XFOIL Version 6.94 Calculated polar for: NREL S-809 WIND TURBINE 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 0.1333 0.02336 0.01721 -0.0355 0.6373 0.5468 0.500 0.1839 0.02328 0.01718 -0.0362 0.6359 0.5501 1.000 0.2360 0.02313 0.01704 -0.0370 0.6328 0.5537 1.500 0.2875 0.02317 0.01714 -0.0377 0.6309 0.5574 2.000 0.3394 0.02325 0.01728 -0.0386 0.6285 0.5609 2.500 0.3910 0.02372 0.01782 -0.0391 0.6232 0.5649 3.000 0.4276 0.02577 0.02000 -0.0381 0.6168 0.5680 3.500 0.4836 0.02517 0.01947 -0.0393 0.6141 0.5701 4.000 0.5562 0.02222 0.01645 -0.0411 0.5965 0.5718 4.500 0.6287 0.01990 0.01379 -0.0426 0.5668 0.5732 5.000 0.6875 0.01863 0.01260 -0.0434 0.5566 0.5744 5.500 0.7430 0.01834 0.01267 -0.0441 0.5493 0.5753 6.000 0.8013 0.01758 0.01202 -0.0448 0.5340 0.5762 6.500 0.8573 0.01622 0.01066 -0.0456 0.5028 0.5790 7.000 0.8801 0.01600 0.00960 -0.0406 0.3565 0.5837 7.500 0.8586 0.01845 0.01123 -0.0296 0.2122 0.5873 8.000 0.8332 0.02168 0.01383 -0.0202 0.1052 0.5903 8.500 0.8070 0.02652 0.01813 -0.0141 0.0220 0.5937 9.000 0.8166 0.02982 0.02173 -0.0116 0.0154 0.5971 9.500 0.8274 0.03318 0.02537 -0.0097 0.0148 0.6007 10.000 0.8345 0.03679 0.02925 -0.0074 0.0146 0.6050 10.500 0.8426 0.04045 0.03315 -0.0053 0.0146 0.6089 11.000 0.8596 0.04372 0.03673 -0.0026 0.0148 0.6118 11.500 0.8961 0.04673 0.04005 0.0004 0.0154 0.6146 12.000 0.9391 0.05077 0.04445 0.0023 0.0169 0.6174 12.500 0.9590 0.05395 0.04786 0.0029 0.0181 0.6198 13.000 0.9229 0.07006 0.06531 0.0060 0.0291 0.6205