XFOIL         Version 6.94
  
 Calculated polar for: AH 88-K-130/20                                  
  
 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   0.4037   0.01556   0.01059  -0.0968  0.7859  0.9126
   0.500   0.4522   0.01529   0.01023  -0.0950  0.7818  0.9260
   1.000   0.5102   0.01537   0.01022  -0.0958  0.7778  0.9342
   1.500   0.5391   0.01570   0.01062  -0.0919  0.7650  0.9457
   2.000   0.5927   0.01555   0.01047  -0.0920  0.7592  0.9530
   2.500   0.6556   0.01505   0.00995  -0.0934  0.7551  0.9597
   3.000   0.7251   0.01440   0.00932  -0.0958  0.7507  0.9652
   3.500   0.7739   0.01377   0.00878  -0.0949  0.7349  0.9746
   4.000   0.8439   0.01310   0.00818  -0.0978  0.7279  0.9794
   4.500   0.9174   0.01229   0.00742  -0.1012  0.7205  0.9832
   5.000   0.9861   0.01101   0.00630  -0.1035  0.6936  0.9911
   5.500   1.0402   0.01098   0.00645  -0.1040  0.6734  1.0000
   6.000   1.0892   0.01090   0.00646  -0.1031  0.6471  1.0000
   6.500   1.1192   0.01149   0.00652  -0.0985  0.5392  1.0000
   7.000   1.1075   0.01440   0.00859  -0.0889  0.3731  1.0000
   7.500   1.0902   0.01904   0.01220  -0.0811  0.1949  1.0000
   8.000   1.0905   0.02307   0.01548  -0.0760  0.0649  1.0000
   8.500   1.0961   0.02703   0.01931  -0.0709  0.0229  1.0000
   9.000   1.1108   0.03035   0.02278  -0.0673  0.0177  1.0000
   9.500   1.1357   0.03285   0.02542  -0.0650  0.0142  1.0000
  10.000   1.1556   0.03589   0.02864  -0.0622  0.0131  1.0000
  11.000   1.1975   0.04346   0.03676  -0.0551  0.0088  1.0000
  11.500   1.2348   0.04840   0.04225  -0.0521  0.0082  1.0000
  12.000   1.2488   0.05546   0.04997  -0.0485  0.0085  1.0000
  12.500   1.2320   0.06462   0.05977  -0.0442  0.0090  1.0000