XFOIL Version 6.94 Calculated polar for: bbl1 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.045 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 0.3828 0.04809 0.03634 -0.0682 0.8521 0.2269 -2.750 0.4400 0.04724 0.03590 -0.0734 0.8408 0.4192 -2.500 0.4628 0.04819 0.03702 -0.0732 0.8238 0.4584 -2.250 0.5075 0.04813 0.03722 -0.0759 0.8117 0.5211 -2.000 0.5374 0.04850 0.03792 -0.0762 0.7971 0.5750 -1.750 0.5574 0.04938 0.03907 -0.0750 0.7823 0.6203 -1.500 0.6010 0.04852 0.03857 -0.0760 0.7717 0.6793 -1.250 0.6096 0.04994 0.04018 -0.0732 0.7563 0.7069 -1.000 0.6608 0.04851 0.03895 -0.0752 0.7468 0.7519 -0.750 0.6632 0.05030 0.04091 -0.0717 0.7308 0.7753 -0.500 0.7079 0.04846 0.03938 -0.0716 0.7223 0.8298 -0.250 0.6978 0.05086 0.04203 -0.0665 0.7060 0.8677 0.000 0.7908 0.04722 0.03829 -0.0757 0.6972 1.0016 0.250 0.7938 0.05022 0.04116 -0.0748 0.6787 1.0016 0.500 0.8080 0.05241 0.04320 -0.0745 0.6630 1.0016 0.750 0.9031 0.04838 0.03884 -0.0819 0.6526 1.0016 1.000 0.8995 0.05124 0.04171 -0.0784 0.6343 1.0016 1.250 0.9240 0.05187 0.04227 -0.0776 0.6171 1.0016 1.500 0.9893 0.04895 0.03916 -0.0801 0.5993 1.0016 1.750 1.0528 0.04614 0.03611 -0.0824 0.5807 1.0016 2.000 1.0794 0.04628 0.03620 -0.0813 0.5611 1.0016 2.250 1.0964 0.04726 0.03720 -0.0793 0.5414 1.0016 2.500 1.1205 0.04773 0.03765 -0.0779 0.5218 1.0016 2.750 1.1506 0.04771 0.03758 -0.0772 0.5016 1.0016 3.000 1.1815 0.04757 0.03734 -0.0764 0.4800 1.0016 3.250 1.2092 0.04771 0.03737 -0.0753 0.4578 1.0016 3.500 1.2334 0.04817 0.03775 -0.0739 0.4353 1.0016 3.750 1.2552 0.04879 0.03834 -0.0722 0.4128 1.0016 4.000 1.2764 0.04948 0.03894 -0.0706 0.3914 1.0016 4.250 1.2987 0.05023 0.03958 -0.0692 0.3719 1.0016 4.500 1.3282 0.05087 0.04000 -0.0690 0.3549 1.0016 4.750 1.3249 0.05371 0.04320 -0.0649 0.3415 1.0016 5.000 1.3386 0.05569 0.04522 -0.0631 0.3298 1.0016