XFOIL Version 6.94 Calculated polar for: GOE 711 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 0.6835 0.01201 0.00410 -0.1348 0.5697 0.0366 0.500 0.7348 0.01220 0.00412 -0.1342 0.5378 0.0412 1.000 0.7858 0.01240 0.00420 -0.1336 0.5063 0.0496 1.500 0.8359 0.01258 0.00433 -0.1329 0.4777 0.0752 2.000 0.8782 0.01085 0.00469 -0.1308 0.4539 1.0000 2.500 0.9272 0.01134 0.00493 -0.1298 0.4316 1.0000 3.000 0.9753 0.01184 0.00524 -0.1288 0.4119 1.0000 3.500 1.0214 0.01243 0.00562 -0.1274 0.3944 1.0000 4.000 1.0703 0.01285 0.00597 -0.1266 0.3819 1.0000 4.500 1.1151 0.01345 0.00648 -0.1251 0.3691 1.0000 5.000 1.1501 0.01410 0.00687 -0.1216 0.3325 1.0000 5.500 1.1852 0.01497 0.00749 -0.1184 0.2978 1.0000 6.000 1.2184 0.01607 0.00828 -0.1151 0.2573 1.0000 6.500 1.2505 0.01731 0.00926 -0.1117 0.2187 1.0000 7.000 1.2312 0.02180 0.01277 -0.1006 0.0576 1.0000 8.000 1.2881 0.02514 0.01608 -0.0941 0.0041 1.0000 8.500 1.3194 0.02668 0.01772 -0.0916 0.0042 1.0000 9.000 1.3480 0.02848 0.01965 -0.0889 0.0044 1.0000 9.500 1.3738 0.03054 0.02186 -0.0861 0.0047 1.0000 10.000 1.3974 0.03289 0.02436 -0.0834 0.0051 1.0000 10.500 1.4159 0.03578 0.02745 -0.0805 0.0057 1.0000 11.000 1.4297 0.03928 0.03114 -0.0777 0.0061 1.0000 11.500 1.4436 0.04306 0.03511 -0.0754 0.0069 1.0000 12.000 1.4431 0.04862 0.04091 -0.0729 0.0075 1.0000 12.500 1.4516 0.05356 0.04605 -0.0716 0.0084 1.0000 13.000 1.4401 0.06124 0.05399 -0.0703 0.0092 1.0000 13.500 1.4332 0.06876 0.06175 -0.0699 0.0100 1.0000 14.000 1.4147 0.07825 0.07154 -0.0702 0.0109 1.0000 14.500 1.3823 0.09008 0.08366 -0.0713 0.0114 1.0000