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 = 0.035 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0625 0.04058 0.02974 -0.0504 0.9999 0.2791 -2.750 -0.0349 0.03954 0.02841 -0.0506 0.9999 0.2938 -2.500 -0.0079 0.03877 0.02741 -0.0505 0.9999 0.3123 -2.250 0.0194 0.03804 0.02662 -0.0504 0.9999 0.3350 -2.000 0.0466 0.03740 0.02609 -0.0503 0.9999 0.3646 -1.750 0.0748 0.03684 0.02573 -0.0503 0.9999 0.4065 -1.500 0.1032 0.03611 0.02552 -0.0503 0.9999 0.4773 -1.250 0.1105 0.03367 0.02534 -0.0447 0.9999 0.9221 -1.000 0.1378 0.03442 0.02497 -0.0453 0.9999 1.0001 -0.750 0.1603 0.03539 0.02559 -0.0452 0.9999 1.0001 -0.500 0.1807 0.03656 0.02660 -0.0452 0.9999 1.0001 -0.250 0.3345 0.03321 0.02288 -0.0681 0.9284 1.0001 0.000 0.4639 0.02793 0.01697 -0.0794 0.7345 1.0001 0.250 0.4960 0.02836 0.01640 -0.0775 0.6433 1.0001 0.750 0.5523 0.03018 0.01701 -0.0762 0.5678 1.0001 1.000 0.5819 0.03113 0.01763 -0.0764 0.5465 1.0001 1.250 0.6124 0.03211 0.01833 -0.0769 0.5312 1.0001 1.500 0.6430 0.03316 0.01928 -0.0776 0.5185 1.0001 1.750 0.6739 0.03422 0.02012 -0.0782 0.5084 1.0001 2.000 0.7039 0.03536 0.02127 -0.0791 0.4982 1.0001 2.250 0.7346 0.03649 0.02217 -0.0797 0.4895 1.0001 2.500 0.7635 0.03781 0.02360 -0.0806 0.4809 1.0001 2.750 0.7929 0.03911 0.02486 -0.0812 0.4740 1.0001 3.000 0.8225 0.04059 0.02625 -0.0820 0.4693 1.0001 3.250 0.8497 0.04240 0.02827 -0.0831 0.4652 1.0001 3.500 0.8763 0.04429 0.03030 -0.0840 0.4609 1.0001 3.750 0.9026 0.04613 0.03223 -0.0847 0.4563 1.0001 4.000 0.9299 0.04783 0.03387 -0.0851 0.4517 1.0001 4.250 0.9526 0.05020 0.03640 -0.0857 0.4478 1.0001 4.500 0.9719 0.05309 0.03955 -0.0865 0.4453 1.0001 4.750 0.9884 0.05643 0.04317 -0.0873 0.4443 1.0001 5.000 1.0001 0.06040 0.04739 -0.0881 0.4441 1.0001