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 =     0.300 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
   0.000   0.2423   0.00889   0.00402  -0.0612  0.8865  0.7289
   0.500   0.2958   0.00880   0.00386  -0.0597  0.8580  0.7396
   1.000   0.3486   0.00867   0.00368  -0.0579  0.8188  0.7493
   1.500   0.4012   0.00868   0.00352  -0.0562  0.7506  0.7602
   2.000   0.4381   0.01079   0.00370  -0.0522  0.3098  0.7700
   2.500   0.4882   0.01212   0.00429  -0.0516  0.1744  0.7819
   3.000   0.5417   0.01267   0.00477  -0.0510  0.1515  0.7933
   3.500   0.5952   0.01329   0.00536  -0.0505  0.1368  0.8073
   4.000   0.6475   0.01399   0.00606  -0.0498  0.1247  0.8228
   4.500   0.6989   0.01465   0.00675  -0.0488  0.1127  0.8431
   5.000   0.7482   0.01517   0.00738  -0.0474  0.1012  0.8737
   5.500   0.7978   0.01557   0.00792  -0.0460  0.0907  1.0000
   6.000   0.8541   0.01637   0.00870  -0.0464  0.0805  1.0000
   6.500   0.9084   0.01714   0.00946  -0.0463  0.0705  1.0000
   7.000   0.9610   0.01805   0.01035  -0.0459  0.0618  1.0000
   8.500   1.1075   0.02292   0.01528  -0.0429  0.0469  1.0000
   9.000   1.1543   0.02481   0.01748  -0.0415  0.0441  1.0000
   9.500   1.1993   0.02702   0.01975  -0.0402  0.0422  1.0000
  10.000   1.2386   0.03036   0.02354  -0.0382  0.0408  1.0000
  10.500   1.2708   0.03423   0.02798  -0.0354  0.0396  1.0000
  11.000   1.2955   0.03843   0.03269  -0.0322  0.0385  1.0000
  11.500   1.3180   0.04200   0.03654  -0.0290  0.0375  1.0000
  12.000   1.3267   0.04687   0.04173  -0.0251  0.0368  1.0000
  12.500   1.2987   0.05367   0.04904  -0.0185  0.0366  1.0000
  13.000   1.2501   0.06253   0.05845  -0.0153  0.0367  1.0000