XFOIL Version 6.94 Calculated polar for: GOE 373 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 0.5154 0.01410 0.00774 -0.0972 0.6383 0.0399 2.000 0.7230 0.01109 0.00379 -0.0937 0.5111 0.0868 2.500 0.7761 0.01056 0.00303 -0.0921 0.4874 0.0486 3.000 0.8269 0.01057 0.00283 -0.0909 0.4445 0.0430 3.500 0.8734 0.01099 0.00279 -0.0893 0.3509 0.0430 4.000 0.9192 0.01176 0.00313 -0.0877 0.2606 0.0458 4.500 0.9599 0.01331 0.00393 -0.0854 0.1274 0.0496 6.000 1.1002 0.01497 0.00684 -0.0811 0.0091 1.0000 6.500 1.1474 0.01575 0.00777 -0.0793 0.0112 1.0000 7.000 1.1921 0.01677 0.00895 -0.0769 0.0132 1.0000 7.500 1.2325 0.01814 0.01050 -0.0740 0.0152 1.0000 8.000 1.2686 0.01971 0.01222 -0.0702 0.0177 1.0000 8.500 1.2960 0.02170 0.01436 -0.0651 0.0209 1.0000 9.000 1.3164 0.02379 0.01660 -0.0591 0.0219 1.0000 9.500 1.3279 0.02654 0.01943 -0.0522 0.0219 1.0000 10.000 1.3563 0.03033 0.02326 -0.0485 0.0205 1.0000 10.500 1.3916 0.03435 0.02742 -0.0463 0.0171 1.0000 11.500 1.4210 0.04561 0.03915 -0.0403 0.0082 1.0000 12.000 1.3884 0.04907 0.04311 -0.0332 0.0075 1.0000 12.500 1.3837 0.05639 0.05074 -0.0312 0.0063 1.0000 13.000 1.3521 0.06289 0.05770 -0.0299 0.0060 1.0000 13.500 1.3256 0.07128 0.06649 -0.0315 0.0056 1.0000 14.000 1.3011 0.08119 0.07676 -0.0349 0.0055 1.0000 14.500 1.2750 0.09260 0.08851 -0.0403 0.0054 1.0000 15.000 1.2476 0.10561 0.10186 -0.0474 0.0055 1.0000 15.500 1.2201 0.12007 0.11664 -0.0562 0.0056 1.0000 16.000 1.1957 0.13518 0.13203 -0.0656 0.0059 1.0000 16.500 1.1768 0.14973 0.14677 -0.0746 0.0062 1.0000 17.000 1.1657 0.16276 0.15992 -0.0823 0.0064 1.0000 17.500 1.1577 0.17548 0.17274 -0.0899 0.0065 1.0000 18.000 1.1501 0.18881 0.18618 -0.0981 0.0065 1.0000 18.500 0.9389 0.24206 0.24024 -0.1305 0.0167 1.0000 19.000 0.9444 0.25118 0.24937 -0.1346 0.0150 1.0000 19.500 0.9514 0.25922 0.25742 -0.1381 0.0128 1.0000 20.000 0.9534 0.27163 0.26984 -0.1433 0.0112 1.0000 20.500 0.9579 0.28238 0.28060 -0.1476 0.0096 1.0000 21.000 0.9640 0.29229 0.29052 -0.1515 0.0092 1.0000 21.500 0.9718 0.30017 0.29843 -0.1544 0.0084 1.0000 22.000 0.9754 0.31301 0.31128 -0.1591 0.0077 1.0000 22.500 0.9815 0.32313 0.32142 -0.1628 0.0073 1.0000 23.000 0.9889 0.33101 0.32932 -0.1657 0.0067 1.0000 23.500 0.9957 0.34017 0.33850 -0.1686 0.0065 1.0000 24.000 0.9992 0.35288 0.35122 -0.1730 0.0064 1.0000 24.500 1.0043 0.36414 0.36251 -0.1767 0.0063 1.0000 25.000 1.0094 0.37538 0.37376 -0.1803 0.0058 1.0000 25.500 1.0145 0.38578 0.38418 -0.1836 0.0056 1.0000 26.000 1.0194 0.39632 0.39475 -0.1869 0.0055 1.0000 26.500 1.0239 0.40684 0.40529 -0.1902 0.0053 1.0000 27.000 1.0283 0.41714 0.41562 -0.1934 0.0051 1.0000