XFOIL Version 6.94 Calculated polar for: HQ 3.5/9 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 0.4680 0.00812 0.00339 -0.1038 0.8385 0.8611 0.500 0.5187 0.00780 0.00310 -0.1020 0.8214 0.9105 1.000 0.5867 0.00766 0.00289 -0.1045 0.8023 1.0000 1.500 0.6436 0.00774 0.00285 -0.1048 0.7828 1.0000 2.000 0.6991 0.00784 0.00284 -0.1047 0.7608 1.0000 2.500 0.7537 0.00797 0.00290 -0.1042 0.7356 1.0000 3.000 0.8071 0.00815 0.00300 -0.1036 0.7051 1.0000 3.500 0.8591 0.00839 0.00316 -0.1026 0.6665 1.0000 4.000 0.9090 0.00876 0.00343 -0.1013 0.6137 1.0000 4.500 0.9498 0.00969 0.00377 -0.0983 0.4929 1.0000 5.000 0.9825 0.01147 0.00459 -0.0946 0.3287 1.0000 5.500 1.0223 0.01290 0.00551 -0.0924 0.2380 1.0000 6.000 1.0679 0.01385 0.00628 -0.0909 0.1850 1.0000 6.500 1.1002 0.01614 0.00772 -0.0878 0.0550 1.0000 7.000 1.1354 0.01822 0.00958 -0.0847 0.0046 1.0000 7.500 1.1765 0.01958 0.01124 -0.0822 0.0041 1.0000 8.000 1.2113 0.02149 0.01360 -0.0788 0.0040 1.0000 8.500 1.2347 0.02405 0.01651 -0.0737 0.0040 1.0000 9.000 1.2466 0.02722 0.01999 -0.0673 0.0042 1.0000 9.500 1.2578 0.03113 0.02422 -0.0615 0.0044 1.0000 10.000 1.2716 0.03610 0.02962 -0.0566 0.0047 1.0000 10.500 1.2800 0.04224 0.03628 -0.0520 0.0050 1.0000 11.000 1.2700 0.05003 0.04467 -0.0472 0.0053 1.0000 11.500 1.2625 0.05694 0.05200 -0.0441 0.0058 1.0000 12.000 1.2574 0.06235 0.05779 -0.0423 0.0063 1.0000 12.500 1.2234 0.07373 0.06984 -0.0429 0.0073 1.0000 13.000 1.1858 0.08657 0.08314 -0.0478 0.0075 1.0000 13.500 1.1503 0.10153 0.09846 -0.0565 0.0075 1.0000 14.000 1.1210 0.11793 0.11512 -0.0674 0.0074 1.0000 14.500 1.0963 0.13549 0.13288 -0.0794 0.0071 1.0000 15.000 1.0727 0.15464 0.15217 -0.0918 0.0068 1.0000