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 = 0.450 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.2733 0.00598 0.00178 -0.0603 0.8372 0.8512 0.500 0.3239 0.00583 0.00169 -0.0587 0.8148 0.8925 1.000 0.3904 0.00567 0.00161 -0.0606 0.7933 0.9756 1.500 0.4504 0.00576 0.00158 -0.0616 0.7689 1.0000 2.000 0.5047 0.00590 0.00164 -0.0612 0.7399 1.0000 2.500 0.5583 0.00609 0.00171 -0.0606 0.7046 1.0000 3.000 0.6025 0.00690 0.00178 -0.0581 0.5346 1.0000 3.500 0.6464 0.00828 0.00225 -0.0564 0.3399 1.0000 4.000 0.6954 0.00927 0.00278 -0.0556 0.2406 1.0000 4.500 0.7416 0.01068 0.00349 -0.0546 0.0986 1.0000 5.000 0.7899 0.01184 0.00430 -0.0536 0.0321 1.0000 5.500 0.8401 0.01274 0.00516 -0.0528 0.0089 1.0000 6.000 0.8906 0.01364 0.00617 -0.0518 0.0044 1.0000 6.500 0.9391 0.01491 0.00773 -0.0504 0.0042 1.0000 7.000 0.9814 0.01711 0.01039 -0.0480 0.0042 1.0000 7.500 1.0180 0.02037 0.01406 -0.0448 0.0043 1.0000 8.000 1.0576 0.02341 0.01746 -0.0423 0.0046 1.0000 8.500 1.0951 0.02700 0.02150 -0.0395 0.0051 1.0000 9.000 1.1139 0.03540 0.03089 -0.0344 0.0063 1.0000 9.500 1.0950 0.04711 0.04363 -0.0271 0.0078 1.0000 10.000 1.0612 0.05473 0.05174 -0.0205 0.0082 1.0000 10.500 1.0199 0.06347 0.06084 -0.0193 0.0085 1.0000 11.000 0.9829 0.07491 0.07257 -0.0250 0.0084 1.0000 11.500 0.9482 0.09238 0.09029 -0.0384 0.0080 1.0000 12.000 0.9138 0.11725 0.11520 -0.0529 0.0076 1.0000 12.500 0.8962 0.13370 0.13159 -0.0613 0.0079 1.0000 13.000 0.8868 0.14784 0.14569 -0.0683 0.0087 1.0000