XFOIL Version 6.94 Calculated polar for: NACA 4412 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.090 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0327 0.02723 0.01794 -0.0872 0.9321 0.1691 -2.750 -0.0032 0.02655 0.01732 -0.0878 0.9205 0.1852 -2.500 0.0449 0.02569 0.01646 -0.0916 0.9149 0.2105 -2.250 0.0737 0.02521 0.01609 -0.0920 0.9030 0.2341 -2.000 0.1205 0.02445 0.01547 -0.0954 0.8974 0.2731 -1.750 0.1489 0.02401 0.01518 -0.0956 0.8855 0.3116 -1.500 0.1950 0.02307 0.01454 -0.0987 0.8799 0.3791 -1.000 0.2700 0.02051 0.01405 -0.0996 0.8642 1.0000 -0.750 0.3209 0.02018 0.01333 -0.1034 0.8594 1.0000 -0.500 0.3453 0.02032 0.01326 -0.1028 0.8456 1.0000 0.000 0.4149 0.02010 0.01268 -0.1046 0.8267 1.0000 0.250 0.4467 0.02006 0.01250 -0.1050 0.8168 1.0000 0.500 0.4812 0.01989 0.01219 -0.1057 0.8077 1.0000 0.750 0.5092 0.01997 0.01217 -0.1054 0.7967 1.0000 1.000 0.5444 0.01975 0.01184 -0.1061 0.7884 1.0000 1.250 0.5697 0.01995 0.01196 -0.1055 0.7767 1.0000 1.500 0.6053 0.01971 0.01162 -0.1062 0.7691 1.0000 1.750 0.6284 0.02003 0.01189 -0.1053 0.7569 1.0000 2.000 0.6642 0.01980 0.01155 -0.1060 0.7497 1.0000 2.250 0.6855 0.02023 0.01197 -0.1049 0.7372 1.0000 2.500 0.7213 0.02002 0.01166 -0.1056 0.7304 1.0000 2.750 0.7413 0.02055 0.01220 -0.1043 0.7178 1.0000 3.000 0.7761 0.02041 0.01197 -0.1049 0.7110 1.0000 3.250 0.7958 0.02098 0.01256 -0.1036 0.6986 1.0000 3.500 0.8268 0.02105 0.01258 -0.1037 0.6908 1.0000 3.750 0.8493 0.02150 0.01307 -0.1028 0.6796 1.0000 4.000 0.8782 0.02169 0.01323 -0.1027 0.6712 1.0000 4.250 0.9020 0.02210 0.01367 -0.1020 0.6607 1.0000 4.500 0.9292 0.02239 0.01397 -0.1016 0.6517 1.0000 4.750 0.9542 0.02275 0.01435 -0.1010 0.6415 1.0000 5.000 0.9796 0.02313 0.01476 -0.1004 0.6320 1.0000