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 = 0.022 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 0.0169 0.03340 0.01990 -0.0274 1.0000 1.0000 -2.750 0.0212 0.03304 0.01953 -0.0257 1.0000 1.0000 -2.500 0.0235 0.03276 0.01925 -0.0237 1.0000 1.0000 -2.250 0.0238 0.03255 0.01904 -0.0213 1.0000 1.0000 -2.000 0.0221 0.03238 0.01888 -0.0187 1.0000 1.0000 -1.750 0.0187 0.03224 0.01874 -0.0159 1.0000 1.0000 -1.500 0.0143 0.03211 0.01859 -0.0129 1.0000 1.0000 -1.250 0.0090 0.03198 0.01845 -0.0098 1.0000 1.0000 -1.000 0.0031 0.03183 0.01829 -0.0066 1.0000 1.0000 -0.750 -0.0032 0.03165 0.01810 -0.0034 1.0000 1.0000 -0.500 -0.0099 0.03145 0.01789 -0.0002 1.0000 1.0000 -0.250 -0.0169 0.03122 0.01765 0.0031 1.0000 1.0000 0.000 -0.0239 0.03097 0.01738 0.0064 1.0000 1.0000 0.250 -0.0298 0.03072 0.01711 0.0096 1.0000 1.0000 0.500 -0.0332 0.03052 0.01686 0.0124 1.0000 1.0000 0.750 -0.0308 0.03047 0.01673 0.0143 1.0000 1.0000 1.000 -0.0224 0.03060 0.01678 0.0152 1.0000 1.0000 1.250 -0.0103 0.03088 0.01698 0.0156 1.0000 1.0000 1.500 0.0038 0.03131 0.01735 0.0155 1.0000 1.0000 1.750 0.0186 0.03189 0.01789 0.0152 1.0000 1.0000 2.000 0.0331 0.03263 0.01862 0.0147 1.0000 1.0000 2.250 0.0465 0.03357 0.01958 0.0140 1.0000 1.0000 2.500 0.0576 0.03482 0.02087 0.0129 1.0000 1.0000 2.750 0.0651 0.03649 0.02261 0.0113 1.0000 1.0000 3.000 0.1800 0.04341 0.02984 -0.0114 0.9139 1.0000 3.250 0.2833 0.04766 0.03437 -0.0252 0.8141 1.0000 3.500 0.3485 0.04983 0.03672 -0.0304 0.7499 1.0000 3.750 0.3775 0.05165 0.03856 -0.0312 0.7087 1.0000 4.000 0.4083 0.05351 0.04049 -0.0321 0.6742 1.0000 4.250 0.4378 0.05524 0.04225 -0.0324 0.6426 1.0000 4.500 0.4559 0.05751 0.04452 -0.0325 0.6209 1.0000 4.750 0.4857 0.05946 0.04650 -0.0328 0.5978 1.0000 5.000 0.4941 0.06197 0.04898 -0.0323 0.5798 1.0000