XFOIL         Version 6.94
  
 Calculated polar for: GOE 450 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   0.5967   0.00916   0.00232  -0.1269  0.7370  0.0473
   0.500   0.6516   0.00901   0.00212  -0.1266  0.7163  0.0540
   1.000   0.7044   0.00803   0.00227  -0.1266  0.6942  0.5093
   1.500   0.7573   0.00717   0.00234  -0.1258  0.6744  1.0000
   2.000   0.8110   0.00743   0.00243  -0.1253  0.6532  1.0000
   2.500   0.8632   0.00771   0.00254  -0.1246  0.6229  1.0000
   3.000   0.9156   0.00802   0.00270  -0.1239  0.5961  1.0000
   3.500   0.9585   0.00876   0.00294  -0.1214  0.4953  1.0000
   4.000   0.9936   0.01039   0.00364  -0.1180  0.3393  1.0000
   4.500   1.0277   0.01236   0.00469  -0.1148  0.1765  1.0000
   5.000   1.0581   0.01466   0.00617  -0.1110  0.0294  1.0000
   5.500   1.1051   0.01534   0.00685  -0.1097  0.0235  1.0000
   6.000   1.1489   0.01624   0.00765  -0.1078  0.0043  1.0000
   6.500   1.1928   0.01705   0.00854  -0.1059  0.0037  1.0000
   7.000   1.2347   0.01793   0.00953  -0.1037  0.0038  1.0000
   7.500   1.2729   0.01891   0.01064  -0.1009  0.0041  1.0000
   8.000   1.3062   0.02007   0.01193  -0.0972  0.0044  1.0000
   8.500   1.3355   0.02152   0.01357  -0.0931  0.0047  1.0000
   9.000   1.3649   0.02299   0.01519  -0.0893  0.0052  1.0000
   9.500   1.3897   0.02487   0.01726  -0.0851  0.0058  1.0000
  10.000   1.4029   0.02771   0.02030  -0.0799  0.0063  1.0000
  10.500   1.4223   0.03023   0.02303  -0.0759  0.0071  1.0000
  11.000   1.4237   0.03445   0.02750  -0.0708  0.0078  1.0000
  11.500   1.4270   0.03886   0.03212  -0.0665  0.0086  1.0000
  12.000   1.4226   0.04450   0.03807  -0.0611  0.0100  1.0000