XFOIL Version 6.94 Calculated polar for: GOE 565 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 0.3869 0.00718 0.00285 -0.0887 0.8797 1.0000 0.500 0.4376 0.00714 0.00268 -0.0870 0.8568 1.0000 1.000 0.4887 0.00713 0.00253 -0.0853 0.8307 1.0000 1.500 0.5378 0.00715 0.00241 -0.0831 0.7913 1.0000 2.000 0.5855 0.00733 0.00230 -0.0806 0.7343 1.0000 3.000 0.6606 0.00906 0.00257 -0.0723 0.4483 1.0000 3.500 0.6906 0.01113 0.00328 -0.0680 0.1950 1.0000 4.000 0.7358 0.01210 0.00387 -0.0662 0.1266 1.0000 4.500 0.7787 0.01337 0.00465 -0.0641 0.0377 1.0000 5.000 0.8265 0.01409 0.00527 -0.0626 0.0198 1.0000 6.000 0.9220 0.01560 0.00689 -0.0594 0.0062 1.0000 6.500 0.9685 0.01650 0.00798 -0.0575 0.0062 1.0000 7.000 1.0138 0.01751 0.00925 -0.0554 0.0063 1.0000 7.500 1.0564 0.01878 0.01084 -0.0529 0.0065 1.0000 8.000 1.0939 0.02052 0.01289 -0.0496 0.0068 1.0000 8.500 1.1236 0.02294 0.01561 -0.0452 0.0072 1.0000 9.000 1.1474 0.02618 0.01927 -0.0401 0.0076 1.0000 9.500 1.1690 0.03079 0.02435 -0.0351 0.0082 1.0000 10.000 1.1801 0.03705 0.03124 -0.0295 0.0088 1.0000 10.500 1.1720 0.04386 0.03865 -0.0228 0.0092 1.0000 11.000 1.1821 0.04680 0.04189 -0.0182 0.0100 1.0000 11.500 1.1451 0.05646 0.05230 -0.0132 0.0112 1.0000 12.000 1.1069 0.06604 0.06236 -0.0132 0.0115 1.0000 12.500 1.0651 0.07840 0.07512 -0.0185 0.0115 1.0000 13.000 1.0229 0.09456 0.09159 -0.0293 0.0115 1.0000 13.500 1.0054 0.10777 0.10495 -0.0387 0.0107 1.0000 14.000 0.9717 0.12799 0.12531 -0.0516 0.0102 1.0000