XFOIL Version 6.94 Calculated polar for: S7075 (9%) 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 0.4170 0.00807 0.00384 -0.0958 0.8849 1.0000 0.500 0.4684 0.00783 0.00347 -0.0944 0.8671 1.0000 1.000 0.5199 0.00767 0.00321 -0.0929 0.8452 1.0000 1.500 0.5737 0.00752 0.00295 -0.0919 0.8189 1.0000 2.000 0.6272 0.00746 0.00279 -0.0908 0.7847 1.0000 2.500 0.6798 0.00755 0.00271 -0.0896 0.7401 1.0000 3.000 0.7312 0.00786 0.00278 -0.0882 0.6864 1.0000 3.500 0.7805 0.00834 0.00300 -0.0866 0.6263 1.0000 4.000 0.8282 0.00893 0.00334 -0.0848 0.5642 1.0000 4.500 0.8744 0.00962 0.00376 -0.0828 0.4957 1.0000 5.000 0.9193 0.01042 0.00427 -0.0808 0.4241 1.0000 5.500 0.9627 0.01137 0.00491 -0.0787 0.3478 1.0000 6.000 1.0055 0.01244 0.00565 -0.0766 0.2727 1.0000 6.500 1.0474 0.01364 0.00655 -0.0745 0.2032 1.0000 7.000 1.0879 0.01500 0.00760 -0.0722 0.1397 1.0000 7.500 1.1263 0.01657 0.00895 -0.0696 0.0903 1.0000 8.000 1.1606 0.01843 0.01069 -0.0663 0.0639 1.0000 8.500 1.1929 0.02033 0.01264 -0.0627 0.0523 1.0000 9.000 1.2263 0.02205 0.01456 -0.0592 0.0475 1.0000 9.500 1.2552 0.02393 0.01660 -0.0552 0.0436 1.0000 10.000 1.2812 0.02727 0.02008 -0.0513 0.0399 1.0000 10.500 1.3133 0.02981 0.02291 -0.0482 0.0381 1.0000 11.000 1.3425 0.03253 0.02597 -0.0450 0.0362 1.0000 11.500 1.3623 0.03474 0.02842 -0.0409 0.0335 1.0000 12.000 1.3769 0.03728 0.03110 -0.0370 0.0308 1.0000 13.500 1.3869 0.04978 0.04468 -0.0271 0.0251 1.0000 14.000 1.3837 0.05475 0.04995 -0.0258 0.0236 1.0000 14.500 1.3771 0.06074 0.05613 -0.0259 0.0221 1.0000 15.000 1.3398 0.07256 0.06839 -0.0284 0.0210 1.0000 15.500 1.3123 0.08394 0.08019 -0.0337 0.0203 1.0000 16.000 1.2793 0.09816 0.09482 -0.0417 0.0198 1.0000 16.500 1.2272 0.11833 0.11538 -0.0540 0.0202 1.0000