XFOIL Version 6.94 Calculated polar for: HQ 2.0/9 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 0.2670 0.00774 0.00302 -0.0591 0.8456 0.8594 0.500 0.3173 0.00756 0.00286 -0.0573 0.8251 0.8920 1.000 0.3743 0.00739 0.00274 -0.0569 0.8041 0.9369 1.500 0.4561 0.00729 0.00261 -0.0623 0.7827 0.9954 2.000 0.5091 0.00737 0.00256 -0.0620 0.7581 1.0000 2.500 0.5618 0.00749 0.00266 -0.0612 0.7297 1.0000 3.000 0.6142 0.00766 0.00274 -0.0604 0.6943 1.0000 3.500 0.6649 0.00791 0.00288 -0.0590 0.6405 1.0000 4.000 0.7055 0.00883 0.00300 -0.0559 0.4770 1.0000 4.500 0.7419 0.01059 0.00369 -0.0530 0.2774 1.0000 5.000 0.7893 0.01152 0.00438 -0.0519 0.2127 1.0000 5.500 0.8329 0.01287 0.00520 -0.0504 0.1091 1.0000 6.000 0.8766 0.01424 0.00625 -0.0487 0.0491 1.0000 6.500 0.9217 0.01542 0.00733 -0.0472 0.0193 1.0000 7.000 0.9637 0.01698 0.00889 -0.0450 0.0050 1.0000 7.500 1.0057 0.01848 0.01065 -0.0428 0.0046 1.0000 8.000 1.0440 0.02031 0.01279 -0.0400 0.0045 1.0000 8.500 1.0757 0.02264 0.01542 -0.0365 0.0045 1.0000 9.000 1.0998 0.02557 0.01862 -0.0321 0.0046 1.0000 9.500 1.1175 0.02933 0.02268 -0.0271 0.0047 1.0000 10.000 1.1350 0.03253 0.02630 -0.0220 0.0049 1.0000 10.500 1.1361 0.03818 0.03272 -0.0158 0.0054 1.0000 11.000 1.1154 0.04510 0.04034 -0.0098 0.0059 1.0000 11.500 1.0877 0.05255 0.04830 -0.0069 0.0061 1.0000 12.000 1.0559 0.06156 0.05772 -0.0077 0.0063 1.0000 12.500 1.0226 0.07280 0.06932 -0.0126 0.0063 1.0000 13.000 0.9879 0.08744 0.08428 -0.0220 0.0063 1.0000 13.500 0.9523 0.10647 0.10359 -0.0349 0.0063 1.0000 14.000 0.9100 0.13098 0.12828 -0.0492 0.0062 1.0000 14.500 0.8829 0.15162 0.14888 -0.0593 0.0065 1.0000