XFOIL Version 6.96 Calculated polar for: NASA/AMES A-01 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 1.000 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -15.750 -1.2981 0.05463 0.05208 -0.0200 1.0000 0.0058 -15.500 -1.3229 0.04753 0.04476 -0.0245 1.0000 0.0057 -15.250 -1.3453 0.04232 0.03932 -0.0256 1.0000 0.0058 -15.000 -1.3603 0.03874 0.03556 -0.0244 1.0000 0.0058 -14.750 -1.3721 0.03589 0.03252 -0.0217 1.0000 0.0059 -14.500 -1.3780 0.03368 0.03014 -0.0186 1.0000 0.0060 -14.250 -1.3747 0.03162 0.02793 -0.0165 1.0000 0.0061 -14.000 -1.3654 0.02991 0.02606 -0.0149 1.0000 0.0063 -13.750 -1.3526 0.02840 0.02443 -0.0136 1.0000 0.0065 -13.500 -1.3371 0.02707 0.02299 -0.0126 1.0000 0.0067 -13.250 -1.3197 0.02587 0.02168 -0.0117 1.0000 0.0069 -13.000 -1.3009 0.02476 0.02045 -0.0109 1.0000 0.0072 -12.750 -1.2806 0.02373 0.01930 -0.0103 1.0000 0.0075 -12.500 -1.2591 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