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 = 0.090 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.4447 0.02242 0.01175 0.0205 1.0000 0.0471 -2.750 -0.4221 0.02195 0.01121 0.0220 1.0000 0.0441 -2.500 -0.4022 0.02071 0.01004 0.0237 1.0000 0.0423 -2.250 -0.3825 0.01990 0.00916 0.0254 1.0000 0.0408 -2.000 -0.3616 0.01931 0.00832 0.0269 1.0000 0.0408 -1.750 -0.3395 0.01878 0.00763 0.0282 1.0000 0.0403 -1.500 -0.3165 0.01834 0.00701 0.0292 1.0000 0.0439 -1.250 -0.2924 0.01774 0.00629 0.0301 1.0000 0.0565 -1.000 -0.2685 0.01674 0.00604 0.0308 1.0000 0.2453 -0.750 -0.2491 0.01587 0.00593 0.0321 1.0000 0.3954 0.000 0.0441 0.01647 0.00780 -0.0070 1.0000 1.0000 0.250 0.0557 0.01653 0.00788 -0.0044 1.0000 1.0000 0.500 0.0669 0.01662 0.00802 -0.0017 1.0000 1.0000 0.750 0.0776 0.01671 0.00817 0.0010 1.0000 1.0000 1.000 0.0880 0.01682 0.00837 0.0038 1.0000 1.0000 1.500 0.2990 0.01803 0.00633 -0.0239 0.0591 1.0000 1.750 0.3232 0.01835 0.00669 -0.0231 0.0505 1.0000 2.000 0.3473 0.01867 0.00716 -0.0222 0.0488 1.0000 2.250 0.3714 0.01906 0.00774 -0.0213 0.0486 1.0000 2.500 0.3952 0.01957 0.00844 -0.0204 0.0489 1.0000 2.750 0.4186 0.02007 0.00912 -0.0193 0.0499 1.0000 3.000 0.4414 0.02055 0.00974 -0.0182 0.0517 1.0000 3.250 0.4638 0.02114 0.01049 -0.0168 0.0542 1.0000 3.500 0.4856 0.02198 0.01140 -0.0153 0.0573 1.0000 3.750 0.5083 0.02306 0.01267 -0.0139 0.0600 1.0000 4.000 0.5344 0.02475 0.01422 -0.0131 0.0635 1.0000 4.250 0.5609 0.02512 0.01491 -0.0114 0.0710 1.0000 4.500 0.5903 0.02683 0.01668 -0.0105 0.0792 1.0000 4.750 0.6205 0.02892 0.01907 -0.0092 0.0922 1.0000 5.000 0.6465 0.02952 0.02032 -0.0068 0.1023 1.0000