XFOIL Version 6.94 Calculated polar for: GOA 1b 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.008 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1796 0.03614 0.01666 -0.0232 1.0000 1.0000 -2.750 -0.1533 0.03596 0.01516 -0.0226 1.0000 1.0000 -2.500 -0.1281 0.03581 0.01395 -0.0219 1.0000 1.0000 -2.250 -0.1033 0.03569 0.01294 -0.0212 1.0000 1.0000 -2.000 -0.0788 0.03560 0.01198 -0.0205 1.0000 1.0000 -1.750 -0.0545 0.03554 0.01125 -0.0198 1.0000 1.0000 -1.500 -0.0303 0.03550 0.01064 -0.0190 1.0000 1.0000 -1.250 -0.0063 0.03549 0.01016 -0.0183 1.0000 1.0000 -1.000 0.0177 0.03550 0.00979 -0.0175 1.0000 1.0000 -0.750 0.0415 0.03554 0.00954 -0.0168 1.0000 1.0000 -0.500 0.0653 0.03560 0.00932 -0.0160 1.0000 1.0000 -0.250 0.0891 0.03569 0.00928 -0.0152 1.0000 1.0000 0.000 0.1128 0.03580 0.00937 -0.0145 1.0000 1.0000 0.250 0.1365 0.03593 0.00957 -0.0137 1.0000 1.0000 0.500 0.1603 0.03608 0.00990 -0.0129 1.0000 1.0000 0.750 0.1843 0.03625 0.01032 -0.0120 1.0000 1.0000 1.000 0.2086 0.03644 0.01093 -0.0112 1.0000 1.0000 1.250 0.2336 0.03664 0.01174 -0.0104 1.0000 1.0000 1.500 0.2602 0.03684 0.01279 -0.0098 1.0000 1.0000 1.750 0.2905 0.03703 0.01422 -0.0098 1.0000 1.0000 2.000 0.3250 0.03738 0.01630 -0.0118 1.0000 1.0000 2.250 0.3451 0.03891 0.01913 -0.0161 1.0000 1.0000 2.500 0.3428 0.04193 0.02214 -0.0197 1.0000 1.0000 2.750 0.3424 0.04506 0.02501 -0.0230 1.0000 1.0000 3.000 0.5239 0.05382 0.03448 -0.0586 0.7765 1.0000 3.250 0.5687 0.05717 0.03796 -0.0628 0.7314 1.0000 3.500 0.5914 0.06058 0.04139 -0.0653 0.7146 1.0000 3.750 0.6010 0.06402 0.04477 -0.0669 0.7082 1.0000 4.000 0.6125 0.06758 0.04830 -0.0688 0.7040 1.0000 4.250 0.6202 0.07114 0.05179 -0.0703 0.7032 1.0000 4.500 0.6225 0.07456 0.05517 -0.0712 0.7053 1.0000 4.750 0.6257 0.07805 0.05857 -0.0723 0.7090 1.0000 5.000 0.6328 0.08177 0.06223 -0.0740 0.7138 1.0000