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 =     0.035 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
  -3.000   0.0986   0.05531   0.04742  -0.0721  0.5470  0.3219
  -2.750   0.1290   0.05332   0.04525  -0.0758  0.5451  0.3164
  -2.500   0.1620   0.05161   0.04325  -0.0802  0.5436  0.3136
  -2.250   0.1933   0.05057   0.04195  -0.0837  0.5427  0.3166
  -2.000   0.2246   0.04981   0.04093  -0.0873  0.5424  0.3195
  -1.750   0.2563   0.04938   0.04017  -0.0908  0.5424  0.3235
  -1.500   0.2803   0.04949   0.04032  -0.0919  0.5425  0.3321
  -1.250   0.3072   0.04985   0.04051  -0.0940  0.5425  0.3430
  -1.000   0.3325   0.05035   0.04089  -0.0958  0.5426  0.3539
  -0.750   0.3582   0.05122   0.04157  -0.0976  0.5428  0.3715
  -0.500   0.3702   0.05341   0.04403  -0.1012  0.5487  0.3843
  -0.250   0.3833   0.05593   0.04657  -0.1047  0.5574  0.4001
   0.000   0.4047   0.05751   0.04815  -0.1063  0.5631  0.4301