XFOIL         Version 6.94
  
 Calculated polar for: P-51D TIP (BL215) AIRFOIL                       
  
 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   0.1272   0.01372   0.00832  -0.0293  0.7760  0.8204
   0.500   0.1836   0.01396   0.00859  -0.0306  0.7732  0.8232
   1.000   0.2404   0.01419   0.00884  -0.0320  0.7708  0.8257
   1.500   0.2976   0.01434   0.00903  -0.0332  0.7683  0.8284
   2.000   0.3526   0.01405   0.00878  -0.0329  0.7586  0.8304
   2.500   0.4078   0.01324   0.00801  -0.0318  0.7435  0.8330
   3.000   0.4640   0.01191   0.00664  -0.0304  0.7201  0.8362
   3.500   0.5183   0.01010   0.00467  -0.0283  0.6562  0.8395
   4.000   0.5258   0.01362   0.00555  -0.0218  0.0825  0.8442
   4.500   0.5715   0.01451   0.00620  -0.0209  0.0311  0.8492
   5.000   0.6197   0.01501   0.00669  -0.0202  0.0155  0.8535
   5.500   0.6655   0.01582   0.00754  -0.0191  0.0061  0.8592
   6.000   0.7095   0.01661   0.00857  -0.0175  0.0055  0.8654
   6.500   0.7499   0.01780   0.01008  -0.0152  0.0053  0.8730
   7.000   0.7839   0.01964   0.01232  -0.0120  0.0053  0.8822
   7.500   0.8200   0.02259   0.01566  -0.0089  0.0056  0.8928
   8.000   0.8686   0.02844   0.02227  -0.0072  0.0063  0.9039
   8.500   0.8929   0.03651   0.03137  -0.0031  0.0071  0.9254
   9.000   0.9194   0.04338   0.03895  -0.0017  0.0082  1.0000
   9.500   0.8838   0.05786   0.05443   0.0062  0.0119  1.0000
  10.000   0.8527   0.06523   0.06210   0.0093  0.0124  1.0000