XFOIL Version 6.94 Calculated polar for: GOE 406c 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.008 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.2543 0.08641 0.07512 0.0371 0.9999 0.7198 -2.750 -0.2647 0.08468 0.07352 0.0372 0.9999 0.7089 -2.500 -0.2585 0.08242 0.07137 0.0354 0.9999 0.6998 -2.250 -0.2521 0.08033 0.06937 0.0325 0.9999 0.6904 -2.000 -0.2360 0.07817 0.06724 0.0283 0.9999 0.6830 -1.750 -0.2135 0.07604 0.06516 0.0234 0.9999 0.6773 -1.500 -0.1854 0.07411 0.06324 0.0174 0.9999 0.6736 -1.250 -0.1527 0.07239 0.06153 0.0108 0.9999 0.6725 -1.000 -0.1176 0.07095 0.06009 0.0042 0.9999 0.6740 -0.750 -0.0814 0.06985 0.05903 -0.0024 0.9999 0.6788 -0.500 -0.0445 0.06921 0.05848 -0.0093 0.9999 0.6875 -0.250 -0.0165 0.06890 0.05840 -0.0137 0.9999 0.6993 0.000 0.0097 0.06947 0.05922 -0.0190 0.9999 0.7133 0.250 0.0251 0.07115 0.06120 -0.0235 0.9999 0.7263 0.500 0.0330 0.07399 0.06429 -0.0281 0.9999 0.7372 0.750 0.0419 0.07677 0.06721 -0.0325 0.9999 0.7524 1.000 0.0538 0.07900 0.06963 -0.0362 0.9999 0.7759 1.250 0.0663 0.08071 0.07163 -0.0394 0.9999 0.8134 1.500 0.0742 0.08188 0.07302 -0.0420 0.9999 0.8834 1.750 0.0807 0.08288 0.07394 -0.0456 0.9999 1.0001 2.000 0.1272 0.08708 0.07733 -0.0579 0.9999 1.0001 2.250 0.1667 0.09105 0.08037 -0.0673 0.9999 1.0001 2.500 0.1987 0.09469 0.08308 -0.0739 0.9999 1.0001 2.750 0.2251 0.09805 0.08553 -0.0783 0.9999 1.0001 3.000 0.2475 0.10119 0.08778 -0.0813 0.9999 1.0001 3.250 0.2671 0.10417 0.08992 -0.0834 0.9999 1.0001 3.500 0.2848 0.10705 0.09199 -0.0848 0.9999 1.0001 3.750 0.3012 0.10986 0.09401 -0.0858 0.9999 1.0001 4.000 0.3165 0.11264 0.09605 -0.0866 0.9999 1.0001 4.250 0.3312 0.11540 0.09812 -0.0874 0.9999 1.0001 4.500 0.3452 0.11816 0.10024 -0.0881 0.9999 1.0001 4.750 0.3589 0.12092 0.10241 -0.0889 0.9999 1.0001 5.000 0.3723 0.12369 0.10465 -0.0897 0.9999 1.0001