XFOIL Version 6.94 Calculated polar for: GOE 505 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 0.5519 0.00987 0.00347 -0.1052 0.7072 0.2390 0.500 0.5783 0.00981 0.00330 -0.0986 0.6732 0.2564 1.000 0.5928 0.00993 0.00323 -0.0896 0.6242 0.2758 1.500 0.6009 0.01042 0.00346 -0.0795 0.5668 0.2953 2.000 0.6090 0.01106 0.00387 -0.0698 0.5064 0.3202 3.000 0.6253 0.01102 0.00491 -0.0513 0.4223 0.8621 4.500 0.9190 0.01424 0.00730 -0.0827 0.3087 1.0000 5.000 0.9483 0.01478 0.00774 -0.0781 0.2930 1.0000 5.500 0.9734 0.01558 0.00833 -0.0731 0.2629 1.0000 6.000 0.9947 0.01668 0.00912 -0.0676 0.2169 1.0000 6.500 1.0040 0.01858 0.01053 -0.0608 0.1493 1.0000 7.500 1.0262 0.02288 0.01425 -0.0488 0.0053 1.0000 8.000 1.0572 0.02403 0.01545 -0.0458 0.0052 1.0000 8.500 1.0862 0.02535 0.01685 -0.0427 0.0053 1.0000 9.000 1.1136 0.02686 0.01848 -0.0396 0.0055 1.0000 9.500 1.1384 0.02861 0.02036 -0.0365 0.0057 1.0000 10.000 1.1622 0.03052 0.02238 -0.0335 0.0060 1.0000 10.500 1.1828 0.03276 0.02478 -0.0304 0.0066 1.0000 11.000 1.1963 0.03568 0.02787 -0.0271 0.0070 1.0000 11.500 1.1985 0.03974 0.03213 -0.0235 0.0073 1.0000 12.000 1.2125 0.04305 0.03559 -0.0211 0.0080 1.0000 12.500 1.2107 0.04815 0.04091 -0.0186 0.0087 1.0000 13.000 1.1955 0.05519 0.04819 -0.0168 0.0092 1.0000 13.500 1.1708 0.06411 0.05734 -0.0163 0.0094 1.0000 14.000 1.1779 0.06958 0.06303 -0.0162 0.0108 1.0000 14.500 1.1546 0.07933 0.07300 -0.0173 0.0112 1.0000 15.000 1.1422 0.08787 0.08174 -0.0184 0.0122 1.0000 15.500 1.1275 0.09669 0.09073 -0.0197 0.0135 1.0000