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 = 0.040 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.3790 0.02496 0.01106 0.0074 1.0000 0.1054 -2.750 -0.3581 0.02390 0.01013 0.0091 1.0000 0.1193 -2.500 -0.3447 0.02241 0.00975 0.0117 1.0000 0.3113 -2.250 -0.3336 0.02128 0.00915 0.0150 1.0000 0.4353 -1.750 -0.0764 0.02176 0.00931 -0.0186 1.0000 1.0000 -1.500 -0.0595 0.02168 0.00903 -0.0169 1.0000 1.0000 -1.250 -0.0431 0.02161 0.00881 -0.0151 1.0000 1.0000 -1.000 -0.0270 0.02157 0.00856 -0.0133 1.0000 1.0000 -0.750 -0.0114 0.02155 0.00843 -0.0114 1.0000 1.0000 -0.500 0.0038 0.02154 0.00836 -0.0095 1.0000 1.0000 -0.250 0.0186 0.02155 0.00834 -0.0074 1.0000 1.0000 0.000 0.0330 0.02158 0.00836 -0.0053 1.0000 1.0000 0.250 0.0470 0.02163 0.00842 -0.0031 1.0000 1.0000 0.500 0.0607 0.02169 0.00855 -0.0009 1.0000 1.0000 0.750 0.0740 0.02177 0.00870 0.0013 1.0000 1.0000 1.000 0.0871 0.02187 0.00888 0.0036 1.0000 1.0000 1.250 0.1000 0.02198 0.00912 0.0059 1.0000 1.0000 1.500 0.1128 0.02212 0.00940 0.0083 1.0000 1.0000 1.750 0.1255 0.02227 0.00973 0.0106 1.0000 1.0000 2.000 0.1381 0.02244 0.01013 0.0129 1.0000 1.0000 2.250 0.1506 0.02264 0.01060 0.0152 1.0000 1.0000 2.500 0.1632 0.02285 0.01120 0.0174 1.0000 1.0000 2.750 0.1758 0.02310 0.01177 0.0196 1.0000 1.0000 3.250 0.4201 0.02499 0.01083 -0.0087 0.1035 1.0000 3.500 0.4411 0.02589 0.01181 -0.0071 0.0964 1.0000 3.750 0.4617 0.02688 0.01294 -0.0055 0.0934 1.0000 4.000 0.4826 0.02782 0.01434 -0.0039 0.0920 1.0000 4.250 0.5045 0.02888 0.01561 -0.0024 0.0917 1.0000 4.500 0.5286 0.03013 0.01704 -0.0010 0.0922 1.0000 4.750 0.5581 0.03150 0.01866 -0.0003 0.0934 1.0000 5.000 0.5898 0.03271 0.02024 0.0005 0.0962 1.0000