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 -0.1219 0.03754 0.02816 -0.0505 1.0001 0.1999 -2.750 -0.0959 0.03618 0.02664 -0.0507 1.0001 0.2097 -2.500 -0.0688 0.03520 0.02542 -0.0509 1.0001 0.2211 -2.250 -0.0426 0.03432 0.02453 -0.0510 1.0001 0.2354 -2.000 -0.0160 0.03362 0.02374 -0.0511 1.0001 0.2520 -1.750 0.0102 0.03310 0.02321 -0.0512 1.0001 0.2725 -1.500 0.0358 0.03273 0.02291 -0.0511 1.0001 0.2985 -1.250 0.0629 0.03245 0.02272 -0.0512 1.0001 0.3318 -1.000 0.1220 0.03097 0.02170 -0.0568 0.9888 0.4136 -0.750 0.1984 0.02646 0.01898 -0.0625 0.9583 0.9999 -0.500 0.3044 0.02367 0.01563 -0.0742 0.9017 0.9999 -0.250 0.3725 0.02120 0.01268 -0.0764 0.7598 0.9999 0.000 0.4033 0.02158 0.01162 -0.0737 0.5819 0.9999 0.250 0.4289 0.02255 0.01175 -0.0727 0.5238 0.9999 0.500 0.4567 0.02339 0.01200 -0.0724 0.4949 0.9999 0.750 0.4861 0.02413 0.01234 -0.0726 0.4747 0.9999 1.000 0.5168 0.02484 0.01274 -0.0730 0.4591 0.9999 1.250 0.5484 0.02554 0.01324 -0.0737 0.4475 0.9999 1.500 0.5798 0.02632 0.01374 -0.0742 0.4387 0.9999 1.750 0.6109 0.02709 0.01440 -0.0749 0.4314 0.9999 2.000 0.6416 0.02789 0.01509 -0.0754 0.4247 0.9999 2.250 0.6721 0.02882 0.01579 -0.0759 0.4187 0.9999 2.500 0.7017 0.02971 0.01668 -0.0764 0.4128 0.9999 2.750 0.7310 0.03063 0.01758 -0.0768 0.4067 0.9999 3.000 0.7604 0.03164 0.01847 -0.0772 0.4015 0.9999 3.250 0.7898 0.03284 0.01954 -0.0777 0.3978 0.9999 3.500 0.8182 0.03399 0.02085 -0.0782 0.3948 0.9999 3.750 0.8463 0.03525 0.02223 -0.0787 0.3918 0.9999 4.000 0.8738 0.03657 0.02366 -0.0791 0.3886 0.9999 4.250 0.9010 0.03792 0.02507 -0.0794 0.3850 0.9999 4.500 0.9282 0.03932 0.02643 -0.0797 0.3814 0.9999 4.750 0.9546 0.04100 0.02810 -0.0800 0.3782 0.9999 5.000 0.9781 0.04282 0.03020 -0.0802 0.3762 0.9999