XFOIL Version 6.96 Calculated polar for: NASA/LANGLEY MS(1)-0317 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 ------ -------- --------- --------- -------- -------- -------- -19.750 -1.2481 0.07949 0.07540 -0.0493 1.0000 0.0149 -19.500 -1.2623 0.07376 0.06955 -0.0521 1.0000 0.0150 -19.250 -1.2722 0.06882 0.06450 -0.0544 1.0000 0.0151 -19.000 -1.2794 0.06437 0.05996 -0.0564 1.0000 0.0152 -18.750 -1.2838 0.06046 0.05595 -0.0579 1.0000 0.0153 -18.500 -1.2860 0.05695 0.05236 -0.0592 1.0000 0.0154 -18.250 -1.2862 0.05377 0.04910 -0.0603 1.0000 0.0156 -18.000 -1.2849 0.05086 0.04612 -0.0612 1.0000 0.0157 -17.750 -1.2824 0.04819 0.04337 -0.0618 1.0000 0.0159 -17.500 -1.2784 0.04576 0.04087 -0.0624 1.0000 0.0160 -17.250 -1.2730 0.04354 0.03859 -0.0627 1.0000 0.0162 -17.000 -1.2669 0.04147 0.03646 -0.0629 1.0000 0.0163 -16.750 -1.2603 0.03953 0.03445 -0.0631 1.0000 0.0165 -16.500 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