XFOIL         Version 6.94
  
 Calculated polar for: HQ 1.5/9 AIRFOIL                                
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.080     Re =     0.450 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
   0.000   0.2052   0.00635   0.00191  -0.0450  0.8078  0.8093
   0.500   0.2581   0.00629   0.00186  -0.0440  0.7863  0.8353
   1.000   0.3111   0.00624   0.00184  -0.0430  0.7644  0.8617
   1.500   0.3630   0.00616   0.00182  -0.0418  0.7414  0.8959
   2.000   0.4237   0.00613   0.00184  -0.0424  0.7154  0.9509
   2.500   0.5014   0.00621   0.00191  -0.0472  0.6801  1.0000
   3.000   0.5510   0.00647   0.00199  -0.0460  0.6298  1.0000
   3.500   0.5916   0.00756   0.00219  -0.0432  0.4353  1.0000
   4.000   0.6348   0.00890   0.00272  -0.0415  0.2692  1.0000
   4.500   0.6819   0.00993   0.00328  -0.0404  0.1686  1.0000
   5.000   0.7293   0.01095   0.00393  -0.0393  0.0917  1.0000
   5.500   0.7755   0.01212   0.00477  -0.0380  0.0336  1.0000
   6.000   0.8220   0.01327   0.00578  -0.0366  0.0045  1.0000
   6.500   0.8698   0.01426   0.00691  -0.0352  0.0034  1.0000
   7.000   0.9157   0.01546   0.00834  -0.0335  0.0032  1.0000
   7.500   0.9582   0.01705   0.01018  -0.0314  0.0032  1.0000
   8.000   0.9956   0.01912   0.01249  -0.0286  0.0033  1.0000
   8.500   1.0308   0.02130   0.01490  -0.0255  0.0034  1.0000
   9.000   1.0611   0.02419   0.01814  -0.0217  0.0037  1.0000
   9.500   1.0810   0.02912   0.02369  -0.0168  0.0043  1.0000
  10.000   1.0840   0.03501   0.03023  -0.0107  0.0048  1.0000
  10.500   1.0732   0.04013   0.03588  -0.0031  0.0054  1.0000