XFOIL         Version 6.94
  
 Calculated polar for: ARA-D BL 12%                                    
  
 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
 ------- -------- --------- --------- -------- ------- -------
  -3.000  -0.1882   0.04308   0.03504  -0.0171  1.0000  0.2976
  -2.750  -0.1629   0.04121   0.03287  -0.0180  1.0000  0.3061
  -2.500  -0.1397   0.03945   0.03101  -0.0182  1.0000  0.3159
  -2.250  -0.1161   0.03817   0.02963  -0.0182  1.0000  0.3294
  -2.000  -0.0918   0.03697   0.02838  -0.0182  1.0000  0.3450
  -1.750  -0.0653   0.03607   0.02738  -0.0186  1.0000  0.3649
  -1.500  -0.0390   0.03536   0.02673  -0.0188  1.0000  0.3893
  -1.000   0.2266   0.02682   0.01830  -0.0309  0.4681  1.0000
  -0.750   0.2547   0.02807   0.01847  -0.0300  0.4306  1.0000
  -0.500   0.2831   0.02924   0.01897  -0.0293  0.4057  1.0000
  -0.250   0.3115   0.03041   0.01961  -0.0288  0.3866  1.0000
   0.000   0.3401   0.03161   0.02035  -0.0283  0.3711  1.0000
   0.250   0.3700   0.03288   0.02151  -0.0283  0.3603  1.0000
   0.500   0.3995   0.03426   0.02268  -0.0282  0.3525  1.0000
   0.750   0.4292   0.03587   0.02412  -0.0283  0.3469  1.0000
   1.000   0.4581   0.03761   0.02598  -0.0287  0.3423  1.0000
   1.250   0.4856   0.03946   0.02786  -0.0290  0.3370  1.0000
   1.500   0.5123   0.04132   0.02960  -0.0289  0.3314  1.0000
   1.750   0.5383   0.04349   0.03160  -0.0288  0.3269  1.0000
   2.000   0.5628   0.04603   0.03435  -0.0295  0.3255  1.0000
   2.250   0.5860   0.04891   0.03738  -0.0302  0.3257  1.0000
   2.500   0.6077   0.05206   0.04064  -0.0310  0.3267  1.0000
   2.750   0.6214   0.05633   0.04537  -0.0332  0.3308  1.0000
   3.000   0.6222   0.06243   0.05190  -0.0367  0.3385  1.0000
   3.250   0.6280   0.06747   0.05704  -0.0387  0.3436  1.0000
   3.500   0.6392   0.07187   0.06141  -0.0398  0.3471  1.0000