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 = 0.040 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0566 0.03926 0.02860 -0.0512 0.9999 0.2526 -2.750 -0.0309 0.03834 0.02757 -0.0510 0.9999 0.2664 -2.500 -0.0050 0.03748 0.02664 -0.0508 0.9999 0.2820 -2.250 0.0220 0.03690 0.02596 -0.0507 0.9999 0.3028 -2.000 0.0499 0.03647 0.02548 -0.0507 0.9999 0.3292 -1.750 0.0777 0.03603 0.02520 -0.0507 0.9999 0.3637 -1.500 0.1061 0.03560 0.02508 -0.0509 0.9999 0.4179 -1.000 0.1409 0.03347 0.02479 -0.0459 0.9999 1.0001 -0.750 0.1636 0.03447 0.02542 -0.0459 0.9999 1.0001 -0.500 0.2772 0.03232 0.02278 -0.0617 0.9548 1.0001 -0.250 0.4203 0.02651 0.01665 -0.0773 0.8117 1.0001 0.000 0.4634 0.02597 0.01488 -0.0757 0.6429 1.0001 0.250 0.4897 0.02691 0.01494 -0.0743 0.5788 1.0001 0.500 0.5176 0.02784 0.01528 -0.0738 0.5445 1.0001 0.750 0.5472 0.02872 0.01577 -0.0740 0.5220 1.0001 1.000 0.5777 0.02961 0.01634 -0.0745 0.5063 1.0001 1.250 0.6087 0.03049 0.01702 -0.0751 0.4935 1.0001 1.500 0.6398 0.03142 0.01781 -0.0758 0.4830 1.0001 1.750 0.6711 0.03237 0.01858 -0.0765 0.4731 1.0001 2.000 0.7017 0.03340 0.01953 -0.0773 0.4644 1.0001 2.250 0.7318 0.03444 0.02050 -0.0779 0.4561 1.0001 2.500 0.7622 0.03561 0.02147 -0.0785 0.4500 1.0001 2.750 0.7916 0.03688 0.02290 -0.0794 0.4450 1.0001 3.000 0.8208 0.03824 0.02434 -0.0802 0.4405 1.0001 3.250 0.8496 0.03964 0.02576 -0.0809 0.4361 1.0001 3.500 0.8782 0.04109 0.02714 -0.0814 0.4317 1.0001 3.750 0.9050 0.04279 0.02899 -0.0821 0.4275 1.0001 4.000 0.9303 0.04463 0.03104 -0.0828 0.4232 1.0001 4.250 0.9554 0.04660 0.03316 -0.0834 0.4198 1.0001 4.500 0.9796 0.04878 0.03552 -0.0840 0.4177 1.0001 4.750 1.0024 0.05118 0.03810 -0.0846 0.4161 1.0001 5.000 1.0244 0.05367 0.04073 -0.0851 0.4141 1.0001