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 = 0.035 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1356 0.04241 0.03248 -0.0482 1.0001 0.2624 -2.750 -0.1105 0.04067 0.03064 -0.0484 1.0001 0.2731 -2.500 -0.0807 0.03924 0.02877 -0.0494 1.0001 0.2863 -2.250 -0.0556 0.03806 0.02754 -0.0493 1.0001 0.3037 -2.000 -0.0291 0.03705 0.02641 -0.0493 1.0001 0.3239 -1.750 -0.0008 0.03623 0.02548 -0.0495 1.0001 0.3505 -1.500 0.0270 0.03554 0.02482 -0.0496 1.0001 0.3856 -1.250 0.0549 0.03486 0.02435 -0.0497 1.0001 0.4328 -1.000 0.0823 0.03403 0.02409 -0.0495 1.0001 0.5097 -0.750 0.0939 0.03178 0.02389 -0.0439 1.0001 0.8133 -0.500 0.1155 0.03218 0.02335 -0.0447 1.0001 0.9999 -0.250 0.1367 0.03326 0.02413 -0.0451 1.0001 0.9999 0.000 0.3419 0.02925 0.01950 -0.0747 0.8889 0.9999 0.250 0.4359 0.02599 0.01546 -0.0794 0.7159 0.9999 0.500 0.4678 0.02649 0.01503 -0.0778 0.6374 0.9999 0.750 0.4965 0.02736 0.01524 -0.0770 0.5968 0.9999 1.000 0.5256 0.02828 0.01572 -0.0769 0.5697 0.9999 1.250 0.5550 0.02920 0.01629 -0.0770 0.5491 0.9999 1.500 0.5847 0.03012 0.01691 -0.0772 0.5324 0.9999 1.750 0.6143 0.03107 0.01763 -0.0775 0.5181 0.9999 2.000 0.6439 0.03213 0.01861 -0.0781 0.5068 0.9999 2.250 0.6750 0.03320 0.01950 -0.0788 0.4983 0.9999 2.500 0.7049 0.03446 0.02075 -0.0797 0.4911 0.9999 2.750 0.7340 0.03576 0.02208 -0.0804 0.4842 0.9999 3.000 0.7634 0.03701 0.02319 -0.0809 0.4778 0.9999 3.250 0.7901 0.03858 0.02485 -0.0816 0.4714 0.9999 3.500 0.8164 0.04017 0.02652 -0.0821 0.4650 0.9999 3.750 0.8437 0.04170 0.02801 -0.0825 0.4599 0.9999 4.000 0.8709 0.04347 0.02973 -0.0830 0.4566 0.9999 4.250 0.8927 0.04588 0.03239 -0.0838 0.4543 0.9999 4.500 0.9121 0.04858 0.03534 -0.0844 0.4521 0.9999 4.750 0.9284 0.05156 0.03854 -0.0849 0.4496 0.9999 5.000 0.9415 0.05484 0.04201 -0.0852 0.4469 0.9999