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 =     0.300 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
   0.000   0.4941   0.00683   0.00222  -0.0879  0.8014  1.0000
   0.500   0.5444   0.00741   0.00208  -0.0860  0.6884  1.0000
   1.000   0.5924   0.00812   0.00221  -0.0841  0.5969  1.0000
   1.500   0.6411   0.00879   0.00239  -0.0826  0.5082  1.0000
   2.000   0.6914   0.00943   0.00266  -0.0815  0.4513  1.0000
   2.500   0.7432   0.01000   0.00300  -0.0808  0.4184  1.0000
   3.000   0.7957   0.01050   0.00339  -0.0803  0.3931  1.0000
   3.500   0.8482   0.01100   0.00382  -0.0798  0.3724  1.0000
   4.000   0.9008   0.01147   0.00429  -0.0793  0.3525  1.0000
   4.500   0.9533   0.01194   0.00480  -0.0788  0.3328  1.0000
   5.000   1.0050   0.01242   0.00534  -0.0781  0.3095  1.0000
   5.500   1.0559   0.01290   0.00585  -0.0774  0.2734  1.0000
   6.000   1.1048   0.01359   0.00641  -0.0765  0.2161  1.0000
   6.500   1.1493   0.01488   0.00744  -0.0750  0.1585  1.0000
   7.000   1.1806   0.01801   0.00971  -0.0716  0.0269  1.0000
   7.500   1.2197   0.02005   0.01207  -0.0687  0.0200  1.0000
   8.000   1.2481   0.02290   0.01523  -0.0645  0.0165  1.0000
   8.500   1.2628   0.02695   0.01952  -0.0584  0.0153  1.0000
   9.000   1.2898   0.02970   0.02251  -0.0542  0.0144  1.0000
   9.500   1.3113   0.03315   0.02625  -0.0494  0.0136  1.0000
  10.000   1.3302   0.03769   0.03118  -0.0445  0.0136  1.0000
  10.500   1.3407   0.04360   0.03758  -0.0394  0.0141  1.0000
  11.000   1.3304   0.05274   0.04733  -0.0337  0.0152  1.0000
  11.500   1.3383   0.05719   0.05205  -0.0298  0.0161  1.0000