XFOIL Version 6.94 Calculated polar for: ARA-D BL 12% 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.022 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0587 0.04246 0.03275 -0.0262 0.9997 0.4170 -2.750 -0.0332 0.04124 0.03146 -0.0264 0.9997 0.4377 -2.500 -0.0089 0.04014 0.03044 -0.0262 0.9997 0.4629 -2.250 0.0154 0.03917 0.02968 -0.0258 0.9997 0.4941 -2.000 0.0387 0.03829 0.02918 -0.0250 0.9997 0.5354 -1.750 0.0618 0.03741 0.02892 -0.0241 0.9997 0.5967 -1.500 0.0770 0.03611 0.02896 -0.0205 0.9997 0.7104 -1.250 0.1063 0.03620 0.02907 -0.0247 0.9997 1.0003 -1.000 0.3400 0.03495 0.02411 -0.0425 0.5341 1.0003 -0.750 0.3685 0.03647 0.02495 -0.0415 0.4988 1.0003 -0.500 0.3988 0.03810 0.02614 -0.0414 0.4754 1.0003 -0.250 0.4286 0.03974 0.02734 -0.0412 0.4574 1.0003 0.000 0.4590 0.04153 0.02898 -0.0417 0.4410 1.0003 0.250 0.4887 0.04349 0.03080 -0.0423 0.4277 1.0003 0.500 0.5179 0.04545 0.03259 -0.0427 0.4171 1.0003 0.750 0.5473 0.04788 0.03498 -0.0438 0.4113 1.0003 1.000 0.5754 0.05094 0.03823 -0.0460 0.4076 1.0003 1.250 0.6021 0.05436 0.04181 -0.0485 0.4051 1.0003 1.500 0.6253 0.05809 0.04567 -0.0508 0.4029 1.0003 1.750 0.6445 0.06217 0.04986 -0.0532 0.4012 1.0003 2.000 0.6585 0.06680 0.05461 -0.0557 0.4011 1.0003 2.250 0.6659 0.07217 0.06009 -0.0587 0.4040 1.0003 2.500 0.6722 0.07756 0.06552 -0.0612 0.4091 1.0003 3.000 0.6565 0.09122 0.07931 -0.0690 0.4346 1.0003 3.250 0.5924 0.10047 0.08870 -0.0747 0.4696 1.0003 3.500 0.5614 0.10761 0.09586 -0.0793 0.5145 1.0003 4.250 0.3822 0.11919 0.10778 -0.0846 0.9027 1.0003 4.500 0.4072 0.12461 0.11304 -0.0880 0.8978 1.0003 5.000 0.4327 0.13046 0.11864 -0.0894 0.8693 1.0003