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 = 0.018 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0185 0.04970 0.03916 -0.0142 0.9996 0.9934 -2.750 -0.0581 0.04932 0.03900 -0.0039 0.9996 0.9784 -2.500 -0.0975 0.04875 0.03864 0.0061 0.9996 0.9647 -2.250 -0.1379 0.04799 0.03811 0.0161 0.9996 0.9522 -2.000 -0.1661 0.04736 0.03764 0.0226 0.9996 0.9327 -1.750 -0.1078 0.04909 0.03929 0.0051 0.9996 0.8601 -1.500 0.1367 0.05458 0.04418 -0.0609 0.9996 0.6636 -1.250 0.2484 0.05591 0.04523 -0.0868 0.9996 0.5823 -1.000 0.3221 0.05624 0.04542 -0.1013 0.9996 0.5345 -0.750 0.3882 0.05669 0.04569 -0.1133 0.9996 0.4940 -0.500 0.4421 0.05692 0.04582 -0.1216 0.9996 0.4662 -0.250 0.4901 0.05725 0.04605 -0.1281 0.9996 0.4498 0.000 0.5325 0.05754 0.04630 -0.1330 0.9996 0.4422 0.250 0.5729 0.05804 0.04675 -0.1372 0.9996 0.4376 0.500 0.6100 0.05862 0.04729 -0.1405 0.9996 0.4345 0.750 0.6443 0.05918 0.04791 -0.1429 0.9996 0.4333 1.000 0.6765 0.05976 0.04864 -0.1448 0.9996 0.4389 1.250 0.7077 0.06041 0.04950 -0.1462 0.9996 0.4513 1.500 0.7390 0.06104 0.05047 -0.1473 0.9996 0.4670 1.750 0.7713 0.06164 0.05153 -0.1482 0.9996 0.4872 2.000 0.8063 0.06198 0.05265 -0.1490 0.9996 0.5252 2.250 1.0626 0.04536 0.03280 -0.1557 0.3567 1.0004 2.500 1.0900 0.04797 0.03444 -0.1544 0.3266 1.0004 2.750 1.1253 0.05041 0.03626 -0.1544 0.3070 1.0004 3.000 1.1691 0.05298 0.03834 -0.1556 0.2948 1.0004 3.250 1.2151 0.05545 0.04091 -0.1574 0.2876 1.0004 3.500 1.2607 0.05826 0.04385 -0.1592 0.2830 1.0004 3.750 1.3024 0.06134 0.04717 -0.1607 0.2807 1.0004 4.000 1.3405 0.06477 0.05093 -0.1620 0.2788 1.0004 4.250 1.3731 0.06850 0.05507 -0.1627 0.2774 1.0004 4.500 1.3984 0.07244 0.05946 -0.1626 0.2769 1.0004 4.750 1.4135 0.07655 0.06415 -0.1616 0.2782 1.0004 5.000 1.4194 0.08116 0.06939 -0.1601 0.2812 1.0004