XFOIL Version 6.94 Calculated polar for: voyager2 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.012 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1799 0.03945 0.01812 -0.0241 1.0000 1.0000 -2.750 -0.1670 0.03907 0.01731 -0.0225 1.0000 1.0000 -2.250 -0.1395 0.03849 0.01574 -0.0191 1.0000 1.0000 -2.000 -0.1253 0.03827 0.01516 -0.0173 1.0000 1.0000 -1.750 -0.1110 0.03808 0.01465 -0.0156 1.0000 1.0000 -1.500 -0.0965 0.03793 0.01421 -0.0138 1.0000 1.0000 -1.250 -0.0820 0.03780 0.01384 -0.0120 1.0000 1.0000 -1.000 -0.0674 0.03770 0.01354 -0.0102 1.0000 1.0000 -0.750 -0.0528 0.03762 0.01330 -0.0084 1.0000 1.0000 -0.500 -0.0382 0.03756 0.01312 -0.0066 1.0000 1.0000 -0.250 -0.0236 0.03752 0.01300 -0.0048 1.0000 1.0000 0.000 -0.0090 0.03751 0.01285 -0.0029 1.0000 1.0000 0.250 0.0055 0.03751 0.01283 -0.0011 1.0000 1.0000 0.500 0.0201 0.03754 0.01287 0.0008 1.0000 1.0000 0.750 0.0347 0.03759 0.01297 0.0026 1.0000 1.0000 1.000 0.0492 0.03765 0.01314 0.0044 1.0000 1.0000 1.250 0.0637 0.03775 0.01337 0.0062 1.0000 1.0000 1.500 0.0782 0.03786 0.01366 0.0080 1.0000 1.0000 1.750 0.0926 0.03801 0.01411 0.0098 1.0000 1.0000 2.000 0.1069 0.03818 0.01450 0.0116 1.0000 1.0000 2.250 0.1210 0.03839 0.01497 0.0134 1.0000 1.0000 2.500 0.1350 0.03863 0.01552 0.0151 1.0000 1.0000 2.750 0.1486 0.03892 0.01615 0.0168 1.0000 1.0000 3.000 0.1619 0.03927 0.01688 0.0186 1.0000 1.0000 3.250 0.1746 0.03968 0.01772 0.0203 1.0000 1.0000 3.500 0.1866 0.04018 0.01867 0.0219 1.0000 1.0000 3.750 0.1974 0.04079 0.01976 0.0235 1.0000 1.0000 4.000 0.2064 0.04157 0.02101 0.0251 1.0000 1.0000 4.250 0.2129 0.04258 0.02245 0.0265 1.0000 1.0000 4.500 0.2161 0.04387 0.02442 0.0278 1.0000 1.0000 4.750 0.2174 0.04541 0.02624 0.0289 1.0000 1.0000 5.000 0.2196 0.04702 0.02809 0.0297 1.0000 1.0000