 0.02502 0.01357 -0.0749 0.5786 0.9999 0.750 0.4914 0.02591 0.01392 -0.0746 0.5469 0.9999 1.000 0.5207 0.02676 0.01437 -0.0746 0.5245 0.9999 1.250 0.5505 0.02760 0.01489 -0.0749 0.5070 0.9999 1.500 0.5808 0.02846 0.01544 -0.0752 0.4932 0.9999 1.750 0.6117 0.02933 0.01618 -0.0759 0.4818 0.9999 2.000 0.6432 0.03030 0.01696 -0.0766 0.4736 0.9999 2.250 0.6737 0.03130 0.01795 -0.0773 0.4660 0.9999 2.500 0.7041 0.03234 0.01884 -0.0779 0.4594 0.9999 2.750 0.7335 0.03351 0.01997 -0.0785 0.4533 0.9999 3.000 0.7620 0.03470 0.02121 -0.0790 0.4466 0.9999 3.250 0.7906 0.03589 0.02231 -0.0794 0.4405 0.9999 3.500 0.8189 0.03728 0.02366 -0.0799 0.4358 0.9999 3.750 0.8455 0.03889 0.02546 -0.0805 0.4324 0.9999 4.000 0.8714 0.04063 0.02735 -0.0812 0.4293 0.9999 4.250 0.8966 0.04247 0.02932 -0.0817 0.4264 0.9999 4.500 0.9213 0.04431 0.03126 -0.0820 0.4231 0.9999 4.750 0.9471 0.04606 0.03298 -0.0822 0.4194 0.9999 5.000 0.9694 0.04829 0.03531 -0.0825 0.4163 0.9999