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 = 0.030 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.2267 0.05839 0.05078 0.0151 0.9996 0.7387 -2.750 -0.2253 0.05515 0.04764 0.0201 0.9996 0.7816 -2.500 -0.2196 0.05223 0.04481 0.0235 0.9996 0.8211 -2.250 -0.2135 0.04896 0.04163 0.0270 0.9996 0.8543 -2.000 -0.1996 0.04626 0.03902 0.0278 0.9996 0.8853 -1.750 -0.1780 0.04361 0.03645 0.0267 0.9996 0.9129 -1.500 -0.1458 0.04158 0.03448 0.0223 0.9996 0.9349 -1.250 -0.1016 0.03999 0.03296 0.0149 0.9996 0.9503 -1.000 -0.0464 0.03938 0.03240 0.0039 0.9996 0.9504 -0.750 0.3325 0.05133 0.04371 -0.0977 0.9996 0.5995 -0.500 0.4440 0.05390 0.04581 -0.1232 0.9996 0.4321 -0.250 0.5037 0.05456 0.04624 -0.1328 0.9996 0.3700 0.000 0.5463 0.05464 0.04631 -0.1376 0.9996 0.3454 0.250 0.5877 0.05503 0.04664 -0.1418 0.9996 0.3249 0.500 0.6266 0.05553 0.04706 -0.1451 0.9996 0.3075 0.750 0.6602 0.05588 0.04754 -0.1469 0.9996 0.2993 1.000 0.9587 0.03985 0.02723 -0.1688 0.2363 0.3176 1.250 0.9779 0.04107 0.02820 -0.1665 0.2238 0.3305 1.500 0.9989 0.04219 0.02921 -0.1645 0.2149 0.3531 1.750 1.0255 0.04317 0.03018 -0.1632 0.2088 0.3909 2.000 1.0610 0.04304 0.03117 -0.1633 0.2040 0.5779 2.250 1.0965 0.04406 0.03125 -0.1621 0.1999 1.0004 2.500 1.1480 0.04584 0.03243 -0.1645 0.1935 1.0004 2.750 1.2128 0.04814 0.03445 -0.1694 0.1886 1.0004 3.000 1.2652 0.05042 0.03679 -0.1722 0.1884 1.0004 3.250 1.3095 0.05320 0.03965 -0.1739 0.1894 1.0004 3.500 1.3438 0.05556 0.04235 -0.1738 0.1916 1.0004 3.750 1.3715 0.05785 0.04530 -0.1725 0.1957 1.0004 4.000 1.3982 0.06098 0.04897 -0.1715 0.2009 1.0004 4.250 1.4239 0.06466 0.05301 -0.1707 0.2057 1.0004 4.500 1.4480 0.06882 0.05740 -0.1699 0.2093 1.0004 4.750 1.4657 0.07229 0.06138 -0.1680 0.2134 1.0004 5.000 1.4701 0.07620 0.06614 -0.1650 0.2213 1.0004