XFOIL Version 6.94 Calculated polar for: GOE 255 (MVA CA.6) AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.450 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.5713 0.01103 0.00379 -0.1170 0.5769 0.4587 0.500 0.6284 0.01110 0.00387 -0.1175 0.5698 0.4699 1.000 0.6856 0.01121 0.00396 -0.1181 0.5638 0.4817 1.500 0.7422 0.01128 0.00408 -0.1185 0.5569 0.4931 2.000 0.7987 0.01138 0.00424 -0.1190 0.5506 0.5053 2.500 0.8550 0.01149 0.00441 -0.1195 0.5440 0.5235 3.000 0.9101 0.01153 0.00466 -0.1198 0.5372 0.5567 3.500 0.9638 0.01148 0.00495 -0.1198 0.5300 0.6504 4.000 0.9977 0.01098 0.00514 -0.1150 0.4913 0.9668 4.500 1.0399 0.01136 0.00531 -0.1129 0.4395 1.0000 5.000 1.0166 0.01419 0.00712 -0.0995 0.2861 1.0000 5.500 1.0417 0.01561 0.00823 -0.0951 0.2380 1.0000 6.000 1.0400 0.01853 0.01058 -0.0875 0.1407 1.0000 7.000 1.0521 0.02491 0.01647 -0.0773 0.0041 1.0000 7.500 1.0841 0.02674 0.01834 -0.0758 0.0040 1.0000 8.000 1.1141 0.02881 0.02046 -0.0742 0.0039 1.0000 8.500 1.1416 0.03113 0.02286 -0.0726 0.0040 1.0000 9.000 1.1666 0.03370 0.02557 -0.0710 0.0041 1.0000 9.500 1.1892 0.03654 0.02855 -0.0693 0.0043 1.0000 10.000 1.2085 0.03977 0.03194 -0.0675 0.0044 1.0000 10.500 1.2229 0.04354 0.03591 -0.0657 0.0046 1.0000 11.000 1.2339 0.04778 0.04035 -0.0640 0.0048 1.0000 11.500 1.2451 0.05218 0.04490 -0.0626 0.0051 1.0000 12.000 1.2541 0.05696 0.04987 -0.0614 0.0055 1.0000 12.500 1.2607 0.06222 0.05529 -0.0604 0.0060 1.0000 13.000 1.2653 0.06788 0.06115 -0.0597 0.0068 1.0000 13.500 1.2631 0.07453 0.06802 -0.0591 0.0075 1.0000 14.000 1.2543 0.08217 0.07587 -0.0589 0.0082 1.0000 14.500 1.2392 0.09082 0.08474 -0.0591 0.0087 1.0000