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 =     0.035 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
  -3.000  -0.0341   0.03722   0.02835  -0.0327  0.9997  0.2921
  -2.750  -0.0115   0.03648   0.02755  -0.0322  0.9997  0.3057
  -2.500   0.0117   0.03578   0.02688  -0.0318  0.9997  0.3214
  -2.250   0.0369   0.03534   0.02644  -0.0317  0.9997  0.3411
  -2.000   0.0623   0.03503   0.02631  -0.0316  0.9997  0.3654
  -1.750   0.2228   0.02824   0.01928  -0.0382  0.4888  0.5527
  -1.500   0.2527   0.02718   0.01798  -0.0351  0.4171  1.0003
  -1.250   0.2836   0.02848   0.01821  -0.0347  0.3834  1.0003
  -1.000   0.3140   0.02967   0.01873  -0.0344  0.3608  1.0003
  -0.750   0.3445   0.03090   0.01941  -0.0342  0.3445  1.0003
  -0.500   0.3765   0.03206   0.02033  -0.0343  0.3332  1.0003
  -0.250   0.4081   0.03347   0.02139  -0.0344  0.3254  1.0003
   0.000   0.4402   0.03477   0.02270  -0.0347  0.3192  1.0003
   0.250   0.4715   0.03623   0.02405  -0.0350  0.3129  1.0003
   0.500   0.5014   0.03795   0.02552  -0.0351  0.3066  1.0003
   0.750   0.5318   0.03964   0.02731  -0.0356  0.3015  1.0003
   1.000   0.5618   0.04147   0.02927  -0.0362  0.2975  1.0003
   1.250   0.5915   0.04357   0.03150  -0.0369  0.2958  1.0003
   1.500   0.6205   0.04593   0.03402  -0.0377  0.2952  1.0003
   1.750   0.6484   0.04858   0.03685  -0.0387  0.2953  1.0003
   2.000   0.6746   0.05147   0.03994  -0.0397  0.2951  1.0003
   2.250   0.6985   0.05469   0.04337  -0.0409  0.2948  1.0003
   2.500   0.7197   0.05838   0.04731  -0.0424  0.2955  1.0003
   2.750   0.7380   0.06271   0.05188  -0.0441  0.2989  1.0003
   3.000   0.7575   0.06695   0.05617  -0.0453  0.3035  1.0003