XFOIL Version 6.94 Calculated polar for: GOE 670 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 0.3767 0.00977 0.00330 -0.0832 0.8490 0.2539 0.500 0.5313 0.00781 0.00329 -0.1035 0.8318 1.0000 1.000 0.5771 0.00776 0.00308 -0.1010 0.8024 1.0000 1.500 0.6225 0.00778 0.00296 -0.0984 0.7691 1.0000 2.000 0.6673 0.00789 0.00290 -0.0957 0.7317 1.0000 2.500 0.7019 0.00838 0.00281 -0.0905 0.6407 1.0000 3.000 0.7353 0.00901 0.00301 -0.0856 0.5564 1.0000 3.500 0.7501 0.01047 0.00343 -0.0774 0.3604 1.0000 4.000 0.7823 0.01151 0.00396 -0.0729 0.2704 1.0000 4.500 0.8162 0.01252 0.00454 -0.0688 0.2014 1.0000 5.000 0.8441 0.01406 0.00529 -0.0640 0.0792 1.0000 5.500 0.8785 0.01525 0.00619 -0.0600 0.0221 1.0000 6.000 0.9197 0.01594 0.00691 -0.0572 0.0080 1.0000 6.500 0.9607 0.01666 0.00773 -0.0543 0.0071 1.0000 7.000 1.0008 0.01741 0.00860 -0.0513 0.0069 1.0000 7.500 1.0379 0.01835 0.00977 -0.0477 0.0068 1.0000 8.000 1.0700 0.01951 0.01115 -0.0434 0.0069 1.0000 8.500 1.0979 0.02097 0.01287 -0.0385 0.0070 1.0000 9.000 1.1193 0.02285 0.01500 -0.0330 0.0072 1.0000 9.500 1.1324 0.02533 0.01773 -0.0267 0.0074 1.0000 10.000 1.1421 0.02829 0.02093 -0.0208 0.0075 1.0000 10.500 1.1521 0.03170 0.02463 -0.0156 0.0077 1.0000 11.000 1.1624 0.03630 0.02952 -0.0109 0.0081 1.0000 11.500 1.1792 0.03950 0.03295 -0.0074 0.0084 1.0000 12.000 1.1930 0.04315 0.03689 -0.0042 0.0089 1.0000 12.500 1.1979 0.04912 0.04339 -0.0007 0.0100 1.0000 13.000 1.1758 0.05850 0.05344 0.0024 0.0117 1.0000 13.500 1.1449 0.06802 0.06343 0.0024 0.0124 1.0000 14.000 1.1121 0.07858 0.07438 -0.0005 0.0126 1.0000 14.500 1.0765 0.09120 0.08735 -0.0063 0.0127 1.0000 15.000 1.0411 0.10604 0.10251 -0.0153 0.0127 1.0000 15.500 1.0073 0.12306 0.11981 -0.0265 0.0124 1.0000 16.000 0.9705 0.14322 0.14019 -0.0395 0.0120 1.0000 16.500 0.9222 0.17117 0.16822 -0.0545 0.0126 1.0000