XFOIL         Version 6.94
  
 Calculated polar for: Curtis C-72                                     
  
 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   0.6765   0.00836   0.00239  -0.0993  0.5935  0.4851
   0.500   0.7284   0.00824   0.00255  -0.0989  0.5730  0.5860
   1.000   0.7723   0.00826   0.00271  -0.0968  0.5142  0.7527
   2.000   0.8800   0.00900   0.00322  -0.0971  0.4089  1.0000
   2.500   0.9300   0.00950   0.00353  -0.0966  0.3749  1.0000
   3.000   0.9748   0.01035   0.00398  -0.0953  0.3253  1.0000
   4.000   1.0778   0.01132   0.00497  -0.0950  0.2956  1.0000
   4.500   1.1307   0.01159   0.00532  -0.0952  0.2940  1.0000
   5.000   1.1822   0.01188   0.00563  -0.0952  0.2831  1.0000
   5.500   1.2266   0.01273   0.00622  -0.0941  0.2411  1.0000
   6.000   1.2620   0.01418   0.00713  -0.0918  0.1719  1.0000
   6.500   1.2945   0.01583   0.00835  -0.0891  0.1268  1.0000
   7.000   1.3308   0.01718   0.00949  -0.0870  0.0911  1.0000
   7.500   1.3193   0.02175   0.01376  -0.0775  0.0134  1.0000
   8.000   1.3248   0.02470   0.01711  -0.0710  0.0122  1.0000
   8.500   1.3266   0.02834   0.02105  -0.0656  0.0121  1.0000
   9.000   1.3324   0.03207   0.02502  -0.0613  0.0122  1.0000
   9.500   1.3328   0.03633   0.02950  -0.0576  0.0125  1.0000
  10.000   1.3437   0.04019   0.03352  -0.0552  0.0130  1.0000
  10.500   1.3530   0.04443   0.03795  -0.0532  0.0140  1.0000
  11.000   1.3516   0.05019   0.04410  -0.0505  0.0161  1.0000