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 = 0.100 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -2.000 -0.3649 0.01888 0.00830 0.0276 1.0000 0.0379 -1.750 -0.3432 0.01836 0.00763 0.0289 1.0000 0.0369 -1.500 -0.3209 0.01792 0.00702 0.0301 1.0000 0.0368 -1.250 -0.2984 0.01750 0.00644 0.0312 1.0000 0.0418 -1.000 -0.2753 0.01703 0.00589 0.0322 1.0000 0.0560 -0.750 -0.2532 0.01594 0.00578 0.0331 1.0000 0.2895 0.000 0.0456 0.01596 0.00775 -0.0072 1.0000 1.0000 0.250 0.0569 0.01603 0.00784 -0.0046 1.0000 1.0000 0.500 0.0676 0.01612 0.00798 -0.0018 1.0000 1.0000 0.750 0.0780 0.01622 0.00814 0.0010 1.0000 1.0000 1.000 0.0880 0.01633 0.00835 0.0038 1.0000 1.0000 1.250 0.2778 0.01724 0.00592 -0.0251 0.0819 1.0000 1.500 0.3019 0.01764 0.00626 -0.0243 0.0476 1.0000 1.750 0.3260 0.01787 0.00663 -0.0235 0.0457 1.0000 2.000 0.3501 0.01817 0.00709 -0.0227 0.0453 1.0000 2.250 0.3741 0.01855 0.00765 -0.0218 0.0457 1.0000 2.500 0.3979 0.01903 0.00832 -0.0208 0.0467 1.0000 2.750 0.4208 0.01966 0.00911 -0.0198 0.0479 1.0000 3.000 0.4425 0.02048 0.01004 -0.0185 0.0493 1.0000 3.250 0.4654 0.02086 0.01050 -0.0172 0.0519 1.0000 3.500 0.4881 0.02158 0.01130 -0.0158 0.0558 1.0000 3.750 0.5121 0.02268 0.01257 -0.0146 0.0595 1.0000 4.000 0.5390 0.02411 0.01392 -0.0137 0.0642 1.0000 4.250 0.5672 0.02500 0.01510 -0.0123 0.0742 1.0000 4.500 0.5963 0.02592 0.01632 -0.0107 0.0861 1.0000 4.750 0.6243 0.02846 0.01903 -0.0098 0.0917 1.0000 5.000 0.6499 0.02918 0.02032 -0.0075 0.0998 1.0000