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 = 0.026 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0822 0.06632 0.06054 0.0248 1.0001 0.8757 -2.750 -0.1341 0.06646 0.06091 0.0347 1.0001 0.8514 -2.500 -0.1853 0.06621 0.06088 0.0441 1.0001 0.8305 -2.250 -0.2318 0.06566 0.06053 0.0510 1.0001 0.8058 -2.000 -0.2558 0.06453 0.05956 0.0496 1.0001 0.7602 -1.750 -0.1245 0.06196 0.05675 -0.0058 1.0001 0.5573 -1.500 0.3607 0.04539 0.03722 -0.1043 0.7095 0.2981 -1.250 0.4070 0.04501 0.03547 -0.1061 0.6263 0.2863 -1.000 0.4422 0.04435 0.03397 -0.1064 0.5856 0.2805 -0.750 0.4791 0.04404 0.03284 -0.1072 0.5606 0.2750 -0.500 0.5157 0.04396 0.03212 -0.1081 0.5416 0.2726 -0.250 0.5505 0.04399 0.03171 -0.1088 0.5266 0.2753 0.000 0.5843 0.04413 0.03135 -0.1091 0.5140 0.2813 0.250 0.6171 0.04444 0.03144 -0.1096 0.5016 0.2881 0.500 0.6484 0.04459 0.03142 -0.1095 0.4916 0.2982 0.750 0.6789 0.04501 0.03195 -0.1099 0.4826 0.3151 1.000 0.7105 0.04535 0.03246 -0.1104 0.4762 0.3478 1.500 0.7641 0.04555 0.03423 -0.1091 0.4683 0.9999 1.750 0.7947 0.04771 0.03576 -0.1095 0.4650 0.9999 2.000 0.8219 0.05001 0.03766 -0.1103 0.4615 0.9999 2.250 0.8477 0.05233 0.03971 -0.1109 0.4576 0.9999 2.500 0.8742 0.05450 0.04157 -0.1113 0.4535 0.9999 2.750 0.8963 0.05729 0.04423 -0.1120 0.4505 0.9999 3.000 0.9112 0.06098 0.04802 -0.1129 0.4493 0.9999 3.250 0.9245 0.06500 0.05209 -0.1138 0.4500 0.9999 3.500 0.9373 0.06919 0.05628 -0.1147 0.4515 0.9999 3.750 0.8738 0.08160 0.06948 -0.1167 0.4660 0.9999 4.000 0.8588 0.08877 0.07674 -0.1181 0.4750 0.9999 4.250 0.7618 0.10372 0.09226 -0.1215 0.5106 0.9999 4.500 0.7551 0.11103 0.09955 -0.1250 0.5369 0.9999