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 = 0.045 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.3891 0.02444 0.01107 0.0092 1.0000 0.0919 -2.750 -0.3665 0.02362 0.01012 0.0107 1.0000 0.0961 -2.500 -0.3451 0.02279 0.00932 0.0124 1.0000 0.1079 -2.250 -0.3311 0.02138 0.00884 0.0148 1.0000 0.3058 -2.000 -0.3207 0.02020 0.00843 0.0182 1.0000 0.4384 -1.250 -0.0407 0.02071 0.00867 -0.0156 1.0000 1.0000 -1.000 -0.0247 0.02067 0.00843 -0.0137 1.0000 1.0000 -0.750 -0.0092 0.02065 0.00831 -0.0118 1.0000 1.0000 -0.500 0.0059 0.02064 0.00825 -0.0098 1.0000 1.0000 -0.250 0.0204 0.02066 0.00823 -0.0077 1.0000 1.0000 0.000 0.0345 0.02069 0.00825 -0.0055 1.0000 1.0000 0.250 0.0482 0.02074 0.00832 -0.0033 1.0000 1.0000 0.500 0.0616 0.02080 0.00845 -0.0010 1.0000 1.0000 0.750 0.0746 0.02089 0.00859 0.0013 1.0000 1.0000 1.000 0.0873 0.02098 0.00878 0.0036 1.0000 1.0000 1.250 0.0999 0.02110 0.00901 0.0060 1.0000 1.0000 1.500 0.1124 0.02123 0.00929 0.0084 1.0000 1.0000 1.750 0.1247 0.02138 0.00962 0.0108 1.0000 1.0000 2.000 0.1370 0.02155 0.01002 0.0131 1.0000 1.0000 2.250 0.1493 0.02174 0.01047 0.0155 1.0000 1.0000 2.500 0.1618 0.02195 0.01105 0.0178 1.0000 1.0000 3.250 0.4262 0.02454 0.01082 -0.0099 0.0876 1.0000 3.500 0.4485 0.02529 0.01177 -0.0085 0.0851 1.0000 3.750 0.4712 0.02613 0.01284 -0.0071 0.0843 1.0000 4.000 0.4934 0.02701 0.01421 -0.0058 0.0838 1.0000 4.250 0.5165 0.02803 0.01545 -0.0043 0.0847 1.0000 4.500 0.5427 0.02920 0.01683 -0.0031 0.0864 1.0000 4.750 0.5724 0.03067 0.01847 -0.0023 0.0890 1.0000 5.000 0.6039 0.03272 0.02061 -0.0018 0.0919 1.0000