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 =     0.050 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -2.000  -0.2394   0.05089   0.04440  -0.0037   1.0000   0.2455
  -1.500  -0.2070   0.04686   0.04019  -0.0030   1.0000   0.2742
   0.000   0.0415   0.03143   0.02189  -0.0211   0.9556   0.2343
   0.500   0.1611   0.02979   0.01939  -0.0339   0.9187   0.2936
   1.000   0.2751   0.02875   0.01825  -0.0463   0.8880   0.3961
   1.500   0.4337   0.02792   0.01788  -0.0683   0.8415   1.0000
   2.000   0.5719   0.02704   0.01660  -0.0830   0.8047   1.0000
   2.500   0.6111   0.02714   0.01675  -0.0801   0.7493   1.0000
   3.000   0.7357   0.02443   0.01369  -0.0893   0.7050   1.0000
   3.500   0.7663   0.02548   0.01462  -0.0854   0.6631   1.0000
   4.000   0.8292   0.02538   0.01428  -0.0858   0.6246   1.0000
   4.500   0.8750   0.02737   0.01614  -0.0851   0.5884   1.0000
   5.000   0.9036   0.02930   0.01813  -0.0817   0.5588   1.0000
   5.500   0.9506   0.03029   0.01908  -0.0802   0.5316   1.0000
   6.000   0.9979   0.03190   0.02048  -0.0783   0.4952   1.0000
   6.500   1.0440   0.02671   0.01399  -0.0698   0.3947   1.0000
   7.500   1.0522   0.02867   0.01478  -0.0503   0.1744   1.0000
   8.000   1.0501   0.03300   0.01848  -0.0420   0.0688   1.0000
   8.500   1.0477   0.03678   0.02261  -0.0338   0.0616   1.0000
   9.000   1.0399   0.04099   0.02722  -0.0264   0.0602   1.0000
   9.500   1.0457   0.04540   0.03213  -0.0201   0.0611   1.0000
  10.000   1.1335   0.05501   0.04307  -0.0178   0.0708   1.0000