XFOIL         Version 6.94
  
 Calculated polar for: GIII BL145 AIRFOIL                              
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.080     Re =     0.300 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
   0.000   0.1198   0.00782   0.00320  -0.0180  0.8653  0.9598
   0.500   0.1938   0.00784   0.00311  -0.0217  0.8451  0.9739
   1.000   0.2651   0.00780   0.00301  -0.0253  0.8251  0.9840
   1.500   0.3381   0.00776   0.00294  -0.0294  0.8058  0.9939
   2.000   0.4024   0.00773   0.00291  -0.0318  0.7852  1.0000
   2.500   0.4516   0.00773   0.00300  -0.0308  0.7617  1.0000
   3.000   0.4954   0.00764   0.00273  -0.0280  0.6940  1.0000
   3.500   0.5368   0.00805   0.00262  -0.0250  0.5537  1.0000
   4.000   0.5670   0.01018   0.00320  -0.0213  0.2327  1.0000
   4.500   0.6029   0.01202   0.00411  -0.0187  0.0532  1.0000
   5.000   0.6478   0.01281   0.00488  -0.0170  0.0297  1.0000
   5.500   0.6940   0.01358   0.00567  -0.0155  0.0174  1.0000
   6.000   0.7399   0.01454   0.00676  -0.0139  0.0088  1.0000
   6.500   0.7834   0.01605   0.00849  -0.0118  0.0053  1.0000
   7.000   0.8235   0.01829   0.01106  -0.0091  0.0049  1.0000
   7.500   0.8592   0.02209   0.01534  -0.0058  0.0050  1.0000
   8.000   0.8907   0.02860   0.02273  -0.0019  0.0054  1.0000
   8.500   0.8887   0.04321   0.03892   0.0055  0.0072  1.0000