XFOIL         Version 6.94
  
 Calculated polar for: GOE 591 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
 ------- -------- --------- --------- -------- ------- -------
   1.500   0.7395   0.00790   0.00261  -0.1115  0.5748  1.0000
   2.000   0.7840   0.00819   0.00272  -0.1092  0.5532  1.0000
   2.500   0.8285   0.00850   0.00287  -0.1069  0.5324  1.0000
   3.000   0.8733   0.00881   0.00306  -0.1047  0.5123  1.0000
   3.500   0.9096   0.00926   0.00325  -0.1007  0.4580  1.0000
   4.000   0.9435   0.00995   0.00355  -0.0965  0.3942  1.0000
   4.500   0.9750   0.01086   0.00401  -0.0921  0.3169  1.0000
   5.000   1.0061   0.01189   0.00464  -0.0877  0.2469  1.0000
   5.500   1.0148   0.01399   0.00594  -0.0797  0.1173  1.0000
   6.000   1.0180   0.01602   0.00747  -0.0706  0.0056  1.0000
   6.500   1.0524   0.01672   0.00821  -0.0669  0.0050  1.0000
   7.000   1.0854   0.01751   0.00910  -0.0632  0.0052  1.0000
   7.500   1.1171   0.01841   0.01009  -0.0595  0.0057  1.0000
   8.000   1.1464   0.01949   0.01131  -0.0556  0.0064  1.0000
   8.500   1.1741   0.02073   0.01268  -0.0517  0.0072  1.0000
   9.000   1.1977   0.02230   0.01443  -0.0476  0.0082  1.0000
   9.500   1.2211   0.02402   0.01630  -0.0438  0.0094  1.0000
  10.000   1.2351   0.02652   0.01899  -0.0396  0.0107  1.0000
  10.500   1.2472   0.02940   0.02207  -0.0357  0.0121  1.0000
  11.000   1.2484   0.03343   0.02628  -0.0319  0.0131  1.0000
  11.500   1.2470   0.03812   0.03120  -0.0288  0.0147  1.0000
  12.000   1.2460   0.04315   0.03644  -0.0264  0.0164  1.0000
  12.500   1.2392   0.04889   0.04228  -0.0238  0.0183  1.0000