XFOIL Version 6.94 Calculated polar for: HQ 2.5/11 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 0.3466 0.00728 0.00245 -0.0755 0.7637 0.7513 0.500 0.4018 0.00728 0.00240 -0.0751 0.7468 0.7679 1.000 0.4568 0.00730 0.00239 -0.0746 0.7298 0.7863 1.500 0.5118 0.00733 0.00241 -0.0743 0.7134 0.8030 2.000 0.5659 0.00733 0.00245 -0.0737 0.6932 0.8199 2.500 0.6185 0.00734 0.00248 -0.0728 0.6636 0.8398 3.500 0.7171 0.00740 0.00266 -0.0695 0.5899 0.9144 4.000 0.7799 0.00774 0.00282 -0.0712 0.5219 1.0000 4.500 0.8274 0.00847 0.00322 -0.0699 0.4532 1.0000 5.000 0.8749 0.00920 0.00372 -0.0687 0.3926 1.0000 5.500 0.9196 0.01009 0.00429 -0.0670 0.3159 1.0000 6.000 0.9627 0.01110 0.00502 -0.0652 0.2542 1.0000 6.500 1.0060 0.01206 0.00573 -0.0633 0.1984 1.0000 7.000 1.0474 0.01311 0.00654 -0.0613 0.1494 1.0000 7.500 1.0876 0.01418 0.00744 -0.0590 0.1131 1.0000 8.000 1.1096 0.01654 0.00916 -0.0540 0.0261 1.0000 8.500 1.1391 0.01813 0.01085 -0.0498 0.0199 1.0000 9.000 1.1595 0.01980 0.01263 -0.0443 0.0179 1.0000 9.500 1.1798 0.02145 0.01444 -0.0391 0.0165 1.0000 10.000 1.1949 0.02350 0.01660 -0.0338 0.0155 1.0000 10.500 1.2007 0.02639 0.01958 -0.0282 0.0149 1.0000 11.000 1.2156 0.02898 0.02236 -0.0243 0.0145 1.0000 11.500 1.2310 0.03178 0.02534 -0.0209 0.0138 1.0000 12.000 1.2447 0.03501 0.02874 -0.0179 0.0133 1.0000 12.500 1.2574 0.03873 0.03253 -0.0156 0.0124 1.0000 13.000 1.2676 0.04284 0.03692 -0.0135 0.0119 1.0000 13.500 1.2731 0.04742 0.04182 -0.0120 0.0113 1.0000 14.000 1.2760 0.05282 0.04753 -0.0110 0.0111 1.0000 14.500 1.2729 0.05917 0.05420 -0.0110 0.0108 1.0000 15.000 1.2636 0.06675 0.06213 -0.0122 0.0107 1.0000 15.500 1.2492 0.07565 0.07137 -0.0150 0.0106 1.0000 16.000 1.2421 0.08398 0.07986 -0.0184 0.0101 1.0000 16.500 1.2260 0.09460 0.09072 -0.0234 0.0099 1.0000