XFOIL Version 6.94 Calculated polar for: ARA-D 13% AIRFOIL 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 -0.2574 0.04938 0.03940 -0.0175 1.0000 0.3834 -2.750 -0.2370 0.04763 0.03742 -0.0189 1.0000 0.4058 -2.500 -0.2170 0.04587 0.03555 -0.0191 1.0000 0.4187 -2.250 -0.1910 0.04441 0.03377 -0.0210 1.0000 0.4289 -2.000 -0.1622 0.04330 0.03223 -0.0233 1.0000 0.4431 -1.750 -0.1360 0.04207 0.03061 -0.0246 1.0000 0.4618 -1.500 -0.1115 0.04136 0.02978 -0.0255 1.0000 0.4890 -1.250 -0.0904 0.04065 0.02924 -0.0253 1.0000 0.5300 -1.000 -0.0720 0.04006 0.02900 -0.0243 1.0000 0.5962 -0.750 -0.0566 0.03912 0.02884 -0.0219 1.0000 0.6796 -0.500 -0.0370 0.03779 0.02812 -0.0230 1.0000 1.0000 -0.250 -0.0015 0.03909 0.02792 -0.0279 1.0000 1.0000 0.000 0.0167 0.04034 0.02858 -0.0286 1.0000 1.0000 0.250 0.0328 0.04166 0.02950 -0.0291 1.0000 1.0000 0.500 0.0480 0.04307 0.03059 -0.0296 1.0000 1.0000 0.750 0.0626 0.04457 0.03183 -0.0301 1.0000 1.0000 1.000 0.0766 0.04617 0.03322 -0.0307 1.0000 1.0000 1.250 0.0902 0.04787 0.03474 -0.0314 1.0000 1.0000 1.500 0.1044 0.04969 0.03641 -0.0324 0.9992 1.0000 1.750 0.1759 0.05315 0.03957 -0.0439 0.9635 1.0000 2.000 0.2160 0.05587 0.04213 -0.0494 0.9409 1.0000 2.250 0.3003 0.06044 0.04646 -0.0614 0.8994 1.0000 2.500 0.3118 0.06145 0.04744 -0.0609 0.8675 1.0000 2.750 0.3546 0.06406 0.04994 -0.0648 0.8310 1.0000 3.000 0.3968 0.06694 0.05275 -0.0689 0.8073 1.0000 3.250 0.4144 0.06994 0.05569 -0.0700 0.7919 1.0000 3.500 0.4131 0.07237 0.05811 -0.0690 0.7828 1.0000 3.750 0.4198 0.07517 0.06089 -0.0692 0.7752 1.0000 4.000 0.4392 0.07823 0.06393 -0.0707 0.7637 1.0000 4.250 0.4665 0.08120 0.06682 -0.0718 0.7380 1.0000 4.500 0.5185 0.08347 0.06898 -0.0735 0.6942 1.0000 4.750 0.5082 0.08572 0.07121 -0.0716 0.6833 1.0000 5.000 0.5101 0.08896 0.07444 -0.0717 0.6806 1.0000