XFOIL Version 6.94 Calculated polar for: pp12 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 -0.1898 0.06385 0.04769 0.0157 0.9999 0.8994 -2.750 -0.2167 0.06292 0.04705 0.0223 0.9999 0.8877 -2.500 -0.2364 0.06190 0.04629 0.0271 0.9999 0.8756 -2.250 -0.2544 0.06096 0.04545 0.0311 0.9999 0.8623 -2.000 -0.2522 0.05979 0.04438 0.0309 0.9999 0.8477 -1.750 -0.2399 0.05876 0.04334 0.0285 0.9999 0.8323 -1.500 -0.2156 0.05796 0.04245 0.0236 0.9999 0.8169 -1.250 -0.1796 0.05722 0.04155 0.0167 0.9999 0.8033 -1.000 -0.1354 0.05658 0.04074 0.0085 0.9999 0.7910 -0.750 -0.0845 0.05633 0.04024 -0.0009 0.9999 0.7799 -0.500 -0.0306 0.05591 0.03957 -0.0104 0.9999 0.7709 -0.250 0.0204 0.05574 0.03921 -0.0188 0.9999 0.7658 0.000 0.0718 0.05567 0.03896 -0.0270 0.9999 0.7650 0.250 0.1257 0.05545 0.03865 -0.0352 0.9999 0.7729 0.500 0.1788 0.05560 0.03874 -0.0433 0.9999 0.7795 0.750 0.2338 0.05550 0.03876 -0.0513 0.9999 0.7986 1.000 0.2870 0.05502 0.03872 -0.0590 0.9999 0.8391 1.250 0.3113 0.05473 0.03875 -0.0611 0.9999 1.0001 1.500 0.3746 0.05671 0.04043 -0.0714 0.9999 1.0001 1.750 0.4359 0.05877 0.04220 -0.0808 0.9999 1.0001 2.000 0.4861 0.06072 0.04393 -0.0877 0.9999 1.0001 2.250 0.5366 0.06275 0.04563 -0.0941 0.9999 1.0001 2.500 0.5781 0.06463 0.04718 -0.0983 0.9999 1.0001 2.750 0.6146 0.06645 0.04861 -0.1012 0.9999 1.0001 3.000 0.6454 0.06817 0.04995 -0.1026 0.9999 1.0001 3.250 0.6729 0.06979 0.05116 -0.1031 0.9999 1.0001 3.500 0.6965 0.07139 0.05264 -0.1031 0.9999 1.0001 3.750 0.7183 0.07302 0.05434 -0.1029 0.9999 1.0001 4.000 0.7385 0.07480 0.05633 -0.1028 0.9999 1.0001 4.250 0.7582 0.07676 0.05884 -0.1027 0.9999 1.0001 4.500 0.7743 0.07930 0.06207 -0.1033 0.9999 1.0001 4.750 0.7803 0.08366 0.06723 -0.1054 0.9999 1.0001 5.000 0.7716 0.09015 0.07412 -0.1094 0.9999 1.0001