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 = 0.022 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1770 0.02828 0.01315 -0.0236 1.0000 1.0000 -2.750 -0.1587 0.02803 0.01241 -0.0222 1.0000 1.0000 -2.500 -0.1405 0.02783 0.01175 -0.0208 1.0000 1.0000 -2.250 -0.1226 0.02765 0.01118 -0.0194 1.0000 1.0000 -2.000 -0.1050 0.02750 0.01069 -0.0179 1.0000 1.0000 -1.750 -0.0877 0.02737 0.01028 -0.0164 1.0000 1.0000 -1.500 -0.0707 0.02727 0.00994 -0.0148 1.0000 1.0000 -1.250 -0.0540 0.02719 0.00967 -0.0132 1.0000 1.0000 -1.000 -0.0376 0.02713 0.00936 -0.0115 1.0000 1.0000 -0.750 -0.0214 0.02710 0.00919 -0.0097 1.0000 1.0000 -0.500 -0.0054 0.02708 0.00908 -0.0079 1.0000 1.0000 -0.250 0.0104 0.02708 0.00903 -0.0061 1.0000 1.0000 0.000 0.0259 0.02710 0.00903 -0.0043 1.0000 1.0000 0.250 0.0412 0.02714 0.00909 -0.0024 1.0000 1.0000 0.500 0.0562 0.02720 0.00919 -0.0004 1.0000 1.0000 0.750 0.0711 0.02728 0.00938 0.0015 1.0000 1.0000 1.000 0.0857 0.02737 0.00958 0.0035 1.0000 1.0000 1.250 0.1001 0.02749 0.00984 0.0055 1.0000 1.0000 1.500 0.1144 0.02763 0.01015 0.0075 1.0000 1.0000 1.750 0.1284 0.02779 0.01052 0.0096 1.0000 1.0000 2.000 0.1423 0.02798 0.01096 0.0116 1.0000 1.0000 2.250 0.1560 0.02820 0.01146 0.0136 1.0000 1.0000 2.500 0.1695 0.02845 0.01215 0.0156 1.0000 1.0000 2.750 0.1827 0.02874 0.01279 0.0175 1.0000 1.0000 3.000 0.1955 0.02908 0.01353 0.0195 1.0000 1.0000 3.250 0.2076 0.02948 0.01437 0.0215 1.0000 1.0000 3.500 0.2187 0.02997 0.01531 0.0234 1.0000 1.0000 3.750 0.2279 0.03062 0.01640 0.0253 1.0000 1.0000 4.000 0.2328 0.03157 0.01777 0.0271 1.0000 1.0000 4.500 0.4995 0.03275 0.01862 0.0062 0.2204 1.0000 4.750 0.5165 0.03438 0.01976 0.0084 0.1884 1.0000 5.000 0.5361 0.03615 0.02126 0.0103 0.1706 1.0000