XFOIL         Version 6.94
  
 Calculated polar for: HQ 1.0/8 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   0.1320   0.00691   0.00246  -0.0294  0.8851  0.9099
   0.500   0.1911   0.00684   0.00235  -0.0292  0.8597  0.9466
   1.000   0.2673   0.00681   0.00226  -0.0332  0.8331  0.9784
   1.500   0.3400   0.00681   0.00223  -0.0370  0.8070  1.0000
   2.000   0.3863   0.00691   0.00225  -0.0352  0.7762  1.0000
   2.500   0.4357   0.00705   0.00230  -0.0336  0.7364  1.0000
   3.000   0.4855   0.00725   0.00242  -0.0320  0.6806  1.0000
   3.500   0.5352   0.00761   0.00259  -0.0305  0.6026  1.0000
   4.000   0.5749   0.00912   0.00294  -0.0277  0.3513  1.0000
   4.500   0.6196   0.01067   0.00372  -0.0265  0.1817  1.0000
   5.000   0.6656   0.01220   0.00461  -0.0255  0.0611  1.0000
   5.500   0.7141   0.01348   0.00576  -0.0245  0.0167  1.0000
   6.000   0.7602   0.01549   0.00806  -0.0225  0.0063  1.0000
   6.500   0.8028   0.01823   0.01116  -0.0201  0.0055  1.0000
   7.000   0.8438   0.02223   0.01572  -0.0173  0.0054  1.0000
   7.500   0.8805   0.02826   0.02258  -0.0141  0.0057  1.0000
   8.000   0.9016   0.03640   0.03172  -0.0102  0.0063  1.0000
   8.500   0.9005   0.04586   0.04203  -0.0062  0.0068  1.0000