XFOIL         Version 6.94
  
 Calculated polar for: HQ 2.5/10 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   0.3370   0.00842   0.00342  -0.0744  0.8198  0.8216
   0.500   0.3891   0.00827   0.00324  -0.0732  0.8000  0.8451
   1.000   0.4406   0.00812   0.00315  -0.0719  0.7817  0.8721
   1.500   0.4914   0.00795   0.00304  -0.0703  0.7620  0.9110
   2.000   0.5654   0.00782   0.00292  -0.0738  0.7390  0.9789
   2.500   0.6246   0.00792   0.00295  -0.0748  0.7141  1.0000
   3.000   0.6778   0.00807   0.00302  -0.0743  0.6833  1.0000
   4.000   0.7731   0.00884   0.00318  -0.0707  0.5334  1.0000
   4.500   0.8088   0.01022   0.00370  -0.0673  0.3733  1.0000
   5.000   0.8511   0.01134   0.00439  -0.0653  0.2820  1.0000
   5.500   0.8930   0.01255   0.00518  -0.0633  0.1962  1.0000
   6.000   0.9345   0.01380   0.00606  -0.0613  0.1238  1.0000
   6.500   0.9627   0.01632   0.00779  -0.0574  0.0076  1.0000
   7.000   1.0040   0.01756   0.00920  -0.0550  0.0052  1.0000
   7.500   1.0422   0.01902   0.01092  -0.0521  0.0048  1.0000
   8.000   1.0744   0.02087   0.01303  -0.0485  0.0047  1.0000
   8.500   1.0957   0.02305   0.01543  -0.0433  0.0047  1.0000
   9.000   1.1104   0.02532   0.01786  -0.0373  0.0048  1.0000
   9.500   1.1256   0.02789   0.02063  -0.0320  0.0050  1.0000
  10.000   1.1418   0.03119   0.02418  -0.0273  0.0054  1.0000
  10.500   1.1627   0.03556   0.02890  -0.0236  0.0060  1.0000
  11.000   1.1960   0.04162   0.03520  -0.0220  0.0066  1.0000
  11.500   1.1820   0.04703   0.04148  -0.0160  0.0080  1.0000