XFOIL         Version 6.94
  
 Calculated polar for: NACA 67,1-215                                   
  
 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
 ------- -------- --------- --------- -------- ------- -------
   0.500   0.0548   0.03244   0.02774  -0.0070  0.8229  0.8838
   1.000   0.0887   0.03238   0.02767  -0.0034  0.8157  0.8919
   1.500   0.1380   0.03282   0.02810  -0.0022  0.8103  0.8978
   3.000   0.2648   0.02957   0.02491   0.0034  0.7763  0.9083
   3.500   0.3444   0.02784   0.02324  -0.0005  0.7674  0.9098
   4.500   0.5204   0.02084   0.01632  -0.0075  0.7029  0.9126
   5.000   0.5997   0.01850   0.01284  -0.0098  0.4950  0.9131
   5.500   0.5683   0.01967   0.01353   0.0053  0.3871  0.9167
   6.000   0.5545   0.02090   0.01424   0.0164  0.2704  0.9212
   6.500   0.5548   0.02240   0.01524   0.0246  0.1647  0.9275
   7.000   0.5578   0.02333   0.01606   0.0327  0.0936  0.9317
   7.500   0.5636   0.02408   0.01663   0.0404  0.0652  0.9351
   8.000   0.5683   0.02480   0.01733   0.0484  0.0467  0.9379
   8.500   0.5742   0.02548   0.01810   0.0561  0.0347  0.9406
   9.000   0.5693   0.02683   0.01960   0.0654  0.0281  0.9435
   9.500   0.5834   0.02842   0.02140   0.0717  0.0258  0.9460
  10.000   0.6374   0.03106   0.02437   0.0720  0.0247  0.9473
  10.500   0.6785   0.03395   0.02760   0.0736  0.0247  0.9489
  11.000   0.7021   0.03621   0.03021   0.0771  0.0251  0.9541
  11.500   0.7316   0.03920   0.03352   0.0782  0.0257  0.9583
  12.000   0.7561   0.04278   0.03746   0.0790  0.0265  0.9623
  12.500   0.7621   0.04726   0.04235   0.0811  0.0271  0.9668
  13.000   0.7521   0.05270   0.04821   0.0836  0.0274  0.9718
  13.500   0.7301   0.05899   0.05485   0.0853  0.0273  0.9773
  14.000   0.6921   0.06764   0.06393   0.0851  0.0281  0.9830
  16.000   0.4750   0.15082   0.14770   0.0463  0.0481  0.9957