XFOIL         Version 6.94
  
 Calculated polar for: HQ 1.0/9 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   0.1341   0.00751   0.00274  -0.0297  0.8512  0.8604
   0.500   0.1839   0.00743   0.00265  -0.0277  0.8280  0.8892
   1.000   0.2356   0.00734   0.00260  -0.0260  0.8042  0.9203
   1.500   0.2995   0.00731   0.00260  -0.0272  0.7802  0.9547
   2.000   0.3781   0.00732   0.00258  -0.0319  0.7531  0.9822
   2.500   0.4466   0.00734   0.00258  -0.0349  0.7166  1.0000
   3.000   0.4925   0.00749   0.00262  -0.0329  0.6728  1.0000
   3.500   0.5409   0.00777   0.00277  -0.0312  0.6080  1.0000
   4.000   0.5885   0.00832   0.00300  -0.0295  0.5099  1.0000
   4.500   0.6304   0.00967   0.00352  -0.0273  0.3287  1.0000
   5.000   0.6754   0.01102   0.00426  -0.0261  0.2050  1.0000
   5.500   0.7215   0.01234   0.00511  -0.0250  0.1032  1.0000
   6.000   0.7686   0.01357   0.00610  -0.0240  0.0517  1.0000
   6.500   0.8161   0.01478   0.00725  -0.0229  0.0234  1.0000
   7.000   0.8592   0.01686   0.00947  -0.0207  0.0071  1.0000
   7.500   0.9004   0.01914   0.01211  -0.0185  0.0056  1.0000
   8.000   0.9357   0.02243   0.01577  -0.0155  0.0051  1.0000
   8.500   0.9683   0.02664   0.02047  -0.0125  0.0050  1.0000
   9.000   0.9944   0.03187   0.02635  -0.0091  0.0051  1.0000
   9.500   1.0069   0.03797   0.03322  -0.0049  0.0054  1.0000
  10.000   0.9815   0.04664   0.04274   0.0016  0.0059  1.0000
  10.500   0.9341   0.05640   0.05307   0.0037  0.0063  1.0000
  11.000   0.8972   0.06769   0.06472  -0.0020  0.0064  1.0000
  12.000   0.8038   0.11489   0.11219  -0.0315  0.0072  1.0000
  12.500   0.7864   0.13052   0.12775  -0.0391  0.0078  1.0000
  13.000   0.7803   0.14315   0.14034  -0.0450  0.0085  1.0000