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 = 0.450 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.500 0.4907 0.00740 0.00243 -0.0851 0.6177 1.0000 1.000 0.5415 0.00758 0.00244 -0.0841 0.5999 1.0000 1.500 0.5932 0.00774 0.00252 -0.0833 0.5851 1.0000 2.000 0.6447 0.00798 0.00261 -0.0824 0.5695 1.0000 2.500 0.6972 0.00817 0.00272 -0.0817 0.5543 1.0000 3.000 0.7502 0.00838 0.00291 -0.0811 0.5382 1.0000 3.500 0.7979 0.00876 0.00296 -0.0796 0.4777 1.0000 4.000 0.8392 0.00981 0.00331 -0.0772 0.3498 1.0000 4.500 0.8845 0.01076 0.00388 -0.0757 0.2801 1.0000 5.000 0.9257 0.01210 0.00466 -0.0736 0.1935 1.0000 5.500 0.9715 0.01300 0.00531 -0.0722 0.1512 1.0000 6.000 0.9974 0.01555 0.00705 -0.0680 0.0051 1.0000 6.500 1.0441 0.01625 0.00780 -0.0666 0.0039 1.0000 7.000 1.0893 0.01701 0.00862 -0.0651 0.0039 1.0000 7.500 1.1324 0.01786 0.00958 -0.0632 0.0041 1.0000 8.000 1.1720 0.01888 0.01072 -0.0608 0.0043 1.0000 8.500 1.2041 0.02009 0.01210 -0.0572 0.0045 1.0000 9.000 1.2360 0.02137 0.01350 -0.0539 0.0049 1.0000 9.500 1.2647 0.02294 0.01523 -0.0505 0.0054 1.0000 10.000 1.2858 0.02518 0.01765 -0.0467 0.0059 1.0000 10.500 1.3035 0.02784 0.02048 -0.0432 0.0063 1.0000 11.000 1.3210 0.03072 0.02357 -0.0402 0.0071 1.0000 11.500 1.3210 0.03541 0.02846 -0.0371 0.0077 1.0000 12.000 1.3313 0.03953 0.03279 -0.0352 0.0084 1.0000 12.500 1.3249 0.04568 0.03917 -0.0336 0.0093 1.0000 13.000 1.3100 0.05291 0.04655 -0.0322 0.0099 1.0000 13.500 1.3161 0.05815 0.05206 -0.0309 0.0115 1.0000