XFOIL Version 6.94 Calculated polar for: GOE 563 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 ------- -------- --------- --------- -------- ------- ------- 0.000 0.3612 0.00748 0.00318 -0.0810 0.8889 1.0000 0.500 0.4077 0.00737 0.00296 -0.0786 0.8660 1.0000 1.000 0.4551 0.00731 0.00280 -0.0763 0.8414 1.0000 1.500 0.5017 0.00730 0.00268 -0.0738 0.8102 1.0000 2.000 0.5476 0.00735 0.00259 -0.0710 0.7675 1.0000 2.500 0.5865 0.00766 0.00242 -0.0666 0.6652 1.0000 3.000 0.6166 0.00871 0.00254 -0.0610 0.4872 1.0000 3.500 0.6431 0.01044 0.00314 -0.0557 0.2733 1.0000 4.000 0.6834 0.01146 0.00363 -0.0530 0.1729 1.0000 4.500 0.7248 0.01248 0.00423 -0.0505 0.1010 1.0000 5.000 0.7637 0.01385 0.00506 -0.0476 0.0210 1.0000 5.500 0.8088 0.01466 0.00584 -0.0455 0.0071 1.0000 6.000 0.8554 0.01532 0.00663 -0.0436 0.0060 1.0000 6.500 0.9007 0.01613 0.00761 -0.0416 0.0058 1.0000 7.000 0.9452 0.01704 0.00877 -0.0393 0.0057 1.0000 7.500 0.9878 0.01813 0.01011 -0.0368 0.0058 1.0000 8.000 1.0272 0.01952 0.01173 -0.0338 0.0059 1.0000 8.500 1.0599 0.02147 0.01405 -0.0299 0.0062 1.0000 9.000 1.0850 0.02403 0.01692 -0.0250 0.0066 1.0000 9.500 1.1061 0.02691 0.02013 -0.0197 0.0068 1.0000 10.000 1.1204 0.03105 0.02471 -0.0140 0.0072 1.0000 10.500 1.1260 0.03611 0.03032 -0.0081 0.0076 1.0000 11.000 1.1134 0.04247 0.03726 -0.0020 0.0079 1.0000 11.500 1.0916 0.04948 0.04477 0.0022 0.0081 1.0000 12.000 1.0813 0.05521 0.05083 0.0037 0.0083 1.0000 12.500 1.0686 0.06234 0.05830 0.0030 0.0089 1.0000 13.000 1.0328 0.07410 0.07042 -0.0026 0.0086 1.0000 13.500 0.9989 0.08882 0.08549 -0.0119 0.0088 1.0000 14.000 0.9511 0.10988 0.10683 -0.0259 0.0085 1.0000 14.500 0.8932 0.13751 0.13464 -0.0417 0.0083 1.0000 15.000 0.8737 0.15549 0.15261 -0.0502 0.0085 1.0000 15.500 0.8524 0.17783 0.17484 -0.0595 0.0111 1.0000