XFOIL Version 6.94 Calculated polar for: GOE 546 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 0.4929 0.00777 0.00321 -0.1101 0.8054 1.0000 0.500 0.5401 0.00780 0.00306 -0.1080 0.7820 1.0000 1.000 0.5883 0.00788 0.00297 -0.1061 0.7577 1.0000 1.500 0.6371 0.00801 0.00292 -0.1043 0.7328 1.0000 2.000 0.6838 0.00816 0.00294 -0.1020 0.7016 1.0000 2.500 0.7312 0.00841 0.00303 -0.0999 0.6683 1.0000 3.000 0.7784 0.00873 0.00323 -0.0979 0.6367 1.0000 3.500 0.8133 0.00927 0.00331 -0.0931 0.5503 1.0000 4.000 0.8464 0.01007 0.00359 -0.0885 0.4448 1.0000 4.500 0.8707 0.01165 0.00430 -0.0828 0.2862 1.0000 5.000 0.9015 0.01315 0.00520 -0.0786 0.1775 1.0000 5.500 0.9210 0.01546 0.00660 -0.0728 0.0352 1.0000 6.000 0.9612 0.01626 0.00741 -0.0700 0.0296 1.0000 6.500 1.0004 0.01704 0.00822 -0.0672 0.0250 1.0000 7.000 1.0371 0.01776 0.00900 -0.0638 0.0209 1.0000 7.500 1.0726 0.01853 0.00986 -0.0603 0.0171 1.0000 8.500 1.1291 0.02101 0.01245 -0.0516 0.0042 1.0000 9.000 1.1554 0.02242 0.01405 -0.0473 0.0042 1.0000 9.500 1.1788 0.02408 0.01590 -0.0429 0.0042 1.0000 10.000 1.1996 0.02602 0.01805 -0.0387 0.0043 1.0000 10.500 1.2186 0.02823 0.02048 -0.0346 0.0044 1.0000 11.000 1.2366 0.03064 0.02312 -0.0310 0.0045 1.0000 11.500 1.2523 0.03339 0.02610 -0.0276 0.0047 1.0000 12.000 1.2640 0.03665 0.02961 -0.0245 0.0049 1.0000 12.500 1.2701 0.04065 0.03388 -0.0218 0.0051 1.0000 13.000 1.2709 0.04551 0.03901 -0.0197 0.0053 1.0000 13.500 1.2664 0.05138 0.04515 -0.0185 0.0056 1.0000 14.000 1.2588 0.05816 0.05220 -0.0183 0.0057 1.0000 14.500 1.2475 0.06598 0.06026 -0.0190 0.0059 1.0000 15.000 1.2419 0.07358 0.06813 -0.0205 0.0060 1.0000 15.500 1.2365 0.08182 0.07677 -0.0229 0.0063 1.0000 16.000 1.2203 0.09237 0.08772 -0.0266 0.0066 1.0000 16.500 1.1987 0.10464 0.10037 -0.0323 0.0067 1.0000 17.000 1.1706 0.11919 0.11532 -0.0400 0.0069 1.0000 17.500 1.1411 0.13523 0.13171 -0.0494 0.0069 1.0000 18.000 1.1083 0.15335 0.15015 -0.0605 0.0069 1.0000 18.500 1.0757 0.17289 0.16991 -0.0724 0.0067 1.0000