XFOIL Version 6.96 Calculated polar for: FX 66-S-161 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 1.000 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -14.000 -0.6261 0.07463 0.07267 -0.0586 1.0000 0.0075 -13.750 -0.6646 0.06352 0.06130 -0.0660 1.0000 0.0073 -13.500 -0.6928 0.05572 0.05335 -0.0712 1.0000 0.0074 -13.250 -0.6968 0.05195 0.04954 -0.0737 1.0000 0.0076 -13.000 -0.7239 0.04529 0.04266 -0.0776 1.0000 0.0075 -12.750 -0.7335 0.04133 0.03859 -0.0797 1.0000 0.0075 -12.500 -0.7352 0.03612 0.03314 -0.0853 0.9603 0.0075 -12.250 -0.6276 0.03136 0.02801 -0.1112 0.8981 0.0080 -12.000 -0.6383 0.02828 0.02434 -0.1113 0.8122 0.0079 -11.750 -0.6403 0.02677 0.02259 -0.1100 0.7755 0.0080 -11.500 -0.6372 0.02590 0.02157 -0.1085 0.7492 0.0082 -11.000 -0.6334 0.02368 0.01899 -0.1038 0.7148 0.0083 -10.750 -0.6238 0.02246 0.01759 -0.1022 0.7016 0.0084 -10.500 -0.6103 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-0.0589 0.0040 1.0000 18.500 1.3337 0.11776 0.11345 -0.0604 0.0039 1.0000