XFOIL Version 6.94 Calculated polar for: E211 (10.96%) 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.450 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.4280 0.00695 0.00205 -0.1051 0.7930 0.7016 0.500 0.4816 0.00712 0.00210 -0.1044 0.7567 0.7214 1.000 0.5340 0.00731 0.00220 -0.1034 0.7229 0.7416 1.500 0.5859 0.00749 0.00233 -0.1024 0.6906 0.7623 2.000 0.6373 0.00768 0.00247 -0.1013 0.6585 0.7844 2.500 0.6878 0.00784 0.00267 -0.1000 0.6286 0.8092 3.000 0.7369 0.00801 0.00287 -0.0984 0.5981 0.8405 3.500 0.7826 0.00816 0.00310 -0.0960 0.5678 0.8848 4.000 0.8314 0.00823 0.00325 -0.0941 0.5348 1.0000 4.500 0.8836 0.00861 0.00355 -0.0936 0.4951 1.0000 5.000 0.9299 0.00924 0.00391 -0.0919 0.4258 1.0000 5.500 0.9747 0.01006 0.00442 -0.0900 0.3510 1.0000 6.000 1.0127 0.01149 0.00527 -0.0872 0.2314 1.0000 6.500 1.0506 0.01302 0.00628 -0.0845 0.1373 1.0000 7.000 1.0886 0.01448 0.00738 -0.0817 0.0705 1.0000 7.500 1.1197 0.01646 0.00903 -0.0776 0.0184 1.0000 8.000 1.1555 0.01793 0.01061 -0.0741 0.0145 1.0000 8.500 1.1877 0.01932 0.01214 -0.0702 0.0127 1.0000 9.000 1.2114 0.02114 0.01413 -0.0649 0.0118 1.0000 9.500 1.2235 0.02378 0.01698 -0.0583 0.0111 1.0000 10.000 1.2371 0.02669 0.02011 -0.0526 0.0108 1.0000 10.500 1.2553 0.02955 0.02321 -0.0480 0.0105 1.0000 11.000 1.2723 0.03305 0.02698 -0.0438 0.0104 1.0000 11.500 1.2869 0.03726 0.03154 -0.0400 0.0103 1.0000 12.000 1.2944 0.04243 0.03711 -0.0363 0.0103 1.0000 12.500 1.2896 0.04898 0.04411 -0.0330 0.0105 1.0000 13.500 1.2602 0.06483 0.06074 -0.0306 0.0111 1.0000 14.000 1.2390 0.07348 0.06974 -0.0324 0.0112 1.0000 14.500 1.2090 0.08478 0.08148 -0.0370 0.0116 1.0000