XFOIL Version 6.94 Calculated polar for: mta1 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.018 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.2957 0.05200 0.04110 -0.0195 1.0000 0.3183 -2.750 -0.2529 0.04916 0.03765 -0.0247 1.0000 0.3020 -2.500 -0.2186 0.04641 0.03468 -0.0264 1.0000 0.2937 -2.250 -0.1782 0.04425 0.03191 -0.0293 1.0000 0.2833 -2.000 -0.1422 0.04222 0.02952 -0.0305 1.0000 0.2775 -1.750 -0.1066 0.04043 0.02740 -0.0313 1.0000 0.2740 -1.500 -0.0717 0.03887 0.02559 -0.0316 1.0000 0.2747 -1.250 -0.0377 0.03752 0.02406 -0.0316 1.0000 0.2791 -1.000 -0.0051 0.03615 0.02281 -0.0312 1.0000 0.2868 -0.750 0.0271 0.03492 0.02184 -0.0305 1.0000 0.2986 -0.500 0.0598 0.03362 0.02117 -0.0297 1.0000 0.3180 -0.250 0.0953 0.03207 0.02081 -0.0292 1.0000 0.3551 0.000 0.1238 0.02851 0.02054 -0.0268 1.0000 0.8678 0.250 0.1413 0.02667 0.01924 -0.0233 0.7696 1.0000 0.500 0.3155 0.03281 0.01934 -0.0425 0.4419 1.0000 0.750 0.3568 0.03454 0.02014 -0.0438 0.4188 1.0000 1.000 0.3981 0.03605 0.02121 -0.0455 0.4032 1.0000 1.250 0.4403 0.03773 0.02256 -0.0475 0.3951 1.0000 1.500 0.4807 0.03935 0.02414 -0.0495 0.3903 1.0000 1.750 0.5194 0.04112 0.02594 -0.0513 0.3880 1.0000 2.000 0.5553 0.04302 0.02794 -0.0529 0.3868 1.0000 2.250 0.5880 0.04504 0.03009 -0.0541 0.3855 1.0000 2.500 0.6181 0.04720 0.03242 -0.0550 0.3846 1.0000 2.750 0.6456 0.04948 0.03492 -0.0559 0.3855 1.0000 3.000 0.6711 0.05196 0.03766 -0.0569 0.3896 1.0000 3.250 0.6978 0.05485 0.04072 -0.0582 0.3960 1.0000 3.500 0.7184 0.05764 0.04387 -0.0600 0.4057 1.0000 3.750 0.7395 0.06100 0.04741 -0.0617 0.4163 1.0000 4.000 0.7513 0.06438 0.05110 -0.0641 0.4307 1.0000 4.250 0.7615 0.06815 0.05507 -0.0668 0.4466 1.0000 4.500 0.7672 0.07224 0.05930 -0.0697 0.4646 1.0000 4.750 0.7617 0.07661 0.06377 -0.0729 0.4863 1.0000 5.000 0.7484 0.08122 0.06841 -0.0763 0.5120 1.0000