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 = 0.060 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1182 0.03527 0.02624 -0.0508 1.0001 0.1759 -2.750 -0.0892 0.03417 0.02467 -0.0512 1.0001 0.1843 -2.500 -0.0643 0.03328 0.02389 -0.0512 1.0001 0.1956 -2.250 -0.0370 0.03253 0.02297 -0.0515 1.0001 0.2082 -2.000 -0.0108 0.03219 0.02248 -0.0515 1.0001 0.2242 -1.750 0.0143 0.03176 0.02216 -0.0514 1.0001 0.2432 -1.500 0.0768 0.03021 0.02077 -0.0578 0.9877 0.2830 -1.250 0.1618 0.02760 0.01855 -0.0674 0.9653 0.3641 -1.000 0.2164 0.02296 0.01647 -0.0691 0.9401 0.9999 -0.750 0.3000 0.02077 0.01351 -0.0766 0.8910 0.9999 -0.500 0.3484 0.01924 0.01145 -0.0764 0.7584 0.9999 -0.250 0.3741 0.01987 0.01052 -0.0731 0.5479 0.9999 0.000 0.3989 0.02085 0.01065 -0.0722 0.4859 0.9999 0.250 0.4269 0.02163 0.01087 -0.0721 0.4583 0.9999 0.500 0.4569 0.02228 0.01114 -0.0724 0.4398 0.9999 0.750 0.4876 0.02289 0.01146 -0.0728 0.4264 0.9999 1.000 0.5184 0.02359 0.01181 -0.0732 0.4171 0.9999 1.250 0.5493 0.02418 0.01227 -0.0737 0.4089 0.9999 1.500 0.5799 0.02488 0.01274 -0.0742 0.4022 0.9999 1.750 0.6105 0.02568 0.01331 -0.0746 0.3966 0.9999 2.000 0.6406 0.02636 0.01396 -0.0751 0.3907 0.9999 2.250 0.6704 0.02713 0.01461 -0.0754 0.3848 0.9999 2.500 0.7002 0.02808 0.01533 -0.0759 0.3797 0.9999 2.750 0.7297 0.02895 0.01619 -0.0763 0.3756 0.9999 3.000 0.7591 0.02983 0.01712 -0.0767 0.3721 0.9999 3.250 0.7883 0.03080 0.01810 -0.0772 0.3688 0.9999 3.500 0.8174 0.03183 0.01913 -0.0776 0.3657 0.9999 3.750 0.8461 0.03294 0.02021 -0.0780 0.3626 0.9999 4.000 0.8745 0.03430 0.02148 -0.0784 0.3594 0.9999 4.250 0.9015 0.03547 0.02283 -0.0787 0.3564 0.9999 4.500 0.9281 0.03672 0.02427 -0.0789 0.3533 0.9999 4.750 0.9543 0.03812 0.02583 -0.0791 0.3510 0.9999 5.000 0.9800 0.03966 0.02752 -0.0794 0.3490 0.9999