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 = 0.026 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1749 0.02648 0.01273 -0.0243 1.0000 1.0000 -2.750 -0.1562 0.02624 0.01202 -0.0229 1.0000 1.0000 -2.500 -0.1378 0.02605 0.01139 -0.0215 1.0000 1.0000 -2.250 -0.1198 0.02588 0.01084 -0.0200 1.0000 1.0000 -2.000 -0.1021 0.02574 0.01037 -0.0186 1.0000 1.0000 -1.750 -0.0848 0.02562 0.00997 -0.0170 1.0000 1.0000 -1.500 -0.0678 0.02552 0.00965 -0.0154 1.0000 1.0000 -1.250 -0.0512 0.02545 0.00940 -0.0137 1.0000 1.0000 -1.000 -0.0349 0.02540 0.00910 -0.0120 1.0000 1.0000 -0.750 -0.0188 0.02536 0.00895 -0.0102 1.0000 1.0000 -0.500 -0.0030 0.02535 0.00885 -0.0083 1.0000 1.0000 -0.250 0.0126 0.02535 0.00881 -0.0065 1.0000 1.0000 0.000 0.0278 0.02538 0.00882 -0.0045 1.0000 1.0000 0.250 0.0427 0.02542 0.00887 -0.0026 1.0000 1.0000 0.500 0.0574 0.02547 0.00898 -0.0005 1.0000 1.0000 0.750 0.0719 0.02555 0.00916 0.0015 1.0000 1.0000 1.000 0.0861 0.02565 0.00935 0.0036 1.0000 1.0000 1.250 0.1002 0.02576 0.00960 0.0057 1.0000 1.0000 1.500 0.1140 0.02590 0.00990 0.0077 1.0000 1.0000 1.750 0.1277 0.02606 0.01026 0.0098 1.0000 1.0000 2.000 0.1413 0.02624 0.01069 0.0119 1.0000 1.0000 2.250 0.1547 0.02645 0.01117 0.0140 1.0000 1.0000 2.500 0.1679 0.02669 0.01184 0.0161 1.0000 1.0000 2.750 0.1809 0.02697 0.01247 0.0181 1.0000 1.0000 3.000 0.1935 0.02730 0.01318 0.0201 1.0000 1.0000 3.250 0.2056 0.02769 0.01399 0.0222 1.0000 1.0000 3.500 0.2165 0.02818 0.01491 0.0242 1.0000 1.0000 3.750 0.2247 0.02886 0.01601 0.0262 1.0000 1.0000 4.000 0.4696 0.02857 0.01504 0.0002 0.2480 1.0000 4.250 0.4822 0.03036 0.01584 0.0027 0.1593 1.0000 4.500 0.4990 0.03195 0.01716 0.0049 0.1422 1.0000 4.750 0.5181 0.03344 0.01867 0.0069 0.1349 1.0000 5.000 0.5404 0.03495 0.02038 0.0086 0.1320 1.0000