XFOIL Version 6.96 Calculated polar for: FX 61-184 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.200 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -15.000 -0.5386 0.09581 0.09185 -0.0687 1.0000 0.0147 -14.750 -0.5758 0.08447 0.08030 -0.0756 1.0000 0.0146 -14.500 -0.6040 0.07617 0.07178 -0.0802 1.0000 0.0145 -14.250 -0.6252 0.06982 0.06525 -0.0833 1.0000 0.0145 -14.000 -0.6453 0.06426 0.05947 -0.0856 1.0000 0.0147 -13.750 -0.6630 0.05954 0.05453 -0.0869 1.0000 0.0148 -13.500 -0.6729 0.05468 0.04941 -0.0888 0.9812 0.0151 -13.250 -0.6651 0.04993 0.04429 -0.0946 0.9378 0.0156 -13.000 -0.6364 0.04701 0.04123 -0.1011 0.9209 0.0160 -12.750 -0.6074 0.04418 0.03818 -0.1077 0.9043 0.0167 -12.500 -0.5819 0.04142 0.03512 -0.1129 0.8873 0.0175 -12.250 -0.5597 0.03895 0.03229 -0.1162 0.8721 0.0182 -12.000 -0.5384 0.03695 0.02995 -0.1182 0.8601 0.0188 -11.500 -0.4989 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8.000 1.2863 0.01845 0.01159 -0.1010 0.4298 0.6877 8.250 1.2980 0.01890 0.01207 -0.0986 0.4159 0.6897 8.500 1.3074 0.01948 0.01266 -0.0959 0.3995 0.6919 8.750 1.3143 0.02019 0.01335 -0.0931 0.3811 0.6944 9.000 1.3189 0.02108 0.01419 -0.0901 0.3620 0.6971 9.250 1.3223 0.02213 0.01519 -0.0873 0.3419 0.6999 9.500 1.3235 0.02336 0.01639 -0.0845 0.3207 0.7020 9.750 1.3222 0.02486 0.01782 -0.0816 0.2993 0.7042 10.000 1.3224 0.02644 0.01937 -0.0793 0.2779 0.7066 10.250 1.3225 0.02815 0.02103 -0.0771 0.2578 0.7090 10.500 1.3231 0.02996 0.02279 -0.0752 0.2398 0.7116 10.750 1.3243 0.03182 0.02461 -0.0735 0.2234 0.7142 11.000 1.3262 0.03371 0.02647 -0.0721 0.2074 0.7167 11.250 1.3281 0.03562 0.02839 -0.0706 0.1923 0.7190 11.500 1.3297 0.03765 0.03041 -0.0693 0.1784 0.7217 11.750 1.3326 0.03966 0.03243 -0.0682 0.1650 0.7247 12.000 1.3358 0.04172 0.03451 -0.0673 0.1518 0.7279 12.250 1.3386 0.04391 0.03671 -0.0665 0.1387 0.7311 12.500 1.3407 0.04621 0.03903 -0.0657 0.1262 0.7338 12.750 1.3425 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