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 = 0.090 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0339 0.03212 0.02293 -0.0519 0.9980 0.1461 -2.750 0.0230 0.03022 0.02108 -0.0577 0.9904 0.1618 -2.500 0.0818 0.02845 0.01930 -0.0636 0.9807 0.1815 -2.000 0.2081 0.02414 0.01534 -0.0764 0.9502 0.2442 -1.750 0.2635 0.02191 0.01346 -0.0802 0.9128 0.2968 -1.500 0.2971 0.01916 0.01245 -0.0792 0.8439 0.6689 -1.250 0.3203 0.01789 0.01037 -0.0736 0.6118 1.0001 -1.000 0.3422 0.01928 0.01027 -0.0720 0.4320 1.0001 -0.750 0.3702 0.01995 0.01039 -0.0720 0.4002 1.0001 -0.500 0.3988 0.02050 0.01054 -0.0721 0.3847 1.0001 -0.250 0.4276 0.02103 0.01073 -0.0722 0.3739 1.0001 0.000 0.4569 0.02154 0.01097 -0.0724 0.3656 1.0001 0.250 0.4863 0.02205 0.01124 -0.0726 0.3579 1.0001 0.500 0.5156 0.02276 0.01162 -0.0729 0.3514 1.0001 0.750 0.5456 0.02325 0.01199 -0.0733 0.3461 1.0001 1.000 0.5755 0.02379 0.01241 -0.0737 0.3414 1.0001 1.250 0.6054 0.02441 0.01287 -0.0742 0.3372 1.0001 1.500 0.6353 0.02512 0.01341 -0.0746 0.3335 1.0001 1.750 0.6654 0.02609 0.01416 -0.0752 0.3302 1.0001 2.000 0.6953 0.02664 0.01473 -0.0756 0.3275 1.0001 2.250 0.7251 0.02726 0.01535 -0.0761 0.3241 1.0001 2.500 0.7547 0.02794 0.01600 -0.0765 0.3204 1.0001 2.750 0.7842 0.02870 0.01670 -0.0770 0.3170 1.0001 3.000 0.8137 0.02958 0.01749 -0.0775 0.3142 1.0001 3.250 0.8431 0.03070 0.01853 -0.0781 0.3119 1.0001 3.500 0.8723 0.03162 0.01952 -0.0785 0.3100 1.0001 3.750 0.9012 0.03244 0.02045 -0.0790 0.3080 1.0001 4.000 0.9299 0.03334 0.02146 -0.0794 0.3058 1.0001 4.250 0.9582 0.03427 0.02248 -0.0797 0.3031 1.0001 4.500 0.9864 0.03526 0.02352 -0.0801 0.3003 1.0001 4.750 1.0146 0.03632 0.02459 -0.0805 0.2978 1.0001 5.000 1.0426 0.03756 0.02584 -0.0809 0.2958 1.0001