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 = 0.300 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.3399 0.00899 0.00378 -0.0745 0.7973 0.7933 0.500 0.3928 0.00891 0.00372 -0.0736 0.7816 0.8135 1.000 0.4461 0.00880 0.00362 -0.0726 0.7638 0.8322 1.500 0.4994 0.00871 0.00353 -0.0717 0.7461 0.8516 2.000 0.5518 0.00864 0.00349 -0.0706 0.7291 0.8746 3.000 0.6648 0.00838 0.00335 -0.0700 0.6741 0.9617 3.500 0.7291 0.00850 0.00341 -0.0721 0.6379 1.0000 4.000 0.7812 0.00876 0.00354 -0.0715 0.5925 1.0000 4.500 0.8296 0.00923 0.00379 -0.0701 0.5272 1.0000 5.000 0.8736 0.01005 0.00426 -0.0680 0.4546 1.0000 5.500 0.9176 0.01093 0.00490 -0.0661 0.3862 1.0000 6.000 0.9580 0.01200 0.00564 -0.0637 0.3110 1.0000 6.500 0.9970 0.01319 0.00654 -0.0612 0.2484 1.0000 7.000 1.0368 0.01428 0.00743 -0.0589 0.1941 1.0000 7.500 1.0750 0.01543 0.00840 -0.0563 0.1523 1.0000 8.000 1.1110 0.01667 0.00948 -0.0535 0.1093 1.0000 8.500 1.1254 0.01929 0.01146 -0.0474 0.0306 1.0000 9.000 1.1452 0.02116 0.01346 -0.0419 0.0252 1.0000 10.000 1.1725 0.02569 0.01827 -0.0309 0.0216 1.0000 10.500 1.1812 0.02864 0.02133 -0.0261 0.0204 1.0000 11.000 1.1961 0.03152 0.02437 -0.0224 0.0194 1.0000 11.500 1.2131 0.03473 0.02771 -0.0193 0.0187 1.0000 12.000 1.2380 0.03821 0.03124 -0.0172 0.0181 1.0000 12.500 1.2712 0.04220 0.03544 -0.0157 0.0176 1.0000 13.000 1.2870 0.04666 0.04031 -0.0134 0.0174 1.0000 13.500 1.2907 0.05209 0.04620 -0.0113 0.0173 1.0000 14.000 1.2836 0.05859 0.05316 -0.0098 0.0172 1.0000 14.500 1.2633 0.06642 0.06145 -0.0098 0.0171 1.0000 15.000 1.2357 0.07582 0.07131 -0.0120 0.0171 1.0000 15.500 1.2062 0.08703 0.08293 -0.0163 0.0172 1.0000 16.000 1.1694 0.10062 0.09691 -0.0239 0.0172 1.0000 16.500 1.1369 0.11547 0.11209 -0.0329 0.0174 1.0000 17.000 1.1127 0.12997 0.12680 -0.0416 0.0177 1.0000