XFOIL Version 6.96 Calculated polar for: NASA/LANGLEY LS(1)-0417MOD 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 = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -19.000 -0.8623 0.12900 0.12536 -0.0266 1.0000 0.0145 -18.500 -0.9175 0.10962 0.10564 -0.0377 1.0000 0.0145 -18.250 -0.9361 0.10211 0.09799 -0.0419 1.0000 0.0146 -18.000 -0.9531 0.09506 0.09082 -0.0458 1.0000 0.0146 -17.750 -0.9674 0.08864 0.08428 -0.0494 1.0000 0.0147 -17.500 -0.9798 0.08261 0.07813 -0.0527 1.0000 0.0147 -17.250 -0.9900 0.07712 0.07252 -0.0557 1.0000 0.0148 -17.000 -0.9981 0.07204 0.06733 -0.0584 1.0000 0.0149 -16.750 -1.0041 0.06742 0.06260 -0.0608 1.0000 0.0150 -16.500 -1.0081 0.06325 0.05834 -0.0629 1.0000 0.0150 -16.250 -1.0106 0.05942 0.05441 -0.0646 1.0000 0.0151 -16.000 -1.0115 0.05596 0.05086 -0.0661 1.0000 0.0152 -15.750 -1.0112 0.05282 0.04763 -0.0673 1.0000 0.0153 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0.02688 -0.0584 0.0988 0.8090 14.500 1.7478 0.03501 0.02860 -0.0572 0.0948 0.8128 14.750 1.7525 0.03686 0.03050 -0.0560 0.0911 0.8168 15.000 1.7556 0.03890 0.03259 -0.0549 0.0872 0.8205 15.250 1.7573 0.04109 0.03487 -0.0538 0.0841 0.8246 15.500 1.7589 0.04340 0.03727 -0.0530 0.0813 0.8289 15.750 1.7569 0.04616 0.04010 -0.0522 0.0790 0.8332 16.250 1.7535 0.05196 0.04612 -0.0514 0.0750 0.8413 16.500 1.7468 0.05568 0.04995 -0.0515 0.0732 0.8456 16.750 1.7389 0.05975 0.05415 -0.0519 0.0717 0.8501 17.250 1.7160 0.06948 0.06416 -0.0539 0.0687 0.8588 17.500 1.6983 0.07558 0.07041 -0.0557 0.0674 0.8633 17.750 1.6762 0.08242 0.07739 -0.0579 0.0664 0.8675 18.000 1.6578 0.08866 0.08378 -0.0599 0.0652 0.8719 18.250 1.6368 0.09528 0.09054 -0.0620 0.0641 0.8764