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

GOE 504 AIRFOIL (goe504-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: GOE 504 AIRFOIL (goe504-il)
Reynolds number: 100,000
Max Cl/Cd: 34.21 at α=5.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe504-il-100000.txt
Download as CSV file: xf-goe504-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 504 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.250  -0.1220   0.10199   0.09770  -0.0818   0.9416   0.1727
  -9.000  -0.1467   0.09861   0.09440  -0.0864   0.9300   0.1801
  -8.750  -0.3716   0.06869   0.06409  -0.1023   0.9021   0.0927
  -8.500  -0.3299   0.06445   0.05999  -0.1043   0.8989   0.0877
  -8.250  -0.4140   0.05708   0.05152  -0.0984   0.8825   0.0797
  -8.000  -0.3925   0.05234   0.04650  -0.1000   0.8787   0.0781
  -7.750  -0.3991   0.05022   0.04417  -0.0960   0.8674   0.0775
  -7.500  -0.3797   0.04670   0.04007  -0.0965   0.8631   0.0781
  -7.250  -0.3831   0.04529   0.03838  -0.0922   0.8521   0.0782
  -7.000  -0.3555   0.04246   0.03502  -0.0929   0.8481   0.0786
  -6.750  -0.3534   0.04130   0.03358  -0.0890   0.8377   0.0787
  -6.500  -0.3220   0.03922   0.03103  -0.0897   0.8335   0.0793
  -6.250  -0.3080   0.03753   0.02927  -0.0879   0.8256   0.0809
  -6.000  -0.2796   0.03610   0.02778  -0.0881   0.8197   0.0839
  -5.750  -0.2367   0.03438   0.02585  -0.0903   0.8168   0.0867
  -5.500  -0.2310   0.03413   0.02551  -0.0868   0.8066   0.0881
  -5.250  -0.1921   0.03293   0.02408  -0.0881   0.8026   0.0918
  -5.000  -0.1471   0.03111   0.02241  -0.0908   0.8006   0.0985
  -4.500  -0.1105   0.03001   0.02126  -0.0874   0.7859   0.1093
  -4.250  -0.1083   0.03011   0.02136  -0.0833   0.7757   0.1159
  -4.000  -0.0788   0.02875   0.02027  -0.0833   0.7714   0.1468
  -3.750  -0.0398   0.02800   0.02000  -0.0846   0.7689   0.3060
  -3.500  -0.0483   0.02898   0.02096  -0.0789   0.7567   0.3211
  -3.250  -0.0056   0.02857   0.02041  -0.0807   0.7540   0.3562
  -3.000  -0.0091   0.02946   0.02133  -0.0759   0.7429   0.3693
  -2.750   0.0291   0.02902   0.02080  -0.0771   0.7394   0.3938
  -2.500   0.0776   0.02831   0.02006  -0.0800   0.7374   0.4210
  -2.250   0.0654   0.02940   0.02115  -0.0739   0.7251   0.4288
  -2.000   0.1107   0.02866   0.02041  -0.0763   0.7226   0.4540
  -1.750   0.1612   0.02780   0.01958  -0.0794   0.7209   0.4865
  -1.500   0.1421   0.02910   0.02089  -0.0723   0.7080   0.4955
  -1.250   0.1929   0.02813   0.01997  -0.0755   0.7060   0.5250
  -1.000   0.1774   0.02945   0.02138  -0.0691   0.6940   0.5358
  -0.750   0.2252   0.02847   0.02049  -0.0718   0.6913   0.5697
  -0.500   0.2838   0.02720   0.01948  -0.0762   0.6897   0.6333
  -0.250   0.4470   0.02531   0.01833  -0.0984   0.6899   0.9270
   0.000   0.5949   0.02436   0.01693  -0.1196   0.6880   0.9840
   0.250   0.6326   0.02499   0.01747  -0.1231   0.6782   1.0000
   0.500   0.6629   0.02466   0.01694  -0.1236   0.6735   1.0000
   0.750   0.6469   0.02567   0.01792  -0.1166   0.6641   1.0000
   1.000   0.6750   0.02550   0.01759  -0.1166   0.6589   1.0000
   1.250   0.6881   0.02582   0.01781  -0.1142   0.6525   1.0000
   1.500   0.6906   0.02640   0.01833  -0.1101   0.6446   1.0000
   1.750   0.7357   0.02585   0.01761  -0.1127   0.6408   1.0000
   2.000   0.7094   0.02735   0.01914  -0.1041   0.6307   1.0000
   2.250   0.7513   0.02687   0.01852  -0.1061   0.6264   1.0000
   2.500   0.7412   0.02798   0.01961  -0.1001   0.6179   1.0000
   2.750   0.7688   0.02793   0.01947  -0.0999   0.6122   1.0000
   3.000   0.8224   0.02717   0.01856  -0.1038   0.6088   1.0000
   3.250   0.7892   0.02897   0.02042  -0.0942   0.5983   1.0000
   3.500   0.8397   0.02825   0.01958  -0.0975   0.5944   1.0000
   3.750   0.8165   0.02983   0.02119  -0.0895   0.5851   1.0000
   4.000   0.8564   0.02942   0.02070  -0.0912   0.5803   1.0000
   4.250   0.9186   0.02842   0.01956  -0.0965   0.5770   1.0000
   4.500   0.8738   0.03064   0.02187  -0.0852   0.5663   1.0000
   4.750   0.9330   0.02965   0.02078  -0.0898   0.5626   1.0000
   5.000   0.8917   0.03180   0.02300  -0.0792   0.5528   1.0000
   5.250   0.9466   0.03098   0.02210  -0.0833   0.5484   1.0000
   5.500   1.0181   0.02976   0.02076  -0.0899   0.5451   1.0000
   5.750   0.9510   0.03260   0.02373  -0.0754   0.5343   1.0000
   6.000   1.0268   0.03120   0.02224  -0.0826   0.5308   1.0000
   6.250   0.9589   0.03439   0.02554  -0.0687   0.5203   1.0000
   6.500   1.0247   0.03300   0.02408  -0.0738   0.5165   1.0000
   6.750   1.0876   0.03216   0.02318  -0.0792   0.5123   1.0000
   7.000   1.0182   0.03542   0.02656  -0.0654   0.5022   1.0000
   7.250   1.1025   0.03345   0.02451  -0.0731   0.4991   1.0000
   7.500   1.0095   0.03851   0.02972  -0.0579   0.4872   1.0000
   7.750   1.0868   0.03612   0.02729  -0.0635   0.4847   1.0000
   8.000   0.7651   0.06465   0.05605  -0.0415   0.4401   1.0000
   8.250   0.7987   0.06374   0.05514  -0.0411   0.4379   1.0000
   8.500   0.8411   0.06172   0.05314  -0.0409   0.4369   1.0000
   8.750   0.8878   0.05909   0.05051  -0.0407   0.4366   1.0000
   9.000   0.9402   0.05600   0.04744  -0.0408   0.4365   1.0000
   9.250   0.9994   0.05255   0.04402  -0.0416   0.4364   1.0000
   9.500   1.0727   0.04842   0.03991  -0.0437   0.4364   1.0000
   9.750   1.1608   0.04420   0.03568  -0.0480   0.4359   1.0000
  10.000   0.7281   0.09365   0.08527  -0.0372   0.3831   1.0000
  10.250   0.7247   0.09758   0.08925  -0.0373   0.3782   1.0000
  10.500   0.7513   0.09776   0.08945  -0.0369   0.3756   1.0000
  10.750   0.7914   0.09613   0.08786  -0.0362   0.3739   1.0000
  11.000   0.7084   0.11032   0.10211  -0.0381   0.3646   1.0000
  11.250   0.7129   0.11359   0.10542  -0.0385   0.3622   1.0000
  11.500   0.7541   0.11203   0.10390  -0.0376   0.3593   1.0000
  11.750   0.6735   0.13430   0.12634  -0.0447   0.4040   1.0000
  12.000   0.6304   0.13824   0.13031  -0.0451   0.4009   1.0000
<< Back to GOE 504 AIRFOIL (goe504-il)

Polar data table (+)

Polar graphs


<< Back to GOE 504 AIRFOIL (goe504-il)