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 = 0.035 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.3709 0.02506 0.01111 0.0055 1.0000 0.1403 -2.750 -0.3596 0.02346 0.01070 0.0086 1.0000 0.3443 -2.250 -0.1144 0.02310 0.01031 -0.0212 1.0000 1.0000 -2.000 -0.0965 0.02297 0.00987 -0.0197 1.0000 1.0000 -1.750 -0.0791 0.02286 0.00950 -0.0181 1.0000 1.0000 -1.500 -0.0622 0.02278 0.00920 -0.0164 1.0000 1.0000 -1.250 -0.0457 0.02271 0.00898 -0.0147 1.0000 1.0000 -1.000 -0.0296 0.02266 0.00871 -0.0129 1.0000 1.0000 -0.750 -0.0138 0.02264 0.00858 -0.0110 1.0000 1.0000 -0.500 0.0016 0.02263 0.00850 -0.0091 1.0000 1.0000 -0.250 0.0167 0.02264 0.00847 -0.0071 1.0000 1.0000 0.000 0.0313 0.02266 0.00849 -0.0051 1.0000 1.0000 0.250 0.0456 0.02271 0.00855 -0.0030 1.0000 1.0000 0.500 0.0596 0.02277 0.00868 -0.0008 1.0000 1.0000 0.750 0.0734 0.02285 0.00882 0.0014 1.0000 1.0000 1.000 0.0868 0.02295 0.00901 0.0036 1.0000 1.0000 1.250 0.1001 0.02306 0.00925 0.0059 1.0000 1.0000 1.500 0.1133 0.02319 0.00954 0.0081 1.0000 1.0000 1.750 0.1263 0.02335 0.00988 0.0103 1.0000 1.0000 2.000 0.1392 0.02352 0.01028 0.0126 1.0000 1.0000 2.250 0.1520 0.02372 0.01075 0.0148 1.0000 1.0000 2.500 0.1648 0.02395 0.01138 0.0170 1.0000 1.0000 2.750 0.1775 0.02421 0.01197 0.0192 1.0000 1.0000 3.000 0.1900 0.02451 0.01263 0.0213 1.0000 1.0000 3.250 0.4206 0.02461 0.01118 -0.0080 0.2815 1.0000 3.500 0.4349 0.02628 0.01181 -0.0058 0.1187 1.0000 3.750 0.4545 0.02742 0.01291 -0.0040 0.1090 1.0000 4.000 0.4745 0.02837 0.01430 -0.0022 0.1051 1.0000 4.250 0.4944 0.02938 0.01552 -0.0003 0.1031 1.0000 4.500 0.5149 0.03051 0.01686 0.0016 0.1024 1.0000 4.750 0.5395 0.03183 0.01840 0.0029 0.1027 1.0000 5.000 0.5700 0.03338 0.02024 0.0036 0.1040 1.0000