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 = 0.035 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1620 0.02695 0.01567 -0.0279 1.0000 1.0000 -2.750 -0.1520 0.02650 0.01480 -0.0261 1.0000 1.0000 -2.500 -0.1393 0.02624 0.01405 -0.0242 1.0000 1.0000 -2.250 -0.1260 0.02606 0.01354 -0.0223 1.0000 1.0000 -2.000 -0.1125 0.02591 0.01311 -0.0203 1.0000 1.0000 -1.750 -0.0992 0.02579 0.01275 -0.0182 1.0000 1.0000 -1.500 -0.0859 0.02570 0.01245 -0.0161 1.0000 1.0000 -1.250 -0.0729 0.02562 0.01220 -0.0140 1.0000 1.0000 -1.000 -0.0599 0.02556 0.01200 -0.0118 1.0000 1.0000 -0.750 -0.0471 0.02551 0.01184 -0.0096 1.0000 1.0000 -0.500 -0.0344 0.02548 0.01172 -0.0073 1.0000 1.0000 -0.250 -0.0218 0.02545 0.01164 -0.0051 1.0000 1.0000 0.000 -0.0092 0.02544 0.01154 -0.0028 1.0000 1.0000 0.250 0.0034 0.02544 0.01152 -0.0005 1.0000 1.0000 0.500 0.0160 0.02545 0.01154 0.0017 1.0000 1.0000 0.750 0.0286 0.02547 0.01160 0.0040 1.0000 1.0000 1.000 0.0414 0.02550 0.01171 0.0062 1.0000 1.0000 1.250 0.0543 0.02555 0.01186 0.0085 1.0000 1.0000 1.500 0.0673 0.02561 0.01207 0.0107 1.0000 1.0000 1.750 0.0805 0.02569 0.01241 0.0129 1.0000 1.0000 2.000 0.0939 0.02580 0.01273 0.0150 1.0000 1.0000 2.250 0.1074 0.02593 0.01311 0.0171 1.0000 1.0000 2.500 0.1209 0.02610 0.01355 0.0192 1.0000 1.0000 2.750 0.3768 0.02701 0.01150 -0.0172 0.1042 1.0000 3.000 0.3955 0.02780 0.01244 -0.0155 0.0928 1.0000 3.250 0.4155 0.02847 0.01326 -0.0137 0.0891 1.0000 3.500 0.4355 0.02917 0.01416 -0.0120 0.0874 1.0000 3.750 0.4549 0.02993 0.01513 -0.0102 0.0870 1.0000 4.000 0.4735 0.03074 0.01615 -0.0082 0.0875 1.0000 4.250 0.4927 0.03165 0.01725 -0.0063 0.0886 1.0000 4.500 0.5131 0.03266 0.01845 -0.0045 0.0903 1.0000 4.750 0.5373 0.03386 0.01982 -0.0032 0.0926 1.0000 5.000 0.5724 0.03553 0.02163 -0.0036 0.0955 1.0000