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 = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -16.000 -1.3252 0.05356 0.05083 -0.0221 1.0000 0.0036 -15.750 -1.3570 0.04621 0.04323 -0.0258 1.0000 0.0036 -15.500 -1.3739 0.04194 0.03876 -0.0259 1.0000 0.0036 -15.250 -1.3855 0.03876 0.03541 -0.0245 1.0000 0.0036 -15.000 -1.3920 0.03635 0.03283 -0.0222 1.0000 0.0036 -14.750 -1.3954 0.03437 0.03070 -0.0192 1.0000 0.0036 -14.500 -1.3909 0.03259 0.02878 -0.0171 1.0000 0.0037 -14.250 -1.3819 0.03095 0.02700 -0.0155 1.0000 0.0037 -14.000 -1.3696 0.02949 0.02539 -0.0142 1.0000 0.0038 -13.750 -1.3551 0.02816 0.02393 -0.0130 0.9829 0.0039 -13.500 -1.3414 0.02704 0.02267 -0.0114 0.9636 0.0039 -13.250 -1.3265 0.02600 0.02149 -0.0098 0.9548 0.0040 -13.000 -1.3101 0.02502 0.02038 -0.0083 0.9485 0.0040 -12.750 -1.2925 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