XFOIL Version 6.96 Calculated polar for: FX 66-182 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 = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -13.750 -0.1528 0.11681 0.11338 -0.0697 0.6780 0.0235 -13.500 -0.1493 0.11351 0.11006 -0.0713 0.6722 0.0242 -13.250 -0.4479 0.05806 0.05469 -0.1114 0.8467 0.0139 -12.500 -0.5068 0.04631 0.04193 -0.1140 0.7387 0.0139 -12.250 -0.5200 0.04331 0.03868 -0.1139 0.7217 0.0139 -12.000 -0.5310 0.04053 0.03567 -0.1135 0.7078 0.0139 -11.750 -0.5431 0.03766 0.03255 -0.1126 0.6963 0.0139 -11.500 -0.5498 0.03502 0.02970 -0.1115 0.6866 0.0141 -11.000 -0.5502 0.03131 0.02566 -0.1085 0.6701 0.0144 -10.750 -0.5461 0.03017 0.02439 -0.1065 0.6631 0.0147 -10.500 -0.5414 0.02882 0.02289 -0.1043 0.6569 0.0148 -10.250 -0.5297 0.02783 0.02179 -0.1030 0.6507 0.0151 -10.000 -0.5176 0.02658 0.02034 -0.1016 0.6455 0.0154 -9.750 -0.5029 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