XFOIL         Version 6.94
  
 Calculated polar for: HQ 1.0/8 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   0.1364   0.00586   0.00165  -0.0301  0.8408  0.8499
   0.500   0.1872   0.00577   0.00160  -0.0285  0.8145  0.8833
   1.000   0.2395   0.00570   0.00159  -0.0272  0.7900  0.9242
   1.500   0.3093   0.00571   0.00161  -0.0299  0.7640  0.9709
   2.000   0.3819   0.00579   0.00162  -0.0337  0.7286  1.0000
   2.500   0.4317   0.00599   0.00166  -0.0324  0.6808  1.0000
   3.000   0.4822   0.00630   0.00181  -0.0312  0.6137  1.0000
   3.500   0.5220   0.00796   0.00214  -0.0287  0.3229  1.0000
   4.000   0.5681   0.00939   0.00272  -0.0277  0.1470  1.0000
   4.500   0.6165   0.01057   0.00345  -0.0269  0.0493  1.0000
   5.000   0.6680   0.01131   0.00416  -0.0263  0.0246  1.0000
   5.500   0.7178   0.01245   0.00529  -0.0252  0.0042  1.0000
   6.000   0.7665   0.01391   0.00706  -0.0238  0.0033  1.0000
   6.500   0.8098   0.01641   0.00993  -0.0215  0.0032  1.0000
   7.000   0.8524   0.01934   0.01327  -0.0192  0.0033  1.0000
   7.500   0.8935   0.02324   0.01772  -0.0167  0.0037  1.0000
   8.000   0.9133   0.03371   0.02949  -0.0113  0.0050  1.0000
   8.500   0.9085   0.04491   0.04164  -0.0064  0.0061  1.0000