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 = 0.070 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1138 0.03339 0.02451 -0.0511 1.0001 0.1582 -2.750 -0.0855 0.03232 0.02310 -0.0514 1.0001 0.1661 -2.500 -0.0607 0.03178 0.02260 -0.0514 1.0001 0.1770 -2.250 -0.0340 0.03114 0.02186 -0.0515 1.0001 0.1886 -2.000 0.0190 0.02994 0.02068 -0.0564 0.9927 0.2110 -1.750 0.0912 0.02796 0.01875 -0.0643 0.9764 0.2482 -1.500 0.1634 0.02561 0.01668 -0.0717 0.9543 0.3039 -1.250 0.2361 0.02273 0.01457 -0.0787 0.9314 0.4202 -1.000 0.2810 0.01898 0.01246 -0.0776 0.8898 0.9999 -0.750 0.3230 0.01785 0.01075 -0.0769 0.7761 0.9999 -0.500 0.3468 0.01850 0.00978 -0.0733 0.5332 0.9999 -0.250 0.3713 0.01953 0.00989 -0.0724 0.4600 0.9999 0.000 0.3998 0.02024 0.01007 -0.0724 0.4325 0.9999 0.250 0.4292 0.02086 0.01028 -0.0726 0.4157 0.9999 0.500 0.4591 0.02140 0.01053 -0.0729 0.4034 0.9999 0.750 0.4893 0.02201 0.01084 -0.0732 0.3949 0.9999 1.000 0.5196 0.02256 0.01120 -0.0735 0.3874 0.9999 1.250 0.5497 0.02323 0.01159 -0.0739 0.3812 0.9999 1.500 0.5800 0.02393 0.01210 -0.0743 0.3760 0.9999 1.750 0.6100 0.02455 0.01264 -0.0746 0.3704 0.9999 2.000 0.6397 0.02526 0.01319 -0.0750 0.3650 0.9999 2.250 0.6696 0.02616 0.01385 -0.0754 0.3602 0.9999 2.500 0.6993 0.02695 0.01461 -0.0759 0.3565 0.9999 2.750 0.7289 0.02771 0.01538 -0.0763 0.3532 0.9999 3.000 0.7583 0.02853 0.01621 -0.0767 0.3501 0.9999 3.250 0.7876 0.02942 0.01709 -0.0771 0.3471 0.9999 3.500 0.8166 0.03038 0.01803 -0.0774 0.3441 0.9999 3.750 0.8453 0.03148 0.01905 -0.0778 0.3409 0.9999 4.000 0.8735 0.03278 0.02032 -0.0782 0.3379 0.9999 4.250 0.9010 0.03371 0.02145 -0.0784 0.3351 0.9999 4.500 0.9283 0.03483 0.02271 -0.0787 0.3327 0.9999 4.750 0.9553 0.03606 0.02408 -0.0789 0.3306 0.9999 5.000 0.9819 0.03739 0.02556 -0.0791 0.3287 0.9999