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GOE 457 AIRFOIL (goe457-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: GOE 457 AIRFOIL (goe457-il)
Reynolds number: 100,000
Max Cl/Cd: 56.89 at α=6.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe457-il-100000-n5.txt
Download as CSV file: xf-goe457-il-100000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 457 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.500  -0.3493   0.09337   0.08872  -0.0228   1.0000   0.0486
  -7.250  -0.3622   0.09193   0.08739  -0.0223   1.0000   0.0499
  -7.000  -0.3692   0.08977   0.08532  -0.0253   1.0000   0.0507
  -6.750  -0.3694   0.08682   0.08241  -0.0283   1.0000   0.0510
  -6.500  -0.3656   0.08345   0.07904  -0.0306   1.0000   0.0511
  -6.250  -0.3595   0.07976   0.07534  -0.0325   1.0000   0.0512
  -5.250  -0.3267   0.06210   0.05758  -0.0355   1.0000   0.0388
  -5.000  -0.2968   0.05689   0.05225  -0.0410   0.9959   0.0379
  -4.750  -0.2546   0.05001   0.04506  -0.0498   0.9901   0.0382
  -4.250  -0.1741   0.03687   0.03086  -0.0627   0.9799   0.0404
  -4.000  -0.1419   0.03530   0.02923  -0.0652   0.9749   0.0432
  -3.750  -0.1072   0.03211   0.02560  -0.0679   0.9692   0.0459
  -3.500  -0.0683   0.02859   0.02146  -0.0708   0.9649   0.0469
  -3.250  -0.0350   0.02621   0.01843  -0.0720   0.9580   0.0502
  -3.000   0.0038   0.02417   0.01580  -0.0741   0.9532   0.0512
  -2.750   0.0351   0.02253   0.01391  -0.0748   0.9455   0.0523
  -2.500   0.0732   0.02141   0.01264  -0.0769   0.9396   0.0546
  -2.250   0.1053   0.02061   0.01169  -0.0775   0.9307   0.0581
  -2.000   0.1450   0.01952   0.01038  -0.0795   0.9244   0.0599
  -1.750   0.1771   0.01865   0.00939  -0.0799   0.9144   0.0614
  -1.500   0.2151   0.01780   0.00846  -0.0814   0.9067   0.0634
  -1.250   0.2492   0.01704   0.00774  -0.0823   0.8959   0.0668
  -1.000   0.2813   0.01647   0.00715  -0.0827   0.8829   0.0716
  -0.750   0.3132   0.01594   0.00661  -0.0831   0.8682   0.0815
  -0.500   0.3455   0.01538   0.00607  -0.0835   0.8512   0.0991
  -0.250   0.3735   0.01478   0.00565  -0.0831   0.8286   0.1416
   0.000   0.4063   0.01404   0.00539  -0.0838   0.8031   0.3139
   0.250   0.4580   0.01222   0.00503  -0.0879   0.7729   1.0000
   0.500   0.4975   0.01213   0.00458  -0.0894   0.7353   1.0000
   0.750   0.5320   0.01221   0.00430  -0.0901   0.6960   1.0000
   1.000   0.5620   0.01242   0.00418  -0.0901   0.6599   1.0000
   1.250   0.5896   0.01268   0.00416  -0.0897   0.6285   1.0000
   1.500   0.6159   0.01296   0.00420  -0.0891   0.6020   1.0000
   1.750   0.6419   0.01325   0.00428  -0.0885   0.5802   1.0000
   2.000   0.6678   0.01355   0.00440  -0.0879   0.5613   1.0000
   2.250   0.6935   0.01385   0.00455  -0.0874   0.5449   1.0000
   2.500   0.7192   0.01416   0.00473  -0.0868   0.5299   1.0000
   2.750   0.7447   0.01447   0.00495  -0.0862   0.5162   1.0000
   3.000   0.7702   0.01479   0.00518  -0.0857   0.5035   1.0000
   3.250   0.7956   0.01512   0.00543  -0.0851   0.4919   1.0000
   3.500   0.8209   0.01543   0.00572  -0.0845   0.4799   1.0000
   3.750   0.8460   0.01574   0.00602  -0.0839   0.4682   1.0000
   4.000   0.8710   0.01607   0.00633  -0.0833   0.4572   1.0000
   4.250   0.8957   0.01640   0.00665  -0.0826   0.4459   1.0000
   4.500   0.9197   0.01671   0.00696  -0.0819   0.4328   1.0000
   4.750   0.9432   0.01701   0.00726  -0.0810   0.4180   1.0000
   5.000   0.9662   0.01731   0.00757  -0.0800   0.4027   1.0000
   5.250   0.9889   0.01761   0.00787  -0.0790   0.3865   1.0000
   5.500   1.0114   0.01790   0.00819  -0.0779   0.3690   1.0000
   5.750   1.0337   0.01823   0.00854  -0.0769   0.3520   1.0000
   6.000   1.0560   0.01857   0.00894  -0.0759   0.3360   1.0000
   6.250   1.0781   0.01895   0.00935  -0.0749   0.3199   1.0000
   6.500   1.0998   0.01935   0.00978  -0.0738   0.3029   1.0000
   6.750   1.1207   0.01979   0.01024  -0.0726   0.2847   1.0000
   7.000   1.1416   0.02025   0.01079  -0.0714   0.2645   1.0000
   7.250   1.1617   0.02079   0.01135  -0.0702   0.2432   1.0000
   7.500   1.1809   0.02141   0.01197  -0.0688   0.2195   1.0000
   7.750   1.1987   0.02215   0.01268  -0.0673   0.1940   1.0000
   8.000   1.2151   0.02303   0.01347  -0.0657   0.1670   1.0000
   8.250   1.2299   0.02405   0.01439  -0.0639   0.1438   1.0000
   8.500   1.2436   0.02515   0.01537  -0.0621   0.1264   1.0000
   8.750   1.2578   0.02617   0.01638  -0.0603   0.1143   1.0000
   9.000   1.2736   0.02703   0.01736  -0.0587   0.1060   1.0000
   9.250   1.2865   0.02806   0.01841  -0.0568   0.0989   1.0000
   9.500   1.3026   0.02883   0.01937  -0.0552   0.0908   1.0000
   9.750   1.3148   0.02980   0.02046  -0.0532   0.0850   1.0000
  10.000   1.3279   0.03066   0.02144  -0.0513   0.0749   1.0000
  10.250   1.3395   0.03160   0.02249  -0.0492   0.0642   1.0000
  10.500   1.3489   0.03269   0.02367  -0.0469   0.0530   1.0000
  11.000   1.3541   0.03593   0.02686  -0.0415   0.0259   1.0000
  11.250   1.3519   0.03799   0.02894  -0.0388   0.0237   1.0000
  11.500   1.3478   0.04033   0.03136  -0.0363   0.0219   1.0000
  11.750   1.3434   0.04283   0.03400  -0.0343   0.0209   1.0000
  12.000   1.3369   0.04571   0.03703  -0.0327   0.0198   1.0000
  12.250   1.3306   0.04880   0.04033  -0.0316   0.0191   1.0000
  12.500   1.3237   0.05223   0.04398  -0.0313   0.0187   1.0000
  12.750   1.3147   0.05624   0.04825  -0.0316   0.0183   1.0000
  13.000   1.3053   0.06065   0.05286  -0.0325   0.0180   1.0000
  13.250   1.2945   0.06561   0.05803  -0.0340   0.0178   1.0000
  13.500   1.2828   0.07099   0.06361  -0.0360   0.0177   1.0000
  13.750   1.2706   0.07667   0.06948  -0.0383   0.0175   1.0000
  14.000   1.2574   0.08266   0.07565  -0.0409   0.0174   1.0000
  14.250   1.2442   0.08881   0.08198  -0.0435   0.0174   1.0000
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