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 = 0.030 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1660 0.02826 0.01584 -0.0274 1.0000 1.0000 -2.750 -0.1545 0.02788 0.01505 -0.0256 1.0000 1.0000 -2.500 -0.1415 0.02763 0.01429 -0.0237 1.0000 1.0000 -2.250 -0.1281 0.02743 0.01377 -0.0218 1.0000 1.0000 -2.000 -0.1145 0.02727 0.01332 -0.0199 1.0000 1.0000 -1.750 -0.1010 0.02715 0.01294 -0.0179 1.0000 1.0000 -1.500 -0.0876 0.02704 0.01263 -0.0158 1.0000 1.0000 -1.250 -0.0743 0.02696 0.01236 -0.0137 1.0000 1.0000 -1.000 -0.0611 0.02689 0.01215 -0.0116 1.0000 1.0000 -0.750 -0.0481 0.02684 0.01198 -0.0094 1.0000 1.0000 -0.500 -0.0350 0.02680 0.01185 -0.0072 1.0000 1.0000 -0.250 -0.0221 0.02677 0.01176 -0.0050 1.0000 1.0000 0.000 -0.0092 0.02676 0.01166 -0.0028 1.0000 1.0000 0.250 0.0037 0.02676 0.01164 -0.0006 1.0000 1.0000 0.500 0.0167 0.02677 0.01167 0.0016 1.0000 1.0000 0.750 0.0296 0.02680 0.01173 0.0038 1.0000 1.0000 1.000 0.0427 0.02683 0.01184 0.0060 1.0000 1.0000 1.250 0.0558 0.02689 0.01200 0.0081 1.0000 1.0000 1.500 0.0691 0.02696 0.01222 0.0103 1.0000 1.0000 1.750 0.0825 0.02705 0.01257 0.0124 1.0000 1.0000 2.000 0.0960 0.02716 0.01290 0.0145 1.0000 1.0000 2.250 0.1096 0.02731 0.01330 0.0166 1.0000 1.0000 2.500 0.1232 0.02749 0.01376 0.0186 1.0000 1.0000 2.750 0.1367 0.02771 0.01430 0.0206 1.0000 1.0000 3.250 0.4037 0.02890 0.01320 -0.0115 0.1175 1.0000 3.500 0.4221 0.02992 0.01413 -0.0096 0.1057 1.0000 3.750 0.4413 0.03075 0.01511 -0.0077 0.1020 1.0000 4.000 0.4600 0.03163 0.01615 -0.0058 0.1003 1.0000 4.250 0.4783 0.03255 0.01725 -0.0037 0.0998 1.0000 4.500 0.4965 0.03351 0.01841 -0.0016 0.1003 1.0000 4.750 0.5157 0.03453 0.01964 0.0005 0.1015 1.0000 5.000 0.5387 0.03571 0.02102 0.0020 0.1035 1.0000