XFOIL Version 6.94 Calculated polar for: voyager2 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 -0.4658 0.02457 0.01266 0.0185 1.0000 0.0744 -2.750 -0.4414 0.02387 0.01176 0.0197 1.0000 0.0753 -2.500 -0.4159 0.02325 0.01092 0.0206 1.0000 0.0786 -2.250 -0.3912 0.02259 0.01016 0.0217 1.0000 0.0870 -2.000 -0.3695 0.02174 0.00983 0.0229 1.0000 0.2477 -1.750 -0.3510 0.02116 0.00964 0.0246 1.0000 0.3194 -1.500 -0.3389 0.01995 0.00932 0.0272 1.0000 0.4475 -0.750 -0.0438 0.02170 0.01153 -0.0101 1.0000 1.0000 -0.500 -0.0322 0.02168 0.01144 -0.0077 1.0000 1.0000 -0.250 -0.0206 0.02166 0.01137 -0.0053 1.0000 1.0000 0.000 -0.0092 0.02165 0.01129 -0.0028 1.0000 1.0000 0.250 0.0022 0.02165 0.01127 -0.0003 1.0000 1.0000 0.500 0.0136 0.02165 0.01127 0.0022 1.0000 1.0000 0.750 0.0251 0.02165 0.01132 0.0046 1.0000 1.0000 1.000 0.0368 0.02167 0.01140 0.0071 1.0000 1.0000 1.250 0.0487 0.02169 0.01154 0.0095 1.0000 1.0000 1.750 0.3369 0.02257 0.00905 -0.0309 0.0575 1.0000 2.000 0.3598 0.02288 0.00952 -0.0298 0.0562 1.0000 2.250 0.3834 0.02323 0.01006 -0.0289 0.0562 1.0000 2.500 0.4074 0.02363 0.01068 -0.0280 0.0571 1.0000 2.750 0.4311 0.02407 0.01150 -0.0270 0.0578 1.0000 3.000 0.4539 0.02463 0.01227 -0.0259 0.0597 1.0000 3.250 0.4753 0.02532 0.01313 -0.0244 0.0619 1.0000 3.500 0.4957 0.02621 0.01413 -0.0227 0.0643 1.0000 3.750 0.5168 0.02721 0.01519 -0.0211 0.0666 1.0000 4.000 0.5399 0.02763 0.01579 -0.0194 0.0710 1.0000 4.250 0.5660 0.02875 0.01698 -0.0182 0.0763 1.0000 4.500 0.5978 0.03071 0.01880 -0.0182 0.0814 1.0000 4.750 0.6233 0.03103 0.01956 -0.0163 0.0893 1.0000 5.000 0.6548 0.03297 0.02155 -0.0159 0.0971 1.0000