XFOIL         Version 6.94
  
 Calculated polar for: NACA 3210                                       
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.020 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
  -3.000  -0.2470   0.04289   0.03114  -0.0282  1.0000  0.3174
  -2.500  -0.1896   0.03862   0.02597  -0.0297  1.0000  0.3252
  -2.000  -0.1318   0.03537   0.02208  -0.0302  1.0000  0.3556
  -1.500  -0.0782   0.03247   0.01936  -0.0295  1.0000  0.4164
  -1.000  -0.0238   0.02811   0.01737  -0.0260  1.0000  1.0000
  -0.500   0.0238   0.02907   0.01621  -0.0245  1.0000  1.0000
   0.000   0.0588   0.03020   0.01626  -0.0230  1.0000  1.0000
   0.500   0.0906   0.03168   0.01706  -0.0223  1.0000  1.0000
   1.000   0.1194   0.03363   0.01852  -0.0223  1.0000  1.0000
   1.500   0.1445   0.03617   0.02075  -0.0230  1.0000  1.0000
   2.000   0.1666   0.03935   0.02370  -0.0245  1.0000  1.0000
   2.500   0.1874   0.04302   0.02717  -0.0265  1.0000  1.0000
   3.000   0.2890   0.04976   0.03362  -0.0437  0.9484  1.0000
   3.500   0.3680   0.05534   0.03904  -0.0543  0.8965  1.0000
   4.000   0.4083   0.06001   0.04362  -0.0579  0.8666  1.0000
   4.500   0.4517   0.06515   0.04871  -0.0619  0.8449  1.0000
   5.000   0.4652   0.06982   0.05333  -0.0620  0.8402  1.0000