XFOIL Version 6.94 Calculated polar for: GOE 562 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.300 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.5454 0.01432 0.00529 -0.0896 0.5286 0.0453 0.500 0.5854 0.01408 0.00509 -0.0862 0.5195 0.0511 1.000 0.6305 0.01395 0.00493 -0.0840 0.5109 0.0616 1.500 0.9067 0.01266 0.00573 -0.1339 0.4937 1.0000 2.000 0.9507 0.01291 0.00581 -0.1317 0.4854 1.0000 2.500 0.9940 0.01319 0.00602 -0.1294 0.4776 1.0000 3.000 1.0349 0.01340 0.00619 -0.1265 0.4689 1.0000 3.500 1.0798 0.01378 0.00639 -0.1246 0.4599 1.0000 5.000 1.1436 0.01424 0.00662 -0.1035 0.3828 1.0000 5.500 1.1609 0.01478 0.00696 -0.0961 0.3469 1.0000 6.000 1.1763 0.01564 0.00759 -0.0887 0.3036 1.0000 6.500 1.1766 0.01723 0.00879 -0.0791 0.2476 1.0000 7.000 1.1899 0.01855 0.00998 -0.0722 0.2233 1.0000 7.500 1.2007 0.02012 0.01144 -0.0654 0.1949 1.0000 8.000 1.1864 0.02316 0.01413 -0.0561 0.1227 1.0000 8.500 1.1464 0.02872 0.01940 -0.0464 0.0210 1.0000 9.000 1.1550 0.03184 0.02261 -0.0429 0.0056 1.0000 9.500 1.1680 0.03491 0.02585 -0.0404 0.0054 1.0000 10.000 1.1783 0.03845 0.02958 -0.0381 0.0054 1.0000 10.500 1.1842 0.04263 0.03397 -0.0362 0.0056 1.0000 11.000 1.1857 0.04743 0.03901 -0.0345 0.0058 1.0000 11.500 1.1824 0.05296 0.04477 -0.0331 0.0059 1.0000 12.000 1.1771 0.05897 0.05100 -0.0320 0.0062 1.0000 12.500 1.1732 0.06508 0.05733 -0.0313 0.0066 1.0000 13.000 1.1640 0.07210 0.06459 -0.0310 0.0068 1.0000 13.500 1.1517 0.07981 0.07252 -0.0311 0.0072 1.0000 14.000 1.1368 0.08812 0.08105 -0.0316 0.0074 1.0000 14.500 1.1186 0.09720 0.09033 -0.0325 0.0076 1.0000 15.000 1.1072 0.10544 0.09874 -0.0335 0.0080 1.0000 15.500 1.1016 0.11296 0.10644 -0.0345 0.0086 1.0000 16.000 1.0948 0.12062 0.11424 -0.0357 0.0088 1.0000 16.500 1.0944 0.12684 0.12047 -0.0363 0.0096 1.0000 17.000 1.1079 0.13109 0.12486 -0.0364 0.0105 1.0000 17.500 1.1703 0.12469 0.11831 -0.0298 0.0145 1.0000