XFOIL Version 6.94 Calculated polar for: GOE 623 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.300 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 1.000 0.5424 0.00878 0.00308 -0.0845 0.6399 1.0000 2.000 0.6451 0.00926 0.00328 -0.0826 0.6041 1.0000 2.500 0.6972 0.00953 0.00344 -0.0819 0.5871 1.0000 3.000 0.7496 0.00984 0.00366 -0.0812 0.5702 1.0000 3.500 0.8013 0.01011 0.00384 -0.0803 0.5484 1.0000 4.000 0.8479 0.01033 0.00386 -0.0784 0.4899 1.0000 4.500 0.8906 0.01105 0.00408 -0.0761 0.3954 1.0000 5.000 0.9294 0.01234 0.00477 -0.0736 0.2959 1.0000 5.500 0.9719 0.01345 0.00554 -0.0717 0.2352 1.0000 6.000 1.0138 0.01459 0.00635 -0.0698 0.1834 1.0000 6.500 1.0585 0.01546 0.00711 -0.0682 0.1570 1.0000 7.000 1.0742 0.01851 0.00926 -0.0627 0.0166 1.0000 7.500 1.1124 0.01967 0.01046 -0.0602 0.0051 1.0000 8.000 1.1487 0.02073 0.01165 -0.0573 0.0049 1.0000 8.500 1.1812 0.02196 0.01306 -0.0540 0.0049 1.0000 9.000 1.2114 0.02343 0.01470 -0.0507 0.0050 1.0000 9.500 1.2383 0.02521 0.01668 -0.0474 0.0051 1.0000 10.000 1.2611 0.02742 0.01911 -0.0442 0.0053 1.0000 10.500 1.2785 0.03021 0.02214 -0.0410 0.0056 1.0000 11.000 1.2892 0.03380 0.02598 -0.0381 0.0058 1.0000 11.500 1.2908 0.03854 0.03095 -0.0356 0.0060 1.0000 12.000 1.2867 0.04429 0.03693 -0.0339 0.0061 1.0000 12.500 1.2925 0.04940 0.04227 -0.0332 0.0064 1.0000 13.000 1.2882 0.05598 0.04909 -0.0330 0.0067 1.0000 13.500 1.2784 0.06355 0.05689 -0.0334 0.0070 1.0000 14.000 1.2680 0.07142 0.06496 -0.0340 0.0073 1.0000 14.500 1.2604 0.07873 0.07238 -0.0342 0.0077 1.0000 15.000 1.2640 0.08480 0.07861 -0.0344 0.0081 1.0000 15.500 1.2703 0.09044 0.08454 -0.0338 0.0090 1.0000 16.000 1.2915 0.09292 0.08709 -0.0302 0.0100 1.0000 17.500 1.2438 0.12471 0.12068 -0.0350 0.0169 1.0000