XFOIL Version 6.94 Calculated polar for: GOE 242 (MVA PR.2) AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.300 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.500 1.0705 0.01590 0.00875 -0.2190 0.5751 0.5098 1.000 1.1278 0.01604 0.00877 -0.2194 0.5667 0.5144 1.500 1.1852 0.01642 0.00910 -0.2201 0.5599 0.5212 2.000 1.2429 0.01661 0.00928 -0.2209 0.5543 0.5286 2.500 1.2990 0.01672 0.00942 -0.2213 0.5475 0.5337 4.500 1.4924 0.01670 0.00898 -0.2166 0.4321 0.5554 5.000 1.5065 0.01873 0.01033 -0.2101 0.3268 0.5606 5.500 1.5129 0.02094 0.01219 -0.2024 0.2701 0.5657 6.000 1.5197 0.02366 0.01457 -0.1958 0.2173 0.5704 6.500 1.5334 0.02624 0.01703 -0.1909 0.1824 0.5751 7.500 1.4834 0.03894 0.02909 -0.1764 0.0050 0.5827 8.000 1.5022 0.04224 0.03254 -0.1740 0.0048 0.5895 8.500 1.5181 0.04593 0.03644 -0.1718 0.0049 0.5955 9.000 1.5306 0.05013 0.04086 -0.1696 0.0050 0.6015 9.500 1.5400 0.05495 0.04592 -0.1676 0.0051 0.6082 10.000 1.5477 0.06015 0.05135 -0.1659 0.0053 0.6144 10.500 1.5548 0.06560 0.05705 -0.1645 0.0056 0.6212 11.000 1.5578 0.07167 0.06336 -0.1631 0.0060 0.6292 11.500 1.5567 0.07839 0.07035 -0.1619 0.0063 0.6367 12.000 1.5499 0.08602 0.07827 -0.1610 0.0066 0.6447 12.500 1.5488 0.09301 0.08549 -0.1604 0.0070 0.6534 13.000 1.5459 0.10044 0.09319 -0.1601 0.0075 0.6627 13.500 1.5346 0.10924 0.10225 -0.1603 0.0080 0.6725 14.000 1.5172 0.11920 0.11246 -0.1613 0.0083 0.6825 14.500 1.5179 0.12653 0.12004 -0.1622 0.0089 0.6988 15.000 1.5094 0.13539 0.12914 -0.1638 0.0097 0.7186 15.500 1.4984 0.14461 0.13856 -0.1660 0.0102 0.7565 16.500 1.5176 0.15544 0.14974 -0.1679 0.0128 1.0000