XFOIL Version 6.94 Calculated polar for: GOE 510 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 0.4909 0.01158 0.00392 -0.0922 0.5974 0.0705 0.500 0.5200 0.01138 0.00367 -0.0864 0.5679 0.0734 1.000 0.5487 0.01144 0.00359 -0.0805 0.5347 0.0767 1.500 0.5764 0.01159 0.00361 -0.0746 0.5036 0.0824 2.000 0.6088 0.01173 0.00369 -0.0697 0.4793 0.0995 2.500 0.6406 0.01174 0.00406 -0.0649 0.4611 0.2506 3.000 0.6803 0.01200 0.00435 -0.0617 0.4473 0.2899 3.500 0.7189 0.01237 0.00464 -0.0583 0.4338 0.3162 4.000 0.7593 0.01250 0.00490 -0.0554 0.4245 0.3482 5.000 1.0211 0.01317 0.00660 -0.0895 0.3476 1.0000 5.500 1.0379 0.01372 0.00694 -0.0822 0.3171 1.0000 6.000 1.0564 0.01435 0.00738 -0.0753 0.2865 1.0000 6.500 1.0685 0.01539 0.00812 -0.0677 0.2419 1.0000 7.000 1.0827 0.01658 0.00904 -0.0608 0.2008 1.0000 8.000 1.0566 0.02245 0.01391 -0.0410 0.0055 1.0000 8.500 1.0826 0.02373 0.01523 -0.0373 0.0051 1.0000 9.000 1.1076 0.02517 0.01674 -0.0338 0.0049 1.0000 9.500 1.1308 0.02685 0.01850 -0.0303 0.0048 1.0000 10.000 1.1521 0.02878 0.02054 -0.0270 0.0048 1.0000 10.500 1.1708 0.03105 0.02293 -0.0238 0.0049 1.0000 11.000 1.1864 0.03373 0.02574 -0.0208 0.0049 1.0000 11.500 1.1983 0.03693 0.02911 -0.0181 0.0050 1.0000 12.000 1.2068 0.04067 0.03302 -0.0157 0.0051 1.0000 12.500 1.2116 0.04503 0.03757 -0.0137 0.0052 1.0000 13.000 1.2107 0.05030 0.04303 -0.0122 0.0052 1.0000 13.500 1.2033 0.05662 0.04956 -0.0113 0.0054 1.0000 14.000 1.1904 0.06400 0.05716 -0.0110 0.0055 1.0000 14.500 1.1749 0.07215 0.06552 -0.0115 0.0055 1.0000 15.000 1.1531 0.08152 0.07511 -0.0127 0.0056 1.0000 15.500 1.1354 0.09061 0.08440 -0.0142 0.0057 1.0000 16.000 1.1148 0.10026 0.09422 -0.0161 0.0057 1.0000 16.500 1.1100 0.10777 0.10187 -0.0176 0.0059 1.0000 17.000 1.1041 0.11552 0.10977 -0.0195 0.0060 1.0000 17.500 1.1014 0.12281 0.11720 -0.0214 0.0063 1.0000 18.000 1.0997 0.12987 0.12438 -0.0233 0.0065 1.0000 18.500 1.1026 0.13597 0.13056 -0.0248 0.0070 1.0000 19.000 1.1139 0.14023 0.13484 -0.0256 0.0073 1.0000 19.500 1.1339 0.14296 0.13756 -0.0257 0.0079 1.0000