XFOIL Version 6.94 Calculated polar for: HQ 2.0/8 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 0.2672 0.00706 0.00262 -0.0594 0.8753 0.9257 0.500 0.3462 0.00688 0.00238 -0.0639 0.8542 0.9950 1.000 0.4003 0.00695 0.00230 -0.0637 0.8316 1.0000 1.500 0.4536 0.00704 0.00230 -0.0631 0.8082 1.0000 2.000 0.5070 0.00713 0.00234 -0.0624 0.7815 1.0000 2.500 0.5602 0.00725 0.00240 -0.0616 0.7505 1.0000 3.000 0.6129 0.00742 0.00255 -0.0606 0.7116 1.0000 3.500 0.6553 0.00813 0.00248 -0.0573 0.5439 1.0000 4.000 0.6971 0.00956 0.00300 -0.0551 0.3537 1.0000 4.500 0.7449 0.01063 0.00371 -0.0541 0.2590 1.0000 5.000 0.7887 0.01227 0.00453 -0.0528 0.1075 1.0000 5.500 0.8361 0.01355 0.00554 -0.0517 0.0472 1.0000 6.000 0.8843 0.01472 0.00665 -0.0507 0.0177 1.0000 6.500 0.9318 0.01608 0.00815 -0.0492 0.0070 1.0000 7.000 0.9769 0.01784 0.01036 -0.0472 0.0062 1.0000 7.500 1.0162 0.02044 0.01336 -0.0445 0.0060 1.0000 8.000 1.0510 0.02431 0.01768 -0.0411 0.0062 1.0000 8.500 1.0828 0.03027 0.02438 -0.0373 0.0067 1.0000 9.000 1.0983 0.03845 0.03354 -0.0324 0.0074 1.0000 9.500 1.0894 0.04740 0.04334 -0.0266 0.0080 1.0000 10.000 1.0597 0.05474 0.05120 -0.0205 0.0083 1.0000 10.500 1.0212 0.06339 0.06025 -0.0193 0.0085 1.0000 11.000 0.9835 0.07495 0.07213 -0.0249 0.0085 1.0000 11.500 0.9468 0.09257 0.09001 -0.0383 0.0084 1.0000 12.000 0.9116 0.11686 0.11435 -0.0526 0.0085 1.0000 12.500 0.8942 0.13317 0.13060 -0.0608 0.0089 1.0000 13.000 0.8844 0.14715 0.14452 -0.0677 0.0094 1.0000