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 = 0.500 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -12.000 -0.6393 0.13696 0.13480 0.0264 1.0000 0.0213 -11.750 -1.0581 0.03936 0.03598 -0.0208 1.0000 0.0134 -11.500 -1.0513 0.03731 0.03380 -0.0197 1.0000 0.0138 -11.250 -1.0279 0.03737 0.03392 -0.0195 1.0000 0.0141 -11.000 -1.0041 0.03740 0.03398 -0.0194 1.0000 0.0144 -10.750 -0.9865 0.03624 0.03271 -0.0190 1.0000 0.0149 -10.500 -0.9777 0.03323 0.02936 -0.0181 1.0000 0.0156 -10.250 -0.9664 0.03021 0.02593 -0.0171 1.0000 0.0163 -10.000 -0.9489 0.02809 0.02346 -0.0166 1.0000 0.0168 -9.750 -0.9333 0.02547 0.02054 -0.0159 1.0000 0.0175 -9.500 -0.9079 0.02496 0.01999 -0.0161 1.0000 0.0180 -9.250 -0.8820 0.02446 0.01946 -0.0163 1.0000 0.0185 -9.000 -0.8568 0.02368 0.01857 -0.0164 1.0000 0.0192 -8.750 -0.8322 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