XFOIL         Version 6.94
  
 Calculated polar for: NACA 65(2)-215                                  
  
 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   0.1675   0.00938   0.00426  -0.0388  0.6998  0.6993
   0.500   0.2246   0.00909   0.00403  -0.0396  0.6960  0.7126
   1.000   0.2815   0.00909   0.00409  -0.0402  0.6909  0.7208
   1.500   0.3385   0.00915   0.00420  -0.0408  0.6847  0.7282
   2.000   0.3956   0.00923   0.00433  -0.0415  0.6782  0.7358
   2.500   0.4529   0.00921   0.00431  -0.0420  0.6685  0.7428
   3.000   0.5093   0.00920   0.00422  -0.0421  0.6467  0.7480
   3.500   0.5650   0.00882   0.00388  -0.0421  0.6273  0.7592
   4.000   0.6178   0.00880   0.00384  -0.0415  0.5978  0.7718
   4.500   0.6678   0.00888   0.00386  -0.0404  0.5532  0.7825
   5.000   0.6963   0.00997   0.00424  -0.0357  0.3885  0.7926
   5.500   0.6921   0.01222   0.00563  -0.0260  0.1767  0.8027
   6.000   0.6717   0.01357   0.00656  -0.0127  0.0747  0.8249
   6.500   0.6748   0.01469   0.00764  -0.0040  0.0430  0.8419
   7.000   0.6783   0.01600   0.00906   0.0042  0.0301  0.8628