XFOIL Version 6.94 Calculated polar for: COUDE2 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.030 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -2.500 -0.2587 0.06404 0.05652 -0.0103 0.9999 0.3465 -2.000 -0.1597 0.05837 0.04993 -0.0234 0.9999 0.2751 -1.750 -0.1187 0.05668 0.04789 -0.0264 0.9999 0.2619 -1.500 -0.0830 0.05498 0.04617 -0.0278 0.9999 0.2551 -1.250 -0.0418 0.05368 0.04482 -0.0299 0.9999 0.2460 -0.750 0.2278 0.04484 0.03194 -0.0663 0.4329 0.2448 -0.500 0.3070 0.04438 0.03103 -0.0755 0.4079 0.2556 -0.250 0.3830 0.04433 0.03074 -0.0844 0.3868 0.2720 0.000 0.4568 0.04424 0.03093 -0.0935 0.3698 0.3217 0.250 0.5373 0.04292 0.03121 -0.1022 0.3597 1.0001 0.500 0.6418 0.04562 0.03285 -0.1175 0.3525 1.0001 0.750 0.7229 0.04807 0.03513 -0.1289 0.3492 1.0001 1.000 0.7802 0.05039 0.03742 -0.1354 0.3452 1.0001 1.250 0.8188 0.05252 0.03951 -0.1378 0.3410 1.0001 1.500 0.8553 0.05482 0.04178 -0.1398 0.3377 1.0001 1.750 0.8826 0.05706 0.04420 -0.1399 0.3397 1.0001 2.000 0.9016 0.05936 0.04672 -0.1384 0.3438 1.0001 2.250 0.9194 0.06191 0.04941 -0.1368 0.3483 1.0001 2.500 0.9412 0.06493 0.05247 -0.1361 0.3527 1.0001 2.750 0.9274 0.06649 0.05446 -0.1282 0.3610 1.0001 3.000 0.9182 0.06908 0.05730 -0.1219 0.3700 1.0001 3.250 0.9263 0.07222 0.06055 -0.1192 0.3792 1.0001 3.500 0.8696 0.07415 0.06289 -0.1050 0.3917 1.0001 3.750 0.8336 0.07726 0.06626 -0.0961 0.4063 1.0001 4.000 0.7916 0.08149 0.07075 -0.0883 0.4264 1.0001 4.250 0.7036 0.08851 0.07815 -0.0795 0.4638 1.0001 4.500 0.2446 0.11035 0.10285 -0.0635 0.9617 1.0001 4.750 0.2562 0.11306 0.10517 -0.0637 0.9616 1.0001 5.000 0.2696 0.11559 0.10735 -0.0644 0.9580 1.0001