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 = 0.060 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.4252 0.02385 0.01181 0.0165 1.0000 0.0576 -2.750 -0.4017 0.02289 0.01072 0.0177 1.0000 0.0602 -2.500 -0.3752 0.02185 0.00953 0.0181 1.0000 0.0679 -2.250 -0.3500 0.02125 0.00862 0.0189 1.0000 0.0742 -2.000 -0.3253 0.02060 0.00785 0.0199 1.0000 0.0808 -1.750 -0.3056 0.01934 0.00743 0.0213 1.0000 0.2581 -1.500 -0.2913 0.01844 0.00718 0.0239 1.0000 0.3826 -1.000 -0.0188 0.01869 0.00815 -0.0147 1.0000 1.0000 -0.750 -0.0036 0.01868 0.00805 -0.0127 1.0000 1.0000 -0.500 0.0110 0.01868 0.00800 -0.0106 1.0000 1.0000 -0.250 0.0250 0.01870 0.00800 -0.0084 1.0000 1.0000 0.000 0.0384 0.01874 0.00803 -0.0061 1.0000 1.0000 0.250 0.0513 0.01879 0.00811 -0.0038 1.0000 1.0000 0.500 0.0638 0.01886 0.00823 -0.0013 1.0000 1.0000 0.750 0.0759 0.01895 0.00838 0.0012 1.0000 1.0000 1.000 0.0878 0.01905 0.00856 0.0037 1.0000 1.0000 1.250 0.0995 0.01917 0.00879 0.0062 1.0000 1.0000 1.500 0.1110 0.01930 0.00907 0.0088 1.0000 1.0000 1.750 0.1225 0.01945 0.00940 0.0113 1.0000 1.0000 2.000 0.1341 0.01961 0.00978 0.0138 1.0000 1.0000 2.500 0.3753 0.02133 0.00842 -0.0169 0.0713 1.0000 2.750 0.3999 0.02184 0.00911 -0.0160 0.0680 1.0000 3.000 0.4247 0.02240 0.00990 -0.0151 0.0670 1.0000 3.250 0.4487 0.02304 0.01076 -0.0140 0.0671 1.0000 3.500 0.4716 0.02375 0.01168 -0.0127 0.0681 1.0000 3.750 0.4938 0.02455 0.01290 -0.0112 0.0691 1.0000 4.000 0.5161 0.02551 0.01401 -0.0096 0.0710 1.0000 4.250 0.5400 0.02675 0.01531 -0.0081 0.0737 1.0000 4.500 0.5672 0.02860 0.01711 -0.0072 0.0766 1.0000 4.750 0.5940 0.02941 0.01822 -0.0057 0.0812 1.0000 5.000 0.6222 0.03114 0.02029 -0.0043 0.0882 1.0000