XFOIL         Version 6.94
  
 Calculated polar for: HQ 2.0/9 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   0.2747   0.00644   0.00202  -0.0603  0.8098  0.8091
   0.500   0.3281   0.00636   0.00194  -0.0594  0.7890  0.8324
   1.000   0.3808   0.00627   0.00189  -0.0584  0.7676  0.8606
   1.500   0.4322   0.00617   0.00186  -0.0571  0.7451  0.9011
   2.000   0.5020   0.00611   0.00184  -0.0597  0.7181  0.9807
   2.500   0.5602   0.00627   0.00193  -0.0604  0.6861  1.0000
   3.000   0.6128   0.00652   0.00204  -0.0597  0.6435  1.0000
   3.500   0.6553   0.00747   0.00222  -0.0572  0.4777  1.0000
   4.000   0.6937   0.00924   0.00286  -0.0548  0.2579  1.0000
   4.500   0.7427   0.01005   0.00336  -0.0539  0.1925  1.0000
   5.000   0.7870   0.01141   0.00413  -0.0524  0.0812  1.0000
   5.500   0.8337   0.01247   0.00496  -0.0512  0.0345  1.0000
   6.000   0.8795   0.01366   0.00599  -0.0497  0.0040  1.0000
   6.500   0.9273   0.01457   0.00704  -0.0483  0.0033  1.0000
   7.000   0.9732   0.01568   0.00835  -0.0466  0.0032  1.0000
   7.500   1.0166   0.01703   0.00994  -0.0446  0.0032  1.0000
   8.000   1.0564   0.01868   0.01181  -0.0420  0.0034  1.0000
   8.500   1.0900   0.02082   0.01420  -0.0386  0.0036  1.0000
   9.000   1.1158   0.02372   0.01739  -0.0341  0.0040  1.0000
   9.500   1.1364   0.02763   0.02167  -0.0293  0.0044  1.0000
  10.000   1.1531   0.03218   0.02659  -0.0244  0.0048  1.0000
  10.500   1.1419   0.03901   0.03438  -0.0160  0.0061  1.0000