XFOIL         Version 6.94
  
 Calculated polar for: RG 14A-1.4/7.0 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   0.1915   0.00692   0.00259  -0.0402  0.9579  1.0000
   0.500   0.2707   0.00665   0.00230  -0.0449  0.9324  1.0000
   1.000   0.3370   0.00644   0.00203  -0.0466  0.8896  1.0000
   1.500   0.3869   0.00647   0.00191  -0.0447  0.8332  1.0000
   2.000   0.4319   0.00668   0.00192  -0.0420  0.7700  1.0000
   2.500   0.4767   0.00702   0.00204  -0.0393  0.7056  1.0000
   3.000   0.5213   0.00744   0.00223  -0.0367  0.6266  1.0000
   3.500   0.5667   0.00801   0.00251  -0.0345  0.5391  1.0000
   4.500   0.6456   0.01122   0.00363  -0.0295  0.1182  1.0000
   5.000   0.6916   0.01258   0.00457  -0.0282  0.0333  1.0000
   5.500   0.7404   0.01365   0.00573  -0.0269  0.0185  1.0000
   6.000   0.7893   0.01467   0.00695  -0.0258  0.0128  1.0000
   6.500   0.8347   0.01640   0.00911  -0.0239  0.0083  1.0000
   7.000   0.8723   0.01986   0.01301  -0.0206  0.0062  1.0000
   7.500   0.9097   0.02461   0.01839  -0.0173  0.0056  1.0000
   8.000   0.9383   0.03142   0.02614  -0.0134  0.0056  1.0000
   8.500   0.9479   0.04007   0.03579  -0.0088  0.0059  1.0000
   9.000   0.9364   0.04920   0.04566  -0.0045  0.0062  1.0000
   9.500   0.9020   0.05732   0.05422  -0.0008  0.0063  1.0000
  10.000   0.8656   0.06796   0.06516  -0.0063  0.0064  1.0000
  10.500   0.8367   0.08663   0.08403  -0.0214  0.0065  1.0000