XFOIL Version 6.96 Calculated polar for: NASA/LANGLEY LS(1)-0421 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.200 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -19.750 -0.7893 0.12202 0.11656 -0.0611 1.0000 0.0321 -19.500 -0.8040 0.11525 0.10966 -0.0652 1.0000 0.0323 -19.250 -0.8156 0.10921 0.10351 -0.0688 1.0000 0.0325 -19.000 -0.8252 0.10367 0.09784 -0.0720 1.0000 0.0327 -18.750 -0.8335 0.09845 0.09251 -0.0749 1.0000 0.0329 -18.500 -0.8409 0.09349 0.08743 -0.0777 1.0000 0.0331 -18.250 -0.8469 0.08884 0.08267 -0.0802 1.0000 0.0334 -18.000 -0.8518 0.08447 0.07818 -0.0825 1.0000 0.0337 -17.750 -0.8556 0.08036 0.07395 -0.0846 1.0000 0.0340 -17.500 -0.8585 0.07647 0.06993 -0.0865 1.0000 0.0344 -17.250 -0.8603 0.07281 0.06615 -0.0882 1.0000 0.0348 -17.000 -0.8614 0.06935 0.06255 -0.0897 1.0000 0.0353 -16.750 -0.8616 0.06611 0.05915 -0.0910 1.0000 0.0358 -16.500 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-0.0489 0.1113 0.8099 13.250 1.3385 0.05593 0.04777 -0.0483 0.1089 0.8126 13.500 1.3493 0.05750 0.04943 -0.0477 0.1064 0.8156 13.750 1.3590 0.05922 0.05121 -0.0472 0.1040 0.8189 14.000 1.3671 0.06117 0.05318 -0.0468 0.1020 0.8221 14.250 1.3764 0.06304 0.05509 -0.0465 0.1000 0.8253 14.500 1.3860 0.06476 0.05692 -0.0460 0.0978 0.8278 14.750 1.3940 0.06668 0.05893 -0.0456 0.0957 0.8305 15.000 1.4008 0.06875 0.06106 -0.0452 0.0939 0.8335 15.250 1.4066 0.07098 0.06331 -0.0449 0.0924 0.8366 15.500 1.4148 0.07303 0.06546 -0.0447 0.0906 0.8400 15.750 1.4228 0.07515 0.06768 -0.0446 0.0888 0.8433 16.000 1.4285 0.07742 0.07005 -0.0444 0.0871 0.8462 16.250 1.4331 0.07982 0.07251 -0.0443 0.0856 0.8494 16.500 1.4370 0.08234 0.07506 -0.0442 0.0843 0.8530 16.750 1.4423 0.08476 0.07758 -0.0443 0.0828 0.8570 17.000 1.4477 0.08727 0.08021 -0.0445 0.0813 0.8611 17.250 1.4511 0.08986 0.08292 -0.0445 0.0799 0.8648 17.500 1.4542 0.09256 0.08571 -0.0447 0.0786 0.8690 17.750 1.4567 0.09541 0.08862 -0.0451 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