XFOIL Version 6.96 Calculated polar for: NASA/LANGLEY NLF(1)-0416 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 ------ -------- --------- --------- -------- -------- -------- -18.500 -0.7425 0.13733 0.13491 -0.0255 1.0000 0.0054 -18.250 -0.7673 0.12784 0.12530 -0.0302 1.0000 0.0054 -18.000 -0.7966 0.11786 0.11521 -0.0352 1.0000 0.0055 -17.750 -0.8168 0.10992 0.10716 -0.0391 1.0000 0.0055 -17.500 -0.8366 0.10226 0.09939 -0.0430 1.0000 0.0055 -17.250 -0.8570 0.09467 0.09169 -0.0468 1.0000 0.0054 -17.000 -0.8742 0.08782 0.08473 -0.0503 1.0000 0.0054 -16.750 -0.8924 0.08094 0.07774 -0.0539 1.0000 0.0054 -16.500 -0.9087 0.07457 0.07125 -0.0572 1.0000 0.0054 -16.250 -0.9240 0.06855 0.06511 -0.0603 1.0000 0.0054 -16.000 -0.9330 0.06367 0.06012 -0.0627 1.0000 0.0054 -15.750 -0.9447 0.05863 0.05496 -0.0651 1.0000 0.0054 -15.500 -0.9558 0.05393 0.05014 -0.0671 1.0000 0.0054 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0.3652 0.6663 4.250 0.9365 0.00898 0.00310 -0.1023 0.3565 0.6673 4.500 0.9638 0.00913 0.00322 -0.1023 0.3457 0.6682 4.750 0.9911 0.00928 0.00334 -0.1022 0.3353 0.6690 5.000 1.0178 0.00947 0.00349 -0.1021 0.3235 0.6697 5.250 1.0442 0.00968 0.00366 -0.1019 0.3114 0.6704 5.500 1.0706 0.00990 0.00383 -0.1017 0.2990 0.6711 5.750 1.0967 0.01012 0.00402 -0.1015 0.2864 0.6719 6.000 1.1226 0.01037 0.00423 -0.1012 0.2740 0.6726 6.250 1.1480 0.01064 0.00445 -0.1009 0.2611 0.6735 6.500 1.1732 0.01093 0.00468 -0.1005 0.2487 0.6744 6.750 1.1983 0.01121 0.00493 -0.1001 0.2369 0.6754 7.000 1.2233 0.01148 0.00517 -0.0997 0.2252 0.6762 7.250 1.2479 0.01179 0.00543 -0.0993 0.2136 0.6769 7.500 1.2720 0.01211 0.00572 -0.0988 0.2025 0.6776 7.750 1.2958 0.01245 0.00601 -0.0982 0.1916 0.6783 8.000 1.3198 0.01276 0.00630 -0.0976 0.1815 0.6790 8.250 1.3429 0.01311 0.00662 -0.0970 0.1714 0.6797 8.500 1.3651 0.01351 0.00697 -0.0962 0.1604 0.6803 8.750 1.3876 0.01388 0.00731 -0.0954 0.1509 0.6809 9.000 1.4095 0.01425 0.00767 -0.0945 0.1422 0.6815 9.250 1.4301 0.01469 0.00807 -0.0935 0.1329 0.6825 9.500 1.4512 0.01506 0.00845 -0.0925 0.1253 0.6834 9.750 1.4707 0.01550 0.00887 -0.0913 0.1173 0.6843 10.000 1.4882 0.01603 0.00936 -0.0898 0.1077 0.6852 10.250 1.5040 0.01653 0.00984 -0.0879 0.0986 0.6861 10.500 1.5169 0.01707 0.01036 -0.0856 0.0897 0.6871 10.750 1.5289 0.01768 0.01095 -0.0832 0.0819 0.6882 11.000 1.5412 0.01829 0.01155 -0.0808 0.0753 0.6893 11.250 1.5540 0.01889 0.01215 -0.0787 0.0704 0.6904 11.500 1.5646 0.01960 0.01287 -0.0763 0.0648 0.6914 11.750 1.5756 0.02031 0.01359 -0.0741 0.0607 0.6923 12.000 1.5853 0.02110 0.01441 -0.0718 0.0564 0.6933 12.250 1.5937 0.02200 0.01532 -0.0696 0.0527 0.6943 12.500 1.6021 0.02295 0.01630 -0.0674 0.0492 0.6952 12.750 1.6070 0.02415 0.01751 -0.0651 0.0456 0.6961 13.000 1.6137 0.02533 0.01874 -0.0632 0.0424 0.6969 13.250 1.6171 0.02682 0.02025 -0.0613 0.0394 0.6977 13.500 1.6217 0.02834 0.02182 -0.0597 0.0369 0.6988 13.750 1.6249 0.03007 0.02361 -0.0583 0.0345 0.7000 14.000 1.6263 0.03210 0.02570 -0.0571 0.0323 0.7011 14.250 1.6290 0.03417 0.02783 -0.0562 0.0303 0.7023 14.500 1.6289 0.03661 0.03033 -0.0555 0.0282 0.7034 14.750 1.6288 0.03920 0.03298 -0.0550 0.0263 0.7046 15.000 1.6283 0.04196 0.03582 -0.0548 0.0246 0.7057 15.250 1.6246 0.04516 0.03907 -0.0547 0.0227 0.7069 15.500 1.6236 0.04818 0.04218 -0.0548 0.0214 0.7080 15.750 1.6198 0.05161 0.04567 -0.0551 0.0198 0.7091 16.000 1.6151 0.05524 0.04938 -0.0556 0.0187 0.7102 16.250 1.6120 0.05879 0.05300 -0.0561 0.0176 0.7112 16.500 1.6070 0.06266 0.05694 -0.0569 0.0165 0.7123 17.000 1.5965 0.07077 0.06521 -0.0588 0.0146 0.7142 17.250 1.5911 0.07501 0.06954 -0.0600 0.0137 0.7156 17.500 1.5847 0.07947 0.07409 -0.0614 0.0130 0.7169 17.750 1.5789 0.08395 0.07866 -0.0628 0.0123 0.7181 18.000 1.5743 0.08832 0.08312 -0.0643 0.0118 0.7195 18.250 1.5685 0.09299 0.08788 -0.0660 0.0112 0.7207 18.500 1.5616 0.09788 0.09286 -0.0679 0.0107 0.7220 18.750 1.5544 0.10290 0.09797 -0.0700 0.0103 0.7232