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 =     0.045 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
   1.000   0.6607   0.02885   0.01589  -0.0708  0.3778  1.0002
   1.250   0.6941   0.02976   0.01658  -0.0720  0.3720  1.0002
   1.500   0.7276   0.03089   0.01745  -0.0734  0.3673  1.0002
   1.750   0.7587   0.03192   0.01842  -0.0744  0.3639  1.0002
   2.000   0.7885   0.03291   0.01946  -0.0752  0.3610  1.0002
   2.250   0.8175   0.03399   0.02060  -0.0759  0.3579  1.0002
   2.500   0.8457   0.03517   0.02181  -0.0765  0.3541  1.0002
   2.750   0.8733   0.03647   0.02310  -0.0771  0.3497  1.0002
   3.000   0.9007   0.03809   0.02462  -0.0777  0.3454  1.0002
   3.250   0.9265   0.03964   0.02625  -0.0781  0.3424  1.0002
   3.500   0.9520   0.04109   0.02789  -0.0785  0.3411  1.0002
   3.750   0.9768   0.04271   0.02973  -0.0788  0.3402  1.0002
   4.000   1.0006   0.04446   0.03169  -0.0791  0.3397  1.0002
   4.250   1.0237   0.04634   0.03379  -0.0793  0.3395  1.0002
   4.500   1.0454   0.04843   0.03609  -0.0794  0.3397  1.0002
   4.750   1.0660   0.05071   0.03860  -0.0794  0.3402  1.0002
   5.000   1.0852   0.05323   0.04139  -0.0794  0.3412  1.0002