XFOIL         Version 6.94
  
 Calculated polar for: RAE(NPL) 5213 AIRFOIL                           
  
 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   0.2377   0.00721   0.00253  -0.0541  0.7952  0.6944
   0.500   0.2924   0.00726   0.00244  -0.0533  0.7581  0.7033
   1.000   0.3468   0.00734   0.00244  -0.0523  0.7190  0.7123
   2.000   0.4501   0.00795   0.00248  -0.0496  0.5508  0.7306
   2.500   0.4905   0.00966   0.00295  -0.0470  0.2710  0.7404
   3.000   0.5374   0.01084   0.00351  -0.0456  0.1395  0.7500
   3.500   0.5892   0.01147   0.00400  -0.0448  0.1179  0.7612
   4.000   0.6405   0.01200   0.00454  -0.0438  0.1046  0.7729
   4.500   0.6920   0.01255   0.00508  -0.0429  0.0934  0.7863
   5.000   0.7421   0.01315   0.00572  -0.0417  0.0842  0.8037
   6.000   0.8395   0.01383   0.00674  -0.0386  0.0688  0.8746
   6.500   0.8945   0.01432   0.00734  -0.0385  0.0608  1.0000
   7.500   0.9917   0.01625   0.00920  -0.0362  0.0473  1.0000
   8.000   1.0395   0.01720   0.01020  -0.0349  0.0429  1.0000
   9.500   1.1681   0.02173   0.01497  -0.0291  0.0359  1.0000
  10.000   1.2087   0.02344   0.01688  -0.0270  0.0345  1.0000
  10.500   1.2469   0.02526   0.01883  -0.0246  0.0334  1.0000
  11.000   1.2817   0.02765   0.02131  -0.0222  0.0323  1.0000
  11.500   1.3096   0.03029   0.02428  -0.0189  0.0315  1.0000
  12.000   1.3273   0.03311   0.02746  -0.0143  0.0308  1.0000
  12.500   1.3322   0.03632   0.03103  -0.0086  0.0303  1.0000
  13.000   1.3286   0.04021   0.03529  -0.0035  0.0299  1.0000
  13.500   1.3145   0.04532   0.04077   0.0004  0.0297  1.0000
  14.000   1.2917   0.05187   0.04770   0.0021  0.0294  1.0000
  14.500   1.2446   0.06267   0.05898  -0.0003  0.0294  1.0000