XFOIL Version 6.94 Calculated polar for: GOE 795 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.300 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.3781 0.00707 0.00255 -0.0845 0.8812 1.0000 0.500 0.4308 0.00706 0.00240 -0.0835 0.8505 1.0000 1.000 0.4801 0.00713 0.00228 -0.0816 0.8154 1.0000 1.500 0.5277 0.00727 0.00226 -0.0794 0.7796 1.0000 2.000 0.5744 0.00747 0.00233 -0.0770 0.7443 1.0000 2.500 0.6206 0.00771 0.00246 -0.0746 0.7081 1.0000 3.000 0.6655 0.00798 0.00263 -0.0719 0.6650 1.0000 3.500 0.7086 0.00830 0.00283 -0.0688 0.6090 1.0000 4.000 0.7480 0.00884 0.00308 -0.0651 0.5219 1.0000 4.500 0.7840 0.00972 0.00355 -0.0610 0.4159 1.0000 5.000 0.8214 0.01071 0.00419 -0.0574 0.3300 1.0000 5.500 0.8563 0.01199 0.00496 -0.0536 0.2058 1.0000 6.000 0.8863 0.01386 0.00620 -0.0491 0.0936 1.0000 6.500 0.9250 0.01501 0.00728 -0.0460 0.0706 1.0000 7.000 0.9643 0.01607 0.00836 -0.0431 0.0558 1.0000 7.500 0.9994 0.01749 0.00986 -0.0396 0.0373 1.0000 8.000 1.0379 0.01870 0.01109 -0.0364 0.0197 1.0000 8.500 1.0647 0.02087 0.01342 -0.0315 0.0140 1.0000 9.000 1.0914 0.02313 0.01595 -0.0266 0.0114 1.0000 9.500 1.1051 0.02714 0.02027 -0.0203 0.0091 1.0000 10.000 1.1204 0.03164 0.02535 -0.0144 0.0081 1.0000 10.500 1.1266 0.03639 0.03064 -0.0081 0.0082 1.0000 11.000 1.1159 0.04207 0.03692 -0.0013 0.0083 1.0000 11.500 1.0918 0.04876 0.04416 0.0038 0.0083 1.0000 12.000 1.0585 0.05721 0.05309 0.0054 0.0084 1.0000 12.500 1.0224 0.06769 0.06396 0.0026 0.0086 1.0000 13.000 0.9841 0.08115 0.07775 -0.0048 0.0087 1.0000 13.500 0.9393 0.09973 0.09660 -0.0167 0.0089 1.0000