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 = 0.005 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1912 0.05887 0.03958 -0.0085 0.9999 1.0001 -2.750 -0.1991 0.05588 0.03686 -0.0096 0.9999 1.0001 -2.500 -0.1980 0.05338 0.03376 -0.0127 0.9999 1.0001 -2.250 -0.1662 0.05229 0.03016 -0.0178 0.9999 1.0001 -2.000 -0.1324 0.05207 0.02774 -0.0187 0.9999 1.0001 -1.750 -0.1044 0.05193 0.02616 -0.0181 0.9999 1.0001 -1.500 -0.0785 0.05182 0.02498 -0.0172 0.9999 1.0001 -1.250 -0.0536 0.05174 0.02407 -0.0163 0.9999 1.0001 -1.000 -0.0293 0.05171 0.02335 -0.0154 0.9999 1.0001 -0.750 -0.0054 0.05170 0.02280 -0.0145 0.9999 1.0001 -0.500 0.0183 0.05173 0.02239 -0.0137 0.9999 1.0001 -0.250 0.0416 0.05180 0.02213 -0.0128 0.9999 1.0001 0.000 0.0643 0.05192 0.02202 -0.0121 0.9999 1.0001 0.250 0.0872 0.05205 0.02195 -0.0113 0.9999 1.0001 0.500 0.1099 0.05223 0.02198 -0.0105 0.9999 1.0001 0.750 0.1322 0.05245 0.02212 -0.0097 0.9999 1.0001 1.000 0.1542 0.05271 0.02238 -0.0089 0.9999 1.0001 1.250 0.1758 0.05303 0.02275 -0.0082 0.9999 1.0001 1.500 0.1969 0.05340 0.02325 -0.0074 0.9999 1.0001 1.750 0.2174 0.05385 0.02387 -0.0068 0.9999 1.0001 2.000 0.2370 0.05439 0.02464 -0.0061 0.9999 1.0001 2.250 0.2557 0.05506 0.02554 -0.0055 0.9999 1.0001 2.500 0.2726 0.05586 0.02670 -0.0051 0.9999 1.0001 2.750 0.2873 0.05695 0.02813 -0.0048 0.9999 1.0001 3.000 0.2983 0.05843 0.02992 -0.0047 0.9999 1.0001 3.250 0.3042 0.06046 0.03215 -0.0048 0.9999 1.0001 3.500 0.3052 0.06304 0.03474 -0.0052 0.9999 1.0001 3.750 0.3047 0.06590 0.03745 -0.0059 0.9999 1.0001 4.000 0.3058 0.06871 0.04007 -0.0067 0.9999 1.0001 4.250 0.3090 0.07139 0.04257 -0.0075 0.9999 1.0001 4.500 0.3139 0.07396 0.04498 -0.0083 0.9999 1.0001 4.750 0.3201 0.07645 0.04734 -0.0090 0.9999 1.0001 5.000 0.3273 0.07889 0.04967 -0.0098 0.9999 1.0001