XFOIL Version 6.94 Calculated polar for: mta3 300.2.14.15.2.c13 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.012 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.2530 0.05140 0.03850 -0.0172 1.0000 0.3979 -2.750 -0.2147 0.04870 0.03536 -0.0216 1.0000 0.3939 -2.500 -0.1786 0.04623 0.03255 -0.0243 1.0000 0.3940 -2.250 -0.1446 0.04391 0.03009 -0.0259 1.0000 0.3979 -2.000 -0.1107 0.04183 0.02787 -0.0272 1.0000 0.4057 -1.750 -0.0762 0.03997 0.02590 -0.0283 1.0000 0.4186 -1.500 -0.0440 0.03817 0.02420 -0.0285 1.0000 0.4389 -1.250 -0.0145 0.03630 0.02278 -0.0280 1.0000 0.4701 -1.000 0.0123 0.03425 0.02164 -0.0266 1.0000 0.5259 -0.750 0.0307 0.03090 0.02058 -0.0222 1.0000 1.0000 -0.500 0.0805 0.03126 0.01945 -0.0250 1.0000 1.0000 -0.250 0.1144 0.03148 0.01882 -0.0239 1.0000 1.0000 0.000 0.1448 0.03163 0.01853 -0.0223 1.0000 1.0000 0.250 0.1759 0.03172 0.01854 -0.0210 1.0000 1.0000 0.500 0.2099 0.03178 0.01891 -0.0207 1.0000 1.0000 0.750 0.2434 0.03216 0.01999 -0.0223 1.0000 1.0000 1.000 0.4127 0.03543 0.02154 -0.0457 0.6335 1.0000 1.250 0.4463 0.03747 0.02262 -0.0454 0.5912 1.0000 1.500 0.4783 0.03956 0.02416 -0.0459 0.5673 1.0000 1.750 0.5090 0.04174 0.02599 -0.0467 0.5495 1.0000 2.000 0.5385 0.04403 0.02815 -0.0479 0.5356 1.0000 2.250 0.5669 0.04652 0.03057 -0.0494 0.5251 1.0000 2.500 0.5945 0.04918 0.03316 -0.0508 0.5171 1.0000 2.750 0.6205 0.05210 0.03607 -0.0526 0.5132 1.0000 3.000 0.6436 0.05533 0.03942 -0.0552 0.5133 1.0000 3.250 0.6623 0.05889 0.04313 -0.0581 0.5160 1.0000 3.500 0.6729 0.06287 0.04727 -0.0616 0.5216 1.0000 3.750 0.6833 0.06694 0.05136 -0.0645 0.5281 1.0000 4.000 0.6892 0.07117 0.05561 -0.0675 0.5357 1.0000 4.250 0.6890 0.07555 0.05994 -0.0704 0.5455 1.0000 4.500 0.6862 0.07985 0.06418 -0.0731 0.5562 1.0000 4.750 0.6894 0.08415 0.06836 -0.0755 0.5672 1.0000 5.000 0.6747 0.08809 0.07219 -0.0772 0.5821 1.0000