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 = 0.080 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -2.750 -0.4185 0.02205 0.01100 0.0212 1.0000 0.0459 -2.500 -0.3978 0.02138 0.01025 0.0229 1.0000 0.0443 -2.250 -0.3780 0.02051 0.00928 0.0245 1.0000 0.0436 -2.000 -0.3564 0.01991 0.00843 0.0259 1.0000 0.0441 -1.750 -0.3321 0.01911 0.00744 0.0267 1.0000 0.0499 -1.500 -0.3071 0.01859 0.00679 0.0275 1.0000 0.0608 -1.250 -0.2825 0.01747 0.00640 0.0280 1.0000 0.2357 -1.000 -0.2634 0.01666 0.00625 0.0295 1.0000 0.3734 -0.500 0.0165 0.01700 0.00780 -0.0115 1.0000 1.0000 -0.250 0.0298 0.01703 0.00781 -0.0092 1.0000 1.0000 0.000 0.0424 0.01707 0.00786 -0.0068 1.0000 1.0000 0.250 0.0544 0.01714 0.00794 -0.0042 1.0000 1.0000 0.500 0.0660 0.01722 0.00807 -0.0016 1.0000 1.0000 0.750 0.0771 0.01731 0.00822 0.0010 1.0000 1.0000 1.000 0.0880 0.01742 0.00842 0.0037 1.0000 1.0000 1.250 0.0987 0.01754 0.00865 0.0064 1.0000 1.0000 1.500 0.1093 0.01767 0.00893 0.0092 1.0000 1.0000 1.750 0.3188 0.01882 0.00675 -0.0224 0.0614 1.0000 2.000 0.3431 0.01914 0.00718 -0.0216 0.0556 1.0000 2.250 0.3674 0.01949 0.00769 -0.0207 0.0538 1.0000 2.500 0.3916 0.01991 0.00829 -0.0197 0.0534 1.0000 2.750 0.4155 0.02038 0.00896 -0.0187 0.0541 1.0000 3.000 0.4388 0.02095 0.00972 -0.0175 0.0554 1.0000 3.250 0.4611 0.02164 0.01058 -0.0162 0.0573 1.0000 3.500 0.4826 0.02252 0.01155 -0.0147 0.0595 1.0000 3.750 0.5043 0.02361 0.01286 -0.0132 0.0613 1.0000 4.000 0.5284 0.02489 0.01409 -0.0120 0.0639 1.0000 4.250 0.5546 0.02550 0.01494 -0.0105 0.0696 1.0000 4.500 0.5839 0.02765 0.01709 -0.0098 0.0763 1.0000 4.750 0.6115 0.02826 0.01817 -0.0079 0.0863 1.0000 5.000 0.6406 0.02978 0.02010 -0.0062 0.0989 1.0000