XFOIL Version 6.94 Calculated polar for: GOE 600 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 0.2193 0.00773 0.00241 -0.0476 0.7282 0.7123 0.500 0.2748 0.00761 0.00248 -0.0474 0.7164 0.7685 1.000 0.3279 0.00749 0.00259 -0.0464 0.7042 0.8419 1.500 0.3790 0.00744 0.00271 -0.0446 0.6912 0.9230 2.000 0.4460 0.00749 0.00278 -0.0465 0.6713 0.9701 2.500 0.5242 0.00757 0.00267 -0.0512 0.6165 0.9924 3.000 0.5909 0.00785 0.00268 -0.0539 0.5398 1.0000 3.500 0.6240 0.00938 0.00314 -0.0503 0.3414 1.0000 4.000 0.6660 0.01038 0.00360 -0.0483 0.2407 1.0000 4.500 0.7099 0.01135 0.00418 -0.0466 0.1621 1.0000 5.000 0.7451 0.01329 0.00531 -0.0438 0.0224 1.0000 5.500 0.7923 0.01412 0.00596 -0.0424 0.0058 1.0000 6.000 0.8429 0.01458 0.00648 -0.0417 0.0044 1.0000 6.500 0.8914 0.01525 0.00721 -0.0407 0.0042 1.0000 7.000 0.9385 0.01601 0.00811 -0.0394 0.0041 1.0000 7.500 0.9843 0.01685 0.00908 -0.0379 0.0041 1.0000 8.000 1.0276 0.01782 0.01021 -0.0360 0.0042 1.0000 8.500 1.0675 0.01898 0.01155 -0.0337 0.0043 1.0000 9.000 1.1021 0.02035 0.01315 -0.0306 0.0045 1.0000 9.500 1.1280 0.02199 0.01498 -0.0262 0.0047 1.0000 10.000 1.1501 0.02407 0.01726 -0.0220 0.0049 1.0000 10.500 1.1669 0.02672 0.02015 -0.0179 0.0051 1.0000 11.000 1.1781 0.03008 0.02377 -0.0141 0.0054 1.0000 11.500 1.1846 0.03417 0.02815 -0.0106 0.0057 1.0000 12.000 1.1876 0.03895 0.03324 -0.0075 0.0060 1.0000 12.500 1.1868 0.04451 0.03916 -0.0049 0.0063 1.0000 13.000 1.1793 0.05116 0.04619 -0.0029 0.0066 1.0000 13.500 1.1597 0.05970 0.05515 -0.0019 0.0069 1.0000 14.000 1.1369 0.06914 0.06497 -0.0026 0.0072 1.0000 14.500 1.1369 0.07591 0.07194 -0.0047 0.0075 1.0000 15.000 1.1249 0.08524 0.08154 -0.0084 0.0077 1.0000 15.500 1.1015 0.09755 0.09418 -0.0142 0.0080 1.0000 16.000 1.0675 0.11316 0.11016 -0.0227 0.0083 1.0000 16.500 1.0345 0.12987 0.12716 -0.0327 0.0082 1.0000 17.000 1.0010 0.14826 0.14578 -0.0442 0.0080 1.0000 17.500 0.9578 0.17201 0.16973 -0.0587 0.0077 1.0000