XFOIL Version 6.94 Calculated polar for: pp14 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.035 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -2.250 -0.0930 0.05279 0.04584 -0.0117 0.9996 0.7198 -2.000 -0.0875 0.04992 0.04309 -0.0074 0.9996 0.7654 -1.750 -0.0756 0.04735 0.04061 -0.0052 0.9996 0.8055 -1.500 -0.0523 0.04519 0.03853 -0.0061 0.9996 0.8414 -1.250 -0.0263 0.04299 0.03642 -0.0077 0.9996 0.8684 -1.000 0.0152 0.04147 0.03498 -0.0136 0.9996 0.8903 -0.750 0.0832 0.04104 0.03461 -0.0264 0.9996 0.8973 -0.500 0.3880 0.05092 0.04402 -0.1072 0.9996 0.5806 -0.250 0.4915 0.05352 0.04608 -0.1302 0.9996 0.3897 0.000 0.5454 0.05399 0.04634 -0.1378 0.9996 0.3258 0.250 0.5867 0.05420 0.04655 -0.1418 0.9996 0.3024 0.750 0.9458 0.03858 0.02600 -0.1723 0.2058 0.2738 1.000 0.9660 0.03957 0.02678 -0.1702 0.1967 0.2806 1.250 0.9864 0.04073 0.02771 -0.1681 0.1904 0.2876 1.500 1.0131 0.04147 0.02844 -0.1668 0.1861 0.3026 1.750 1.0456 0.04212 0.02924 -0.1664 0.1813 0.3363 2.000 1.0894 0.04249 0.03013 -0.1680 0.1754 0.4476 2.250 1.1402 0.04327 0.03036 -0.1696 0.1705 1.0004 2.500 1.2065 0.04549 0.03210 -0.1746 0.1688 1.0004 2.750 1.2522 0.04760 0.03424 -0.1764 0.1694 1.0004 3.000 1.2885 0.04944 0.03627 -0.1765 0.1711 1.0004 3.250 1.3215 0.05150 0.03872 -0.1760 0.1742 1.0004 3.500 1.3523 0.05409 0.04174 -0.1753 0.1786 1.0004 3.750 1.3813 0.05716 0.04519 -0.1746 0.1827 1.0004 4.000 1.4079 0.06058 0.04888 -0.1738 0.1857 1.0004 4.250 1.4328 0.06445 0.05295 -0.1729 0.1882 1.0004 4.500 1.4534 0.06741 0.05642 -0.1711 0.1924 1.0004 4.750 1.4648 0.07100 0.06089 -0.1681 0.2025 1.0004 5.000 1.4892 0.07682 0.06677 -0.1679 0.2107 1.0004