XFOIL         Version 6.94
  
 Calculated polar for: TSAGI 12% AIRFOIL                               
  
 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
 ------- -------- --------- --------- -------- ------- -------
   0.500   0.2906   0.00949   0.00438  -0.0298  0.6181  0.9464
   1.000   0.3779   0.01003   0.00479  -0.0361  0.5941  0.9601
   1.500   0.4480   0.01004   0.00471  -0.0395  0.5688  0.9650
   2.500   0.5600   0.01047   0.00450  -0.0404  0.4292  0.9775
   3.000   0.6217   0.01075   0.00450  -0.0423  0.3679  0.9826
   3.500   0.6721   0.01146   0.00475  -0.0420  0.2674  0.9891
   4.000   0.7196   0.01354   0.00563  -0.0423  0.0239  0.9945
   4.500   0.7804   0.01394   0.00602  -0.0439  0.0060  0.9989
   5.000   0.8274   0.01435   0.00654  -0.0426  0.0064  1.0000
   5.500   0.8655   0.01487   0.00718  -0.0393  0.0073  1.0000
   6.000   0.9007   0.01559   0.00805  -0.0355  0.0082  1.0000
   6.500   0.9332   0.01644   0.00907  -0.0311  0.0094  1.0000
   7.000   0.9581   0.01769   0.01049  -0.0254  0.0109  1.0000
   7.500   0.9682   0.01953   0.01250  -0.0173  0.0120  1.0000
   8.000   0.9716   0.02139   0.01448  -0.0080  0.0130  1.0000
   8.500   0.9613   0.02373   0.01693   0.0036  0.0146  1.0000