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 =     0.300 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
   0.000   0.2356   0.00874   0.00386  -0.0543  0.8508  0.7280
   0.500   0.2886   0.00863   0.00361  -0.0526  0.8092  0.7386
   1.000   0.3425   0.00858   0.00346  -0.0514  0.7729  0.7483
   1.500   0.3959   0.00864   0.00341  -0.0501  0.7295  0.7592
   2.000   0.4475   0.00875   0.00335  -0.0485  0.6660  0.7693
   2.500   0.4934   0.00935   0.00337  -0.0459  0.5127  0.7811
   3.000   0.5267   0.01158   0.00406  -0.0424  0.1866  0.7927
   3.500   0.5754   0.01248   0.00470  -0.0411  0.1402  0.8069
   4.000   0.6248   0.01314   0.00535  -0.0398  0.1228  0.8226
   5.000   0.7185   0.01441   0.00675  -0.0361  0.0979  0.8766
   5.500   0.7742   0.01512   0.00756  -0.0362  0.0873  1.0000
   7.000   0.9264   0.01806   0.01040  -0.0343  0.0616  1.0000
   8.000   1.0204   0.02097   0.01342  -0.0316  0.0508  1.0000
   8.500   1.0661   0.02276   0.01515  -0.0303  0.0473  1.0000
   9.000   1.1105   0.02456   0.01726  -0.0286  0.0444  1.0000
   9.500   1.1541   0.02660   0.01936  -0.0271  0.0425  1.0000
  10.000   1.1925   0.02986   0.02299  -0.0251  0.0411  1.0000
  10.500   1.2222   0.03361   0.02729  -0.0221  0.0399  1.0000
  11.000   1.2429   0.03772   0.03190  -0.0184  0.0388  1.0000
  11.500   1.2619   0.04101   0.03545  -0.0148  0.0379  1.0000
  12.000   1.2686   0.04473   0.03941  -0.0102  0.0372  1.0000
  12.500   1.2467   0.05029   0.04537  -0.0040  0.0369  1.0000
  13.000   1.2050   0.05775   0.05330  -0.0004  0.0369  1.0000