XFOIL Version 6.94 Calculated polar for: ARA-D BL 13% 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.035 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.2011 0.04481 0.03743 -0.0162 1.0000 0.2690 -2.750 -0.1777 0.04308 0.03554 -0.0168 1.0000 0.2747 -2.500 -0.1514 0.04159 0.03376 -0.0179 1.0000 0.2820 -2.250 -0.1266 0.04015 0.03230 -0.0184 1.0000 0.2910 -2.000 -0.0413 0.03759 0.02968 -0.0283 0.9726 0.3170 -1.750 0.1436 0.02997 0.02104 -0.0353 0.5320 0.3991 -1.500 0.1655 0.03043 0.02035 -0.0323 0.4314 0.4303 -1.250 0.1921 0.03038 0.02004 -0.0308 0.3951 0.4753 -1.000 0.2175 0.02991 0.01990 -0.0289 0.3730 0.5522 -0.750 0.2713 0.02877 0.01943 -0.0298 0.3507 1.0000 -0.500 0.3033 0.02984 0.01984 -0.0299 0.3373 1.0000 -0.250 0.3340 0.03106 0.02050 -0.0296 0.3282 1.0000 0.000 0.3646 0.03225 0.02154 -0.0296 0.3219 1.0000 0.250 0.3943 0.03354 0.02269 -0.0295 0.3164 1.0000 0.500 0.4233 0.03497 0.02395 -0.0293 0.3119 1.0000 0.750 0.4513 0.03665 0.02537 -0.0291 0.3073 1.0000 1.000 0.4785 0.03834 0.02713 -0.0292 0.3034 1.0000 1.250 0.5047 0.04014 0.02905 -0.0292 0.2993 1.0000 1.500 0.5302 0.04217 0.03116 -0.0294 0.2965 1.0000 1.750 0.5549 0.04452 0.03363 -0.0296 0.2960 1.0000 2.000 0.5779 0.04719 0.03646 -0.0300 0.2967 1.0000 2.250 0.5985 0.05026 0.03972 -0.0306 0.2985 1.0000 2.500 0.6166 0.05369 0.04331 -0.0312 0.3009 1.0000 2.750 0.6325 0.05733 0.04707 -0.0319 0.3029 1.0000 3.000 0.6470 0.06115 0.05097 -0.0325 0.3047 1.0000 3.250 0.6592 0.06524 0.05517 -0.0333 0.3074 1.0000 3.500 0.6240 0.07555 0.06612 -0.0398 0.3299 1.0000 4.250 0.3966 0.11271 0.10414 -0.0770 0.6488 1.0000 4.750 0.4048 0.11676 0.10795 -0.0748 0.6187 1.0000 5.000 0.4319 0.12107 0.11212 -0.0758 0.6044 1.0000