XFOIL         Version 6.94
  
 Calculated polar for: tilt                                            
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.035 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
   0.500   0.5889   0.02823   0.01548  -0.0674  0.4647  1.0002
   0.750   0.6181   0.02926   0.01603  -0.0677  0.4495  1.0002
   1.000   0.6490   0.03021   0.01670  -0.0684  0.4381  1.0002
   1.250   0.6817   0.03128   0.01743  -0.0695  0.4294  1.0002
   1.500   0.7144   0.03226   0.01833  -0.0709  0.4219  1.0002
   1.750   0.7479   0.03335   0.01930  -0.0724  0.4154  1.0002
   2.000   0.7824   0.03463   0.02041  -0.0740  0.4102  1.0002
   2.250   0.8129   0.03581   0.02167  -0.0751  0.4065  1.0002
   2.500   0.8424   0.03709   0.02306  -0.0761  0.4033  1.0002
   2.750   0.8709   0.03850   0.02460  -0.0770  0.4005  1.0002
   3.000   0.8982   0.04001   0.02620  -0.0777  0.3970  1.0002
   3.250   0.9252   0.04164   0.02783  -0.0783  0.3927  1.0002
   3.500   0.9523   0.04360   0.02970  -0.0790  0.3881  1.0002
   3.750   0.9737   0.04537   0.03174  -0.0791  0.3849  1.0002
   4.000   0.9951   0.04738   0.03403  -0.0793  0.3830  1.0002
   4.250   1.0162   0.04963   0.03652  -0.0796  0.3827  1.0002
   4.500   1.0354   0.05211   0.03924  -0.0797  0.3829  1.0002
   4.750   1.0520   0.05488   0.04225  -0.0797  0.3836  1.0002
   5.000   1.0660   0.05797   0.04559  -0.0796  0.3849  1.0002