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 =     0.045 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
  -3.000   0.1125   0.04905   0.04087  -0.0759  0.4912  0.2663
  -2.750   0.1452   0.04703   0.03857  -0.0799  0.4897  0.2646
  -2.500   0.1776   0.04541   0.03664  -0.0831  0.4886  0.2629
  -2.250   0.2076   0.04444   0.03546  -0.0852  0.4875  0.2645
  -2.000   0.2364   0.04395   0.03478  -0.0867  0.4863  0.2703
  -1.750   0.2662   0.04356   0.03410  -0.0884  0.4849  0.2758
  -1.500   0.2943   0.04334   0.03370  -0.0897  0.4837  0.2812
  -1.250   0.3192   0.04361   0.03402  -0.0902  0.4830  0.2913
  -1.000   0.3461   0.04410   0.03434  -0.0915  0.4833  0.3037
  -0.750   0.3696   0.04499   0.03534  -0.0925  0.4852  0.3168
  -0.500   0.3930   0.04610   0.03651  -0.0937  0.4878  0.3348
  -0.250   0.4168   0.04737   0.03781  -0.0949  0.4908  0.3598
   0.000   0.4411   0.04859   0.03908  -0.0959  0.4937  0.3970
   0.250   0.4652   0.04961   0.04031  -0.0960  0.4961  0.4609