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 -0.0468 0.03555 0.02567 -0.0513 0.9999 0.1887 -2.750 -0.0215 0.03491 0.02507 -0.0512 0.9999 0.2002 -2.500 0.0056 0.03431 0.02436 -0.0513 0.9999 0.2125 -2.250 0.0314 0.03412 0.02407 -0.0512 0.9999 0.2285 -2.000 0.0570 0.03380 0.02386 -0.0512 0.9999 0.2472 -1.750 0.0857 0.03350 0.02368 -0.0516 0.9989 0.2711 -1.500 0.1702 0.03079 0.02140 -0.0617 0.9804 0.3444 -1.250 0.2312 0.02605 0.01883 -0.0655 0.9568 1.0001 -1.000 0.3201 0.02325 0.01548 -0.0748 0.9112 1.0001 -0.750 0.3745 0.02120 0.01295 -0.0758 0.7782 1.0001 -0.500 0.4000 0.02174 0.01188 -0.0722 0.5475 1.0001 -0.250 0.4247 0.02277 0.01204 -0.0713 0.4824 1.0001 0.000 0.4529 0.02353 0.01230 -0.0713 0.4557 1.0001 0.250 0.4832 0.02418 0.01260 -0.0717 0.4389 1.0001 0.500 0.5141 0.02485 0.01294 -0.0722 0.4271 1.0001 0.750 0.5449 0.02546 0.01333 -0.0727 0.4167 1.0001 1.000 0.5756 0.02619 0.01379 -0.0732 0.4083 1.0001 1.250 0.6061 0.02684 0.01431 -0.0737 0.4002 1.0001 1.500 0.6365 0.02763 0.01487 -0.0742 0.3933 1.0001 1.750 0.6670 0.02846 0.01556 -0.0748 0.3882 1.0001 2.000 0.6973 0.02923 0.01632 -0.0754 0.3836 1.0001 2.250 0.7275 0.03009 0.01711 -0.0759 0.3793 1.0001 2.500 0.7574 0.03102 0.01796 -0.0765 0.3754 1.0001 2.750 0.7873 0.03217 0.01895 -0.0770 0.3716 1.0001 3.000 0.8166 0.03313 0.01999 -0.0776 0.3680 1.0001 3.250 0.8456 0.03412 0.02107 -0.0781 0.3639 1.0001 3.500 0.8745 0.03520 0.02221 -0.0786 0.3602 1.0001 3.750 0.9032 0.03639 0.02342 -0.0791 0.3574 1.0001 4.000 0.9317 0.03767 0.02474 -0.0796 0.3549 1.0001 4.250 0.9599 0.03913 0.02620 -0.0801 0.3527 1.0001 4.500 0.9873 0.04083 0.02796 -0.0806 0.3504 1.0001 4.750 1.0132 0.04225 0.02966 -0.0810 0.3481 1.0001 5.000 1.0383 0.04385 0.03151 -0.0813 0.3454 1.0001