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 =     0.040 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -1.000  -0.1131   0.03961   0.03030  -0.0114   1.0000   0.1316
   0.000  -0.0539   0.03320   0.02324  -0.0051   1.0000   0.2212
   0.500   0.0079   0.03190   0.02138  -0.0089   0.9929   0.2788
   1.000   0.1867   0.03209   0.02079  -0.0347   0.9355   0.3928
   1.500   0.3733   0.03267   0.02167  -0.0625   0.8906   1.0000
   2.000   0.4348   0.03364   0.02224  -0.0660   0.8343   1.0000
   2.500   0.5526   0.03525   0.02358  -0.0784   0.7971   1.0000
   3.000   0.5551   0.03716   0.02560  -0.0726   0.7553   1.0000
   3.500   0.6696   0.03558   0.02385  -0.0801   0.7156   1.0000
   4.000   0.7137   0.03694   0.02509  -0.0786   0.6748   1.0000
   4.500   0.7775   0.03644   0.02447  -0.0784   0.6323   1.0000
   5.000   0.8532   0.03855   0.02648  -0.0813   0.6003   1.0000
   5.500   0.8405   0.04295   0.03096  -0.0750   0.5821   1.0000
   6.000   0.8144   0.04885   0.03694  -0.0685   0.5623   1.0000
   6.500   0.8230   0.05323   0.04141  -0.0658   0.5414   1.0000
   7.500   0.8984   0.05951   0.04788  -0.0608   0.4883   1.0000
   8.000   0.8370   0.07034   0.05875  -0.0588   0.4625   1.0000
   8.500   0.8186   0.07893   0.06739  -0.0587   0.4482   1.0000
   9.000   0.7989   0.08850   0.07702  -0.0601   0.4498   1.0000
   9.500   0.7837   0.09890   0.08750  -0.0627   0.4678   1.0000