XFOIL         Version 6.94
  
 Calculated polar for: RAE 5215 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   0.2401   0.00735   0.00277  -0.0599  0.8420  0.6952
   0.500   0.2959   0.00733   0.00266  -0.0592  0.8084  0.7040
   1.000   0.3506   0.00739   0.00260  -0.0582  0.7567  0.7131
   2.000   0.4441   0.01048   0.00321  -0.0550  0.1805  0.7312
   2.500   0.4991   0.01108   0.00359  -0.0548  0.1454  0.7410
   3.000   0.5539   0.01152   0.00400  -0.0544  0.1307  0.7506
   3.500   0.6091   0.01197   0.00445  -0.0541  0.1202  0.7618
   4.000   0.6632   0.01242   0.00493  -0.0537  0.1096  0.7734
   4.500   0.7177   0.01287   0.00539  -0.0534  0.0987  0.7868
   5.000   0.7715   0.01331   0.00588  -0.0529  0.0890  0.8039
   5.500   0.8246   0.01371   0.00637  -0.0522  0.0796  0.8259
   6.000   0.8751   0.01399   0.00682  -0.0510  0.0704  0.8682
   6.500   0.9246   0.01433   0.00729  -0.0495  0.0616  1.0000
   7.000   0.9770   0.01531   0.00818  -0.0491  0.0531  1.0000
   7.500   1.0278   0.01643   0.00930  -0.0483  0.0468  1.0000
   8.000   1.0779   0.01754   0.01045  -0.0474  0.0424  1.0000
   9.000   1.1724   0.02045   0.01348  -0.0449  0.0371  1.0000
  10.500   1.3026   0.02603   0.01959  -0.0397  0.0330  1.0000
  11.000   1.3383   0.02904   0.02274  -0.0375  0.0319  1.0000
  11.500   1.3712   0.03153   0.02563  -0.0346  0.0312  1.0000
  12.000   1.3956   0.03479   0.02930  -0.0311  0.0305  1.0000
  12.500   1.4088   0.03854   0.03346  -0.0266  0.0299  1.0000
  13.000   1.4037   0.04270   0.03801  -0.0206  0.0296  1.0000
  13.500   1.3881   0.04818   0.04389  -0.0160  0.0293  1.0000
  14.000   1.3577   0.05615   0.05230  -0.0144  0.0292  1.0000
  14.500   1.2921   0.07139   0.06814  -0.0204  0.0293  1.0000