XFOIL         Version 6.94
  
 Calculated polar for: UNIVERSITY OF ILLINOIS UI-1720 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
 ------- -------- --------- --------- -------- ------- -------
  -3.000   0.1042   0.05224   0.04426  -0.0730  0.5161  0.2928
  -2.750   0.1375   0.05004   0.04181  -0.0779  0.5146  0.2882
  -2.500   0.1728   0.04807   0.03947  -0.0829  0.5135  0.2836
  -2.250   0.2014   0.04729   0.03856  -0.0850  0.5128  0.2872
  -2.000   0.2304   0.04677   0.03785  -0.0872  0.5122  0.2924
  -1.750   0.2603   0.04638   0.03721  -0.0895  0.5114  0.2967
  -1.500   0.2891   0.04628   0.03687  -0.0916  0.5107  0.3032
  -1.250   0.3126   0.04669   0.03735  -0.0924  0.5102  0.3137
  -1.000   0.3396   0.04724   0.03766  -0.0943  0.5102  0.3256
  -0.750   0.3619   0.04811   0.03862  -0.0953  0.5115  0.3390
  -0.500   0.3858   0.04926   0.03973  -0.0968  0.5142  0.3584
  -0.250   0.4102   0.05044   0.04088  -0.0980  0.5170  0.3841
   0.000   0.4318   0.05209   0.04260  -0.1002  0.5223  0.4183
   0.250   0.4381   0.05600   0.04681  -0.1054  0.5382  0.4421
   0.500   0.4619   0.05705   0.04824  -0.1055  0.5437  0.5226