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 = 1.000 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -19.500 -0.9388 0.12166 0.11873 -0.0313 1.0000 0.0146 -19.250 -0.9663 0.11230 0.10920 -0.0368 1.0000 0.0146 -19.000 -0.9902 0.10388 0.10064 -0.0416 1.0000 0.0147 -18.750 -1.0107 0.09637 0.09298 -0.0459 1.0000 0.0147 -18.500 -1.0285 0.08951 0.08599 -0.0497 1.0000 0.0147 -18.250 -1.0437 0.08320 0.07956 -0.0531 1.0000 0.0148 -18.000 -1.0558 0.07755 0.07381 -0.0560 1.0000 0.0149 -17.750 -1.0657 0.07234 0.06849 -0.0587 1.0000 0.0149 -17.500 -1.0728 0.06766 0.06371 -0.0611 1.0000 0.0150 -17.250 -1.0777 0.06342 0.05939 -0.0631 1.0000 0.0151 -17.000 -1.0808 0.05956 0.05544 -0.0649 1.0000 0.0151 -16.750 -1.0821 0.05604 0.05184 -0.0663 1.0000 0.0152 -16.500 -1.0818 0.05286 0.04859 -0.0675 1.0000 0.0153 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0.01389 -0.0793 0.1258 0.7912 12.750 1.7653 0.02059 0.01455 -0.0777 0.1210 0.7948 13.000 1.7792 0.02134 0.01530 -0.0760 0.1165 0.7981 13.250 1.7933 0.02209 0.01607 -0.0744 0.1119 0.8015 13.500 1.8061 0.02294 0.01695 -0.0728 0.1076 0.8054 13.750 1.8188 0.02380 0.01785 -0.0712 0.1037 0.8092 14.000 1.8299 0.02482 0.01889 -0.0695 0.1002 0.8130 14.250 1.8425 0.02576 0.01987 -0.0681 0.0969 0.8166 14.500 1.8512 0.02699 0.02111 -0.0664 0.0935 0.8203 14.750 1.8623 0.02807 0.02227 -0.0651 0.0906 0.8245 15.000 1.8707 0.02940 0.02364 -0.0636 0.0874 0.8287 15.250 1.8784 0.03083 0.02511 -0.0622 0.0847 0.8330 15.500 1.8860 0.03230 0.02663 -0.0609 0.0820 0.8369 15.750 1.8901 0.03408 0.02847 -0.0595 0.0794 0.8416 16.000 1.8966 0.03572 0.03019 -0.0583 0.0772 0.8464 16.250 1.9001 0.03766 0.03219 -0.0572 0.0749 0.8515 16.500 1.8996 0.04003 0.03462 -0.0560 0.0727 0.8565 16.750 1.9024 0.04217 0.03685 -0.0551 0.0709 0.8623 17.000 1.9022 0.04469 0.03946 -0.0544 0.0689 0.8684 17.250 1.8974 0.04778 0.04263 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