XFOIL         Version 6.94
  
 Calculated polar for: HQ 3.0/9 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   0.4018   0.00809   0.00341  -0.0895  0.8443  0.8568
   0.500   0.4516   0.00780   0.00312  -0.0874  0.8241  0.8967
   1.000   0.5225   0.00757   0.00292  -0.0903  0.8064  1.0000
   1.500   0.5788   0.00761   0.00282  -0.0905  0.7843  1.0000
   2.000   0.6339   0.00768   0.00276  -0.0902  0.7602  1.0000
   2.500   0.6885   0.00781   0.00278  -0.0898  0.7348  1.0000
   3.000   0.7416   0.00797   0.00296  -0.0891  0.7030  1.0000
   3.500   0.7937   0.00821   0.00309  -0.0882  0.6638  1.0000
   4.000   0.8356   0.00887   0.00318  -0.0850  0.5417  1.0000
   4.500   0.8707   0.01035   0.00376  -0.0815  0.3751  1.0000
   5.000   0.9135   0.01148   0.00449  -0.0795  0.2845  1.0000
   5.500   0.9533   0.01292   0.00536  -0.0773  0.1784  1.0000
   6.000   0.9931   0.01442   0.00638  -0.0751  0.0927  1.0000
   6.500   1.0308   0.01612   0.00769  -0.0725  0.0270  1.0000
   7.000   1.0692   0.01774   0.00928  -0.0697  0.0047  1.0000
   7.500   1.1089   0.01915   0.01093  -0.0671  0.0043  1.0000
   8.500   1.1718   0.02302   0.01539  -0.0595  0.0043  1.0000
   9.000   1.1890   0.02531   0.01792  -0.0536  0.0045  1.0000
   9.500   1.2015   0.02811   0.02098  -0.0475  0.0047  1.0000
  10.000   1.2133   0.03196   0.02515  -0.0421  0.0051  1.0000
  10.500   1.2286   0.03703   0.03065  -0.0377  0.0056  1.0000
  11.000   1.2421   0.04324   0.03726  -0.0342  0.0061  1.0000
  11.500   1.2196   0.05127   0.04630  -0.0284  0.0074  1.0000
  12.000   1.1978   0.05932   0.05478  -0.0262  0.0078  1.0000