XFOIL         Version 6.94
  
 Calculated polar for: HQ 1.5/9 B 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   0.2075   0.00638   0.00190  -0.0455  0.8025  0.8020
   0.500   0.2610   0.00635   0.00184  -0.0446  0.7798  0.8252
   1.000   0.3142   0.00631   0.00185  -0.0438  0.7597  0.8532
   1.500   0.3664   0.00625   0.00188  -0.0427  0.7391  0.8839
   2.000   0.4189   0.00620   0.00191  -0.0415  0.7161  0.9305
   3.000   0.5374   0.00728   0.00198  -0.0433  0.4582  1.0000
   3.500   0.5809   0.00868   0.00250  -0.0418  0.2703  1.0000
   4.000   0.6264   0.00999   0.00306  -0.0407  0.1266  1.0000
   5.000   0.7231   0.01202   0.00451  -0.0388  0.0171  1.0000
   5.500   0.7746   0.01259   0.00515  -0.0381  0.0139  1.0000
   6.000   0.8242   0.01340   0.00595  -0.0372  0.0036  1.0000
   6.500   0.8732   0.01427   0.00695  -0.0361  0.0033  1.0000
   7.000   0.9208   0.01531   0.00819  -0.0347  0.0033  1.0000
   7.500   0.9663   0.01657   0.00967  -0.0330  0.0035  1.0000
   8.000   1.0081   0.01825   0.01161  -0.0307  0.0038  1.0000
   8.500   1.0416   0.02086   0.01455  -0.0273  0.0042  1.0000
   9.000   1.0660   0.02495   0.01907  -0.0228  0.0047  1.0000
   9.500   1.0762   0.03345   0.02865  -0.0164  0.0059  1.0000
  10.000   1.0669   0.04140   0.03729  -0.0096  0.0071  1.0000