XFOIL Version 6.96 Calculated polar for: FX 66-17AII-182 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 ------ -------- --------- --------- -------- -------- -------- -16.250 -0.7174 0.08076 0.07844 -0.0652 1.0000 0.0076 -16.000 -0.7465 0.07184 0.06932 -0.0703 1.0000 0.0074 -15.750 -0.7539 0.06698 0.06437 -0.0731 1.0000 0.0073 -15.500 -0.7832 0.05924 0.05643 -0.0772 1.0000 0.0074 -15.250 -0.7960 0.05444 0.05150 -0.0795 1.0000 0.0074 -15.000 -0.8116 0.04976 0.04670 -0.0813 1.0000 0.0076 -14.750 -0.8204 0.04623 0.04305 -0.0824 1.0000 0.0077 -14.500 -0.8124 0.04256 0.03926 -0.0862 0.9921 0.0078 -14.250 -0.7303 0.03668 0.03305 -0.1081 0.9220 0.0080 -14.000 -0.6861 0.03392 0.02963 -0.1179 0.8050 0.0082 -13.750 -0.6838 0.03258 0.02806 -0.1169 0.7633 0.0083 -13.500 -0.6771 0.03147 0.02678 -0.1162 0.7343 0.0085 -13.250 -0.6742 0.02995 0.02509 -0.1154 0.7127 0.0086 -13.000 -0.6731 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1.3022 0.08094 0.07571 -0.0486 0.0090 1.0000 16.000 1.2928 0.08509 0.07994 -0.0486 0.0084 1.0000