XFOIL         Version 6.94
  
 Calculated polar for: GOE 529 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   0.5752   0.01141   0.00359  -0.1096  0.6406  0.1063
   0.500   0.6279   0.01124   0.00365  -0.1091  0.6277  0.1887
   1.000   0.7013   0.00954   0.00376  -0.1131  0.6142  1.0000
   1.500   0.7523   0.00976   0.00374  -0.1120  0.5991  1.0000
   2.000   0.8021   0.00990   0.00376  -0.1106  0.5812  1.0000
   2.500   0.8526   0.01011   0.00383  -0.1094  0.5649  1.0000
   3.000   0.9035   0.01036   0.00391  -0.1084  0.5494  1.0000
   3.500   0.9532   0.01051   0.00403  -0.1072  0.5314  1.0000
   4.000   0.9906   0.01070   0.00383  -0.1033  0.4539  1.0000
   4.500   1.0328   0.01135   0.00421  -0.1009  0.4074  1.0000
   5.000   1.0723   0.01218   0.00472  -0.0981  0.3497  1.0000
   5.500   1.1033   0.01354   0.00553  -0.0941  0.2591  1.0000
   6.000   1.1147   0.01612   0.00723  -0.0873  0.1159  1.0000
   6.500   1.1325   0.01795   0.00871  -0.0813  0.0409  1.0000
   7.000   1.1513   0.01941   0.01011  -0.0753  0.0065  1.0000
   7.500   1.1792   0.02051   0.01135  -0.0710  0.0065  1.0000
   8.000   1.2035   0.02188   0.01290  -0.0665  0.0068  1.0000
   8.500   1.2265   0.02343   0.01465  -0.0623  0.0073  1.0000
   9.000   1.2445   0.02547   0.01689  -0.0580  0.0079  1.0000
   9.500   1.2549   0.02829   0.01996  -0.0537  0.0085  1.0000
  10.000   1.2691   0.03116   0.02305  -0.0504  0.0095  1.0000
  10.500   1.2691   0.03557   0.02769  -0.0471  0.0102  1.0000
  11.000   1.2673   0.04063   0.03296  -0.0448  0.0110  1.0000
  11.500   1.2631   0.04636   0.03892  -0.0430  0.0122  1.0000
  12.000   1.2516   0.05302   0.04568  -0.0415  0.0129  1.0000
  12.500   1.2617   0.05750   0.05030  -0.0394  0.0153  1.0000