XFOIL         Version 6.94
  
 Calculated polar for: NACA 2410                                       
  
 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
 ------- -------- --------- --------- -------- ------- -------
   0.500   0.2932   0.00817   0.00134  -0.0513  0.5619  0.3677
   1.000   0.3399   0.00783   0.00145  -0.0502  0.5013  0.5708
   1.500   0.3874   0.00733   0.00161  -0.0488  0.4874  0.7731
   2.000   0.4450   0.00749   0.00185  -0.0492  0.4363  0.9423
   2.500   0.5230   0.00793   0.00203  -0.0549  0.3715  0.9972
   3.500   0.6009   0.01118   0.00332  -0.0504  0.0387  1.0000
   4.000   0.6504   0.01159   0.00380  -0.0496  0.0375  1.0000
   4.500   0.7001   0.01206   0.00435  -0.0488  0.0373  1.0000
   5.000   0.7492   0.01266   0.00506  -0.0479  0.0375  1.0000
   5.500   0.7971   0.01346   0.00596  -0.0468  0.0379  1.0000
   6.000   0.8431   0.01449   0.00711  -0.0454  0.0385  1.0000
   6.500   0.8874   0.01577   0.00853  -0.0437  0.0391  1.0000
   7.000   0.9333   0.01676   0.00954  -0.0427  0.0360  1.0000
   7.500   0.9771   0.01807   0.01093  -0.0413  0.0347  1.0000
   8.000   1.0202   0.01951   0.01245  -0.0398  0.0333  1.0000
   8.500   1.0624   0.02095   0.01396  -0.0383  0.0314  1.0000
   9.000   1.1021   0.02266   0.01574  -0.0366  0.0287  1.0000
   9.500   1.1328   0.02646   0.01983  -0.0337  0.0248  1.0000
  10.500   1.1569   0.03813   0.03306  -0.0220  0.0152  1.0000
  11.000   1.1229   0.04734   0.04298  -0.0132  0.0165  1.0000
  11.500   1.0827   0.05588   0.05200  -0.0087  0.0171  1.0000
  12.000   1.0433   0.06566   0.06215  -0.0098  0.0173  1.0000
  12.500   1.0075   0.07717   0.07394  -0.0153  0.0173  1.0000
  13.000   0.9758   0.09001   0.08700  -0.0232  0.0170  1.0000
  13.500   0.9500   0.10442   0.10158  -0.0326  0.0165  1.0000
  14.000   0.9309   0.11966   0.11694  -0.0420  0.0159  1.0000