XFOIL Version 6.96 Calculated polar for: NASA/LANGLEY RC-12(B)3 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 ------ -------- --------- --------- -------- -------- -------- -14.500 -1.0168 0.04947 0.04701 -0.0564 1.0000 0.0135 -14.250 -1.0421 0.04385 0.04117 -0.0579 1.0000 0.0135 -14.000 -1.0648 0.03965 0.03675 -0.0563 1.0000 0.0136 -13.750 -1.0801 0.03656 0.03347 -0.0535 1.0000 0.0137 -13.500 -1.0859 0.03441 0.03119 -0.0505 1.0000 0.0138 -13.250 -1.0849 0.03284 0.02952 -0.0478 1.0000 0.0140 -13.000 -1.0795 0.03161 0.02821 -0.0452 1.0000 0.0141 -12.750 -1.0719 0.03053 0.02705 -0.0428 1.0000 0.0143 -12.500 -1.0618 0.02943 0.02585 -0.0407 1.0000 0.0145 -12.250 -1.0501 0.02820 0.02451 -0.0390 1.0000 0.0147 -12.000 -1.0372 0.02700 0.02318 -0.0373 1.0000 0.0149 -11.750 -1.0162 0.02570 0.02171 -0.0372 0.9868 0.0151 -11.500 -0.9964 0.02453 0.02039 -0.0366 0.9760 0.0154 -11.250 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0.0096 1.0000 18.750 1.3196 0.08020 0.07609 0.0299 0.0095 1.0000