XFOIL         Version 6.94
  
 Calculated polar for: RAE 5214 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   0.2302   0.00872   0.00392  -0.0541  0.8707  0.7281
   0.500   0.2851   0.00859   0.00369  -0.0527  0.8389  0.7388
   1.000   0.3386   0.00847   0.00353  -0.0513  0.8017  0.7485
   1.500   0.3912   0.00848   0.00341  -0.0497  0.7472  0.7594
   2.000   0.4395   0.00875   0.00327  -0.0472  0.6263  0.7694
   2.500   0.4718   0.01083   0.00370  -0.0429  0.2713  0.7813
   3.000   0.5182   0.01190   0.00427  -0.0413  0.1739  0.7928
   3.500   0.5682   0.01262   0.00483  -0.0402  0.1453  0.8071
   4.000   0.6181   0.01324   0.00545  -0.0389  0.1285  0.8229
   4.500   0.6671   0.01381   0.00606  -0.0374  0.1147  0.8438
   5.000   0.7137   0.01431   0.00667  -0.0354  0.1022  0.8771
   5.500   0.7703   0.01493   0.00741  -0.0356  0.0907  1.0000
   6.000   0.8232   0.01589   0.00831  -0.0355  0.0808  1.0000
   6.500   0.8744   0.01680   0.00920  -0.0348  0.0714  1.0000
   7.000   0.9242   0.01776   0.01015  -0.0339  0.0628  1.0000
   7.500   0.9722   0.01902   0.01144  -0.0326  0.0562  1.0000
   8.000   1.0189   0.02060   0.01306  -0.0312  0.0514  1.0000
   8.500   1.0650   0.02217   0.01461  -0.0299  0.0477  1.0000
  10.000   1.1920   0.02928   0.02239  -0.0246  0.0412  1.0000
  10.500   1.2225   0.03281   0.02647  -0.0216  0.0400  1.0000
  11.000   1.2440   0.03686   0.03103  -0.0180  0.0390  1.0000
  11.500   1.2622   0.04032   0.03480  -0.0142  0.0380  1.0000
  12.000   1.2704   0.04388   0.03859  -0.0097  0.0372  1.0000
  12.500   1.2519   0.04909   0.04417  -0.0035  0.0369  1.0000
  13.000   1.2125   0.05624   0.05179   0.0006  0.0368  1.0000