Airfoil Tools
Search 1636 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

GOE 6K AIRFOIL (goe06k-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: GOE 6K AIRFOIL (goe06k-il)
Reynolds number: 200,000
Max Cl/Cd: 74.36 at α=3.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe06k-il-200000.txt
Download as CSV file: xf-goe06k-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 6K AIRFOIL                                  
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.4897   0.09762   0.09432  -0.0255   1.0000   0.0322
  -7.750  -0.5025   0.09564   0.09240  -0.0239   1.0000   0.0322
  -7.500  -0.5178   0.09374   0.09056  -0.0235   1.0000   0.0325
  -7.250  -0.5237   0.09084   0.08767  -0.0243   1.0000   0.0326
  -7.000  -0.5267   0.08767   0.08448  -0.0264   1.0000   0.0329
  -6.750  -0.5267   0.08439   0.08116  -0.0273   1.0000   0.0330
  -6.500  -0.5246   0.08106   0.07775  -0.0278   1.0000   0.0331
  -6.250  -0.5213   0.07757   0.07417  -0.0281   1.0000   0.0332
  -6.000  -0.5167   0.07408   0.07056  -0.0279   1.0000   0.0333
  -5.750  -0.5227   0.06699   0.06354  -0.0272   1.0000   0.0342
  -5.500  -0.5107   0.06324   0.05980  -0.0277   0.9982   0.0353
  -5.250  -0.4885   0.05952   0.05599  -0.0303   0.9951   0.0372
  -5.000  -0.4644   0.05548   0.05178  -0.0333   0.9911   0.0396
  -4.750  -0.4239   0.05287   0.04851  -0.0371   0.9867   0.0454
  -4.250  -0.3894   0.04338   0.03886  -0.0393   0.9809   0.0500
  -4.000  -0.3656   0.04057   0.03581  -0.0398   0.9772   0.0538
  -3.750  -0.3348   0.03771   0.03221  -0.0405   0.9739   0.0603
  -3.500  -0.3090   0.03466   0.02916  -0.0417   0.9722   0.0633
  -3.250  -0.2890   0.03288   0.02713  -0.0404   0.9681   0.0692
  -3.000  -0.2663   0.03065   0.02462  -0.0400   0.9649   0.0779
  -2.750  -0.2396   0.02912   0.02287  -0.0404   0.9621   0.0941
  -2.500  -0.1996   0.02518   0.01772  -0.0378   0.9608   0.0447
  -2.250  -0.1673   0.02323   0.01541  -0.0383   0.9594   0.0441
  -2.000  -0.1479   0.02229   0.01423  -0.0363   0.9548   0.0452
  -1.750  -0.1219   0.02085   0.01271  -0.0361   0.9521   0.0503
  -1.500  -0.0917   0.02018   0.01195  -0.0364   0.9496   0.0555
  -1.250  -0.0603   0.01941   0.01113  -0.0371   0.9477   0.0624
  -1.000  -0.0378   0.01910   0.01080  -0.0362   0.9437   0.0727
  -0.750  -0.0161   0.01859   0.01035  -0.0352   0.9397   0.0849
  -0.500   0.0971   0.01584   0.01047  -0.0547   0.9499   1.0000
  -0.250   0.1312   0.01609   0.01053  -0.0563   0.9477   1.0000
   0.000   0.1536   0.01625   0.01058  -0.0555   0.9427   1.0000
   0.250   0.1805   0.01643   0.01067  -0.0557   0.9386   1.0000
   0.500   0.2136   0.01664   0.01080  -0.0572   0.9359   1.0000
   0.750   0.2513   0.01686   0.01094  -0.0595   0.9338   1.0000
   1.000   0.2715   0.01689   0.01093  -0.0582   0.9255   1.0000
   1.250   0.3130   0.01692   0.01093  -0.0611   0.9220   1.0000
   1.500   0.3367   0.01704   0.01106  -0.0605   0.9155   1.0000
   1.750   0.3682   0.01714   0.01117  -0.0615   0.9111   1.0000
   2.000   0.4066   0.01720   0.01125  -0.0638   0.9084   1.0000
   2.250   0.4269   0.01738   0.01146  -0.0625   0.9013   1.0000
   2.500   0.4690   0.01717   0.01131  -0.0653   0.8963   1.0000
   2.750   0.5064   0.01684   0.01108  -0.0668   0.8882   1.0000
   3.000   0.5828   0.01520   0.00961  -0.0751   0.8811   1.0000
   3.250   0.6552   0.01312   0.00776  -0.0821   0.8685   1.0000
   3.500   0.7296   0.01098   0.00581  -0.0893   0.8384   1.0000
   3.750   0.7607   0.01023   0.00506  -0.0883   0.7783   1.0000
   4.000   0.7739   0.01130   0.00458  -0.0837   0.4700   1.0000
   4.250   0.7521   0.01328   0.00534  -0.0738   0.2368   1.0000
   4.500   0.7477   0.01513   0.00626  -0.0678   0.0885   1.0000
   4.750   0.7571   0.01623   0.00728  -0.0640   0.0644   1.0000
   5.000   0.7677   0.01735   0.00838  -0.0605   0.0539   1.0000
   5.250   0.7829   0.01828   0.00928  -0.0581   0.0456   1.0000
   5.500   0.8034   0.01965   0.01071  -0.0566   0.0413   1.0000
   5.750   0.8304   0.02103   0.01216  -0.0563   0.0377   1.0000
   6.000   0.8681   0.02384   0.01498  -0.0587   0.0323   1.0000
   6.250   0.8978   0.02547   0.01687  -0.0587   0.0313   1.0000
   6.500   0.9295   0.02820   0.01992  -0.0589   0.0310   1.0000
   6.750   0.9561   0.03216   0.02427  -0.0583   0.0317   1.0000
   7.000   0.9797   0.03482   0.02718  -0.0569   0.0332   1.0000
   7.250   0.9922   0.03896   0.03170  -0.0541   0.0328   1.0000
   8.250   1.0047   0.05468   0.04958  -0.0330   0.0519   1.0000
   8.500   1.0034   0.05742   0.05255  -0.0289   0.0493   1.0000
   8.750   1.0020   0.06030   0.05561  -0.0254   0.0479   1.0000
   9.000   1.0012   0.06330   0.05872  -0.0223   0.0468   1.0000
   9.250   1.0031   0.06699   0.06244  -0.0201   0.0456   1.0000
  14.000   0.6391   0.14753   0.14412  -0.0281   0.0420   1.0000
  14.250   0.6353   0.15089   0.14748  -0.0306   0.0404   1.0000
<< Back to GOE 6K AIRFOIL (goe06k-il)

Polar data table (+)

Polar graphs


<< Back to GOE 6K AIRFOIL (goe06k-il)