XFOIL Version 6.94 Calculated polar for: pp13 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.008 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1812 0.06320 0.04689 0.0133 0.9998 0.9016 -2.750 -0.2057 0.06224 0.04621 0.0193 0.9998 0.8894 -2.500 -0.2303 0.06128 0.04548 0.0252 0.9998 0.8776 -2.250 -0.2487 0.06028 0.04468 0.0295 0.9998 0.8648 -2.000 -0.2506 0.05913 0.04360 0.0302 0.9998 0.8503 -1.750 -0.2432 0.05812 0.04261 0.0288 0.9998 0.8355 -1.500 -0.2233 0.05733 0.04172 0.0248 0.9998 0.8204 -1.250 -0.1896 0.05643 0.04071 0.0184 0.9998 0.8066 -1.000 -0.1487 0.05582 0.03989 0.0108 0.9998 0.7945 -0.750 -0.1006 0.05539 0.03926 0.0020 0.9998 0.7831 -0.500 -0.0491 0.05511 0.03874 -0.0071 0.9998 0.7729 -0.250 0.0023 0.05490 0.03832 -0.0158 0.9998 0.7658 0.000 0.0533 0.05472 0.03797 -0.0239 0.9998 0.7647 0.250 0.1055 0.05453 0.03765 -0.0319 0.9998 0.7689 0.500 0.1588 0.05454 0.03759 -0.0400 0.9998 0.7763 0.750 0.2126 0.05444 0.03757 -0.0479 0.9998 0.7902 1.000 0.2670 0.05418 0.03760 -0.0559 0.9998 0.8200 1.250 0.3052 0.05349 0.03745 -0.0610 0.9998 0.8901 1.500 0.3556 0.05500 0.03875 -0.0682 0.9998 1.0002 1.750 0.4155 0.05697 0.04046 -0.0775 0.9998 1.0002 2.000 0.4713 0.05896 0.04217 -0.0856 0.9998 1.0002 2.250 0.5219 0.06094 0.04386 -0.0921 0.9998 1.0002 2.500 0.5659 0.06283 0.04540 -0.0970 0.9998 1.0002 2.750 0.6046 0.06463 0.04682 -0.1003 0.9998 1.0002 3.000 0.6382 0.06634 0.04809 -0.1022 0.9998 1.0002 3.250 0.6666 0.06794 0.04930 -0.1030 0.9998 1.0002 3.500 0.6914 0.06952 0.05064 -0.1031 0.9998 1.0002 3.750 0.7143 0.07113 0.05228 -0.1030 0.9998 1.0002 4.000 0.7360 0.07281 0.05421 -0.1029 0.9998 1.0002 4.250 0.7568 0.07467 0.05649 -0.1030 0.9998 1.0002 4.500 0.7757 0.07699 0.05949 -0.1035 0.9998 1.0002 4.750 0.7869 0.08077 0.06413 -0.1053 0.9998 1.0002 5.000 0.7795 0.08785 0.07179 -0.1100 0.9998 1.0002