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 = 0.450 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.5108 0.01209 0.00370 -0.0812 0.5047 0.0325 0.500 0.5546 0.01205 0.00372 -0.0786 0.4968 0.0400 1.000 0.5968 0.01201 0.00363 -0.0757 0.4880 0.0458 1.500 0.6396 0.01201 0.00361 -0.0731 0.4790 0.0573 2.000 0.9471 0.01110 0.00477 -0.1306 0.4614 1.0000 2.500 0.9902 0.01127 0.00490 -0.1282 0.4541 1.0000 3.000 1.0323 0.01146 0.00503 -0.1256 0.4460 1.0000 3.500 1.0738 0.01180 0.00525 -0.1229 0.4372 1.0000 4.500 1.1159 0.01234 0.00543 -0.1089 0.3607 1.0000 5.000 1.1400 0.01294 0.00583 -0.1028 0.3240 1.0000 5.500 1.1416 0.01450 0.00684 -0.0928 0.2449 1.0000 6.000 1.1614 0.01545 0.00765 -0.0864 0.2205 1.0000 6.500 1.1753 0.01670 0.00871 -0.0793 0.1871 1.0000 7.500 1.1411 0.02243 0.01380 -0.0576 0.0044 1.0000 8.000 1.1612 0.02405 0.01552 -0.0535 0.0043 1.0000 8.500 1.1788 0.02605 0.01766 -0.0497 0.0045 1.0000 9.000 1.1938 0.02849 0.02024 -0.0462 0.0047 1.0000 9.500 1.2069 0.03136 0.02327 -0.0432 0.0050 1.0000 10.000 1.2167 0.03480 0.02687 -0.0407 0.0053 1.0000 10.500 1.2213 0.03897 0.03124 -0.0384 0.0058 1.0000 11.000 1.2249 0.04349 0.03595 -0.0365 0.0064 1.0000 11.500 1.2252 0.04848 0.04112 -0.0348 0.0069 1.0000 12.000 1.2171 0.05463 0.04749 -0.0334 0.0075 1.0000 12.500 1.2142 0.06046 0.05351 -0.0325 0.0084 1.0000 13.000 1.2034 0.06751 0.06077 -0.0319 0.0091 1.0000 13.500 1.1818 0.07625 0.06971 -0.0318 0.0096 1.0000 14.000 1.1788 0.08285 0.07651 -0.0318 0.0106 1.0000 14.500 1.1591 0.09189 0.08575 -0.0324 0.0114 1.0000 15.000 1.1491 0.09983 0.09389 -0.0330 0.0128 1.0000 15.500 1.1339 0.10841 0.10261 -0.0338 0.0141 1.0000 16.000 1.1380 0.11408 0.10843 -0.0338 0.0161 1.0000