XFOIL Version 6.94 Calculated polar for: Voyager1 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.050 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.4016 0.02400 0.01117 0.0115 1.0000 0.0823 -2.750 -0.3771 0.02322 0.01020 0.0127 1.0000 0.0854 -2.500 -0.3531 0.02257 0.00934 0.0139 1.0000 0.0887 -2.250 -0.3317 0.02176 0.00847 0.0155 1.0000 0.1014 -2.000 -0.3172 0.02044 0.00819 0.0178 1.0000 0.3086 -1.750 -0.3069 0.01926 0.00781 0.0212 1.0000 0.4470 -0.750 -0.0072 0.01989 0.00821 -0.0121 1.0000 1.0000 -0.500 0.0077 0.01989 0.00815 -0.0101 1.0000 1.0000 -0.250 0.0221 0.01991 0.00814 -0.0080 1.0000 1.0000 0.000 0.0359 0.01994 0.00817 -0.0058 1.0000 1.0000 0.250 0.0493 0.01999 0.00824 -0.0035 1.0000 1.0000 0.500 0.0624 0.02006 0.00837 -0.0011 1.0000 1.0000 0.750 0.0751 0.02014 0.00851 0.0012 1.0000 1.0000 1.000 0.0875 0.02024 0.00869 0.0037 1.0000 1.0000 1.250 0.0998 0.02035 0.00892 0.0061 1.0000 1.0000 1.500 0.1119 0.02049 0.00920 0.0085 1.0000 1.0000 1.750 0.1240 0.02064 0.00953 0.0110 1.0000 1.0000 2.000 0.1360 0.02080 0.00992 0.0134 1.0000 1.0000 2.250 0.1482 0.02099 0.01037 0.0158 1.0000 1.0000 2.750 0.3876 0.02267 0.00915 -0.0138 0.0868 1.0000 3.000 0.4102 0.02337 0.00997 -0.0125 0.0803 1.0000 3.250 0.4340 0.02416 0.01094 -0.0114 0.0777 1.0000 3.500 0.4574 0.02496 0.01195 -0.0103 0.0771 1.0000 3.750 0.4802 0.02588 0.01306 -0.0090 0.0772 1.0000 4.000 0.5029 0.02683 0.01449 -0.0076 0.0775 1.0000 4.250 0.5269 0.02769 0.01553 -0.0061 0.0790 1.0000 4.500 0.5531 0.02867 0.01673 -0.0048 0.0816 1.0000 4.750 0.5818 0.03006 0.01830 -0.0037 0.0853 1.0000 5.000 0.6109 0.03196 0.02038 -0.0027 0.0896 1.0000