XFOIL         Version 6.94
  
 Calculated polar for: NACA 63-212 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   0.1701   0.00915   0.00354  -0.0396  0.7490  0.7457
   0.500   0.2253   0.00914   0.00356  -0.0392  0.7329  0.7622
   1.000   0.2806   0.00918   0.00365  -0.0388  0.7170  0.7795
   1.500   0.3358   0.00925   0.00375  -0.0385  0.7015  0.7970
   2.000   0.3906   0.00933   0.00384  -0.0379  0.6851  0.8155
   2.500   0.4449   0.00937   0.00398  -0.0374  0.6667  0.8342
   3.000   0.4976   0.00940   0.00405  -0.0363  0.6438  0.8537
   3.500   0.5486   0.00941   0.00411  -0.0348  0.6158  0.8736
   4.000   0.5961   0.00940   0.00413  -0.0326  0.5762  0.8951
   4.500   0.6412   0.00943   0.00425  -0.0299  0.5372  0.9194
   5.000   0.6845   0.00959   0.00437  -0.0269  0.4767  0.9508
   6.000   0.7722   0.01407   0.00673  -0.0268  0.0713  1.0000
   6.500   0.8152   0.01579   0.00830  -0.0257  0.0470  1.0000
   8.000   0.9314   0.02118   0.01365  -0.0198  0.0305  1.0000
   9.500   1.0667   0.02906   0.02194  -0.0166  0.0268  1.0000
  10.000   1.1127   0.03311   0.02629  -0.0160  0.0265  1.0000
  10.500   1.1455   0.03807   0.03171  -0.0141  0.0260  1.0000
  11.000   1.1599   0.04305   0.03728  -0.0103  0.0259  1.0000
  11.500   1.1631   0.04872   0.04344  -0.0061  0.0260  1.0000
  12.500   1.1245   0.06068   0.05616   0.0029  0.0262  1.0000