XFOIL Version 6.96 Calculated polar for: NASA/LANGLEY 64-012 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 -1.0425 0.08111 0.07882 -0.0110 1.0000 0.0098 -15.500 -1.0658 0.07277 0.07031 -0.0163 1.0000 0.0098 -15.250 -1.0851 0.06597 0.06335 -0.0201 1.0000 0.0098 -15.000 -1.1082 0.05916 0.05633 -0.0234 1.0000 0.0098 -14.750 -1.1188 0.05469 0.05175 -0.0253 1.0000 0.0098 -14.500 -1.1331 0.04996 0.04684 -0.0268 1.0000 0.0099 -14.250 -1.1395 0.04653 0.04329 -0.0276 1.0000 0.0100 -14.000 -1.1410 0.04387 0.04056 -0.0279 1.0000 0.0101 -13.750 -1.1500 0.04040 0.03690 -0.0279 1.0000 0.0101 -13.500 -1.1458 0.03868 0.03513 -0.0276 1.0000 0.0103 -13.250 -1.1423 0.03697 0.03335 -0.0271 1.0000 0.0104 -13.000 -1.1414 0.03499 0.03124 -0.0262 1.0000 0.0106 -12.750 -1.1408 0.03301 0.02912 -0.0250 1.0000 0.0106 -12.500 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