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 = 0.022 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1955 0.07175 0.06538 0.0406 1.0001 0.8041 -2.750 -0.2267 0.07069 0.06453 0.0442 1.0001 0.7698 -2.500 -0.2607 0.06994 0.06398 0.0455 1.0001 0.7351 -2.250 -0.2399 0.06771 0.06179 0.0312 1.0001 0.6595 -2.000 -0.1163 0.06461 0.05824 -0.0145 1.0001 0.5078 -1.750 -0.0436 0.06243 0.05586 -0.0334 1.0001 0.4384 -1.500 0.0109 0.06163 0.05494 -0.0462 1.0001 0.3954 -1.250 0.3984 0.04771 0.03882 -0.1109 0.7348 0.3164 -1.000 0.4482 0.04713 0.03707 -0.1128 0.6719 0.3101 -0.750 0.4848 0.04680 0.03600 -0.1131 0.6329 0.3107 -0.500 0.5188 0.04658 0.03526 -0.1134 0.6058 0.3143 -0.250 0.5531 0.04663 0.03487 -0.1139 0.5868 0.3192 0.000 0.5871 0.04691 0.03473 -0.1145 0.5722 0.3252 0.250 0.6198 0.04709 0.03464 -0.1146 0.5598 0.3366 0.500 0.6496 0.04775 0.03533 -0.1152 0.5482 0.3552 0.750 0.6807 0.04783 0.03546 -0.1150 0.5379 0.3873 1.000 0.7073 0.04845 0.03691 -0.1158 0.5283 0.4512 1.250 0.7268 0.04814 0.03748 -0.1131 0.5210 0.9999 1.500 0.7599 0.05011 0.03845 -0.1130 0.5149 0.9999 1.750 0.7837 0.05313 0.04103 -0.1140 0.5113 0.9999 2.000 0.8044 0.05649 0.04411 -0.1153 0.5087 0.9999 2.250 0.8205 0.06031 0.04779 -0.1167 0.5073 0.9999 2.500 0.8296 0.06479 0.05223 -0.1179 0.5068 0.9999 2.750 0.8307 0.06995 0.05741 -0.1189 0.5072 0.9999 3.000 0.8250 0.07567 0.06316 -0.1196 0.5088 0.9999 3.250 0.8161 0.08158 0.06907 -0.1202 0.5113 0.9999 3.500 0.8127 0.08701 0.07443 -0.1209 0.5140 0.9999 4.000 0.7624 0.10186 0.08945 -0.1232 0.5372 0.9999 4.250 0.6981 0.11135 0.09936 -0.1251 0.5732 0.9999 4.500 0.6751 0.11886 0.10702 -0.1287 0.6202 0.9999