XFOIL         Version 6.94
  
 Calculated polar for: USA 25 AIRFOIL                                  
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.080     Re =     0.450 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
   0.500   0.4715   0.01091   0.00278  -0.0731  0.4570  0.0946
   1.000   0.5226   0.01113   0.00300  -0.0721  0.4488  0.1218
   1.500   0.5735   0.01118   0.00310  -0.0710  0.4418  0.1405
   2.000   0.6236   0.01128   0.00324  -0.0699  0.4344  0.1644
   2.500   0.6746   0.01156   0.00350  -0.0690  0.4254  0.1942
   3.000   0.7241   0.01155   0.00365  -0.0677  0.4199  0.2310
   3.500   0.7660   0.01140   0.00355  -0.0649  0.3965  0.2876
   4.000   0.8947   0.01058   0.00387  -0.0813  0.3111  0.9950
   4.500   0.9418   0.01243   0.00473  -0.0808  0.1488  1.0000
   5.000   0.9548   0.01431   0.00594  -0.0728  0.0063  1.0000
   5.500   0.9894   0.01478   0.00647  -0.0685  0.0064  1.0000
   6.000   1.0240   0.01530   0.00705  -0.0642  0.0070  1.0000
   6.500   1.0574   0.01595   0.00781  -0.0597  0.0079  1.0000
   7.000   1.0894   0.01664   0.00862  -0.0549  0.0090  1.0000
   7.500   1.1145   0.01750   0.00961  -0.0488  0.0103  1.0000
   8.000   1.1388   0.01850   0.01078  -0.0428  0.0123  1.0000
   8.500   1.1628   0.01963   0.01205  -0.0369  0.0146  1.0000
   9.000   1.1816   0.02110   0.01369  -0.0305  0.0171  1.0000
   9.500   1.1951   0.02294   0.01572  -0.0238  0.0198  1.0000
  10.000   1.2015   0.02541   0.01838  -0.0169  0.0230  1.0000
  10.500   1.2030   0.02866   0.02179  -0.0104  0.0268  1.0000
  11.000   1.2184   0.03185   0.02510  -0.0053  0.0304  1.0000
  11.500   1.3723   0.03733   0.03043  -0.0146  0.0347  1.0000
  12.000   1.3544   0.03923   0.03281  -0.0044  0.0331  1.0000