XFOIL Version 6.94 Calculated polar for: HQ 2.5/9 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 0.3358 0.00784 0.00309 -0.0748 0.8485 0.8646 0.500 0.3852 0.00760 0.00287 -0.0727 0.8270 0.9029 1.000 0.4545 0.00739 0.00268 -0.0750 0.8080 0.9711 1.500 0.5161 0.00742 0.00261 -0.0765 0.7855 1.0000 2.000 0.5707 0.00752 0.00259 -0.0762 0.7603 1.0000 2.500 0.6247 0.00765 0.00268 -0.0757 0.7323 1.0000 3.000 0.6777 0.00781 0.00276 -0.0749 0.6969 1.0000 3.500 0.7300 0.00805 0.00292 -0.0740 0.6555 1.0000 4.000 0.7770 0.00855 0.00311 -0.0720 0.5625 1.0000 4.500 0.8163 0.00986 0.00360 -0.0691 0.4019 1.0000 5.000 0.8612 0.01096 0.00430 -0.0676 0.3067 1.0000 5.500 0.9069 0.01204 0.00511 -0.0662 0.2342 1.0000 6.000 0.9502 0.01341 0.00603 -0.0646 0.1434 1.0000 6.500 0.9845 0.01591 0.00778 -0.0618 0.0173 1.0000 7.000 1.0274 0.01754 0.00963 -0.0595 0.0087 1.0000 7.500 1.0672 0.01943 0.01188 -0.0568 0.0078 1.0000 8.000 1.0994 0.02202 0.01479 -0.0530 0.0077 1.0000 8.500 1.1273 0.02536 0.01844 -0.0488 0.0078 1.0000 9.000 1.1570 0.02980 0.02330 -0.0449 0.0081 1.0000 9.500 1.1818 0.03328 0.02721 -0.0413 0.0067 1.0000 10.000 1.1912 0.03767 0.03201 -0.0362 0.0059 1.0000 10.500 1.1874 0.04322 0.03811 -0.0304 0.0059 1.0000 11.000 1.1724 0.04959 0.04506 -0.0258 0.0061 1.0000 11.500 1.1456 0.05759 0.05355 -0.0237 0.0061 1.0000 12.000 1.1093 0.06884 0.06537 -0.0255 0.0068 1.0000 12.500 1.0696 0.08263 0.07957 -0.0325 0.0070 1.0000 13.000 1.0340 0.09920 0.09645 -0.0436 0.0070 1.0000 14.000 0.9430 0.15166 0.14913 -0.0739 0.0099 1.0000 14.500 0.9374 0.16429 0.16173 -0.0802 0.0115 1.0000