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 =     0.045 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
  -3.000   0.4674   0.03450   0.02262  -0.0267  0.2789  0.9542
  -2.750   0.5137   0.03494   0.02295  -0.0312  0.2766  0.9684
  -2.500   0.5655   0.03549   0.02337  -0.0368  0.2745  0.9865
  -2.250   0.6005   0.03554   0.02330  -0.0405  0.2731  1.0019
  -2.000   0.6361   0.03603   0.02353  -0.0450  0.2720  1.0019
  -1.750   0.6766   0.03696   0.02420  -0.0488  0.2710  1.0019
  -1.500   0.7137   0.03798   0.02500  -0.0514  0.2704  1.0019
  -1.250   0.7486   0.03905   0.02592  -0.0532  0.2700  1.0019
  -1.000   0.7817   0.04019   0.02694  -0.0546  0.2696  1.0019
  -0.750   0.8134   0.04140   0.02808  -0.0557  0.2690  1.0019
  -0.500   0.8438   0.04269   0.02932  -0.0566  0.2682  1.0019
  -0.250   0.8731   0.04405   0.03068  -0.0573  0.2673  1.0019
   0.000   0.9013   0.04550   0.03214  -0.0578  0.2666  1.0019
   0.250   0.9283   0.04703   0.03372  -0.0581  0.2659  1.0019
   0.500   0.9538   0.04864   0.03542  -0.0583  0.2658  1.0019
   0.750   0.9773   0.05036   0.03730  -0.0582  0.2666  1.0019
   1.000   0.9987   0.05225   0.03939  -0.0578  0.2680  1.0019
   1.250   1.0183   0.05436   0.04170  -0.0573  0.2696  1.0019
   1.500   1.0362   0.05669   0.04425  -0.0567  0.2714  1.0019
   1.750   1.0523   0.05928   0.04703  -0.0561  0.2733  1.0019
   2.000   1.0672   0.06214   0.05004  -0.0554  0.2753  1.0019
   2.250   1.0834   0.06526   0.05326  -0.0550  0.2773  1.0019
   2.500   1.1039   0.06864   0.05668  -0.0553  0.2791  1.0019
   2.750   0.8531   0.09041   0.08024  -0.0394  0.2979  1.0019
   3.000   0.8898   0.09210   0.08185  -0.0400  0.3002  1.0019