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GOE 113 AIRFOIL (goe113-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: GOE 113 AIRFOIL (goe113-il)
Reynolds number: 50,000
Max Cl/Cd: 40.04 at α=5.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe113-il-50000-n5.txt
Download as CSV file: xf-goe113-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 113 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.4494   0.10959   0.10293  -0.0107   1.0000   0.0693
  -8.000  -0.4524   0.10775   0.10121  -0.0146   1.0000   0.0701
  -7.750  -0.4503   0.10531   0.09884  -0.0205   1.0000   0.0706
  -7.500  -0.4363   0.09859   0.09214  -0.0133   1.0000   0.0745
  -7.250  -0.4305   0.09522   0.08882  -0.0146   1.0000   0.0782
  -7.000  -0.4258   0.09223   0.08590  -0.0191   1.0000   0.0826
  -6.750  -0.4184   0.09008   0.08372  -0.0294   1.0000   0.0848
  -6.500  -0.4109   0.08456   0.07833  -0.0232   1.0000   0.0886
  -6.250  -0.4015   0.08117   0.07495  -0.0253   1.0000   0.0953
  -6.000  -0.3919   0.07737   0.07117  -0.0291   1.0000   0.1016
  -5.500  -0.3661   0.07157   0.06512  -0.0368   1.0000   0.1281
  -5.250  -0.3566   0.06749   0.06114  -0.0359   1.0000   0.1432
  -5.000  -0.3465   0.06369   0.05741  -0.0348   1.0000   0.1590
  -4.750  -0.3344   0.06022   0.05397  -0.0343   1.0000   0.1752
  -4.500  -0.3207   0.05699   0.05074  -0.0341   1.0000   0.1946
  -4.250  -0.2610   0.04918   0.04200  -0.0436   1.0000   0.0658
  -4.000  -0.2329   0.04509   0.03745  -0.0450   1.0000   0.0543
  -3.750  -0.2099   0.04162   0.03371  -0.0455   1.0000   0.0525
  -3.500  -0.1843   0.03861   0.03026  -0.0460   1.0000   0.0541
  -3.250  -0.1582   0.03580   0.02696  -0.0462   1.0000   0.0550
  -3.000  -0.1325   0.03304   0.02374  -0.0461   1.0000   0.0545
  -2.750  -0.1060   0.03063   0.02084  -0.0458   1.0000   0.0548
  -2.500  -0.0814   0.02853   0.01845  -0.0457   1.0000   0.0606
  -2.250  -0.0557   0.02678   0.01635  -0.0452   1.0000   0.0639
  -2.000  -0.0296   0.02521   0.01440  -0.0445   1.0000   0.0658
  -1.750  -0.0049   0.02403   0.01293  -0.0440   1.0000   0.0738
  -1.500   0.0190   0.02302   0.01172  -0.0432   1.0000   0.0781
  -1.250   0.0439   0.02224   0.01069  -0.0425   1.0000   0.0795
  -1.000   0.0689   0.02168   0.00990  -0.0422   0.9995   0.0810
  -0.750   0.1129   0.02105   0.00897  -0.0455   0.9882   0.0839
  -0.500   0.1557   0.02048   0.00821  -0.0489   0.9751   0.0892
  -0.250   0.1988   0.02002   0.00766  -0.0522   0.9619   0.1018
   0.000   0.2343   0.01701   0.00732  -0.0538   0.9492   1.0000
   0.250   0.2748   0.01716   0.00697  -0.0563   0.9328   1.0000
   0.500   0.3152   0.01728   0.00679  -0.0589   0.9174   1.0000
   0.750   0.3538   0.01738   0.00668  -0.0611   0.9017   1.0000
   1.000   0.3902   0.01748   0.00662  -0.0627   0.8853   1.0000
   1.250   0.4245   0.01761   0.00664  -0.0639   0.8681   1.0000
   1.500   0.4582   0.01773   0.00669  -0.0650   0.8516   1.0000
   1.750   0.4912   0.01785   0.00676  -0.0658   0.8350   1.0000
   2.000   0.5236   0.01793   0.00679  -0.0662   0.8171   1.0000
   2.250   0.5522   0.01806   0.00688  -0.0658   0.7949   1.0000
   2.500   0.5807   0.01813   0.00693  -0.0651   0.7713   1.0000
   2.750   0.6071   0.01825   0.00699  -0.0639   0.7457   1.0000
   3.000   0.6324   0.01845   0.00719  -0.0629   0.7214   1.0000
   3.250   0.6585   0.01868   0.00747  -0.0620   0.7011   1.0000
   3.500   0.6834   0.01897   0.00783  -0.0611   0.6795   1.0000
   3.750   0.7086   0.01922   0.00813  -0.0601   0.6578   1.0000
   4.000   0.7324   0.01949   0.00849  -0.0589   0.6316   1.0000
   4.250   0.7561   0.01977   0.00888  -0.0576   0.6030   1.0000
   4.500   0.7796   0.02005   0.00924  -0.0563   0.5723   1.0000
   4.750   0.8028   0.02038   0.00967  -0.0550   0.5399   1.0000
   5.000   0.8258   0.02076   0.01020  -0.0538   0.5048   1.0000
   5.250   0.8485   0.02119   0.01081  -0.0525   0.4670   1.0000
   5.500   0.8668   0.02170   0.01115  -0.0505   0.3972   1.0000
   5.750   0.8867   0.02243   0.01173  -0.0490   0.3402   1.0000
   6.000   0.9036   0.02359   0.01247  -0.0475   0.2712   1.0000
   6.250   0.9241   0.02463   0.01353  -0.0465   0.2345   1.0000
   6.500   0.9407   0.02624   0.01467  -0.0455   0.1319   1.0000
   7.000   0.9695   0.03086   0.01876  -0.0424   0.0385   1.0000
   7.250   0.9854   0.03276   0.02090  -0.0409   0.0344   1.0000
   7.500   1.0005   0.03465   0.02304  -0.0394   0.0307   1.0000
   7.750   1.0121   0.03685   0.02542  -0.0377   0.0284   1.0000
   8.000   1.0251   0.03886   0.02773  -0.0358   0.0275   1.0000
   8.250   1.0381   0.04099   0.03014  -0.0339   0.0268   1.0000
   8.500   1.0523   0.04325   0.03267  -0.0321   0.0262   1.0000
   8.750   1.0682   0.04572   0.03544  -0.0304   0.0257   1.0000
   9.000   1.0836   0.04846   0.03850  -0.0288   0.0249   1.0000
   9.250   1.0961   0.05144   0.04183  -0.0273   0.0240   1.0000
   9.500   1.1046   0.05462   0.04535  -0.0258   0.0231   1.0000
   9.750   1.1090   0.05800   0.04921  -0.0242   0.0226   1.0000
  10.000   1.1087   0.06151   0.05307  -0.0225   0.0223   1.0000
  10.250   1.1028   0.06505   0.05693  -0.0207   0.0223   1.0000
  10.500   1.0933   0.06889   0.06109  -0.0196   0.0223   1.0000
  10.750   1.0805   0.07318   0.06568  -0.0195   0.0225   1.0000
  11.000   1.0657   0.07808   0.07084  -0.0207   0.0226   1.0000
  11.250   1.0494   0.08369   0.07670  -0.0233   0.0228   1.0000
  11.500   1.0322   0.09012   0.08334  -0.0271   0.0230   1.0000
  11.750   1.0143   0.09746   0.09085  -0.0319   0.0233   1.0000
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