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GOE 113 AIRFOIL (goe113-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: GOE 113 AIRFOIL (goe113-il)
Reynolds number: 100,000
Max Cl/Cd: 54.65 at α=5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe113-il-100000.txt
Download as CSV file: xf-goe113-il-100000.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 =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.750  -0.3703   0.11314   0.10840  -0.0088   1.0000   0.0468
  -9.500  -0.3689   0.10989   0.10517  -0.0099   1.0000   0.0482
  -9.250  -0.4769   0.11694   0.11195   0.0006   1.0000   0.0426
  -9.000  -0.4717   0.11352   0.10856  -0.0005   1.0000   0.0440
  -8.750  -0.4670   0.11017   0.10524  -0.0019   1.0000   0.0454
  -8.500  -0.4627   0.10687   0.10197  -0.0035   1.0000   0.0469
  -8.250  -0.4591   0.10369   0.09884  -0.0055   1.0000   0.0484
  -8.000  -0.4569   0.10082   0.09602  -0.0083   1.0000   0.0497
  -7.750  -0.4572   0.09836   0.09364  -0.0124   1.0000   0.0506
  -7.500  -0.4514   0.09544   0.09076  -0.0201   1.0000   0.0511
  -7.250  -0.4411   0.09220   0.08747  -0.0275   1.0000   0.0515
  -7.000  -0.4376   0.08599   0.08138  -0.0229   1.0000   0.0528
  -6.750  -0.4302   0.08224   0.07766  -0.0200   1.0000   0.0551
  -6.500  -0.4207   0.07865   0.07408  -0.0218   1.0000   0.0576
  -6.250  -0.4089   0.07490   0.07032  -0.0254   1.0000   0.0608
  -6.000  -0.3843   0.07204   0.06715  -0.0365   1.0000   0.0646
  -5.750  -0.3769   0.06652   0.06174  -0.0358   1.0000   0.0661
  -5.500  -0.3688   0.06306   0.05837  -0.0335   1.0000   0.0695
  -5.250  -0.3406   0.06093   0.05576  -0.0401   1.0000   0.0783
  -5.000  -0.3353   0.05568   0.05079  -0.0378   1.0000   0.0820
  -4.750  -0.3137   0.05251   0.04730  -0.0406   1.0000   0.0937
  -4.500  -0.2945   0.04979   0.04437  -0.0417   1.0000   0.1070
  -4.250  -0.2820   0.04625   0.04101  -0.0402   1.0000   0.1162
  -4.000  -0.2612   0.04328   0.03780  -0.0415   1.0000   0.1357
  -3.750  -0.2455   0.04037   0.03495  -0.0407   1.0000   0.1532
  -3.500  -0.2281   0.03786   0.03239  -0.0403   1.0000   0.1815
  -2.500  -0.1816   0.02994   0.02499  -0.0297   1.0000   0.4129
  -2.250  -0.0832   0.02527   0.01756  -0.0413   1.0000   0.1348
  -2.000  -0.0525   0.02271   0.01454  -0.0403   1.0000   0.1006
  -1.750  -0.0265   0.02129   0.01282  -0.0397   1.0000   0.0996
  -1.500  -0.0008   0.02073   0.01177  -0.0385   1.0000   0.0923
  -1.250   0.0246   0.01936   0.01032  -0.0383   1.0000   0.0908
  -1.000   0.0496   0.01862   0.00944  -0.0382   1.0000   0.0909
  -0.750   0.0752   0.01792   0.00881  -0.0387   0.9994   0.0970
  -0.500   0.1264   0.01701   0.00795  -0.0435   0.9893   0.0991
  -0.250   0.1772   0.01639   0.00729  -0.0482   0.9773   0.1040
   0.000   0.2273   0.01585   0.00671  -0.0528   0.9641   0.1178
   0.250   0.2734   0.01311   0.00638  -0.0559   0.9564   1.0000
   0.500   0.3247   0.01306   0.00598  -0.0604   0.9413   1.0000
   0.750   0.3686   0.01293   0.00569  -0.0633   0.9218   1.0000
   1.000   0.4110   0.01274   0.00536  -0.0656   0.9030   1.0000
   1.250   0.4460   0.01263   0.00515  -0.0664   0.8814   1.0000
   1.500   0.4787   0.01253   0.00496  -0.0665   0.8608   1.0000
   1.750   0.5064   0.01256   0.00492  -0.0657   0.8372   1.0000
   2.000   0.5331   0.01263   0.00489  -0.0647   0.8146   1.0000
   2.250   0.5589   0.01276   0.00493  -0.0636   0.7927   1.0000
   2.500   0.5836   0.01294   0.00503  -0.0625   0.7694   1.0000
   2.750   0.6082   0.01311   0.00512  -0.0611   0.7459   1.0000
   3.000   0.6328   0.01330   0.00527  -0.0599   0.7229   1.0000
   3.250   0.6575   0.01352   0.00548  -0.0589   0.7002   1.0000
   3.500   0.6827   0.01375   0.00567  -0.0579   0.6799   1.0000
   3.750   0.7075   0.01398   0.00592  -0.0570   0.6571   1.0000
   4.000   0.7322   0.01420   0.00616  -0.0560   0.6329   1.0000
   4.250   0.7566   0.01441   0.00644  -0.0549   0.6057   1.0000
   4.500   0.7805   0.01463   0.00666  -0.0536   0.5739   1.0000
   4.750   0.8025   0.01480   0.00675  -0.0520   0.5256   1.0000
   5.000   0.8236   0.01507   0.00685  -0.0503   0.4614   1.0000
   5.250   0.8446   0.01556   0.00708  -0.0489   0.3904   1.0000
   5.500   0.8660   0.01628   0.00758  -0.0478   0.3300   1.0000
   5.750   0.8871   0.01713   0.00818  -0.0468   0.2825   1.0000
   6.000   0.9084   0.01800   0.00889  -0.0460   0.2339   1.0000
   6.250   0.9309   0.01882   0.00955  -0.0451   0.1658   1.0000
   6.500   0.9461   0.02101   0.01078  -0.0435   0.0594   1.0000
   6.750   0.9644   0.02273   0.01258  -0.0419   0.0482   1.0000
   7.000   0.9834   0.02423   0.01435  -0.0403   0.0450   1.0000
   7.250   1.0022   0.02573   0.01605  -0.0387   0.0430   1.0000
   7.500   1.0205   0.02744   0.01789  -0.0370   0.0417   1.0000
   7.750   1.0402   0.02938   0.01993  -0.0354   0.0409   1.0000
   8.000   1.0623   0.03160   0.02228  -0.0341   0.0404   1.0000
   8.250   1.0839   0.03399   0.02482  -0.0330   0.0386   1.0000
   8.500   1.1046   0.03702   0.02796  -0.0321   0.0368   1.0000
   8.750   1.1252   0.03986   0.03121  -0.0307   0.0374   1.0000
   9.000   1.1414   0.04325   0.03530  -0.0286   0.0392   1.0000
   9.250   1.1501   0.04747   0.04019  -0.0264   0.0415   1.0000
   9.500   1.1545   0.05215   0.04537  -0.0243   0.0436   1.0000
   9.750   1.1558   0.05759   0.05114  -0.0226   0.0455   1.0000
  10.000   1.1557   0.06105   0.05522  -0.0200   0.0496   1.0000
  10.250   1.1298   0.06664   0.06140  -0.0177   0.0522   1.0000
  10.500   1.1066   0.07125   0.06627  -0.0159   0.0534   1.0000
  10.750   1.0831   0.07615   0.07136  -0.0162   0.0539   1.0000
  11.000   1.0582   0.08199   0.07738  -0.0191   0.0538   1.0000
  11.250   1.0322   0.08930   0.08476  -0.0246   0.0532   1.0000
  11.500   1.0056   0.09861   0.09415  -0.0321   0.0526   1.0000
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