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 = 0.050 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0514 0.03714 0.02704 -0.0513 0.9999 0.2145 -2.750 -0.0245 0.03648 0.02612 -0.0513 0.9999 0.2267 -2.500 0.0011 0.03572 0.02540 -0.0512 0.9999 0.2410 -2.250 0.0276 0.03520 0.02485 -0.0512 0.9999 0.2577 -2.000 0.0537 0.03486 0.02454 -0.0510 0.9999 0.2791 -1.750 0.0809 0.03469 0.02442 -0.0510 0.9999 0.3070 -1.500 0.1085 0.03449 0.02446 -0.0512 0.9999 0.3444 -1.250 0.1377 0.03417 0.02467 -0.0518 0.9999 0.4124 -1.000 0.2122 0.02980 0.02169 -0.0582 0.9734 1.0001 -0.750 0.3210 0.02654 0.01797 -0.0716 0.9249 1.0001 -0.500 0.3987 0.02324 0.01428 -0.0760 0.7875 1.0001 -0.250 0.4299 0.02343 0.01298 -0.0730 0.5857 1.0001 0.000 0.4550 0.02446 0.01312 -0.0719 0.5210 1.0001 0.250 0.4829 0.02532 0.01343 -0.0717 0.4918 1.0001 0.500 0.5127 0.02607 0.01381 -0.0720 0.4729 1.0001 0.750 0.5437 0.02679 0.01427 -0.0726 0.4591 1.0001 1.000 0.5755 0.02757 0.01475 -0.0733 0.4485 1.0001 1.250 0.6068 0.02830 0.01535 -0.0740 0.4384 1.0001 1.500 0.6378 0.02918 0.01593 -0.0745 0.4301 1.0001 1.750 0.6683 0.02995 0.01670 -0.0752 0.4220 1.0001 2.000 0.6987 0.03083 0.01745 -0.0757 0.4154 1.0001 2.250 0.7293 0.03190 0.01831 -0.0763 0.4106 1.0001 2.500 0.7593 0.03286 0.01935 -0.0770 0.4063 1.0001 2.750 0.7891 0.03391 0.02045 -0.0777 0.4022 1.0001 3.000 0.8185 0.03501 0.02155 -0.0783 0.3978 1.0001 3.250 0.8477 0.03620 0.02266 -0.0788 0.3935 1.0001 3.500 0.8763 0.03755 0.02400 -0.0793 0.3895 1.0001 3.750 0.9041 0.03886 0.02550 -0.0799 0.3857 1.0001 4.000 0.9317 0.04032 0.02711 -0.0805 0.3829 1.0001 4.250 0.9587 0.04190 0.02883 -0.0810 0.3805 1.0001 4.500 0.9852 0.04359 0.03067 -0.0815 0.3783 1.0001 4.750 1.0114 0.04530 0.03248 -0.0820 0.3758 1.0001 5.000 1.0378 0.04708 0.03425 -0.0823 0.3729 1.0001