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 =     0.022 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
   0.000   0.2560   0.04128   0.02969  -0.0341  0.9998  1.0002
   0.250   0.2742   0.04270   0.03126  -0.0344  0.9998  1.0002
   0.500   0.5535   0.03717   0.02522  -0.0735  0.7257  1.0002
   0.750   0.6091   0.03650   0.02392  -0.0759  0.6675  1.0002
   1.000   0.6485   0.03688   0.02374  -0.0768  0.6304  1.0002
   1.250   0.6836   0.03745   0.02377  -0.0771  0.6019  1.0002
   1.500   0.7148   0.03831   0.02414  -0.0772  0.5782  1.0002
   1.750   0.7424   0.03951   0.02503  -0.0772  0.5585  1.0002
   2.000   0.7672   0.04112   0.02651  -0.0776  0.5440  1.0002
   2.250   0.7966   0.04264   0.02783  -0.0784  0.5345  1.0002
   2.500   0.8175   0.04486   0.03016  -0.0791  0.5269  1.0002
   2.750   0.8413   0.04695   0.03226  -0.0798  0.5204  1.0002
   3.000   0.8686   0.04901   0.03425  -0.0809  0.5153  1.0002
   3.250   0.8809   0.05238   0.03787  -0.0815  0.5124  1.0002
   3.500   0.8888   0.05626   0.04195  -0.0819  0.5107  1.0002
   3.750   0.8899   0.06082   0.04670  -0.0820  0.5106  1.0002
   4.000   0.8832   0.06617   0.05220  -0.0820  0.5121  1.0002
   4.250   0.8764   0.07166   0.05777  -0.0820  0.5144  1.0002
   4.750   0.7809   0.08966   0.07593  -0.0809  0.5341  1.0002