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 =     0.060 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -1.500  -0.1693   0.04600   0.03933  -0.0116   0.9942   0.2144
  -0.500   0.0430   0.03118   0.02228  -0.0274   0.9416   0.1725
   0.000   0.1308   0.02801   0.01848  -0.0339   0.9132   0.2258
   0.500   0.2421   0.02682   0.01687  -0.0456   0.8930   0.2928
   1.000   0.3028   0.02565   0.01571  -0.0474   0.8537   0.3627
   1.500   0.5112   0.02334   0.01394  -0.0775   0.8225   1.0000
   2.000   0.6108   0.02187   0.01238  -0.0845   0.7792   1.0000
   2.500   0.6949   0.02011   0.01034  -0.0879   0.7247   1.0000
   3.000   0.7552   0.02032   0.01026  -0.0887   0.6740   1.0000
   3.500   0.8087   0.02052   0.01014  -0.0880   0.6289   1.0000
   4.000   0.8582   0.02178   0.01116  -0.0877   0.5884   1.0000
   4.500   0.8976   0.02275   0.01208  -0.0853   0.5540   1.0000
   5.000   0.9458   0.02372   0.01282  -0.0839   0.5209   1.0000
   5.500   0.9846   0.02639   0.01552  -0.0824   0.4944   1.0000
   6.000   1.0219   0.02359   0.01133  -0.0754   0.3988   1.0000
   6.500   1.0474   0.02337   0.01134  -0.0691   0.3582   1.0000
   7.500   1.0600   0.02889   0.01473  -0.0523   0.0634   1.0000
   8.000   1.0670   0.03239   0.01853  -0.0442   0.0523   1.0000
   8.500   1.0608   0.03589   0.02255  -0.0349   0.0517   1.0000
   9.000   1.0576   0.03973   0.02680  -0.0271   0.0529   1.0000
   9.500   1.1418   0.04600   0.03414  -0.0249   0.0641   1.0000
  10.000   1.1613   0.05702   0.04618  -0.0207   0.0728   1.0000