XFOIL Version 6.96 Calculated polar for: FX 61-147 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.200 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -10.000 -0.4443 0.07016 0.06676 -0.0588 1.0000 0.0089 -9.750 -0.4692 0.06435 0.06090 -0.0611 1.0000 0.0088 -9.500 -0.4858 0.06152 0.05807 -0.0609 1.0000 0.0087 -9.250 -0.4988 0.05558 0.05195 -0.0664 0.9954 0.0087 -9.000 -0.5103 0.04925 0.04531 -0.0723 0.9844 0.0086 -8.750 -0.5090 0.04283 0.03844 -0.0776 0.9729 0.0087 -8.500 -0.5032 0.03643 0.03136 -0.0806 0.9631 0.0088 -8.250 -0.4893 0.03257 0.02704 -0.0815 0.9537 0.0090 -8.000 -0.4617 0.03019 0.02438 -0.0839 0.9491 0.0094 -7.750 -0.4398 0.02838 0.02235 -0.0843 0.9403 0.0097 -7.500 -0.4092 0.02647 0.02017 -0.0861 0.9358 0.0105 -7.250 -0.3845 0.02479 0.01819 -0.0862 0.9272 0.0115 -7.000 -0.3525 0.02322 0.01640 -0.0878 0.9220 0.0130 -6.750 -0.3239 0.02217 0.01522 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0.8612 13.500 1.3668 0.06245 0.05568 -0.0662 0.0127 0.8677 14.000 1.3544 0.07025 0.06375 -0.0669 0.0097 0.8868 14.250 1.3446 0.07362 0.06732 -0.0661 0.0089 0.9376 14.500 1.3400 0.07787 0.07171 -0.0673 0.0081 1.0000 14.750 1.3356 0.08238 0.07634 -0.0687 0.0075 1.0000 15.000 1.3303 0.08711 0.08118 -0.0703 0.0070 1.0000