XFOIL Version 6.94 Calculated polar for: HQ 3.0/10 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.450 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.4122 0.00701 0.00235 -0.0901 0.7858 0.7673 0.500 0.4671 0.00699 0.00229 -0.0896 0.7695 0.7877 1.000 0.5215 0.00698 0.00225 -0.0891 0.7515 0.8083 1.500 0.5744 0.00693 0.00224 -0.0882 0.7287 0.8329 2.000 0.6262 0.00689 0.00224 -0.0871 0.7060 0.8656 2.500 0.6818 0.00674 0.00225 -0.0867 0.6806 0.9528 3.000 0.7415 0.00695 0.00236 -0.0877 0.6494 1.0000 3.500 0.7926 0.00728 0.00249 -0.0868 0.5940 1.0000 4.000 0.8339 0.00823 0.00281 -0.0841 0.4667 1.0000 4.500 0.8708 0.00973 0.00347 -0.0812 0.3136 1.0000 5.000 0.9154 0.01071 0.00410 -0.0796 0.2401 1.0000 5.500 0.9565 0.01199 0.00486 -0.0775 0.1513 1.0000 6.000 0.9989 0.01311 0.00567 -0.0756 0.0929 1.0000 6.500 1.0340 0.01482 0.00692 -0.0726 0.0211 1.0000 7.000 1.0735 0.01609 0.00815 -0.0700 0.0031 1.0000 7.500 1.1153 0.01705 0.00927 -0.0677 0.0029 1.0000 8.000 1.1537 0.01818 0.01057 -0.0649 0.0029 1.0000 8.500 1.1870 0.01947 0.01206 -0.0613 0.0029 1.0000 9.000 1.2126 0.02097 0.01376 -0.0565 0.0030 1.0000 9.500 1.2322 0.02285 0.01587 -0.0512 0.0031 1.0000 10.000 1.2445 0.02526 0.01850 -0.0454 0.0033 1.0000 10.500 1.2511 0.02824 0.02171 -0.0398 0.0034 1.0000 11.000 1.2538 0.03190 0.02560 -0.0347 0.0036 1.0000 11.500 1.2547 0.03633 0.03025 -0.0305 0.0038 1.0000 12.000 1.2582 0.04139 0.03553 -0.0274 0.0040 1.0000 12.500 1.2665 0.04573 0.04030 -0.0251 0.0045 1.0000 13.000 1.2563 0.05343 0.04855 -0.0232 0.0051 1.0000 13.500 1.2397 0.06226 0.05777 -0.0232 0.0054 1.0000