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 = 0.035 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 0.0370 0.02923 0.01869 -0.0282 1.0000 1.0000 -2.750 0.0410 0.02892 0.01838 -0.0265 1.0000 1.0000 -2.500 0.0421 0.02873 0.01820 -0.0244 1.0000 1.0000 -2.250 0.0403 0.02862 0.01811 -0.0218 1.0000 1.0000 -2.000 0.0361 0.02856 0.01806 -0.0189 1.0000 1.0000 -1.750 0.0305 0.02852 0.01801 -0.0159 1.0000 1.0000 -1.500 0.0244 0.02847 0.01795 -0.0128 1.0000 1.0000 -1.250 0.0178 0.02838 0.01787 -0.0097 1.0000 1.0000 -1.000 0.0109 0.02826 0.01774 -0.0066 1.0000 1.0000 -0.750 0.0036 0.02810 0.01757 -0.0035 1.0000 1.0000 -0.500 -0.0043 0.02789 0.01736 -0.0003 1.0000 1.0000 -0.250 -0.0128 0.02763 0.01709 0.0029 1.0000 1.0000 0.000 -0.0220 0.02731 0.01677 0.0063 1.0000 1.0000 0.250 -0.0316 0.02695 0.01640 0.0097 1.0000 1.0000 0.500 -0.0401 0.02656 0.01599 0.0130 1.0000 1.0000 0.750 -0.0434 0.02629 0.01568 0.0155 1.0000 1.0000 1.000 -0.0364 0.02631 0.01563 0.0164 1.0000 1.0000 1.250 -0.0230 0.02658 0.01583 0.0163 1.0000 1.0000 1.500 -0.0069 0.02704 0.01624 0.0157 1.0000 1.0000 1.750 0.0097 0.02770 0.01688 0.0148 1.0000 1.0000 2.000 0.1294 0.03102 0.02036 -0.0039 0.9383 1.0000 2.250 0.2802 0.03257 0.02231 -0.0224 0.8456 1.0000 2.500 0.4511 0.02985 0.02000 -0.0349 0.7316 1.0000 2.750 0.4842 0.02877 0.01883 -0.0312 0.6684 1.0000 3.000 0.5062 0.02824 0.01809 -0.0273 0.6194 1.0000 3.250 0.5261 0.02813 0.01774 -0.0240 0.5789 1.0000 3.500 0.5461 0.02842 0.01781 -0.0215 0.5441 1.0000 3.750 0.5681 0.02883 0.01799 -0.0194 0.5132 1.0000 4.000 0.5897 0.02959 0.01858 -0.0179 0.4858 1.0000 4.250 0.6136 0.03058 0.01941 -0.0170 0.4635 1.0000 4.500 0.6349 0.03182 0.02060 -0.0161 0.4418 1.0000 4.750 0.6579 0.03307 0.02175 -0.0153 0.4219 1.0000 5.000 0.6803 0.03465 0.02336 -0.0147 0.4057 1.0000