XFOIL         Version 6.94
  
 Calculated polar for: GOE 325 (PFALZ 54) 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   0.4717   0.01015   0.00241  -0.0891  0.5414  0.0378
   0.500   0.5271   0.01020   0.00233  -0.0891  0.5168  0.0533
   1.000   0.5827   0.01013   0.00225  -0.0890  0.4934  0.0719
   1.500   0.6319   0.00811   0.00237  -0.0881  0.4705  1.0000
   2.000   0.6872   0.00838   0.00243  -0.0880  0.4447  1.0000
   2.500   0.7419   0.00870   0.00254  -0.0879  0.4165  1.0000
   3.000   0.7960   0.00909   0.00271  -0.0878  0.3898  1.0000
   3.500   0.8486   0.00962   0.00291  -0.0875  0.3438  1.0000
   4.000   0.8996   0.01037   0.00328  -0.0870  0.2716  1.0000
   4.500   0.9512   0.01110   0.00370  -0.0865  0.2196  1.0000
   5.000   0.9891   0.01390   0.00541  -0.0844  0.0064  1.0000
   5.500   1.0413   0.01449   0.00608  -0.0838  0.0067  1.0000
   6.000   1.0926   0.01518   0.00691  -0.0829  0.0078  1.0000
   6.500   1.1429   0.01600   0.00788  -0.0818  0.0096  1.0000
   7.000   1.1906   0.01712   0.00921  -0.0802  0.0105  1.0000
   7.500   1.2315   0.01894   0.01125  -0.0777  0.0113  1.0000
   8.000   1.2627   0.02140   0.01389  -0.0738  0.0121  1.0000
   8.500   1.2873   0.02407   0.01676  -0.0687  0.0140  1.0000
   9.000   1.3045   0.02729   0.02013  -0.0625  0.0168  1.0000
   9.500   1.3399   0.03087   0.02384  -0.0585  0.0195  1.0000
  10.000   1.3903   0.03578   0.02877  -0.0580  0.0169  1.0000
  10.500   1.3921   0.03864   0.03206  -0.0512  0.0134  1.0000
  11.000   1.3939   0.04326   0.03687  -0.0465  0.0093  1.0000
  11.500   1.2909   0.03982   0.03418  -0.0301  0.0110  1.0000