XFOIL Version 6.96 Calculated polar for: FX 63-145 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.100 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -10.000 -0.2822 0.10951 0.10488 -0.0399 1.0000 0.0485 -9.750 -0.2818 0.10594 0.10138 -0.0410 1.0000 0.0482 -9.500 -0.2989 0.09670 0.09222 -0.0468 1.0000 0.0323 -9.000 -0.2959 0.08959 0.08522 -0.0466 1.0000 0.0261 -8.750 -0.2968 0.08662 0.08234 -0.0466 1.0000 0.0255 -8.500 -0.3020 0.08305 0.07888 -0.0471 1.0000 0.0248 -8.250 -0.3094 0.07844 0.07437 -0.0490 0.9990 0.0239 -8.000 -0.3315 0.06370 0.05944 -0.0663 0.9657 0.0213 -7.750 -0.3226 0.05802 0.05364 -0.0744 0.9372 0.0210 -7.500 -0.3076 0.05237 0.04772 -0.0825 0.9129 0.0205 -7.250 -0.2841 0.04676 0.04169 -0.0902 0.8921 0.0199 -7.000 -0.2543 0.04152 0.03590 -0.0968 0.8717 0.0191 -6.750 -0.2219 0.03697 0.03071 -0.1018 0.8491 0.0184 -6.500 -0.1859 0.03303 0.02600 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1.0000 13.750 1.4060 0.07545 0.06870 -0.0793 0.1116 1.0000 14.000 1.4016 0.07970 0.07307 -0.0805 0.0998 1.0000 14.250 1.3884 0.08531 0.07860 -0.0823 0.0952 1.0000 14.500 1.3757 0.09100 0.08422 -0.0841 0.0916 1.0000