XFOIL         Version 6.94
  
 Calculated polar for: HQ 1.5/8.5 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   0.1989   0.00752   0.00285  -0.0455  0.8500  0.8605
   0.500   0.2473   0.00738   0.00276  -0.0431  0.8296  0.8969
   1.000   0.3032   0.00729   0.00270  -0.0423  0.8110  0.9425
   1.500   0.3805   0.00726   0.00266  -0.0467  0.7918  0.9818
   2.000   0.4412   0.00730   0.00265  -0.0480  0.7680  1.0000
   4.000   0.6243   0.01026   0.00327  -0.0402  0.2395  1.0000
   4.500   0.6689   0.01183   0.00405  -0.0391  0.0924  1.0000
   5.000   0.7171   0.01297   0.00491  -0.0381  0.0367  1.0000
   5.500   0.7676   0.01378   0.00574  -0.0372  0.0232  1.0000
   6.000   0.8158   0.01491   0.00684  -0.0360  0.0058  1.0000
   6.500   0.8636   0.01612   0.00826  -0.0346  0.0053  1.0000
   7.000   0.9088   0.01765   0.01007  -0.0328  0.0053  1.0000
   7.500   0.9494   0.01975   0.01245  -0.0304  0.0054  1.0000
   8.000   0.9876   0.02210   0.01508  -0.0277  0.0057  1.0000
   8.500   1.0202   0.02605   0.01950  -0.0241  0.0064  1.0000
   9.000   1.0453   0.03210   0.02626  -0.0201  0.0072  1.0000
   9.500   1.0368   0.04296   0.03843  -0.0130  0.0089  1.0000
  10.000   1.0172   0.04922   0.04513  -0.0067  0.0095  1.0000