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 = 0.500 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -19.750 -0.9849 0.09412 0.08922 -0.0680 1.0000 0.0240 -19.500 -0.9896 0.08978 0.08478 -0.0702 1.0000 0.0241 -19.250 -0.9932 0.08567 0.08057 -0.0722 1.0000 0.0243 -19.000 -0.9955 0.08176 0.07657 -0.0741 1.0000 0.0245 -18.750 -0.9987 0.07779 0.07253 -0.0760 1.0000 0.0246 -18.500 -1.0015 0.07395 0.06863 -0.0777 1.0000 0.0248 -18.250 -1.0029 0.07033 0.06494 -0.0794 1.0000 0.0250 -18.000 -1.0030 0.06696 0.06151 -0.0809 0.9951 0.0252 -17.750 -1.0001 0.06375 0.05821 -0.0829 0.9659 0.0255 -17.250 -0.9953 0.05828 0.05254 -0.0844 0.9328 0.0259 -17.000 -0.9923 0.05581 0.04997 -0.0847 0.9213 0.0262 -16.750 -0.9886 0.05346 0.04751 -0.0850 0.9111 0.0265 -16.500 -0.9829 0.05128 0.04523 -0.0853 0.9033 0.0268 -16.250 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