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 = 0.005 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1556 0.05981 0.04173 -0.0076 1.0000 1.0000 -2.750 -0.1579 0.05780 0.04008 -0.0070 1.0000 1.0000 -2.500 -0.1621 0.05576 0.03845 -0.0062 1.0000 1.0000 -2.250 -0.1675 0.05372 0.03677 -0.0055 1.0000 1.0000 -2.000 -0.1672 0.05175 0.03487 -0.0066 1.0000 1.0000 -1.750 -0.1366 0.05019 0.03231 -0.0146 1.0000 1.0000 -1.500 -0.0816 0.04982 0.02967 -0.0233 1.0000 1.0000 -1.250 -0.0396 0.04990 0.02779 -0.0255 1.0000 1.0000 -1.000 -0.0071 0.04998 0.02642 -0.0252 1.0000 1.0000 -0.750 0.0215 0.05005 0.02533 -0.0243 1.0000 1.0000 -0.500 0.0483 0.05011 0.02455 -0.0233 1.0000 1.0000 -0.250 0.0743 0.05018 0.02398 -0.0224 1.0000 1.0000 0.000 0.1000 0.05027 0.02361 -0.0215 1.0000 1.0000 0.250 0.1258 0.05037 0.02338 -0.0207 1.0000 1.0000 0.500 0.1518 0.05050 0.02339 -0.0200 1.0000 1.0000 0.750 0.1784 0.05065 0.02359 -0.0194 1.0000 1.0000 1.000 0.2059 0.05082 0.02403 -0.0191 1.0000 1.0000 1.250 0.2349 0.05102 0.02474 -0.0192 1.0000 1.0000 1.500 0.2652 0.05133 0.02576 -0.0201 1.0000 1.0000 1.750 0.2940 0.05198 0.02732 -0.0221 1.0000 1.0000 2.000 0.3138 0.05344 0.02955 -0.0249 1.0000 1.0000 2.250 0.3187 0.05597 0.03236 -0.0274 1.0000 1.0000 2.500 0.3159 0.05903 0.03522 -0.0292 1.0000 1.0000 2.750 0.3150 0.06206 0.03792 -0.0308 1.0000 1.0000 3.000 0.3170 0.06496 0.04050 -0.0325 1.0000 1.0000 3.250 0.3211 0.06776 0.04300 -0.0342 1.0000 1.0000 3.500 0.3265 0.07051 0.04545 -0.0358 1.0000 1.0000 3.750 0.3330 0.07321 0.04788 -0.0373 1.0000 1.0000 4.000 0.3403 0.07591 0.05030 -0.0387 1.0000 1.0000 4.250 0.3482 0.07861 0.05272 -0.0402 1.0000 1.0000 4.500 0.3565 0.08131 0.05516 -0.0415 1.0000 1.0000 4.750 0.3653 0.08402 0.05762 -0.0429 1.0000 1.0000 5.000 0.3743 0.08674 0.06010 -0.0442 1.0000 1.0000