XFOIL Version 6.96 Calculated polar for: RUTAN WING AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.100 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -5.000 -0.5618 0.02775 0.01816 -0.0142 1.0000 0.1154 -4.500 -0.5191 0.02578 0.01606 -0.0126 1.0000 0.1230 -4.000 -0.4734 0.02395 0.01410 -0.0115 1.0000 0.1336 -3.500 -0.4275 0.02251 0.01264 -0.0103 1.0000 0.1479 -3.000 -0.3819 0.02133 0.01161 -0.0093 1.0000 0.1685 -2.500 -0.3241 0.02019 0.01078 -0.0108 0.9966 0.2061 -2.000 -0.2643 0.01724 0.01067 -0.0119 0.9910 0.6939 -1.500 -0.2251 0.01840 0.01205 -0.0047 0.9803 0.8586 -1.000 -0.1710 0.01938 0.01295 -0.0012 0.9708 0.9353 -0.500 0.0030 0.02079 0.01411 -0.0227 0.9724 0.9912 0.000 0.0902 0.02105 0.01426 -0.0314 0.9618 1.0000 0.500 0.1443 0.02130 0.01447 -0.0338 0.9468 1.0000 1.000 0.1912 0.02169 0.01485 -0.0346 0.9312 1.0000 1.500 0.3447 0.02086 0.01413 -0.0512 0.9018 1.0000 2.000 0.4443 0.01922 0.01262 -0.0559 0.8653 1.0000 2.500 0.4924 0.01774 0.01122 -0.0509 0.8252 1.0000 3.000 0.5287 0.01631 0.00982 -0.0437 0.7813 1.0000 3.500 0.5584 0.01495 0.00848 -0.0356 0.7044 1.0000 4.000 0.5579 0.01696 0.00790 -0.0232 0.2570 1.0000 4.500 0.5853 0.01906 0.00932 -0.0187 0.1951 1.0000 5.000 0.6284 0.02088 0.01084 -0.0166 0.1654 1.0000 5.500 0.6783 0.02294 0.01265 -0.0157 0.1458 1.0000 6.000 0.7311 0.02517 0.01486 -0.0151 0.1317 1.0000 6.500 0.7842 0.02787 0.01769 -0.0146 0.1211 1.0000 7.000 0.8351 0.03108 0.02099 -0.0142 0.1125 1.0000 7.500 0.8815 0.03445 0.02494 -0.0126 0.1062 1.0000 8.000 0.9274 0.03912 0.02966 -0.0121 0.1008 1.0000 8.500 0.9567 0.04422 0.03578 -0.0088 0.0980 1.0000 9.000 0.9752 0.05132 0.04371 -0.0053 0.0975 1.0000 9.500 0.9807 0.05971 0.05278 -0.0018 0.0988 1.0000