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 =     0.040 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
  -3.000   0.5286   0.03606   0.02382  -0.0371  0.2967  0.9892
  -2.750   0.5761   0.03653   0.02417  -0.0422  0.2934  1.0018
  -2.500   0.5668   0.03547   0.02311  -0.0385  0.2927  1.0019
  -2.250   0.5919   0.03557   0.02300  -0.0410  0.2909  1.0019
  -2.000   0.6310   0.03635   0.02349  -0.0447  0.2889  1.0019
  -1.750   0.6691   0.03727   0.02416  -0.0476  0.2873  1.0019
  -1.500   0.7057   0.03828   0.02494  -0.0498  0.2860  1.0019
  -1.250   0.7409   0.03936   0.02584  -0.0516  0.2851  1.0019
  -1.000   0.7747   0.04051   0.02686  -0.0530  0.2845  1.0019
  -0.750   0.8073   0.04172   0.02799  -0.0543  0.2842  1.0019
  -0.500   0.8386   0.04302   0.02924  -0.0553  0.2840  1.0019
  -0.250   0.8688   0.04443   0.03064  -0.0561  0.2836  1.0019
   0.000   0.8978   0.04594   0.03216  -0.0567  0.2831  1.0019
   0.250   0.9255   0.04760   0.03384  -0.0573  0.2824  1.0019
   0.500   0.9519   0.04937   0.03565  -0.0576  0.2818  1.0019
   0.750   0.9769   0.05127   0.03762  -0.0578  0.2813  1.0019
   1.000   0.9996   0.05322   0.03971  -0.0577  0.2813  1.0019
   1.250   1.0189   0.05525   0.04195  -0.0572  0.2820  1.0019
   1.500   1.0293   0.05767   0.04480  -0.0557  0.2839  1.0019
   1.750   1.0292   0.06100   0.04861  -0.0534  0.2867  1.0019
   2.000   1.0196   0.06534   0.05340  -0.0507  0.2900  1.0019
   2.250   1.0004   0.07067   0.05910  -0.0477  0.2938  1.0019
   2.500   0.9881   0.07593   0.06454  -0.0457  0.2973  1.0019
   2.750   0.9816   0.08108   0.06981  -0.0445  0.3005  1.0019