XFOIL         Version 6.94
  
 Calculated polar for: bbl2                                            
  
 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   0.5422   0.03599   0.02369  -0.0406  0.3240  1.0019
  -2.750   0.5354   0.03516   0.02276  -0.0371  0.3220  1.0019
  -2.500   0.5529   0.03521   0.02250  -0.0379  0.3188  1.0019
  -2.250   0.5863   0.03588   0.02291  -0.0407  0.3153  1.0019
  -2.000   0.6215   0.03674   0.02351  -0.0431  0.3123  1.0019
  -1.750   0.6569   0.03768   0.02421  -0.0453  0.3097  1.0019
  -1.500   0.6926   0.03869   0.02501  -0.0472  0.3076  1.0019
  -1.250   0.7281   0.03978   0.02592  -0.0491  0.3059  1.0019
  -1.000   0.7629   0.04095   0.02695  -0.0507  0.3046  1.0019
  -0.750   0.7966   0.04221   0.02811  -0.0521  0.3037  1.0019
  -0.500   0.8289   0.04357   0.02941  -0.0533  0.3031  1.0019
  -0.250   0.8596   0.04503   0.03088  -0.0542  0.3028  1.0019
   0.000   0.8882   0.04660   0.03250  -0.0549  0.3029  1.0019
   0.250   0.9151   0.04833   0.03433  -0.0554  0.3029  1.0019
   0.500   0.9400   0.05022   0.03633  -0.0556  0.3029  1.0019
   0.750   0.9626   0.05231   0.03856  -0.0557  0.3029  1.0019
   1.000   0.9821   0.05463   0.04106  -0.0554  0.3028  1.0019
   1.250   0.9973   0.05727   0.04392  -0.0547  0.3031  1.0019
   1.500   1.0031   0.06055   0.04755  -0.0533  0.3040  1.0019
   1.750   0.9716   0.06667   0.05437  -0.0491  0.3074  1.0019
   2.000   0.8736   0.07920   0.06763  -0.0428  0.3154  1.0019
   2.250   0.8413   0.08704   0.07557  -0.0415  0.3196  1.0019