XFOIL         Version 6.94
  
 Calculated polar for: NACA 66-018                                     
  
 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   0.0621   0.02390   0.01866  -0.0014  0.7858  0.7901
   1.500   0.0092   0.02406   0.01880   0.0239  0.7585  0.8064
   2.000   0.1271   0.02301   0.01778   0.0132  0.7637  0.8068
   3.500   0.3119   0.02085   0.01582   0.0099  0.7458  0.8124
   4.000   0.3496   0.02014   0.01521   0.0130  0.7309  0.8161
   4.500   0.4741   0.01609   0.01113   0.0041  0.7113  0.8162
   5.000   0.5319   0.01472   0.00979   0.0042  0.6894  0.8192
   5.500   0.5642   0.01382   0.00897   0.0085  0.6507  0.8235
   6.000   0.5626   0.01346   0.00802   0.0197  0.5102  0.8277
   6.500   0.5210   0.01511   0.00903   0.0361  0.3826  0.8337
   7.000   0.4921   0.01760   0.01075   0.0479  0.2348  0.8413
   7.500   0.4778   0.02031   0.01251   0.0565  0.0783  0.8469
   8.000   0.4970   0.02183   0.01395   0.0609  0.0572  0.8517
   8.500   0.5196   0.02332   0.01544   0.0644  0.0487  0.8573
   9.000   0.5485   0.02460   0.01679   0.0669  0.0444  0.8632
   9.500   0.5732   0.02624   0.01843   0.0700  0.0411  0.8689
  10.000   0.6079   0.02754   0.01986   0.0721  0.0384  0.8751
  10.500   0.6450   0.02889   0.02125   0.0735  0.0356  0.8823
  11.000   0.7044   0.03068   0.02306   0.0728  0.0336  0.8879
  11.500   0.7534   0.03250   0.02515   0.0732  0.0328  0.8951
  12.500   0.8304   0.03796   0.03135   0.0757  0.0310  0.9124
  13.000   0.8482   0.04138   0.03517   0.0787  0.0303  0.9247
  13.500   0.8521   0.04595   0.04028   0.0822  0.0305  0.9405
  14.000   0.8393   0.05359   0.04863   0.0836  0.0314  0.9580
  14.500   0.8198   0.06344   0.05908   0.0808  0.0325  0.9742