XFOIL         Version 6.94
  
 Calculated polar for: NASA/LANGLEY SC(2)-0714 SUPERCR                 
  
 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   0.4478   0.01615   0.01172  -0.1056  0.8977  0.7011
   0.500   0.5092   0.01541   0.01106  -0.1048  0.8775  0.7041
   1.000   0.5716   0.01474   0.01041  -0.1045  0.8410  0.7076
   1.500   0.6355   0.01436   0.00976  -0.1049  0.7453  0.7121
   2.000   0.6723   0.01607   0.00986  -0.1013  0.4435  0.7168
   2.500   0.7204   0.01800   0.01046  -0.1018  0.2040  0.7224
   3.000   0.7638   0.01897   0.01103  -0.0995  0.1357  0.7248
   3.500   0.8111   0.01977   0.01167  -0.0979  0.1088  0.7270
   4.000   0.8570   0.02073   0.01252  -0.0959  0.0941  0.7298
   4.500   0.9058   0.02167   0.01341  -0.0947  0.0844  0.7332
   5.000   0.9590   0.02250   0.01423  -0.0946  0.0769  0.7366
   6.000   1.0708   0.02482   0.01652  -0.0958  0.0655  0.7440
   6.500   1.1197   0.02598   0.01767  -0.0948  0.0613  0.7468
   8.000   1.2607   0.03068   0.02275  -0.0902  0.0514  0.7554
   8.500   1.3073   0.03236   0.02456  -0.0890  0.0489  0.7590
   9.000   1.3548   0.03484   0.02716  -0.0883  0.0467  0.7628
   9.500   1.3944   0.03712   0.02977  -0.0861  0.0444  0.7665
  10.000   1.4264   0.03923   0.03208  -0.0826  0.0429  0.7698
  10.500   1.4580   0.04287   0.03591  -0.0796  0.0416  0.7730
  11.000   1.4540   0.04637   0.04004  -0.0709  0.0408  0.7763
  11.500   1.4451   0.05127   0.04552  -0.0633  0.0399  0.7797
  12.000   1.4253   0.05729   0.05209  -0.0563  0.0393  0.7831
  12.500   1.3961   0.06441   0.05970  -0.0511  0.0387  0.7860
  13.000   1.3498   0.07427   0.07008  -0.0486  0.0384  0.7881
  13.500   1.1959   0.10470   0.10153  -0.0625  0.0397  0.7877