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 =     0.040 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
   0.500   0.5884   0.02770   0.01504  -0.0672  0.4222  1.0002
   0.750   0.6203   0.02864   0.01559  -0.0681  0.4119  1.0002
   1.000   0.6537   0.02952   0.01624  -0.0693  0.4040  1.0002
   1.250   0.6872   0.03042   0.01697  -0.0706  0.3969  1.0002
   1.500   0.7218   0.03147   0.01779  -0.0722  0.3909  1.0002
   1.750   0.7552   0.03260   0.01878  -0.0736  0.3864  1.0002
   2.000   0.7857   0.03362   0.01986  -0.0746  0.3828  1.0002
   2.250   0.8157   0.03475   0.02104  -0.0755  0.3796  1.0002
   2.500   0.8449   0.03599   0.02233  -0.0763  0.3765  1.0002
   2.750   0.8732   0.03734   0.02370  -0.0770  0.3726  1.0002
   3.000   0.9009   0.03890   0.02520  -0.0777  0.3683  1.0002
   3.250   0.9268   0.04057   0.02692  -0.0782  0.3644  1.0002
   3.500   0.9507   0.04205   0.02863  -0.0784  0.3612  1.0002
   3.750   0.9749   0.04377   0.03055  -0.0787  0.3598  1.0002
   4.000   0.9982   0.04566   0.03267  -0.0790  0.3591  1.0002
   4.250   1.0205   0.04769   0.03494  -0.0793  0.3589  1.0002
   4.500   1.0409   0.04995   0.03743  -0.0794  0.3591  1.0002
   4.750   1.0595   0.05245   0.04017  -0.0794  0.3599  1.0002
   5.000   1.0759   0.05522   0.04324  -0.0792  0.3612  1.0002