XFOIL         Version 6.94
  
 Calculated polar for: HQ 1.5/8 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   0.1936   0.00686   0.00237  -0.0425  0.8720  0.9177
   1.000   0.3419   0.00675   0.00215  -0.0502  0.8290  1.0000
   1.500   0.3918   0.00685   0.00217  -0.0490  0.8032  1.0000
   2.000   0.4433   0.00697   0.00220  -0.0479  0.7745  1.0000
   2.500   0.4951   0.00712   0.00232  -0.0468  0.7403  1.0000
   3.000   0.5447   0.00734   0.00234  -0.0451  0.6756  1.0000
   3.500   0.5845   0.00845   0.00242  -0.0417  0.4553  1.0000
   4.000   0.6291   0.00973   0.00299  -0.0401  0.3103  1.0000
   4.500   0.6792   0.01050   0.00364  -0.0393  0.2393  1.0000
   5.000   0.7237   0.01207   0.00446  -0.0381  0.0954  1.0000
   5.500   0.7720   0.01320   0.00547  -0.0370  0.0477  1.0000
   6.000   0.8179   0.01479   0.00693  -0.0354  0.0087  1.0000
   6.500   0.8637   0.01652   0.00900  -0.0334  0.0067  1.0000
   7.000   0.9042   0.01912   0.01205  -0.0307  0.0063  1.0000
   7.500   0.9420   0.02292   0.01630  -0.0276  0.0064  1.0000
   8.000   0.9764   0.02895   0.02311  -0.0240  0.0069  1.0000
   8.500   0.9952   0.03706   0.03221  -0.0194  0.0076  1.0000
   9.000   0.9923   0.04599   0.04199  -0.0144  0.0083  1.0000
   9.500   0.9667   0.05413   0.05069  -0.0091  0.0088  1.0000
  10.000   0.9286   0.06229   0.05920  -0.0076  0.0090  1.0000
  10.500   0.8908   0.07413   0.07134  -0.0144  0.0090  1.0000
  11.000   0.8577   0.09470   0.09210  -0.0300  0.0089  1.0000
  11.500   0.8345   0.11319   0.11053  -0.0398  0.0092  1.0000