XFOIL Version 6.94 Calculated polar for: AG35 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 0.4275 0.00765 0.00196 -0.0469 0.6239 1.0000 0.500 0.4807 0.00811 0.00213 -0.0463 0.5899 1.0000 1.000 0.5347 0.00853 0.00240 -0.0460 0.5623 1.0000 1.500 0.5891 0.00890 0.00268 -0.0457 0.5362 1.0000 2.000 0.6434 0.00924 0.00297 -0.0455 0.5094 1.0000 2.500 0.6977 0.00954 0.00327 -0.0452 0.4787 1.0000 3.000 0.7517 0.00984 0.00356 -0.0450 0.4408 1.0000 3.500 0.8045 0.01028 0.00391 -0.0446 0.3927 1.0000 4.000 0.8561 0.01093 0.00441 -0.0442 0.3339 1.0000 4.500 0.9063 0.01181 0.00509 -0.0437 0.2748 1.0000 5.000 0.9559 0.01279 0.00592 -0.0432 0.2224 1.0000 5.500 1.0045 0.01387 0.00685 -0.0426 0.1751 1.0000 6.000 1.0523 0.01502 0.00789 -0.0420 0.1289 1.0000 6.500 1.0981 0.01639 0.00915 -0.0411 0.0874 1.0000 7.000 1.1393 0.01832 0.01093 -0.0397 0.0513 1.0000 7.500 1.1747 0.02083 0.01346 -0.0374 0.0325 1.0000 8.000 1.2076 0.02334 0.01615 -0.0348 0.0251 1.0000 8.500 1.2377 0.02589 0.01888 -0.0320 0.0214 1.0000 9.500 1.2765 0.03233 0.02583 -0.0244 0.0173 1.0000 10.000 1.2876 0.03600 0.02970 -0.0206 0.0162 1.0000 10.500 1.2892 0.04206 0.03609 -0.0169 0.0154 1.0000 11.000 1.2852 0.04757 0.04209 -0.0146 0.0151 1.0000 11.500 1.2712 0.05489 0.04989 -0.0142 0.0149 1.0000 12.000 1.2479 0.06423 0.05971 -0.0163 0.0148 1.0000 12.500 1.2174 0.07587 0.07178 -0.0213 0.0149 1.0000 13.000 1.1814 0.08948 0.08576 -0.0283 0.0150 1.0000 13.500 1.1418 0.10525 0.10185 -0.0372 0.0152 1.0000 14.000 1.1006 0.12319 0.12005 -0.0475 0.0154 1.0000