XFOIL Version 6.94 Calculated polar for: GOE 279 (DAIMLER X) AIRFOIL 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 ------- -------- --------- --------- -------- ------- ------- 0.500 0.6120 0.00886 0.00239 -0.1055 0.7541 0.0325 1.000 0.6635 0.00909 0.00227 -0.1041 0.6748 0.0407 2.000 0.7672 0.00960 0.00228 -0.1025 0.5605 0.0645 2.500 0.8128 0.00808 0.00253 -0.1008 0.5096 1.0000 3.000 0.8632 0.00873 0.00277 -0.0999 0.4383 1.0000 3.500 0.9128 0.00948 0.00311 -0.0990 0.3738 1.0000 4.000 0.9636 0.01012 0.00350 -0.0983 0.3374 1.0000 4.500 1.0131 0.01087 0.00391 -0.0975 0.2818 1.0000 5.000 1.0533 0.01270 0.00484 -0.0956 0.1297 1.0000 5.500 1.0938 0.01457 0.00611 -0.0936 0.0336 1.0000 6.000 1.1425 0.01533 0.00687 -0.0926 0.0262 1.0000 6.500 1.1909 0.01608 0.00762 -0.0915 0.0184 1.0000 7.000 1.2368 0.01707 0.00861 -0.0901 0.0040 1.0000 7.500 1.2835 0.01789 0.00955 -0.0887 0.0044 1.0000 8.000 1.3289 0.01879 0.01060 -0.0870 0.0051 1.0000 8.500 1.3704 0.02001 0.01204 -0.0847 0.0052 1.0000 9.000 1.4065 0.02160 0.01388 -0.0816 0.0050 1.0000 9.500 1.4346 0.02365 0.01620 -0.0773 0.0048 1.0000 10.000 1.4502 0.02604 0.01883 -0.0712 0.0047 1.0000 10.500 1.4571 0.02906 0.02209 -0.0648 0.0046 1.0000 11.000 1.4577 0.03282 0.02610 -0.0588 0.0045 1.0000 11.500 1.4538 0.03745 0.03097 -0.0539 0.0045 1.0000 12.000 1.4475 0.04291 0.03666 -0.0502 0.0044 1.0000 12.500 1.4400 0.04912 0.04313 -0.0477 0.0043 1.0000 13.000 1.4311 0.05610 0.05038 -0.0465 0.0042 1.0000 13.500 1.4204 0.06389 0.05846 -0.0464 0.0041 1.0000 14.000 1.4078 0.07246 0.06733 -0.0473 0.0041 1.0000 14.500 1.3945 0.08132 0.07641 -0.0486 0.0039 1.0000 15.000 1.3806 0.09033 0.08562 -0.0496 0.0037 1.0000