XFOIL         Version 6.94
  
 Calculated polar for: I.S.A. 571                                      
  
 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
 ------- -------- --------- --------- -------- ------- -------
   0.000   0.4949   0.00825   0.00160  -0.0873  0.6612  0.2472
   0.500   0.5393   0.00695   0.00179  -0.0849  0.5810  1.0000
   1.000   0.5898   0.00752   0.00190  -0.0838  0.4959  1.0000
   1.500   0.6409   0.00812   0.00209  -0.0829  0.4336  1.0000
   2.000   0.6938   0.00858   0.00234  -0.0824  0.3995  1.0000
   2.500   0.7471   0.00899   0.00261  -0.0820  0.3749  1.0000
   3.000   0.8008   0.00936   0.00290  -0.0817  0.3554  1.0000
   3.500   0.8541   0.00975   0.00324  -0.0813  0.3388  1.0000
   4.000   0.9077   0.01010   0.00361  -0.0810  0.3219  1.0000
   4.500   0.9606   0.01052   0.00403  -0.0805  0.3028  1.0000
   5.000   1.0133   0.01092   0.00442  -0.0801  0.2715  1.0000
   5.500   1.0635   0.01161   0.00492  -0.0794  0.2137  1.0000
   6.000   1.1108   0.01266   0.00574  -0.0783  0.1645  1.0000
   6.500   1.1527   0.01443   0.00700  -0.0765  0.0745  1.0000
   7.000   1.1899   0.01689   0.00930  -0.0735  0.0167  1.0000
   7.500   1.2304   0.01869   0.01137  -0.0709  0.0136  1.0000
   8.000   1.2652   0.02087   0.01380  -0.0677  0.0119  1.0000
   8.500   1.2809   0.02473   0.01791  -0.0617  0.0105  1.0000
   9.000   1.3069   0.02744   0.02086  -0.0573  0.0099  1.0000
   9.500   1.3271   0.03077   0.02443  -0.0523  0.0095  1.0000
  10.000   1.3424   0.03473   0.02871  -0.0468  0.0093  1.0000
  10.500   1.3528   0.03863   0.03295  -0.0414  0.0089  1.0000
  11.000   1.3534   0.04371   0.03843  -0.0360  0.0088  1.0000
  11.500   1.3379   0.05104   0.04627  -0.0311  0.0094  1.0000