XFOIL Version 6.96 Calculated polar for: NASA/LANGLEY LS(1)-0421MOD 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 ------ -------- --------- --------- -------- -------- -------- -19.750 -1.1342 0.08234 0.07799 -0.0506 1.0000 0.0236 -19.500 -1.1377 0.07837 0.07395 -0.0523 1.0000 0.0238 -19.250 -1.1404 0.07455 0.07005 -0.0539 1.0000 0.0239 -19.000 -1.1421 0.07088 0.06631 -0.0553 1.0000 0.0240 -18.750 -1.1425 0.06747 0.06282 -0.0566 1.0000 0.0242 -18.500 -1.1416 0.06429 0.05958 -0.0577 1.0000 0.0244 -18.250 -1.1399 0.06125 0.05647 -0.0588 1.0000 0.0245 -18.000 -1.1373 0.05835 0.05351 -0.0597 1.0000 0.0247 -17.750 -1.1332 0.05570 0.05078 -0.0605 1.0000 0.0248 -17.500 -1.1277 0.05325 0.04828 -0.0611 1.0000 0.0250 -17.250 -1.1216 0.05091 0.04587 -0.0617 1.0000 0.0251 -17.000 -1.1136 0.04881 0.04372 -0.0622 1.0000 0.0253 -16.750 -1.1136 0.04600 0.04084 -0.0626 1.0000 0.0255 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1.3533 0.01451 0.00810 -0.1003 0.1960 0.7242 7.500 1.3760 0.01489 0.00844 -0.0996 0.1874 0.7263 7.750 1.3968 0.01533 0.00884 -0.0985 0.1790 0.7283 8.000 1.4168 0.01577 0.00924 -0.0974 0.1718 0.7300 8.250 1.4360 0.01621 0.00966 -0.0961 0.1656 0.7321 8.500 1.4531 0.01661 0.01008 -0.0945 0.1606 0.7354 8.750 1.4645 0.01717 0.01063 -0.0919 0.1555 0.7382 9.000 1.4820 0.01766 0.01114 -0.0904 0.1517 0.7408 9.250 1.4987 0.01823 0.01173 -0.0889 0.1477 0.7433 9.500 1.5131 0.01896 0.01245 -0.0873 0.1433 0.7457 9.750 1.5300 0.01959 0.01309 -0.0860 0.1401 0.7480 10.000 1.5464 0.02028 0.01380 -0.0848 0.1368 0.7501 10.250 1.5603 0.02114 0.01465 -0.0834 0.1332 0.7518 10.500 1.5740 0.02205 0.01558 -0.0821 0.1298 0.7545 10.750 1.5906 0.02281 0.01639 -0.0812 0.1275 0.7575 11.000 1.6047 0.02376 0.01737 -0.0801 0.1249 0.7601 11.250 1.6166 0.02489 0.01852 -0.0789 0.1220 0.7627 11.500 1.6276 0.02613 0.01979 -0.0777 0.1193 0.7652 11.750 1.6423 0.02714 0.02084 -0.0769 0.1175 0.7677 12.000 1.6549 0.02833 0.02207 -0.0760 0.1154 0.7701 12.250 1.6656 0.02969 0.02344 -0.0750 0.1133 0.7723 12.500 1.6735 0.03131 0.02509 -0.0740 0.1109 0.7747 12.750 1.6822 0.03290 0.02673 -0.0731 0.1090 0.7780 13.000 1.6941 0.03423 0.02813 -0.0724 0.1074 0.7809 13.250 1.7029 0.03584 0.02980 -0.0716 0.1056 0.7836 13.500 1.7098 0.03766 0.03165 -0.0707 0.1038 0.7863 13.750 1.7144 0.03972 0.03374 -0.0699 0.1019 0.7889 14.000 1.7165 0.04209 0.03615 -0.0691 0.1000 0.7913 14.250 1.7253 0.04388 0.03800 -0.0686 0.0990 0.7934 14.500 1.7330 0.04577 0.03997 -0.0682 0.0976 0.7965 14.750 1.7375 0.04803 0.04229 -0.0677 0.0962 0.7995 15.000 1.7408 0.05045 0.04476 -0.0673 0.0947 0.8024 15.250 1.7414 0.05319 0.04754 -0.0669 0.0932 0.8053 15.500 1.7368 0.05655 0.05096 -0.0665 0.0916 0.8080 15.750 1.7422 0.05886 0.05335 -0.0664 0.0907 0.8108 16.000 1.7467 0.06131 0.05586 -0.0663 0.0897 0.8130 16.250 1.7487 0.06407 0.05870 -0.0662 0.0885 0.8163 16.500 1.7496 0.06690 0.06159 -0.0661 0.0873 0.8193 16.750 1.7471 0.07021 0.06497 -0.0662 0.0862 0.8222 17.000 1.7424 0.07381 0.06862 -0.0663 0.0850 0.8251 17.250 1.7342 0.07787 0.07274 -0.0665 0.0836 0.8278 17.500 1.7345 0.08096 0.07591 -0.0668 0.0829 0.8304 17.750 1.7360 0.08386 0.07888 -0.0671 0.0822 0.8330 18.000 1.7341 0.08722 0.08234 -0.0675 0.0814 0.8359 18.250 1.7332 0.09045 0.08565 -0.0679 0.0805 0.8387 18.500 1.7294 0.09412 0.08939 -0.0685 0.0796 0.8415 18.750 1.7251 0.09783 0.09316 -0.0692 0.0788 0.8444 19.000 1.7192 0.10178 0.09716 -0.0700 0.0779 0.8471 19.250 1.7101 0.10622 0.10165 -0.0710 0.0769 0.8494