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 =     0.030 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
   0.500   0.5950   0.02897   0.01647  -0.0693  0.5359  1.0002
   0.750   0.6220   0.03006   0.01693  -0.0690  0.5068  1.0002
   1.000   0.6504   0.03117   0.01753  -0.0692  0.4867  1.0002
   1.250   0.6813   0.03232   0.01822  -0.0699  0.4735  1.0002
   1.500   0.7118   0.03340   0.01921  -0.0709  0.4632  1.0002
   1.750   0.7438   0.03460   0.02018  -0.0720  0.4548  1.0002
   2.000   0.7751   0.03585   0.02140  -0.0733  0.4483  1.0002
   2.250   0.8060   0.03717   0.02277  -0.0746  0.4426  1.0002
   2.500   0.8369   0.03859   0.02422  -0.0759  0.4378  1.0002
   2.750   0.8674   0.04015   0.02577  -0.0772  0.4340  1.0002
   3.000   0.8970   0.04190   0.02752  -0.0784  0.4312  1.0002
   3.250   0.9215   0.04375   0.02962  -0.0790  0.4290  1.0002
   3.500   0.9435   0.04579   0.03189  -0.0795  0.4259  1.0002
   3.750   0.9639   0.04797   0.03428  -0.0797  0.4220  1.0002
   4.000   0.9840   0.05021   0.03669  -0.0798  0.4177  1.0002
   4.250   1.0049   0.05254   0.03911  -0.0800  0.4142  1.0002
   4.500   1.0210   0.05539   0.04219  -0.0801  0.4135  1.0002
   4.750   1.0270   0.05911   0.04626  -0.0798  0.4151  1.0002
   5.000   1.0198   0.06405   0.05158  -0.0788  0.4187  1.0002