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 =     0.450 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
   0.000   0.1880   0.00551   0.00193  -0.0377  0.9170  0.9769
   0.500   0.2720   0.00549   0.00179  -0.0432  0.8805  0.9880
   1.000   0.3548   0.00554   0.00164  -0.0488  0.8253  0.9973
   1.500   0.4051   0.00569   0.00153  -0.0476  0.7610  1.0000
   2.000   0.4436   0.00590   0.00153  -0.0441  0.6997  1.0000
   2.500   0.4824   0.00623   0.00159  -0.0406  0.6250  1.0000
   3.000   0.5237   0.00671   0.00185  -0.0376  0.5454  1.0000
   3.500   0.5602   0.00787   0.00214  -0.0341  0.3535  1.0000
   4.000   0.6033   0.00897   0.00263  -0.0322  0.2100  1.0000
   4.500   0.6479   0.01017   0.00324  -0.0307  0.0932  1.0000
   5.000   0.6936   0.01137   0.00405  -0.0293  0.0212  1.0000
   5.500   0.7428   0.01216   0.00489  -0.0282  0.0118  1.0000
   6.000   0.7922   0.01295   0.00583  -0.0271  0.0060  1.0000
   6.500   0.8395   0.01411   0.00715  -0.0256  0.0029  1.0000
   7.000   0.8833   0.01590   0.00929  -0.0234  0.0027  1.0000
   7.500   0.9195   0.01897   0.01281  -0.0200  0.0028  1.0000
   8.000   0.9486   0.02428   0.01883  -0.0155  0.0031  1.0000
   8.500   0.9673   0.03112   0.02654  -0.0106  0.0034  1.0000
   9.000   0.9358   0.04512   0.04189  -0.0014  0.0050  1.0000
   9.500   0.9106   0.05079   0.04774   0.0044  0.0054  1.0000