XFOIL Version 6.94 Calculated polar for: NACA 4412 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.045 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.2566 0.04105 0.03182 -0.0508 1.0000 0.2114 -2.750 -0.2331 0.03961 0.03016 -0.0515 1.0000 0.2195 -2.500 -0.2074 0.03818 0.02835 -0.0526 1.0000 0.2305 -2.250 -0.1837 0.03729 0.02719 -0.0532 1.0000 0.2436 -2.000 -0.1616 0.03655 0.02639 -0.0534 1.0000 0.2587 -1.750 -0.1389 0.03596 0.02565 -0.0538 1.0000 0.2764 -1.500 -0.1164 0.03557 0.02516 -0.0542 1.0000 0.2989 -1.250 -0.0761 0.03528 0.02480 -0.0574 0.9942 0.3351 -1.000 -0.0238 0.03506 0.02459 -0.0625 0.9833 0.3946 -0.750 0.0225 0.03470 0.02461 -0.0664 0.9723 0.4817 -0.500 0.0581 0.03328 0.02483 -0.0663 0.9639 0.7321 -0.250 0.1027 0.03370 0.02467 -0.0705 0.9492 1.0000 0.000 0.1352 0.03462 0.02517 -0.0727 0.9359 1.0000 0.250 0.1682 0.03564 0.02585 -0.0750 0.9232 1.0000 0.500 0.2093 0.03678 0.02668 -0.0784 0.9118 1.0000 0.750 0.2391 0.03775 0.02744 -0.0800 0.8996 1.0000 1.000 0.2656 0.03878 0.02828 -0.0811 0.8878 1.0000 1.250 0.3110 0.03990 0.02920 -0.0849 0.8776 1.0000 1.500 0.3244 0.04091 0.03011 -0.0839 0.8655 1.0000 1.750 0.3547 0.04202 0.03109 -0.0854 0.8549 1.0000 2.000 0.3826 0.04312 0.03208 -0.0865 0.8443 1.0000 2.250 0.4024 0.04433 0.03322 -0.0865 0.8338 1.0000 2.500 0.4396 0.04540 0.03420 -0.0888 0.8238 1.0000 2.750 0.4493 0.04678 0.03554 -0.0875 0.8133 1.0000 3.250 0.4944 0.04938 0.03807 -0.0882 0.7933 1.0000 3.500 0.5299 0.05056 0.03922 -0.0901 0.7839 1.0000 3.750 0.5376 0.05215 0.04081 -0.0887 0.7734 1.0000 4.000 0.5653 0.05351 0.04216 -0.0896 0.7639 1.0000 4.250 0.5784 0.05512 0.04378 -0.0890 0.7537 1.0000 4.500 0.5992 0.05667 0.04536 -0.0891 0.7439 1.0000 4.750 0.6182 0.05825 0.04696 -0.0891 0.7339 1.0000 5.000 0.6314 0.06007 0.04881 -0.0885 0.7238 1.0000