XFOIL Version 6.96 Calculated polar for: WORTMANN FX 62-K-131 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 ------ -------- --------- --------- -------- -------- -------- -14.250 -0.1797 0.13856 0.13640 -0.0652 0.9894 0.0098 -14.000 -0.1737 0.13575 0.13357 -0.0677 0.9881 0.0101 -13.750 -0.1676 0.13274 0.13055 -0.0704 0.9869 0.0102 -13.500 -0.1619 0.12955 0.12736 -0.0732 0.9851 0.0103 -13.250 -0.1567 0.12639 0.12420 -0.0752 0.9797 0.0104 -13.000 -0.1491 0.12287 0.12067 -0.0777 0.9761 0.0105 -12.750 -0.1414 0.11919 0.11698 -0.0802 0.9725 0.0105 -12.500 -0.1344 0.11545 0.11324 -0.0825 0.9676 0.0105 -12.250 -0.1264 0.11141 0.10919 -0.0852 0.9640 0.0105 -12.000 -0.1183 0.10730 0.10507 -0.0880 0.9607 0.0106 -11.750 -0.1112 0.10311 0.10087 -0.0906 0.9558 0.0106 -11.500 -0.1104 0.09739 0.09516 -0.0925 0.9517 0.0111 -11.250 -0.1006 0.09372 0.09147 -0.0946 0.9469 0.0114 -11.000 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8.750 1.3727 0.02697 0.01971 -0.1249 0.0059 0.8522 9.000 1.3786 0.02863 0.02148 -0.1226 0.0051 0.8548 9.250 1.3861 0.03023 0.02319 -0.1205 0.0049 0.8572 9.500 1.3939 0.03167 0.02475 -0.1184 0.0047 0.8591 9.750 1.3992 0.03336 0.02657 -0.1162 0.0046 0.8611 10.000 1.4044 0.03513 0.02846 -0.1140 0.0044 0.8634 10.250 1.4093 0.03700 0.03045 -0.1119 0.0042 0.8657 10.500 1.4138 0.03895 0.03251 -0.1099 0.0041 0.8681 10.750 1.4178 0.04103 0.03471 -0.1080 0.0039 0.8701 11.000 1.4207 0.04316 0.03696 -0.1060 0.0038 0.8719 11.250 1.4228 0.04535 0.03931 -0.1039 0.0037 0.8739 11.500 1.4247 0.04762 0.04172 -0.1019 0.0036 0.8759 11.750 1.4269 0.04998 0.04419 -0.1001 0.0034 0.8780 12.000 1.4281 0.05252 0.04687 -0.0982 0.0034 0.8800 12.250 1.4298 0.05511 0.04962 -0.0964 0.0034 0.8819 12.500 1.4303 0.05799 0.05267 -0.0946 0.0033 0.8838 12.750 1.4301 0.06095 0.05584 -0.0927 0.0033 0.8856 13.000 1.4277 0.06428 0.05942 -0.0907 0.0034 0.8872 13.250 1.4194 0.06879 0.06430 -0.0881 0.0037 0.8886 13.500 1.4036 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