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 = 0.012 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.3791 0.05794 0.04320 -0.0058 0.9999 0.3173 -2.750 -0.3481 0.05513 0.03994 -0.0072 0.9999 0.3187 -2.500 -0.3144 0.05254 0.03684 -0.0086 0.9999 0.3233 -2.250 -0.2811 0.04993 0.03405 -0.0094 0.9999 0.3335 -2.000 -0.2457 0.04758 0.03147 -0.0103 0.9999 0.3503 -1.750 -0.2108 0.04514 0.02914 -0.0108 0.9999 0.3777 -1.500 -0.1758 0.04257 0.02700 -0.0110 0.9999 0.4245 -1.250 -0.1428 0.03890 0.02498 -0.0101 0.9999 0.5349 -1.000 -0.0388 0.03688 0.02140 -0.0201 0.9999 1.0001 -0.750 -0.0139 0.03695 0.02079 -0.0184 0.9999 1.0001 -0.500 0.0102 0.03703 0.02046 -0.0169 0.9999 1.0001 -0.250 0.0342 0.03713 0.02030 -0.0154 0.9999 1.0001 0.000 0.0583 0.03724 0.02027 -0.0141 0.9999 1.0001 0.250 0.0828 0.03736 0.02033 -0.0129 0.9999 1.0001 0.500 0.1078 0.03748 0.02050 -0.0119 0.9999 1.0001 0.750 0.1336 0.03762 0.02078 -0.0112 0.9999 1.0001 1.000 0.1601 0.03779 0.02121 -0.0108 0.9999 1.0001 1.250 0.1865 0.03807 0.02188 -0.0109 0.9999 1.0001 1.500 0.2095 0.03871 0.02297 -0.0113 0.9999 1.0001 1.750 0.2197 0.04037 0.02503 -0.0117 0.9999 1.0001 2.000 0.2049 0.04394 0.02858 -0.0115 0.9999 1.0001 2.250 0.2887 0.04744 0.03201 -0.0294 0.9308 1.0001 2.500 0.3365 0.04958 0.03413 -0.0366 0.8935 1.0001 2.750 0.3693 0.05162 0.03615 -0.0404 0.8732 1.0001 3.000 0.3910 0.05370 0.03820 -0.0421 0.8607 1.0001 3.250 0.4112 0.05579 0.04026 -0.0435 0.8535 1.0001 3.500 0.4302 0.05792 0.04235 -0.0446 0.8500 1.0001 3.750 0.4429 0.05999 0.04425 -0.0444 0.8445 1.0001 4.000 0.6245 0.04719 0.03159 -0.0368 0.6091 1.0001 4.250 0.6456 0.04870 0.03251 -0.0314 0.5482 1.0001 4.500 0.6656 0.05124 0.03487 -0.0282 0.4991 1.0001 4.750 0.6826 0.05442 0.03771 -0.0242 0.4414 1.0001 5.000 0.7033 0.05868 0.04201 -0.0224 0.3995 1.0001