XFOIL Version 6.94 Calculated polar for: HQ 3.0/10 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 0.4038 0.00868 0.00370 -0.0890 0.8204 0.8130 0.500 0.4561 0.00847 0.00347 -0.0878 0.8018 0.8358 1.000 0.5079 0.00832 0.00336 -0.0866 0.7849 0.8645 1.500 0.5585 0.00811 0.00323 -0.0850 0.7663 0.9069 2.000 0.6297 0.00794 0.00305 -0.0880 0.7425 1.0000 2.500 0.6858 0.00807 0.00305 -0.0882 0.7183 1.0000 3.000 0.7404 0.00824 0.00314 -0.0879 0.6912 1.0000 3.500 0.7930 0.00844 0.00330 -0.0872 0.6571 1.0000 4.000 0.8423 0.00876 0.00344 -0.0857 0.6017 1.0000 4.500 0.8830 0.00958 0.00369 -0.0827 0.4866 1.0000 5.000 0.9167 0.01108 0.00440 -0.0791 0.3437 1.0000 5.500 0.9591 0.01213 0.00513 -0.0771 0.2726 1.0000 6.000 0.9975 0.01348 0.00601 -0.0746 0.1867 1.0000 6.500 1.0377 0.01469 0.00692 -0.0723 0.1278 1.0000 7.000 1.0731 0.01625 0.00809 -0.0695 0.0629 1.0000 7.500 1.1062 0.01793 0.00953 -0.0661 0.0200 1.0000 8.000 1.1371 0.01966 0.01126 -0.0623 0.0045 1.0000 8.500 1.1673 0.02107 0.01289 -0.0583 0.0041 1.0000 9.000 1.1918 0.02277 0.01485 -0.0535 0.0040 1.0000 9.500 1.2113 0.02484 0.01721 -0.0485 0.0039 1.0000 10.000 1.2246 0.02738 0.02001 -0.0433 0.0039 1.0000 10.500 1.2331 0.03038 0.02325 -0.0383 0.0040 1.0000 11.000 1.2385 0.03391 0.02702 -0.0340 0.0040 1.0000 11.500 1.2427 0.03793 0.03131 -0.0304 0.0041 1.0000 12.000 1.2448 0.04266 0.03633 -0.0275 0.0043 1.0000 12.500 1.2438 0.04842 0.04245 -0.0253 0.0045 1.0000 13.000 1.2384 0.05529 0.04975 -0.0242 0.0048 1.0000 13.500 1.2254 0.06359 0.05851 -0.0246 0.0050 1.0000 14.000 1.2048 0.07371 0.06907 -0.0271 0.0052 1.0000 14.500 1.1802 0.08556 0.08132 -0.0319 0.0053 1.0000 15.000 1.1536 0.09917 0.09530 -0.0392 0.0054 1.0000 15.500 1.1258 0.11452 0.11099 -0.0483 0.0054 1.0000 16.000 1.0988 0.13111 0.12787 -0.0588 0.0055 1.0000 16.500 1.0730 0.14867 0.14568 -0.0700 0.0055 1.0000 17.000 1.0491 0.16689 0.16408 -0.0813 0.0056 1.0000