XFOIL Version 6.96 Calculated polar for: FX LIII-142 K 25 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 ------ -------- --------- --------- -------- -------- -------- -19.750 -1.0891 0.14144 0.13864 0.0232 1.0000 0.0044 -19.500 -1.1176 0.13171 0.12879 0.0181 1.0000 0.0046 -19.250 -1.1496 0.12164 0.11857 0.0129 1.0000 0.0045 -19.000 -1.1732 0.11336 0.11017 0.0088 1.0000 0.0044 -18.750 -1.2047 0.10388 0.10051 0.0041 1.0000 0.0043 -18.500 -1.2259 0.09644 0.09295 0.0005 1.0000 0.0044 -18.250 -1.2460 0.08937 0.08574 -0.0029 1.0000 0.0044 -18.000 -1.2647 0.08271 0.07895 -0.0060 1.0000 0.0044 -17.750 -1.2816 0.07648 0.07259 -0.0088 1.0000 0.0044 -17.500 -1.2919 0.07145 0.06746 -0.0110 1.0000 0.0044 -17.250 -1.3037 0.06630 0.06219 -0.0132 1.0000 0.0045 -17.000 -1.3105 0.06204 0.05782 -0.0149 1.0000 0.0044 -16.750 -1.3194 0.05766 0.05334 -0.0165 1.0000 0.0047 -16.500 -1.3270 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