XFOIL Version 6.94 Calculated polar for: nabad 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 -0.3773 0.04972 0.04123 -0.0003 0.9999 0.1842 -2.750 -0.3464 0.04580 0.03688 -0.0008 0.9999 0.1608 -2.500 -0.3133 0.04271 0.03301 -0.0007 0.9999 0.1422 -2.250 -0.2827 0.03965 0.02966 -0.0001 0.9999 0.1312 -2.000 -0.2512 0.03690 0.02676 0.0003 0.9999 0.1233 -1.750 -0.2183 0.03455 0.02421 0.0009 0.9999 0.1166 -1.500 -0.1687 0.03227 0.02153 -0.0011 0.9099 0.1099 -1.250 -0.1308 0.03207 0.01934 0.0016 0.4734 0.1052 -1.000 -0.0956 0.03167 0.01802 0.0012 0.4105 0.1028 -0.750 -0.0608 0.03102 0.01694 0.0010 0.3925 0.1012 -0.500 -0.0283 0.03050 0.01616 0.0011 0.3817 0.1009 -0.250 0.0032 0.03023 0.01565 0.0012 0.3746 0.1025 0.000 0.0356 0.02991 0.01528 0.0013 0.3701 0.1062 0.250 0.0684 0.02963 0.01510 0.0012 0.3667 0.1195 0.500 0.1557 0.02764 0.01583 -0.0103 0.3628 1.0001 0.750 0.1850 0.02809 0.01605 -0.0096 0.3619 1.0001 1.000 0.2144 0.02855 0.01637 -0.0090 0.3617 1.0001 1.250 0.2440 0.02903 0.01675 -0.0085 0.3622 1.0001 1.500 0.2735 0.02953 0.01715 -0.0081 0.3623 1.0001 1.750 0.2996 0.03025 0.01751 -0.0073 0.3426 1.0001 2.000 0.3231 0.03080 0.01756 -0.0064 0.3035 1.0001 2.250 0.3471 0.03128 0.01771 -0.0058 0.2759 1.0001 2.500 0.3719 0.03167 0.01788 -0.0054 0.2563 1.0001 2.750 0.3964 0.03213 0.01800 -0.0048 0.2413 1.0001 3.000 0.4221 0.03201 0.01796 -0.0044 0.2284 1.0001 3.250 0.4467 0.03205 0.01783 -0.0039 0.2143 1.0001 3.500 0.4713 0.03196 0.01756 -0.0034 0.1994 1.0001 3.750 0.4970 0.03181 0.01737 -0.0030 0.1870 1.0001 4.000 0.5236 0.03172 0.01734 -0.0024 0.1768 1.0001 4.250 0.5505 0.03173 0.01740 -0.0018 0.1664 1.0001 4.500 0.5777 0.03191 0.01768 -0.0011 0.1572 1.0001 4.750 0.6049 0.03225 0.01815 -0.0003 0.1452 1.0001 5.000 0.6321 0.03268 0.01872 0.0005 0.1255 1.0001