XFOIL Version 6.94 Calculated polar for: manu04 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.035 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.3593 0.05890 0.05302 0.0534 0.9999 0.8392 -2.750 -0.4206 0.05648 0.05072 0.0652 0.9999 0.8208 -2.500 -0.4653 0.05368 0.04804 0.0726 0.9999 0.8023 -2.250 -0.4742 0.05112 0.04551 0.0720 0.9999 0.7759 -2.000 -0.4028 0.05176 0.04584 0.0473 0.9999 0.6560 -1.750 0.0028 0.04226 0.03162 -0.0407 0.4731 0.2872 -1.500 0.0282 0.04146 0.03017 -0.0404 0.4522 0.2861 -1.250 0.0606 0.04050 0.02858 -0.0413 0.4382 0.2843 -1.000 0.0991 0.03943 0.02700 -0.0433 0.4270 0.2846 -0.750 0.1549 0.03868 0.02541 -0.0487 0.4170 0.2896 -0.500 0.2148 0.03782 0.02408 -0.0550 0.4082 0.3032 -0.250 0.2861 0.03723 0.02291 -0.0637 0.3981 0.3364 0.000 0.3845 0.03557 0.02131 -0.0780 0.3876 0.4429 0.250 0.4526 0.03401 0.02076 -0.0862 0.3815 0.6272 0.500 0.5878 0.03492 0.02162 -0.1095 0.3766 1.0001 0.750 0.6346 0.03585 0.02219 -0.1134 0.3752 1.0001 1.000 0.6747 0.03680 0.02291 -0.1158 0.3746 1.0001 1.250 0.7100 0.03777 0.02376 -0.1173 0.3745 1.0001 1.500 0.7419 0.03874 0.02467 -0.1181 0.3749 1.0001 1.750 0.7717 0.03974 0.02566 -0.1186 0.3758 1.0001 2.000 0.7992 0.04078 0.02678 -0.1186 0.3771 1.0001 2.250 0.8251 0.04189 0.02802 -0.1183 0.3791 1.0001 2.500 0.8498 0.04315 0.02941 -0.1180 0.3815 1.0001 2.750 0.8736 0.04455 0.03092 -0.1175 0.3842 1.0001 3.000 0.8964 0.04606 0.03254 -0.1169 0.3865 1.0001 3.250 0.9178 0.04765 0.03423 -0.1162 0.3886 1.0001 3.500 0.9377 0.04936 0.03608 -0.1152 0.3904 1.0001 3.750 0.9567 0.05120 0.03801 -0.1141 0.3922 1.0001 4.000 0.9750 0.05319 0.04009 -0.1130 0.3937 1.0001 4.250 0.9949 0.05540 0.04234 -0.1123 0.3956 1.0001 4.500 0.9963 0.05701 0.04448 -0.1091 0.4027 1.0001 4.750 1.0034 0.05960 0.04735 -0.1071 0.4099 1.0001 5.000 1.0199 0.06238 0.05020 -0.1064 0.4154 1.0001