XFOIL         Version 6.96
  
 Calculated polar for: USA 5 AIRFOIL
                                  
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.055 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -1.500  -0.2019   0.04561   0.03909  -0.0055   1.0000   0.2356
  -0.500  -0.0098   0.03330   0.02456  -0.0195   0.9615   0.1776
   0.000   0.0849   0.02968   0.02014  -0.0272   0.9295   0.2274
   0.500   0.2197   0.02838   0.01818  -0.0431   0.9043   0.2983
   1.000   0.2839   0.02712   0.01694  -0.0458   0.8685   0.3734
   1.500   0.4716   0.02572   0.01603  -0.0728   0.8302   1.0000
   2.000   0.5973   0.02427   0.01443  -0.0847   0.7918   1.0000
   2.500   0.6637   0.02292   0.01288  -0.0851   0.7361   1.0000
   3.000   0.7468   0.02211   0.01171  -0.0889   0.6884   1.0000
   3.500   0.7908   0.02260   0.01205  -0.0869   0.6454   1.0000
   4.000   0.8546   0.02359   0.01263  -0.0884   0.5991   1.0000
   4.500   0.8861   0.02484   0.01394  -0.0850   0.5710   1.0000
   5.000   0.9292   0.02591   0.01495  -0.0830   0.5392   1.0000
   5.500   0.9786   0.02765   0.01651  -0.0822   0.5087   1.0000
   6.000   1.0125   0.02449   0.01228  -0.0742   0.4270   1.0000
   6.500   1.0451   0.02482   0.01249  -0.0693   0.3795   1.0000
   7.000   1.0521   0.02537   0.01238  -0.0596   0.2724   1.0000
   7.500   1.0601   0.02826   0.01446  -0.0515   0.1228   1.0000
   8.000   1.0599   0.03270   0.01847  -0.0433   0.0586   1.0000
   8.500   1.0555   0.03627   0.02250  -0.0345   0.0556   1.0000
   9.000   1.0490   0.04029   0.02696  -0.0269   0.0558   1.0000
   9.500   1.0739   0.04432   0.03156  -0.0206   0.0594   1.0000
  10.000   1.1497   0.05587   0.04452  -0.0193   0.0715   1.0000