XFOIL         Version 6.94
  
 Calculated polar for: HQ 3.0/8 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   0.4112   0.00732   0.00294  -0.0919  0.8712  0.9673
   0.500   0.4716   0.00727   0.00270  -0.0928  0.8523  1.0000
   1.000   0.5269   0.00733   0.00262  -0.0926  0.8327  1.0000
   1.500   0.5808   0.00736   0.00254  -0.0919  0.8079  1.0000
   2.000   0.6348   0.00744   0.00254  -0.0912  0.7821  1.0000
   2.500   0.6886   0.00757   0.00259  -0.0906  0.7541  1.0000
   3.000   0.7416   0.00774   0.00276  -0.0897  0.7194  1.0000
   3.500   0.7936   0.00799   0.00293  -0.0886  0.6756  1.0000
   4.000   0.8437   0.00839   0.00320  -0.0873  0.6144  1.0000
   4.500   0.8905   0.00909   0.00365  -0.0854  0.5237  1.0000
   5.000   0.9261   0.01086   0.00442  -0.0821  0.3387  1.0000
   5.500   0.9681   0.01235   0.00532  -0.0803  0.2226  1.0000
   6.000   1.0109   0.01384   0.00629  -0.0787  0.1255  1.0000
   6.500   1.0474   0.01624   0.00801  -0.0762  0.0175  1.0000
   7.000   1.0900   0.01796   0.00987  -0.0740  0.0041  1.0000
   7.500   1.1304   0.01995   0.01227  -0.0713  0.0036  1.0000
   8.000   1.1625   0.02285   0.01559  -0.0674  0.0036  1.0000
   8.500   1.1883   0.02690   0.02008  -0.0628  0.0037  1.0000
   9.000   1.2129   0.03244   0.02624  -0.0582  0.0040  1.0000
   9.500   1.2242   0.03940   0.03398  -0.0528  0.0044  1.0000
  10.000   1.2081   0.04660   0.04186  -0.0452  0.0048  1.0000
  10.500   1.1757   0.05454   0.05037  -0.0393  0.0050  1.0000
  11.000   1.1395   0.06383   0.06011  -0.0378  0.0051  1.0000