XFOIL         Version 6.94
  
 Calculated polar for: NPL 9510 AIRFOIL                                
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.080     Re =     0.450 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
   1.000   0.3265   0.01208   0.00401  -0.0812  0.0671  0.5892
   1.500   0.3814   0.01265   0.00456  -0.0807  0.0562  0.5966
   2.000   0.4361   0.01330   0.00518  -0.0802  0.0502  0.6039
   2.500   0.4902   0.01413   0.00598  -0.0795  0.0463  0.6112
   3.000   0.5439   0.01496   0.00679  -0.0789  0.0427  0.6189
   3.500   0.5977   0.01592   0.00779  -0.0781  0.0400  0.6264
   4.500   0.7063   0.01858   0.01058  -0.0768  0.0352  0.6413
   5.000   0.7605   0.02013   0.01241  -0.0760  0.0332  0.6509
   5.500   0.8132   0.02213   0.01467  -0.0751  0.0313  0.6592
   6.000   0.8635   0.02528   0.01834  -0.0735  0.0300  0.6674
   6.500   0.9076   0.02983   0.02355  -0.0711  0.0295  0.6759
   7.000   0.9515   0.03290   0.02679  -0.0696  0.0287  0.6862
   7.500   0.9800   0.03989   0.03447  -0.0657  0.0287  0.6951
   8.500   0.9100   0.07554   0.07255  -0.0516  0.0328  0.7054
   9.000   0.8920   0.08264   0.07984  -0.0489  0.0325  0.7142
   9.500   0.8637   0.09199   0.08932  -0.0505  0.0323  0.7226
  10.000   0.8469   0.10353   0.10093  -0.0579  0.0322  0.7309