XFOIL Version 6.94 Calculated polar for: tilt 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.005 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0678 0.05852 0.03973 -0.0277 0.9998 1.0002 -2.750 -0.0303 0.05858 0.03798 -0.0318 0.9998 1.0002 -2.500 0.0002 0.05889 0.03684 -0.0329 0.9998 1.0002 -2.250 0.0267 0.05925 0.03604 -0.0329 0.9998 1.0002 -2.000 0.0510 0.05966 0.03548 -0.0324 0.9998 1.0002 -1.750 0.0739 0.06010 0.03509 -0.0318 0.9998 1.0002 -1.500 0.0960 0.06057 0.03485 -0.0310 0.9998 1.0002 -1.250 0.1175 0.06109 0.03475 -0.0303 0.9998 1.0002 -1.000 0.1385 0.06166 0.03477 -0.0295 0.9998 1.0002 -0.750 0.1590 0.06227 0.03491 -0.0289 0.9998 1.0002 -0.500 0.1792 0.06294 0.03505 -0.0282 0.9998 1.0002 -0.250 0.1989 0.06367 0.03542 -0.0276 0.9998 1.0002 0.000 0.2182 0.06446 0.03592 -0.0271 0.9998 1.0002 0.250 0.2371 0.06533 0.03656 -0.0267 0.9998 1.0002 0.500 0.2553 0.06628 0.03733 -0.0263 0.9998 1.0002 0.750 0.2729 0.06735 0.03828 -0.0261 0.9998 1.0002 1.000 0.2896 0.06854 0.03941 -0.0259 0.9998 1.0002 1.250 0.3051 0.06991 0.04072 -0.0259 0.9998 1.0002 1.500 0.3189 0.07152 0.04237 -0.0261 0.9998 1.0002 1.750 0.3303 0.07345 0.04440 -0.0264 0.9998 1.0002 2.000 0.3383 0.07584 0.04691 -0.0271 0.9998 1.0002 2.250 0.3423 0.07878 0.04993 -0.0280 0.9998 1.0002 2.500 0.3434 0.08216 0.05332 -0.0293 0.9998 1.0002 2.750 0.3442 0.08566 0.05676 -0.0307 0.9998 1.0002 3.000 0.3466 0.08908 0.06007 -0.0321 0.9998 1.0002 3.250 0.3506 0.09237 0.06322 -0.0335 0.9998 1.0002 3.500 0.3559 0.09553 0.06618 -0.0349 0.9998 1.0002 3.750 0.3621 0.09860 0.06909 -0.0362 0.9998 1.0002 4.000 0.3692 0.10159 0.07193 -0.0375 0.9998 1.0002 4.250 0.3769 0.10453 0.07471 -0.0387 0.9998 1.0002 4.500 0.3851 0.10743 0.07745 -0.0400 0.9998 1.0002 4.750 0.3937 0.11029 0.08016 -0.0412 0.9998 1.0002 5.000 0.4026 0.11313 0.08285 -0.0425 0.9998 1.0002