XFOIL Version 6.94 Calculated polar for: AG37 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 0.3843 0.00720 0.00187 -0.0459 0.6879 1.0000 0.500 0.4366 0.00763 0.00191 -0.0449 0.6213 1.0000 1.000 0.4900 0.00812 0.00213 -0.0443 0.5781 1.0000 1.500 0.5440 0.00856 0.00240 -0.0440 0.5418 1.0000 2.000 0.5983 0.00895 0.00269 -0.0436 0.5052 1.0000 2.500 0.6524 0.00932 0.00298 -0.0433 0.4647 1.0000 3.500 0.7590 0.01026 0.00367 -0.0427 0.3607 1.0000 4.000 0.8114 0.01092 0.00419 -0.0423 0.3067 1.0000 4.500 0.8630 0.01171 0.00482 -0.0419 0.2495 1.0000 5.000 0.9134 0.01269 0.00560 -0.0415 0.1821 1.0000 5.500 0.9607 0.01421 0.00675 -0.0408 0.1054 1.0000 6.000 1.0064 0.01598 0.00836 -0.0397 0.0661 1.0000 6.500 1.0507 0.01783 0.01027 -0.0382 0.0505 1.0000 7.000 1.0930 0.01985 0.01242 -0.0366 0.0416 1.0000 7.500 1.1313 0.02243 0.01509 -0.0345 0.0352 1.0000 8.000 1.1732 0.02437 0.01730 -0.0328 0.0307 1.0000 8.500 1.2068 0.02799 0.02116 -0.0304 0.0272 1.0000 9.000 1.2417 0.03073 0.02430 -0.0280 0.0242 1.0000 9.500 1.2677 0.03485 0.02864 -0.0254 0.0217 1.0000 10.000 1.2799 0.03979 0.03426 -0.0213 0.0203 1.0000 10.500 1.2731 0.04511 0.04017 -0.0157 0.0193 1.0000 11.000 1.2532 0.05137 0.04692 -0.0121 0.0186 1.0000 11.500 1.2280 0.05964 0.05563 -0.0130 0.0182 1.0000 12.000 1.1875 0.07266 0.06913 -0.0197 0.0183 1.0000 12.500 1.1267 0.09243 0.08936 -0.0324 0.0192 1.0000 13.000 1.0653 0.11530 0.11249 -0.0467 0.0203 1.0000 13.500 1.0066 0.14062 0.13791 -0.0605 0.0214 1.0000