XFOIL Version 6.96 Calculated polar for: FX 60-160 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 ------ -------- --------- --------- -------- -------- -------- -16.250 -0.7825 0.07449 0.07153 -0.0826 1.0000 0.0153 -16.000 -0.8198 0.06522 0.06204 -0.0878 1.0000 0.0153 -15.750 -0.8372 0.05988 0.05657 -0.0902 1.0000 0.0154 -15.500 -0.8532 0.05530 0.05186 -0.0918 1.0000 0.0156 -15.250 -0.8671 0.05150 0.04792 -0.0926 1.0000 0.0157 -15.000 -0.8796 0.04821 0.04452 -0.0927 1.0000 0.0159 -14.750 -0.8894 0.04551 0.04171 -0.0921 1.0000 0.0161 -14.500 -0.8949 0.04284 0.03891 -0.0922 0.9996 0.0163 -14.250 -0.8857 0.04009 0.03598 -0.0948 0.9978 0.0167 -14.000 -0.8758 0.03752 0.03322 -0.0971 0.9953 0.0171 -13.750 -0.8635 0.03547 0.03103 -0.0988 0.9925 0.0175 -13.500 -0.8493 0.03374 0.02925 -0.1006 0.9893 0.0179 -13.250 -0.8294 0.03247 0.02793 -0.1026 0.9865 0.0183 -13.000 -0.8092 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0.06856 0.06269 -0.0340 0.0084 1.0000 16.750 1.3803 0.07163 0.06584 -0.0342 0.0080 1.0000 17.000 1.3808 0.07474 0.06904 -0.0345 0.0077 1.0000 17.250 1.3804 0.07804 0.07242 -0.0349 0.0074 1.0000 17.500 1.3794 0.08147 0.07595 -0.0354 0.0071 1.0000 17.750 1.3795 0.08483 0.07940 -0.0361 0.0071 1.0000 18.000 1.3803 0.08811 0.08279 -0.0368 0.0069 1.0000 18.250 1.3796 0.09168 0.08647 -0.0377 0.0068 1.0000 18.500 1.3783 0.09538 0.09027 -0.0387 0.0066 1.0000 18.750 1.3768 0.09917 0.09417 -0.0398 0.0065 1.0000 19.000 1.3742 0.10317 0.09829 -0.0411 0.0063 1.0000 19.250 1.3713 0.10728 0.10251 -0.0426 0.0063 1.0000