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 = 0.045 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0532 0.03813 0.02774 -0.0514 0.9999 0.2319 -2.750 -0.0277 0.03729 0.02679 -0.0512 0.9999 0.2441 -2.500 -0.0016 0.03652 0.02597 -0.0511 0.9999 0.2594 -2.250 0.0248 0.03597 0.02535 -0.0510 0.9999 0.2776 -2.000 0.0521 0.03560 0.02497 -0.0509 0.9999 0.3009 -1.750 0.0797 0.03536 0.02476 -0.0510 0.9999 0.3321 -1.500 0.1076 0.03506 0.02474 -0.0511 0.9999 0.3764 -1.250 0.1376 0.03432 0.02490 -0.0517 0.9999 0.4731 -1.000 0.1433 0.03271 0.02471 -0.0465 0.9999 1.0001 -0.750 0.2446 0.03100 0.02215 -0.0598 0.9666 1.0001 -0.500 0.3711 0.02645 0.01729 -0.0749 0.8879 1.0001 -0.250 0.4332 0.02415 0.01410 -0.0755 0.6821 1.0001 0.000 0.4584 0.02505 0.01377 -0.0731 0.5727 1.0001 0.250 0.4851 0.02601 0.01401 -0.0725 0.5296 1.0001 0.500 0.5143 0.02685 0.01443 -0.0726 0.5043 1.0001 0.750 0.5445 0.02767 0.01488 -0.0729 0.4875 1.0001 1.000 0.5755 0.02846 0.01545 -0.0735 0.4742 1.0001 1.250 0.6073 0.02928 0.01608 -0.0743 0.4635 1.0001 1.500 0.6386 0.03012 0.01673 -0.0750 0.4537 1.0001 1.750 0.6695 0.03103 0.01751 -0.0757 0.4451 1.0001 2.000 0.6999 0.03194 0.01833 -0.0763 0.4370 1.0001 2.250 0.7304 0.03300 0.01918 -0.0769 0.4308 1.0001 2.500 0.7605 0.03406 0.02031 -0.0776 0.4260 1.0001 2.750 0.7902 0.03520 0.02151 -0.0783 0.4214 1.0001 3.000 0.8197 0.03640 0.02272 -0.0790 0.4171 1.0001 3.250 0.8489 0.03767 0.02392 -0.0795 0.4128 1.0001 3.500 0.8772 0.03910 0.02537 -0.0801 0.4087 1.0001 3.750 0.9044 0.04056 0.02704 -0.0808 0.4044 1.0001 4.000 0.9314 0.04213 0.02873 -0.0813 0.4006 1.0001 4.250 0.9580 0.04384 0.03057 -0.0819 0.3981 1.0001 4.500 0.9841 0.04568 0.03254 -0.0825 0.3959 1.0001 4.750 1.0098 0.04761 0.03455 -0.0829 0.3938 1.0001 5.000 1.0358 0.04963 0.03658 -0.0833 0.3912 1.0001