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GOE 57 AIRFOIL (goe57-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: GOE 57 AIRFOIL (goe57-il)
Reynolds number: 500,000
Max Cl/Cd: 93.02 at α=2.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe57-il-500000.txt
Download as CSV file: xf-goe57-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 57 AIRFOIL                                  
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.3163   0.10184   0.09962  -0.0268   1.0000   0.0200
  -8.250  -0.3195   0.09974   0.09758  -0.0270   1.0000   0.0201
  -8.000  -0.3264   0.09799   0.09588  -0.0258   1.0000   0.0201
  -7.250  -0.3199   0.08843   0.08642  -0.0236   0.9977   0.0207
  -7.000  -0.2968   0.08480   0.08278  -0.0280   0.9952   0.0212
  -6.750  -0.2713   0.08117   0.07913  -0.0337   0.9915   0.0223
  -6.500  -0.2413   0.07688   0.07482  -0.0415   0.9869   0.0243
  -6.250  -0.1865   0.07118   0.06903  -0.0606   0.9808   0.0256
  -6.000  -0.1504   0.06622   0.06401  -0.0693   0.9756   0.0257
  -2.000   0.3312   0.01987   0.01551  -0.1145   0.8304   0.0486
  -1.750   0.3612   0.01917   0.01456  -0.1142   0.8200   0.0537
  -1.500   0.3905   0.01906   0.01419  -0.1138   0.8087   0.0549
  -1.250   0.4179   0.01555   0.01038  -0.1147   0.7973   0.0576
  -1.000   0.4448   0.01476   0.00951  -0.1146   0.7838   0.0594
  -0.750   0.4744   0.01303   0.00743  -0.1142   0.7702   0.0511
  -0.500   0.5030   0.01130   0.00528  -0.1138   0.7538   0.0464
  -0.250   0.5300   0.01061   0.00440  -0.1134   0.7316   0.0461
   0.000   0.5566   0.01007   0.00369  -0.1128   0.7082   0.0460
   0.250   0.5829   0.00977   0.00324  -0.1123   0.6815   0.0469
   0.500   0.6087   0.00962   0.00293  -0.1117   0.6533   0.0485
   0.750   0.6344   0.00955   0.00270  -0.1111   0.6238   0.0501
   1.000   0.6604   0.00949   0.00253  -0.1107   0.5972   0.0518
   1.250   0.6865   0.00950   0.00241  -0.1102   0.5694   0.0537
   1.500   0.7123   0.00946   0.00225  -0.1097   0.5366   0.0574
   1.750   0.7376   0.00957   0.00220  -0.1092   0.4944   0.0630
   2.000   0.7622   0.00983   0.00223  -0.1086   0.4478   0.0687
   2.250   0.7871   0.01005   0.00229  -0.1080   0.4064   0.0840
   2.500   0.8124   0.01007   0.00248  -0.1078   0.3693   0.2285
   2.750   0.8335   0.00896   0.00273  -0.1067   0.3333   1.0000
   3.000   0.8580   0.00941   0.00291  -0.1061   0.2899   1.0000
   3.250   0.8820   0.00992   0.00314  -0.1055   0.2436   1.0000
   3.500   0.9069   0.01031   0.00336  -0.1050   0.2193   1.0000
   3.750   0.9324   0.01063   0.00357  -0.1046   0.2075   1.0000
   4.000   0.9576   0.01096   0.00379  -0.1042   0.1959   1.0000
   4.250   0.9827   0.01129   0.00405  -0.1037   0.1871   1.0000
   4.500   1.0086   0.01153   0.00427  -0.1034   0.1829   1.0000
   4.750   1.0344   0.01176   0.00449  -0.1030   0.1787   1.0000
   5.000   1.0592   0.01211   0.00477  -0.1026   0.1723   1.0000
   5.250   1.0852   0.01229   0.00498  -0.1023   0.1658   1.0000
   5.500   1.1100   0.01263   0.00526  -0.1018   0.1577   1.0000
   5.750   1.1361   0.01279   0.00541  -0.1015   0.1480   1.0000
   6.000   1.1614   0.01306   0.00559  -0.1011   0.1314   1.0000
   6.250   1.1779   0.01440   0.00645  -0.0995   0.0448   1.0000
   6.500   1.1994   0.01516   0.00710  -0.0985   0.0206   1.0000
   6.750   1.2227   0.01567   0.00767  -0.0977   0.0188   1.0000
   7.000   1.2454   0.01624   0.00833  -0.0967   0.0172   1.0000
   7.250   1.2669   0.01695   0.00915  -0.0957   0.0159   1.0000
   7.500   1.2886   0.01760   0.00990  -0.0946   0.0154   1.0000
   7.750   1.3092   0.01835   0.01074  -0.0935   0.0149   1.0000
   8.000   1.3286   0.01920   0.01168  -0.0921   0.0144   1.0000
   8.250   1.3466   0.02014   0.01271  -0.0906   0.0140   1.0000
   8.500   1.3628   0.02119   0.01385  -0.0888   0.0136   1.0000
   8.750   1.3774   0.02232   0.01509  -0.0867   0.0133   1.0000
   9.000   1.3900   0.02355   0.01640  -0.0845   0.0130   1.0000
   9.250   1.3996   0.02494   0.01786  -0.0819   0.0125   1.0000
   9.500   1.4033   0.02666   0.01963  -0.0785   0.0121   1.0000
   9.750   1.4058   0.02851   0.02153  -0.0750   0.0118   1.0000
  10.000   1.4123   0.03026   0.02334  -0.0721   0.0118   1.0000
  10.250   1.4219   0.03201   0.02514  -0.0698   0.0117   1.0000
  10.500   1.4342   0.03381   0.02702  -0.0679   0.0117   1.0000
  10.750   1.4478   0.03551   0.02878  -0.0661   0.0118   1.0000
  11.000   1.4595   0.03700   0.03037  -0.0642   0.0119   1.0000
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