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 = 0.012 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0122 0.04099 0.02233 -0.0256 1.0000 1.0000 -2.750 -0.0075 0.04058 0.02188 -0.0239 1.0000 1.0000 -2.500 -0.0042 0.04022 0.02151 -0.0219 1.0000 1.0000 -2.250 -0.0020 0.03992 0.02119 -0.0198 1.0000 1.0000 -2.000 -0.0011 0.03966 0.02091 -0.0174 1.0000 1.0000 -1.750 -0.0014 0.03942 0.02067 -0.0148 1.0000 1.0000 -1.500 -0.0027 0.03922 0.02045 -0.0121 1.0000 1.0000 -1.250 -0.0048 0.03903 0.02022 -0.0092 1.0000 1.0000 -1.000 -0.0076 0.03884 0.02002 -0.0062 1.0000 1.0000 -0.750 -0.0108 0.03867 0.01981 -0.0032 1.0000 1.0000 -0.500 -0.0140 0.03849 0.01961 -0.0002 1.0000 1.0000 -0.250 -0.0168 0.03833 0.01941 0.0028 1.0000 1.0000 0.000 -0.0186 0.03820 0.01923 0.0057 1.0000 1.0000 0.250 -0.0189 0.03812 0.01907 0.0083 1.0000 1.0000 0.500 -0.0163 0.03811 0.01897 0.0104 1.0000 1.0000 0.750 -0.0104 0.03820 0.01896 0.0121 1.0000 1.0000 1.000 -0.0019 0.03839 0.01904 0.0132 1.0000 1.0000 1.250 0.0088 0.03869 0.01925 0.0140 1.0000 1.0000 1.500 0.0207 0.03909 0.01959 0.0146 1.0000 1.0000 1.750 0.0334 0.03959 0.02005 0.0149 1.0000 1.0000 2.000 0.0463 0.04022 0.02063 0.0149 1.0000 1.0000 2.250 0.0590 0.04097 0.02139 0.0148 1.0000 1.0000 2.500 0.0710 0.04188 0.02232 0.0145 1.0000 1.0000 2.750 0.0819 0.04299 0.02347 0.0140 1.0000 1.0000 3.000 0.0912 0.04433 0.02487 0.0132 1.0000 1.0000 3.250 0.0986 0.04593 0.02653 0.0121 1.0000 1.0000 3.500 0.1049 0.04778 0.02841 0.0107 1.0000 1.0000 3.750 0.1108 0.04980 0.03045 0.0091 1.0000 1.0000 4.000 0.1172 0.05191 0.03258 0.0075 1.0000 1.0000 4.250 0.1243 0.05408 0.03477 0.0057 1.0000 1.0000 4.500 0.1320 0.05630 0.03699 0.0040 1.0000 1.0000 4.750 0.1403 0.05856 0.03925 0.0023 1.0000 1.0000 5.000 0.1492 0.06087 0.04157 0.0007 1.0000 1.0000