XFOIL         Version 6.94
  
 Calculated polar for: AG38                                            
  
 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   0.3605   0.00657   0.00185  -0.0481  0.8188  1.0000
   0.500   0.4113   0.00702   0.00188  -0.0463  0.7324  1.0000
   1.000   0.4631   0.00750   0.00196  -0.0452  0.6446  1.0000
   1.500   0.5159   0.00807   0.00218  -0.0444  0.5859  1.0000
   2.000   0.5697   0.00856   0.00248  -0.0440  0.5397  1.0000
   2.500   0.6237   0.00900   0.00280  -0.0437  0.4909  1.0000
   3.000   0.6773   0.00944   0.00310  -0.0433  0.4249  1.0000
   3.500   0.7294   0.01018   0.00351  -0.0428  0.3388  1.0000
   4.000   0.7808   0.01110   0.00413  -0.0425  0.2551  1.0000
   4.500   0.8317   0.01216   0.00488  -0.0421  0.1786  1.0000
   5.000   0.8812   0.01350   0.00590  -0.0416  0.1083  1.0000
   5.500   0.9297   0.01503   0.00730  -0.0407  0.0690  1.0000
   6.000   0.9762   0.01679   0.00906  -0.0395  0.0525  1.0000
   6.500   1.0217   0.01867   0.01106  -0.0382  0.0433  1.0000
   7.000   1.0659   0.02074   0.01323  -0.0367  0.0373  1.0000
   7.500   1.1092   0.02325   0.01604  -0.0349  0.0331  1.0000
   8.000   1.1500   0.02615   0.01909  -0.0333  0.0293  1.0000
   8.500   1.1875   0.02998   0.02348  -0.0310  0.0270  1.0000
   9.000   1.2214   0.03329   0.02725  -0.0287  0.0243  1.0000
   9.500   1.2444   0.03847   0.03282  -0.0258  0.0227  1.0000
  10.000   1.2413   0.04550   0.04069  -0.0208  0.0216  1.0000
  10.500   1.2182   0.05190   0.04773  -0.0146  0.0211  1.0000
  11.000   1.1790   0.06062   0.05696  -0.0130  0.0210  1.0000
  11.500   1.1355   0.07367   0.07044  -0.0201  0.0212  1.0000
  12.000   1.0879   0.09151   0.08856  -0.0326  0.0218  1.0000