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 = 0.050 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -14.500 -0.5024 0.11241 0.10406 -0.0736 1.0000 0.0821 -14.250 -0.5791 0.09830 0.08977 -0.0789 1.0000 0.0817 -14.000 -0.6367 0.09029 0.08161 -0.0800 1.0000 0.0812 -13.750 -0.6840 0.08478 0.07592 -0.0792 1.0000 0.0810 -13.500 -0.7240 0.08052 0.07151 -0.0774 1.0000 0.0809 -13.250 -0.7592 0.07699 0.06781 -0.0750 1.0000 0.0810 -13.000 -0.7805 0.07206 0.06251 -0.0762 0.9951 0.0818 -12.750 -0.7842 0.06857 0.05879 -0.0774 0.9904 0.0831 -12.500 -0.7723 0.06658 0.05675 -0.0781 0.9856 0.0849 -12.250 -0.7662 0.06403 0.05402 -0.0790 0.9808 0.0868 -12.000 -0.7619 0.06137 0.05108 -0.0795 0.9762 0.0888 -11.750 -0.7611 0.05889 0.04824 -0.0790 0.9704 0.0906 -11.500 -0.7378 0.05767 0.04711 -0.0797 0.9669 0.0932 -11.250 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3.500 0.2743 0.05308 0.04236 -0.0107 0.6704 0.8577 3.750 0.3092 0.05292 0.04222 -0.0123 0.6637 0.8612 4.000 0.3236 0.05323 0.04258 -0.0111 0.6523 0.8644 4.250 0.3546 0.05295 0.04233 -0.0118 0.6447 0.8672 4.500 0.3712 0.05332 0.04275 -0.0113 0.6327 0.8704 4.750 0.4054 0.05290 0.04237 -0.0122 0.6251 0.8738 5.000 0.4223 0.05341 0.04293 -0.0120 0.6117 0.8774 5.250 0.4581 0.05258 0.04216 -0.0126 0.6049 0.8805 5.500 0.4712 0.05305 0.04270 -0.0116 0.5901 0.8837 6.000 0.5278 0.05224 0.04201 -0.0120 0.5683 0.8898 6.250 0.5473 0.05249 0.04233 -0.0117 0.5533 0.8934 6.500 0.5869 0.05081 0.04071 -0.0119 0.5468 0.8969 6.750 0.6020 0.05106 0.04104 -0.0108 0.5305 0.9008 7.250 0.6523 0.05032 0.04045 -0.0104 0.5030 0.9081 7.750 0.7054 0.04913 0.03937 -0.0095 0.4738 0.9152 8.000 0.7306 0.04874 0.03901 -0.0091 0.4571 0.9195 8.250 0.7594 0.04817 0.03843 -0.0089 0.4400 0.9242 8.500 0.7860 0.04753 0.03777 -0.0083 0.4218 0.9290 8.750 0.8117 0.04719 0.03736 -0.0078 0.4024 0.9341 9.250 0.8552 0.04748 0.03748 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1.0000 14.250 1.1533 0.07946 0.06988 -0.0127 0.1662 1.0000 14.500 1.1583 0.08242 0.07295 -0.0134 0.1621 1.0000 14.750 1.1887 0.08261 0.07300 -0.0139 0.1582 1.0000 15.000 1.1628 0.08921 0.08000 -0.0151 0.1554 1.0000 15.250 1.1340 0.09667 0.08782 -0.0171 0.1525 1.0000 15.500 1.1108 0.10385 0.09524 -0.0196 0.1495 1.0000 15.750 1.1311 0.10482 0.09618 -0.0201 0.1464 1.0000 16.000 1.1001 0.11368 0.10528 -0.0239 0.1439 1.0000 16.500 0.9394 0.15630 0.14849 -0.0473 0.1333 1.0000