XFOIL Version 6.94 Calculated polar for: K3311 (original) 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.300 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.500 0.4831 0.00802 0.00268 -0.0882 0.6890 1.0000 1.000 0.5314 0.00827 0.00270 -0.0867 0.6700 1.0000 1.500 0.5806 0.00857 0.00282 -0.0853 0.6510 1.0000 2.000 0.6297 0.00881 0.00298 -0.0839 0.6319 1.0000 2.500 0.6786 0.00899 0.00311 -0.0824 0.6101 1.0000 3.000 0.7285 0.00919 0.00325 -0.0811 0.5910 1.0000 3.500 0.7785 0.00945 0.00346 -0.0798 0.5722 1.0000 4.000 0.8284 0.00977 0.00372 -0.0786 0.5530 1.0000 4.500 0.8775 0.01011 0.00405 -0.0772 0.5300 1.0000 5.000 0.9259 0.01046 0.00441 -0.0757 0.5055 1.0000 5.500 0.9734 0.01086 0.00481 -0.0740 0.4788 1.0000 6.000 1.0207 0.01128 0.00528 -0.0724 0.4524 1.0000 6.500 1.0650 0.01174 0.00575 -0.0702 0.4151 1.0000 7.000 1.1041 0.01242 0.00629 -0.0672 0.3564 1.0000 7.500 1.1385 0.01353 0.00714 -0.0636 0.2929 1.0000 8.000 1.1668 0.01503 0.00829 -0.0593 0.2109 1.0000 8.500 1.1737 0.01783 0.01030 -0.0521 0.0880 1.0000 9.000 1.1719 0.02068 0.01281 -0.0432 0.0320 1.0000 10.000 1.1912 0.02561 0.01809 -0.0316 0.0227 1.0000 10.500 1.1963 0.02882 0.02149 -0.0268 0.0211 1.0000 11.000 1.1892 0.03360 0.02644 -0.0221 0.0195 1.0000 11.500 1.2002 0.03735 0.03039 -0.0192 0.0186 1.0000 12.000 1.2140 0.04105 0.03431 -0.0169 0.0173 1.0000 12.500 1.2274 0.04516 0.03863 -0.0148 0.0162 1.0000 13.000 1.2408 0.04968 0.04345 -0.0128 0.0158 1.0000 13.500 1.2491 0.05496 0.04904 -0.0112 0.0153 1.0000 14.000 1.2486 0.06140 0.05584 -0.0104 0.0150 1.0000 14.500 1.2380 0.06942 0.06431 -0.0105 0.0152 1.0000 15.000 1.2088 0.08024 0.07567 -0.0128 0.0155 1.0000 15.500 1.1711 0.09332 0.08923 -0.0178 0.0159 1.0000 16.000 1.1288 0.10875 0.10509 -0.0257 0.0164 1.0000 16.500 1.0844 0.12667 0.12338 -0.0362 0.0170 1.0000 17.000 1.0366 0.14815 0.14515 -0.0498 0.0177 1.0000 17.500 0.9997 0.16884 0.16595 -0.0611 0.0194 1.0000