XFOIL Version 6.94 Calculated polar for: GOE 509 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 0.4463 0.01000 0.00242 -0.0710 0.5755 0.0952 0.500 0.4786 0.00997 0.00258 -0.0662 0.5080 0.2739 1.000 0.5185 0.01021 0.00283 -0.0629 0.4742 0.3598 1.500 0.5608 0.01030 0.00295 -0.0602 0.4574 0.4105 2.500 0.8499 0.01053 0.00410 -0.0999 0.4134 1.0000 3.000 0.8922 0.01078 0.00428 -0.0973 0.4026 1.0000 3.500 0.9342 0.01104 0.00450 -0.0946 0.3912 1.0000 4.000 0.9761 0.01130 0.00471 -0.0919 0.3798 1.0000 4.500 1.0181 0.01151 0.00494 -0.0892 0.3659 1.0000 5.000 1.0573 0.01177 0.00511 -0.0860 0.3382 1.0000 5.500 1.0884 0.01241 0.00544 -0.0814 0.2738 1.0000 6.000 1.0949 0.01437 0.00660 -0.0726 0.1631 1.0000 6.500 1.0862 0.01662 0.00813 -0.0610 0.0406 1.0000 7.000 1.1066 0.01758 0.00901 -0.0546 0.0228 1.0000 7.500 1.1361 0.01824 0.00972 -0.0501 0.0195 1.0000 8.000 1.1589 0.01933 0.01076 -0.0448 0.0022 1.0000 8.500 1.1872 0.02025 0.01177 -0.0406 0.0021 1.0000 9.000 1.2129 0.02139 0.01303 -0.0362 0.0020 1.0000 9.500 1.2374 0.02267 0.01442 -0.0320 0.0020 1.0000 10.000 1.2593 0.02423 0.01613 -0.0278 0.0021 1.0000 10.500 1.2793 0.02607 0.01816 -0.0239 0.0022 1.0000 11.000 1.2955 0.02837 0.02071 -0.0201 0.0023 1.0000 11.500 1.3072 0.03125 0.02382 -0.0166 0.0023 1.0000 12.000 1.3135 0.03489 0.02768 -0.0134 0.0026 1.0000 12.500 1.3175 0.03901 0.03201 -0.0110 0.0027 1.0000 13.000 1.3158 0.04408 0.03731 -0.0092 0.0027 1.0000 13.500 1.3098 0.05008 0.04354 -0.0082 0.0027 1.0000 14.000 1.2989 0.05724 0.05097 -0.0081 0.0033 1.0000 14.500 1.2857 0.06532 0.05931 -0.0089 0.0036 1.0000 15.000 1.2709 0.07409 0.06837 -0.0104 0.0039 1.0000 15.500 1.2511 0.08373 0.07826 -0.0125 0.0041 1.0000 16.000 1.2334 0.09314 0.08788 -0.0148 0.0040 1.0000 16.500 1.2136 0.10305 0.09804 -0.0171 0.0046 1.0000 17.000 1.1986 0.11219 0.10737 -0.0194 0.0048 1.0000 17.500 1.1897 0.11987 0.11523 -0.0207 0.0051 1.0000 18.000 1.1922 0.12657 0.12207 -0.0228 0.0055 1.0000 18.500 1.1939 0.13329 0.12896 -0.0249 0.0058 1.0000