XFOIL         Version 6.94
  
 Calculated polar for: CHEN AIRFOIL                                    
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.045 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
   0.000  -0.2029   0.10429   0.09867  -0.0253  0.8004  0.1776
   0.250  -0.1840   0.10083   0.09526  -0.0255  0.7932  0.1874
   0.500  -0.1761   0.09962   0.09384  -0.0235  0.7771  0.1976
   0.750  -0.1618   0.10036   0.09429  -0.0224  0.7639  0.2122
   1.000  -0.1394   0.09942   0.09319  -0.0224  0.7550  0.2296
   1.250  -0.1406   0.09624   0.09002  -0.0191  0.7415  0.2394
   1.500  -0.1185   0.09569   0.08930  -0.0188  0.7346  0.2715
   2.500  -0.0802   0.09146   0.08443  -0.0092  0.6947  0.4002
   2.750  -0.0834   0.08928   0.08222  -0.0058  0.6826  0.4128
   3.000  -0.0549   0.08969   0.08242  -0.0059  0.6761  0.4428
   3.250  -0.0488   0.08904   0.08163  -0.0042  0.6665  0.4549
   3.500  -0.0127   0.08986   0.08223  -0.0065  0.6579  0.4681
   3.750   0.0264   0.09164   0.08365  -0.0108  0.6508  0.4688
   4.000   0.0583   0.09297   0.08455  -0.0129  0.6389  0.4582
   4.500   0.1189   0.09685   0.08773  -0.0168  0.6205  0.4037
   4.750   0.1600   0.10128   0.09172  -0.0178  0.6154  0.3504
   5.000   0.1529   0.10063   0.09094  -0.0147  0.6057  0.3355