XFOIL Version 6.94 Calculated polar for: pp13 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 -0.0220 0.06674 0.04500 -0.0238 0.9998 1.0002 -2.750 -0.0224 0.06558 0.04411 -0.0223 0.9998 1.0002 -2.500 -0.0234 0.06444 0.04322 -0.0207 0.9998 1.0002 -2.250 -0.0252 0.06330 0.04235 -0.0190 0.9998 1.0002 -2.000 -0.0277 0.06215 0.04149 -0.0172 0.9998 1.0002 -1.750 -0.0306 0.06101 0.04062 -0.0153 0.9998 1.0002 -1.500 -0.0334 0.05988 0.03973 -0.0135 0.9998 1.0002 -1.250 -0.0348 0.05881 0.03887 -0.0120 0.9998 1.0002 -1.000 -0.0336 0.05784 0.03805 -0.0110 0.9998 1.0002 -0.750 -0.0263 0.05705 0.03732 -0.0112 0.9998 1.0002 -0.500 -0.0100 0.05652 0.03675 -0.0133 0.9998 1.0002 -0.250 0.0165 0.05634 0.03641 -0.0173 0.9998 1.0002 0.000 0.0523 0.05654 0.03639 -0.0231 0.9998 1.0002 0.250 0.0967 0.05712 0.03666 -0.0306 0.9998 1.0002 0.500 0.1469 0.05805 0.03724 -0.0390 0.9998 1.0002 0.750 0.2015 0.05931 0.03811 -0.0481 0.9998 1.0002 1.000 0.2585 0.06083 0.03920 -0.0574 0.9998 1.0002 1.250 0.3158 0.06256 0.04048 -0.0665 0.9998 1.0002 1.500 0.3715 0.06441 0.04184 -0.0749 0.9998 1.0002 1.750 0.4234 0.06630 0.04324 -0.0821 0.9998 1.0002 2.000 0.4708 0.06817 0.04456 -0.0879 0.9998 1.0002 2.250 0.5132 0.06997 0.04578 -0.0923 0.9998 1.0002 2.500 0.5495 0.07167 0.04691 -0.0951 0.9998 1.0002 2.750 0.5807 0.07327 0.04796 -0.0966 0.9998 1.0002 3.000 0.6077 0.07480 0.04904 -0.0973 0.9998 1.0002 3.250 0.6317 0.07631 0.05029 -0.0976 0.9998 1.0002 3.500 0.6542 0.07786 0.05173 -0.0977 0.9998 1.0002 3.750 0.6756 0.07947 0.05338 -0.0977 0.9998 1.0002 4.000 0.6961 0.08117 0.05529 -0.0977 0.9998 1.0002 4.250 0.7159 0.08296 0.05736 -0.0978 0.9998 1.0002 4.500 0.7350 0.08490 0.05970 -0.0979 0.9998 1.0002 4.750 0.7531 0.08702 0.06234 -0.0982 0.9998 1.0002 5.000 0.7697 0.08946 0.06546 -0.0987 0.9998 1.0002