XFOIL Version 6.94 Calculated polar for: SA7035 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 0.3426 0.00691 0.00243 -0.0700 0.8216 1.0000 0.500 0.3884 0.00695 0.00225 -0.0676 0.7867 1.0000 1.000 0.4365 0.00708 0.00215 -0.0657 0.7501 1.0000 1.500 0.4862 0.00730 0.00214 -0.0642 0.7120 1.0000 2.000 0.5366 0.00757 0.00221 -0.0629 0.6714 1.0000 2.500 0.5874 0.00789 0.00234 -0.0617 0.6280 1.0000 3.000 0.6383 0.00827 0.00256 -0.0606 0.5845 1.0000 3.500 0.6891 0.00870 0.00281 -0.0595 0.5414 1.0000 4.000 0.7396 0.00918 0.00316 -0.0585 0.4965 1.0000 4.500 0.7897 0.00971 0.00355 -0.0574 0.4500 1.0000 5.000 0.8378 0.01037 0.00397 -0.0561 0.3877 1.0000 5.500 0.8847 0.01117 0.00447 -0.0548 0.3142 1.0000 6.000 0.9308 0.01213 0.00517 -0.0535 0.2419 1.0000 6.500 0.9754 0.01329 0.00598 -0.0521 0.1706 1.0000 7.000 1.0208 0.01439 0.00695 -0.0507 0.1235 1.0000 7.500 1.0581 0.01635 0.00842 -0.0483 0.0436 1.0000 8.000 1.0935 0.01851 0.01045 -0.0454 0.0085 1.0000 8.500 1.1322 0.02016 0.01252 -0.0427 0.0076 1.0000 9.000 1.1631 0.02232 0.01503 -0.0391 0.0073 1.0000 9.500 1.1827 0.02505 0.01810 -0.0341 0.0071 1.0000 10.000 1.1905 0.02822 0.02159 -0.0278 0.0071 1.0000 10.500 1.1970 0.03211 0.02583 -0.0224 0.0073 1.0000 11.000 1.2008 0.03695 0.03109 -0.0179 0.0074 1.0000 11.500 1.1979 0.04278 0.03739 -0.0144 0.0076 1.0000 12.000 1.1842 0.04996 0.04506 -0.0124 0.0078 1.0000 12.500 1.1607 0.05890 0.05446 -0.0130 0.0080 1.0000 13.000 1.1311 0.06980 0.06579 -0.0167 0.0081 1.0000 13.500 1.0973 0.08303 0.07941 -0.0236 0.0082 1.0000 14.000 1.0621 0.09868 0.09540 -0.0332 0.0083 1.0000 14.500 1.0271 0.11652 0.11352 -0.0445 0.0084 1.0000 15.000 0.9891 0.13769 0.13492 -0.0576 0.0085 1.0000