XFOIL Version 6.94 Calculated polar for: GOE 406c 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.030 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 0.0216 0.06132 0.05605 0.0074 1.0001 0.9433 -2.750 -0.0312 0.06185 0.05682 0.0186 1.0001 0.9150 -2.500 -0.0857 0.06223 0.05743 0.0296 1.0001 0.8918 -2.250 -0.1391 0.06220 0.05760 0.0400 1.0001 0.8713 -2.000 -0.1907 0.06174 0.05735 0.0495 1.0001 0.8524 -1.500 0.3628 0.04401 0.03512 -0.1009 0.5961 0.2749 -1.250 0.4031 0.04378 0.03367 -0.1026 0.5515 0.2631 -1.000 0.4385 0.04317 0.03240 -0.1036 0.5256 0.2580 -0.750 0.4759 0.04283 0.03141 -0.1049 0.5071 0.2524 -0.500 0.5133 0.04276 0.03061 -0.1059 0.4922 0.2470 -0.250 0.5491 0.04276 0.03015 -0.1066 0.4786 0.2446 0.000 0.5825 0.04267 0.02970 -0.1069 0.4674 0.2460 0.250 0.6145 0.04262 0.02948 -0.1071 0.4574 0.2525 0.500 0.6474 0.04286 0.02941 -0.1072 0.4504 0.2618 0.750 0.6792 0.04308 0.02977 -0.1077 0.4449 0.2727 1.000 0.7112 0.04349 0.03027 -0.1082 0.4401 0.2901 1.250 0.7440 0.04381 0.03087 -0.1090 0.4360 0.3299 1.500 0.7663 0.04235 0.03117 -0.1071 0.4326 0.9999 1.750 0.7991 0.04408 0.03216 -0.1069 0.4289 0.9999 2.000 0.8276 0.04596 0.03362 -0.1073 0.4248 0.9999 2.250 0.8546 0.04791 0.03533 -0.1078 0.4206 0.9999 2.500 0.8812 0.04992 0.03714 -0.1082 0.4171 0.9999 2.750 0.9065 0.05227 0.03940 -0.1089 0.4155 0.9999 3.000 0.9289 0.05512 0.04229 -0.1097 0.4153 0.9999 3.250 0.9463 0.05866 0.04598 -0.1107 0.4164 0.9999 3.500 0.9553 0.06330 0.05089 -0.1117 0.4193 0.9999 3.750 0.9552 0.06897 0.05681 -0.1127 0.4237 0.9999 4.000 0.9557 0.07446 0.06241 -0.1134 0.4276 0.9999 4.250 0.9608 0.07943 0.06740 -0.1140 0.4305 0.9999