XFOIL Version 6.94 Calculated polar for: GOE 584 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 ------- -------- --------- --------- -------- ------- ------- 0.000 0.5010 0.00999 0.00277 -0.0945 0.6253 0.1630 0.500 0.5504 0.00939 0.00280 -0.0935 0.6142 0.3844 1.500 0.7204 0.00824 0.00304 -0.1058 0.5899 1.0000 2.000 0.7692 0.00843 0.00312 -0.1043 0.5783 1.0000 3.000 0.8575 0.00876 0.00317 -0.0993 0.5190 1.0000 3.500 0.9022 0.00906 0.00332 -0.0970 0.4900 1.0000 4.000 0.9352 0.00981 0.00362 -0.0926 0.4074 1.0000 4.500 0.9717 0.01060 0.00409 -0.0891 0.3511 1.0000 5.000 0.9803 0.01286 0.00537 -0.0812 0.1945 1.0000 5.500 0.9649 0.01601 0.00752 -0.0695 0.0086 1.0000 6.000 0.9993 0.01669 0.00816 -0.0659 0.0056 1.0000 6.500 1.0335 0.01746 0.00900 -0.0626 0.0054 1.0000 7.000 1.0662 0.01838 0.00999 -0.0592 0.0054 1.0000 7.500 1.0980 0.01944 0.01115 -0.0561 0.0056 1.0000 8.000 1.1279 0.02071 0.01254 -0.0529 0.0059 1.0000 8.500 1.1558 0.02222 0.01417 -0.0499 0.0062 1.0000 9.000 1.1822 0.02396 0.01606 -0.0470 0.0067 1.0000 9.500 1.2066 0.02597 0.01822 -0.0443 0.0070 1.0000 10.000 1.2286 0.02831 0.02073 -0.0417 0.0077 1.0000 10.500 1.2456 0.03119 0.02382 -0.0390 0.0084 1.0000 11.000 1.2558 0.03482 0.02766 -0.0365 0.0086 1.0000 11.500 1.2681 0.03846 0.03146 -0.0345 0.0096 1.0000 12.000 1.2763 0.04268 0.03586 -0.0328 0.0101 1.0000 12.500 1.2878 0.04687 0.04028 -0.0314 0.0118 1.0000 13.000 1.2895 0.05234 0.04596 -0.0304 0.0130 1.0000 13.500 1.2781 0.05956 0.05340 -0.0297 0.0141 1.0000 14.000 1.2742 0.06621 0.06025 -0.0297 0.0135 1.0000