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 =     0.300 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
   0.000   0.1999   0.00758   0.00281  -0.0441  0.8499  0.8628
   0.500   0.2502   0.00744   0.00268  -0.0422  0.8279  0.8926
   1.000   0.3048   0.00731   0.00259  -0.0413  0.8046  0.9302
   1.500   0.3803   0.00726   0.00255  -0.0452  0.7821  0.9722
   2.000   0.4510   0.00728   0.00250  -0.0486  0.7561  1.0000
   2.500   0.5010   0.00740   0.00258  -0.0475  0.7259  1.0000
   3.000   0.5522   0.00757   0.00265  -0.0464  0.6868  1.0000
   3.500   0.6022   0.00784   0.00278  -0.0449  0.6247  1.0000
   4.000   0.6409   0.00907   0.00297  -0.0416  0.4126  1.0000
   4.500   0.6861   0.01025   0.00358  -0.0402  0.2885  1.0000
   5.000   0.7350   0.01113   0.00423  -0.0393  0.2263  1.0000
   5.500   0.7795   0.01258   0.00510  -0.0380  0.1062  1.0000
   6.000   0.8263   0.01379   0.00609  -0.0369  0.0562  1.0000
   6.500   0.8736   0.01493   0.00714  -0.0357  0.0279  1.0000
   7.000   0.9184   0.01645   0.00864  -0.0341  0.0074  1.0000
   7.500   0.9610   0.01836   0.01089  -0.0318  0.0058  1.0000
   8.000   0.9987   0.02081   0.01377  -0.0291  0.0055  1.0000
   8.500   1.0310   0.02411   0.01747  -0.0257  0.0055  1.0000
   9.000   1.0593   0.02851   0.02240  -0.0221  0.0055  1.0000
   9.500   1.0771   0.03427   0.02886  -0.0179  0.0056  1.0000
  10.000   1.0716   0.04114   0.03647  -0.0121  0.0058  1.0000
  10.500   1.0446   0.04791   0.04378  -0.0063  0.0059  1.0000
  11.000   1.0118   0.05615   0.05247  -0.0053  0.0060  1.0000
  11.500   0.9719   0.06781   0.06453  -0.0100  0.0061  1.0000