XFOIL Version 6.94 Calculated polar for: HQ 2.0/10 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 0.2689 0.00833 0.00332 -0.0589 0.8177 0.8216 0.500 0.3214 0.00819 0.00317 -0.0578 0.7987 0.8441 1.000 0.3731 0.00804 0.00305 -0.0564 0.7787 0.8697 1.500 0.4238 0.00788 0.00296 -0.0548 0.7585 0.9017 2.000 0.4851 0.00779 0.00291 -0.0555 0.7371 0.9478 2.500 0.5665 0.00778 0.00289 -0.0609 0.7109 1.0000 3.000 0.6165 0.00792 0.00297 -0.0600 0.6796 1.0000 3.500 0.6681 0.00813 0.00310 -0.0591 0.6417 1.0000 4.000 0.7163 0.00850 0.00322 -0.0574 0.5686 1.0000 4.500 0.7532 0.00975 0.00359 -0.0540 0.3983 1.0000 5.500 0.8392 0.01205 0.00499 -0.0504 0.2178 1.0000 6.000 0.8834 0.01314 0.00572 -0.0489 0.1424 1.0000 6.500 0.9242 0.01456 0.00676 -0.0469 0.0742 1.0000 7.000 0.9662 0.01583 0.00786 -0.0450 0.0401 1.0000 7.500 1.0087 0.01700 0.00899 -0.0430 0.0199 1.0000 8.000 1.0447 0.01873 0.01069 -0.0401 0.0051 1.0000 8.500 1.0812 0.02029 0.01248 -0.0372 0.0045 1.0000 9.000 1.1134 0.02207 0.01454 -0.0338 0.0043 1.0000 9.500 1.1372 0.02407 0.01683 -0.0292 0.0042 1.0000 10.000 1.1502 0.02643 0.01945 -0.0233 0.0042 1.0000 10.500 1.1577 0.02930 0.02259 -0.0177 0.0042 1.0000 11.000 1.1605 0.03286 0.02643 -0.0128 0.0042 1.0000 11.500 1.1606 0.03714 0.03099 -0.0089 0.0043 1.0000 12.000 1.1571 0.04233 0.03650 -0.0061 0.0043 1.0000 12.500 1.1497 0.04847 0.04300 -0.0045 0.0044 1.0000 13.000 1.1411 0.05514 0.05004 -0.0046 0.0044 1.0000 13.500 1.1258 0.06348 0.05881 -0.0065 0.0045 1.0000 14.000 1.1019 0.07444 0.07022 -0.0112 0.0046 1.0000 14.500 1.0695 0.08892 0.08515 -0.0193 0.0048 1.0000 15.000 1.0366 0.10569 0.10229 -0.0296 0.0049 1.0000 15.500 0.9982 0.12578 0.12271 -0.0420 0.0050 1.0000 16.000 0.9538 0.14980 0.14699 -0.0560 0.0051 1.0000 16.500 0.8953 0.18284 0.18004 -0.0717 0.0060 1.0000