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 = 0.100 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -2.750 0.0630 0.02766 0.01872 -0.0646 0.9803 0.1599 -2.500 0.1189 0.02591 0.01701 -0.0698 0.9687 0.1800 -2.250 0.1792 0.02403 0.01518 -0.0757 0.9543 0.2069 -2.000 0.2361 0.02188 0.01334 -0.0803 0.9250 0.2419 -1.750 0.2764 0.02040 0.01206 -0.0809 0.8615 0.2864 -1.500 0.3031 0.01917 0.01105 -0.0782 0.6894 0.4434 -1.250 0.3145 0.01830 0.00998 -0.0723 0.4381 1.0001 -1.000 0.3417 0.01908 0.01003 -0.0721 0.3929 1.0001 -0.750 0.3702 0.01962 0.01014 -0.0721 0.3755 1.0001 -0.500 0.3990 0.02011 0.01029 -0.0722 0.3647 1.0001 -0.250 0.4279 0.02060 0.01049 -0.0724 0.3560 1.0001 0.000 0.4568 0.02119 0.01076 -0.0726 0.3489 1.0001 0.250 0.4865 0.02162 0.01103 -0.0729 0.3424 1.0001 0.500 0.5160 0.02214 0.01134 -0.0732 0.3366 1.0001 0.750 0.5455 0.02280 0.01176 -0.0735 0.3320 1.0001 1.000 0.5754 0.02342 0.01222 -0.0740 0.3284 1.0001 1.250 0.6054 0.02391 0.01265 -0.0744 0.3249 1.0001 1.500 0.6353 0.02447 0.01312 -0.0748 0.3215 1.0001 1.750 0.6652 0.02509 0.01364 -0.0753 0.3183 1.0001 2.000 0.6950 0.02578 0.01420 -0.0758 0.3150 1.0001 2.250 0.7248 0.02673 0.01497 -0.0763 0.3118 1.0001 2.500 0.7544 0.02736 0.01561 -0.0768 0.3090 1.0001 2.750 0.7839 0.02798 0.01626 -0.0772 0.3061 1.0001 3.000 0.8134 0.02869 0.01698 -0.0776 0.3037 1.0001 3.250 0.8428 0.02945 0.01774 -0.0781 0.3013 1.0001 3.500 0.8720 0.03026 0.01856 -0.0785 0.2991 1.0001 3.750 0.9011 0.03116 0.01944 -0.0790 0.2970 1.0001 4.000 0.9301 0.03225 0.02047 -0.0795 0.2947 1.0001 4.250 0.9585 0.03339 0.02166 -0.0800 0.2925 1.0001 4.500 0.9865 0.03417 0.02262 -0.0803 0.2901 1.0001 4.750 1.0144 0.03513 0.02371 -0.0806 0.2878 1.0001 5.000 1.0421 0.03621 0.02490 -0.0809 0.2860 1.0001