XFOIL Version 6.94 Calculated polar for: nacb632 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 0.0244 0.03176 0.01941 -0.0277 1.0000 1.0000 -2.750 0.0286 0.03142 0.01906 -0.0260 1.0000 1.0000 -2.500 0.0305 0.03116 0.01881 -0.0240 1.0000 1.0000 -2.250 0.0302 0.03098 0.01864 -0.0216 1.0000 1.0000 -2.000 0.0277 0.03085 0.01851 -0.0189 1.0000 1.0000 -1.750 0.0235 0.03075 0.01841 -0.0160 1.0000 1.0000 -1.500 0.0183 0.03064 0.01829 -0.0129 1.0000 1.0000 -1.250 0.0124 0.03053 0.01817 -0.0098 1.0000 1.0000 -1.000 0.0061 0.03038 0.01802 -0.0067 1.0000 1.0000 -0.750 -0.0008 0.03021 0.01784 -0.0035 1.0000 1.0000 -0.500 -0.0081 0.03001 0.01763 -0.0002 1.0000 1.0000 -0.250 -0.0159 0.02976 0.01737 0.0031 1.0000 1.0000 0.000 -0.0238 0.02948 0.01708 0.0064 1.0000 1.0000 0.250 -0.0313 0.02919 0.01677 0.0097 1.0000 1.0000 0.500 -0.0366 0.02892 0.01647 0.0127 1.0000 1.0000 0.750 -0.0358 0.02881 0.01629 0.0148 1.0000 1.0000 1.000 -0.0277 0.02890 0.01631 0.0157 1.0000 1.0000 1.250 -0.0151 0.02918 0.01651 0.0159 1.0000 1.0000 1.500 -0.0003 0.02963 0.01690 0.0156 1.0000 1.0000 1.750 0.0152 0.03023 0.01747 0.0151 1.0000 1.0000 2.000 0.0301 0.03102 0.01825 0.0144 1.0000 1.0000 2.250 0.0434 0.03205 0.01930 0.0135 1.0000 1.0000 2.500 0.0532 0.03348 0.02079 0.0120 1.0000 1.0000 2.750 0.2556 0.04054 0.02840 -0.0210 0.8411 1.0000 3.000 0.3669 0.04167 0.02986 -0.0306 0.7454 1.0000 3.250 0.4200 0.04196 0.03024 -0.0315 0.6893 1.0000 3.500 0.4668 0.04196 0.03029 -0.0310 0.6460 1.0000 3.750 0.4969 0.04244 0.03075 -0.0295 0.6109 1.0000 4.000 0.5331 0.04273 0.03101 -0.0283 0.5817 1.0000 4.250 0.5450 0.04458 0.03284 -0.0271 0.5571 1.0000 4.500 0.5824 0.04463 0.03286 -0.0255 0.5316 1.0000 4.750 0.5978 0.04685 0.03508 -0.0250 0.5147 1.0000 5.000 0.6124 0.04906 0.03730 -0.0243 0.4978 1.0000