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 =     0.450 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
   0.000   0.1961   0.00589   0.00163  -0.0429  0.8270  0.8506
   0.500   0.2479   0.00576   0.00156  -0.0416  0.8053  0.8866
   1.000   0.3105   0.00564   0.00152  -0.0426  0.7823  0.9455
   1.500   0.3870   0.00569   0.00151  -0.0472  0.7567  1.0000
   2.000   0.4393   0.00584   0.00155  -0.0465  0.7254  1.0000
   2.500   0.4915   0.00605   0.00166  -0.0456  0.6833  1.0000
   3.000   0.5303   0.00742   0.00180  -0.0424  0.4179  1.0000
   3.500   0.5777   0.00850   0.00228  -0.0413  0.2992  1.0000
   4.000   0.6290   0.00914   0.00268  -0.0407  0.2200  1.0000
   4.500   0.6763   0.01034   0.00335  -0.0397  0.0998  1.0000
   5.000   0.7247   0.01144   0.00413  -0.0387  0.0362  1.0000
   5.500   0.7736   0.01252   0.00512  -0.0376  0.0062  1.0000
   6.000   0.8229   0.01366   0.00652  -0.0362  0.0043  1.0000
   6.500   0.8670   0.01566   0.00887  -0.0340  0.0042  1.0000
   7.000   0.9050   0.01886   0.01255  -0.0310  0.0043  1.0000
   7.500   0.9457   0.02202   0.01608  -0.0285  0.0045  1.0000
   8.000   0.9859   0.02540   0.01989  -0.0259  0.0049  1.0000
   8.500   1.0100   0.03327   0.02874  -0.0212  0.0061  1.0000
   9.000   0.9971   0.04543   0.04200  -0.0145  0.0079  1.0000
   9.500   0.9662   0.05415   0.05126  -0.0088  0.0087  1.0000
  10.000   0.9256   0.06231   0.05975  -0.0072  0.0089  1.0000
  10.500   0.8886   0.07414   0.07184  -0.0142  0.0088  1.0000
  11.000   0.8578   0.09596   0.09384  -0.0303  0.0082  1.0000
  11.500   0.8356   0.11354   0.11136  -0.0398  0.0084  1.0000
  12.000   0.8223   0.12760   0.12537  -0.0468  0.0090  1.0000