XFOIL         Version 6.94
  
 Calculated polar for: GOE 344 (PFALZ 71) 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   0.2386   0.00967   0.00338  -0.0459  0.9439  0.2547
   0.500   0.3482   0.00762   0.00342  -0.0567  0.9238  1.0000
   1.000   0.4243   0.00731   0.00288  -0.0600  0.8561  1.0000
   1.500   0.4754   0.00768   0.00257  -0.0579  0.7188  1.0000
   2.000   0.5216   0.00828   0.00266  -0.0554  0.6220  1.0000
   2.500   0.5671   0.00890   0.00282  -0.0530  0.5278  1.0000
   4.500   0.7495   0.01204   0.00410  -0.0451  0.1706  1.0000
   5.000   0.7943   0.01327   0.00481  -0.0432  0.0795  1.0000
   5.500   0.8409   0.01433   0.00563  -0.0416  0.0386  1.0000
   6.000   0.8875   0.01545   0.00659  -0.0399  0.0068  1.0000
   6.500   0.9359   0.01632   0.00760  -0.0383  0.0058  1.0000
   7.000   0.9837   0.01728   0.00875  -0.0366  0.0058  1.0000
   7.500   1.0304   0.01837   0.01006  -0.0348  0.0061  1.0000
   8.000   1.0752   0.01972   0.01166  -0.0327  0.0065  1.0000
   8.500   1.1164   0.02146   0.01370  -0.0301  0.0069  1.0000
   9.000   1.1552   0.02331   0.01580  -0.0272  0.0074  1.0000
   9.500   1.1874   0.02567   0.01845  -0.0233  0.0082  1.0000
  10.000   1.2079   0.02908   0.02204  -0.0181  0.0089  1.0000
  10.500   1.2321   0.03226   0.02564  -0.0132  0.0101  1.0000
  11.000   1.2445   0.03796   0.03163  -0.0078  0.0112  1.0000
  11.500   1.2283   0.04434   0.03901   0.0007  0.0129  1.0000
  13.000   1.1215   0.07612   0.07248  -0.0031  0.0145  1.0000
  13.500   1.0917   0.09011   0.08673  -0.0126  0.0143  1.0000
  14.000   1.0656   0.10421   0.10099  -0.0214  0.0141  1.0000