XFOIL Version 6.94 Calculated polar for: mta2 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 -0.3167 0.05559 0.04307 -0.0122 1.0000 0.4018 -2.750 -0.2751 0.05256 0.03957 -0.0177 1.0000 0.3944 -2.500 -0.2364 0.04981 0.03650 -0.0212 1.0000 0.3919 -2.250 -0.1969 0.04740 0.03367 -0.0244 1.0000 0.3940 -2.000 -0.1597 0.04518 0.03125 -0.0264 1.0000 0.4006 -1.750 -0.1226 0.04321 0.02911 -0.0279 1.0000 0.4118 -1.500 -0.0872 0.04132 0.02730 -0.0287 1.0000 0.4287 -1.250 -0.0540 0.03944 0.02581 -0.0287 1.0000 0.4569 -1.000 -0.0229 0.03741 0.02460 -0.0279 1.0000 0.5063 -0.750 0.0006 0.03463 0.02373 -0.0247 1.0000 0.6276 -0.500 0.0452 0.03357 0.02244 -0.0256 1.0000 1.0000 -0.250 0.0874 0.03370 0.02184 -0.0254 1.0000 1.0000 0.000 0.1280 0.03365 0.02155 -0.0251 1.0000 1.0000 0.250 0.1719 0.03354 0.02154 -0.0266 1.0000 1.0000 0.500 0.2114 0.03286 0.02145 -0.0303 0.8427 1.0000 0.750 0.3848 0.03783 0.02184 -0.0518 0.5593 1.0000 1.000 0.4234 0.03979 0.02311 -0.0535 0.5383 1.0000 1.250 0.4597 0.04164 0.02455 -0.0551 0.5225 1.0000 1.500 0.4941 0.04350 0.02616 -0.0566 0.5101 1.0000 1.750 0.5276 0.04557 0.02796 -0.0580 0.4997 1.0000 2.000 0.5585 0.04751 0.02991 -0.0595 0.4922 1.0000 2.250 0.5887 0.04969 0.03205 -0.0608 0.4867 1.0000 2.500 0.6176 0.05201 0.03437 -0.0622 0.4840 1.0000 2.750 0.6439 0.05436 0.03687 -0.0638 0.4845 1.0000 3.000 0.6660 0.05683 0.03957 -0.0656 0.4881 1.0000 3.250 0.6854 0.05958 0.04250 -0.0674 0.4929 1.0000 3.500 0.7038 0.06260 0.04562 -0.0690 0.4983 1.0000 3.750 0.7153 0.06578 0.04898 -0.0708 0.5054 1.0000 4.000 0.7229 0.06929 0.05259 -0.0727 0.5140 1.0000 4.250 0.7270 0.07297 0.05635 -0.0746 0.5235 1.0000 4.500 0.7299 0.07687 0.06026 -0.0766 0.5345 1.0000 4.750 0.7193 0.08088 0.06428 -0.0784 0.5477 1.0000 5.000 0.7109 0.08493 0.06827 -0.0801 0.5620 1.0000