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 = 0.022 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1715 0.03142 0.01643 -0.0263 1.0000 1.0000 -2.750 -0.1591 0.03106 0.01567 -0.0246 1.0000 1.0000 -2.500 -0.1459 0.03079 0.01487 -0.0228 1.0000 1.0000 -2.250 -0.1323 0.03058 0.01431 -0.0209 1.0000 1.0000 -2.000 -0.1185 0.03040 0.01383 -0.0190 1.0000 1.0000 -1.750 -0.1047 0.03025 0.01341 -0.0171 1.0000 1.0000 -1.500 -0.0909 0.03013 0.01305 -0.0151 1.0000 1.0000 -1.250 -0.0772 0.03003 0.01276 -0.0131 1.0000 1.0000 -1.000 -0.0635 0.02995 0.01251 -0.0111 1.0000 1.0000 -0.750 -0.0498 0.02989 0.01232 -0.0091 1.0000 1.0000 -0.500 -0.0362 0.02985 0.01218 -0.0070 1.0000 1.0000 -0.250 -0.0227 0.02982 0.01208 -0.0049 1.0000 1.0000 0.000 -0.0091 0.02980 0.01197 -0.0029 1.0000 1.0000 0.250 0.0044 0.02980 0.01195 -0.0008 1.0000 1.0000 0.500 0.0180 0.02982 0.01198 0.0013 1.0000 1.0000 0.750 0.0315 0.02985 0.01206 0.0034 1.0000 1.0000 1.000 0.0451 0.02990 0.01218 0.0055 1.0000 1.0000 1.250 0.0588 0.02997 0.01236 0.0075 1.0000 1.0000 1.500 0.0725 0.03005 0.01260 0.0095 1.0000 1.0000 1.750 0.0863 0.03016 0.01298 0.0115 1.0000 1.0000 2.000 0.1001 0.03030 0.01332 0.0135 1.0000 1.0000 2.250 0.1139 0.03046 0.01374 0.0155 1.0000 1.0000 2.500 0.1276 0.03066 0.01424 0.0174 1.0000 1.0000 2.750 0.1411 0.03091 0.01481 0.0193 1.0000 1.0000 3.000 0.1541 0.03121 0.01548 0.0212 1.0000 1.0000 3.250 0.1662 0.03160 0.01626 0.0231 1.0000 1.0000 3.500 0.1765 0.03214 0.01721 0.0249 1.0000 1.0000 3.750 0.3637 0.03085 0.01822 0.0004 0.8040 1.0000 4.000 0.3774 0.03033 0.01801 0.0052 0.6608 1.0000 4.500 0.4769 0.03398 0.01793 0.0031 0.1619 1.0000 4.750 0.4929 0.03557 0.01929 0.0052 0.1475 1.0000 5.000 0.5103 0.03717 0.02080 0.0072 0.1394 1.0000