XFOIL Version 6.94 Calculated polar for: GOE 573 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 0.5262 0.01232 0.00378 -0.0824 0.4751 0.0335 0.500 0.5744 0.01234 0.00383 -0.0810 0.4692 0.0416 1.000 0.6215 0.01235 0.00382 -0.0793 0.4632 0.0493 1.500 0.6713 0.01246 0.00396 -0.0783 0.4561 0.0801 2.500 1.0031 0.01185 0.00534 -0.1282 0.4443 1.0000 3.000 1.0477 0.01213 0.00554 -0.1263 0.4396 1.0000 3.500 1.0959 0.01265 0.00594 -0.1252 0.4333 1.0000 4.000 1.1186 0.01251 0.00572 -0.1185 0.4113 1.0000 4.500 1.1454 0.01259 0.00570 -0.1127 0.3894 1.0000 5.000 1.1810 0.01275 0.00588 -0.1090 0.3754 1.0000 5.500 1.2009 0.01313 0.00609 -0.1020 0.3459 1.0000 6.000 1.2198 0.01363 0.00647 -0.0950 0.3147 1.0000 6.500 1.2050 0.01538 0.00771 -0.0822 0.2318 1.0000 7.000 1.1880 0.01756 0.00953 -0.0701 0.1629 1.0000 7.500 1.1410 0.02150 0.01305 -0.0552 0.0517 1.0000 8.000 1.1355 0.02440 0.01591 -0.0482 0.0041 1.0000 8.500 1.1501 0.02656 0.01818 -0.0443 0.0040 1.0000 9.000 1.1623 0.02920 0.02095 -0.0409 0.0041 1.0000 9.500 1.1713 0.03240 0.02430 -0.0378 0.0043 1.0000 10.000 1.1771 0.03617 0.02826 -0.0352 0.0045 1.0000 10.500 1.1806 0.04041 0.03266 -0.0330 0.0047 1.0000 11.000 1.1820 0.04507 0.03749 -0.0311 0.0051 1.0000 11.500 1.1789 0.05039 0.04301 -0.0295 0.0054 1.0000 12.000 1.1718 0.05641 0.04922 -0.0282 0.0057 1.0000 12.500 1.1664 0.06249 0.05547 -0.0272 0.0061 1.0000 13.000 1.1633 0.06855 0.06172 -0.0266 0.0067 1.0000 13.500 1.1521 0.07578 0.06913 -0.0263 0.0072 1.0000 14.000 1.1376 0.08366 0.07719 -0.0263 0.0075 1.0000 14.500 1.1386 0.08970 0.08338 -0.0264 0.0084 1.0000 15.000 1.1255 0.09778 0.09163 -0.0268 0.0093 1.0000 15.500 1.1180 0.10511 0.09908 -0.0274 0.0098 1.0000 16.000 1.1164 0.11156 0.10568 -0.0276 0.0116 1.0000 16.500 1.1365 0.11403 0.10820 -0.0260 0.0148 1.0000