XFOIL Version 6.94 Calculated polar for: AH 63-K-127/24 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.450 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.4168 0.01196 0.00738 -0.0917 0.7267 0.8636 0.500 0.4710 0.01201 0.00735 -0.0913 0.7232 0.8771 1.000 0.5207 0.01200 0.00738 -0.0901 0.7184 0.8893 1.500 0.5733 0.01183 0.00724 -0.0897 0.7118 0.9000 2.000 0.6273 0.01163 0.00706 -0.0895 0.7057 0.9076 2.500 0.6846 0.01134 0.00678 -0.0901 0.7008 0.9153 3.000 0.7396 0.01086 0.00635 -0.0897 0.6964 0.9246 3.500 0.7965 0.01052 0.00602 -0.0901 0.6902 0.9326 4.000 0.8495 0.01004 0.00567 -0.0897 0.6793 0.9406 4.500 0.9065 0.00939 0.00507 -0.0899 0.6684 0.9503 5.000 0.9619 0.00890 0.00475 -0.0897 0.6484 0.9607 5.500 1.0223 0.00867 0.00455 -0.0909 0.6259 0.9721 6.500 1.1306 0.00945 0.00511 -0.0918 0.5185 1.0000 7.000 1.1610 0.01084 0.00611 -0.0882 0.4234 1.0000 7.500 1.1386 0.01454 0.00859 -0.0768 0.2205 1.0000 8.000 1.1330 0.01791 0.01139 -0.0700 0.1205 1.0000 8.500 1.1442 0.02099 0.01411 -0.0666 0.0637 1.0000 9.000 1.1617 0.02354 0.01653 -0.0634 0.0342 1.0000 9.500 1.1775 0.02623 0.01916 -0.0600 0.0189 1.0000 10.000 1.1799 0.03005 0.02306 -0.0551 0.0063 1.0000 10.500 1.1974 0.03281 0.02611 -0.0523 0.0083 1.0000 11.000 1.1969 0.03716 0.03067 -0.0477 0.0056 1.0000 11.500 1.2013 0.04141 0.03518 -0.0440 0.0053 1.0000 12.000 1.2167 0.04492 0.03892 -0.0412 0.0054 1.0000 12.500 1.2289 0.04938 0.04376 -0.0375 0.0055 1.0000 13.000 1.2404 0.05554 0.05053 -0.0335 0.0061 1.0000 13.500 1.2255 0.06500 0.06071 -0.0301 0.0070 1.0000 14.000 1.1982 0.07447 0.07066 -0.0290 0.0073 1.0000 14.500 1.1583 0.08627 0.08293 -0.0301 0.0078 1.0000 15.000 1.0903 0.10489 0.10207 -0.0361 0.0088 1.0000 15.500 1.0436 0.12224 0.11974 -0.0453 0.0091 1.0000 16.000 1.0373 0.13291 0.13062 -0.0524 0.0084 1.0000 16.500 0.9894 0.15889 0.15686 -0.0685 0.0085 1.0000