XFOIL Version 6.94 Calculated polar for: nabab 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.045 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -2.750 -0.2703 0.04158 0.03069 -0.0169 0.3329 0.1149 -2.500 -0.2427 0.03924 0.02775 -0.0158 0.3205 0.1056 -2.250 -0.2131 0.03699 0.02475 -0.0146 0.3130 0.0970 -2.000 -0.1836 0.03495 0.02234 -0.0138 0.3072 0.0910 -1.750 -0.1520 0.03339 0.02025 -0.0129 0.3026 0.0858 -1.500 -0.1183 0.03209 0.01860 -0.0127 0.2991 0.0816 -1.250 -0.0811 0.03136 0.01740 -0.0129 0.2965 0.0780 -1.000 -0.0475 0.03068 0.01654 -0.0129 0.2950 0.0777 -0.750 -0.0153 0.02997 0.01582 -0.0128 0.2941 0.0807 -0.500 0.0165 0.02952 0.01538 -0.0127 0.2936 0.0863 -0.250 0.0486 0.02919 0.01519 -0.0127 0.2914 0.0953 0.000 0.0779 0.02902 0.01491 -0.0126 0.2613 0.1109 0.500 0.2051 0.02728 0.01526 -0.0276 0.1865 1.0001 0.750 0.2310 0.02742 0.01511 -0.0268 0.1706 1.0001 1.000 0.2569 0.02745 0.01485 -0.0261 0.1547 1.0001 1.250 0.2831 0.02737 0.01449 -0.0254 0.1397 1.0001 1.500 0.3098 0.02723 0.01409 -0.0247 0.1260 1.0001 1.750 0.3369 0.02707 0.01374 -0.0241 0.1131 1.0001 2.000 0.3644 0.02692 0.01364 -0.0232 0.1007 1.0001 2.750 0.4452 0.02778 0.01391 -0.0209 0.0678 1.0001 3.000 0.4727 0.02824 0.01422 -0.0201 0.0670 1.0001 3.250 0.5004 0.02872 0.01465 -0.0195 0.0668 1.0001 3.500 0.5282 0.02921 0.01514 -0.0189 0.0670 1.0001 3.750 0.5561 0.02971 0.01567 -0.0182 0.0674 1.0001 4.000 0.5838 0.03026 0.01624 -0.0175 0.0682 1.0001 4.250 0.6115 0.03086 0.01688 -0.0168 0.0691 1.0001 4.500 0.6390 0.03155 0.01763 -0.0161 0.0701 1.0001 4.750 0.6662 0.03240 0.01851 -0.0153 0.0712 1.0001 5.000 0.6939 0.03295 0.01933 -0.0144 0.0726 1.0001