XFOIL Version 6.94 Calculated polar for: NASA/LANGLEY LS(1)-0413 (GA(W)- 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 0.4828 0.01193 0.00693 -0.1087 0.8421 0.7740 0.500 0.5404 0.01181 0.00677 -0.1086 0.8263 0.7844 1.000 0.5951 0.01153 0.00648 -0.1076 0.8095 0.7909 1.500 0.6518 0.01128 0.00620 -0.1071 0.7895 0.7972 2.000 0.7114 0.01112 0.00600 -0.1076 0.7700 0.8040 2.500 0.7670 0.01095 0.00587 -0.1072 0.7459 0.8097 3.000 0.8211 0.01086 0.00583 -0.1064 0.7118 0.8154 3.500 0.8751 0.01092 0.00577 -0.1057 0.6548 0.8229 4.000 0.9149 0.01188 0.00597 -0.1024 0.4959 0.8291 4.500 0.9431 0.01384 0.00700 -0.0977 0.3187 0.8363 5.500 1.0200 0.01667 0.00898 -0.0926 0.1658 0.8514 6.500 1.1005 0.01889 0.01104 -0.0878 0.1237 0.8693 7.000 1.1384 0.01999 0.01216 -0.0849 0.1114 0.8796 7.500 1.1723 0.02112 0.01332 -0.0814 0.1016 0.8904 8.000 1.2069 0.02218 0.01449 -0.0781 0.0933 0.9035 9.000 1.2669 0.02473 0.01726 -0.0703 0.0794 0.9428 9.500 1.2974 0.02627 0.01882 -0.0672 0.0733 1.0000 10.000 1.3370 0.02792 0.02059 -0.0663 0.0670 1.0000 12.000 1.4539 0.03665 0.02976 -0.0578 0.0479 1.0000 12.500 1.4744 0.03961 0.03295 -0.0551 0.0447 1.0000 13.000 1.4921 0.04273 0.03615 -0.0526 0.0422 1.0000 13.500 1.5050 0.04646 0.04017 -0.0501 0.0396 1.0000 14.000 1.5159 0.05029 0.04413 -0.0481 0.0375 1.0000 14.500 1.5202 0.05509 0.04921 -0.0463 0.0355 1.0000 15.000 1.5227 0.06029 0.05467 -0.0452 0.0337 1.0000 15.500 1.5230 0.06601 0.06051 -0.0446 0.0325 1.0000 16.000 1.5138 0.07347 0.06834 -0.0448 0.0315 1.0000 16.500 1.4999 0.08219 0.07744 -0.0464 0.0305 1.0000 17.000 1.4858 0.09156 0.08710 -0.0495 0.0295 1.0000 17.500 1.4709 0.10146 0.09722 -0.0535 0.0288 1.0000 18.000 1.4573 0.11132 0.10724 -0.0578 0.0281 1.0000 18.500 1.4234 0.12612 0.12245 -0.0662 0.0278 1.0000 19.000 1.3802 0.14410 0.14086 -0.0779 0.0276 1.0000 19.500 1.3071 0.17139 0.16865 -0.0975 0.0278 1.0000