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GOE 55 AIRFOIL (goe55-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: GOE 55 AIRFOIL (goe55-il)
Reynolds number: 50,000
Max Cl/Cd: 26.29 at α=6°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe55-il-50000.txt
Download as CSV file: xf-goe55-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 55 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.6256   0.12123   0.11417   0.0423   1.0000   0.1673
  -8.500  -0.6227   0.11834   0.11132   0.0410   1.0000   0.1749
  -8.250  -0.6408   0.11822   0.11133   0.0349   1.0000   0.1778
  -8.000  -0.6105   0.11080   0.10386   0.0395   1.0000   0.1859
  -7.750  -0.6199   0.10933   0.10249   0.0351   1.0000   0.1926
  -7.500  -0.6048   0.10409   0.09726   0.0366   1.0000   0.1983
  -7.250  -0.6049   0.10170   0.09493   0.0321   1.0000   0.2079
  -6.750  -0.5890   0.09400   0.08729   0.0275   1.0000   0.2258
  -6.500  -0.5807   0.09108   0.08438   0.0228   1.0000   0.2404
  -6.000  -0.5570   0.08343   0.07679   0.0223   1.0000   0.2722
  -5.750  -0.5441   0.07947   0.07287   0.0236   1.0000   0.2889
  -5.500  -0.5308   0.07572   0.06916   0.0257   1.0000   0.3072
  -5.250  -0.5191   0.07280   0.06626   0.0237   1.0000   0.3365
  -5.000  -0.5057   0.06924   0.06278   0.0285   1.0000   0.3600
  -4.750  -0.4963   0.06639   0.06000   0.0302   1.0000   0.4015
  -4.500  -0.4876   0.06366   0.05738   0.0359   1.0000   0.4492
  -4.250  -0.4779   0.06093   0.05473   0.0443   1.0000   0.5035
  -3.750  -0.1866   0.04154   0.03490   0.0362   1.0000   1.0000
  -3.500  -0.1878   0.04007   0.03351   0.0382   1.0000   0.9936
  -3.250  -0.2457   0.04133   0.03499   0.0520   1.0000   0.9445
  -3.000  -0.2184   0.03551   0.02676  -0.0247   1.0000   0.2171
  -2.750  -0.1767   0.03232   0.02286  -0.0276   1.0000   0.1895
  -2.500  -0.1397   0.02991   0.01981  -0.0290   1.0000   0.1770
  -2.250  -0.1083   0.02780   0.01750  -0.0295   1.0000   0.1749
  -2.000  -0.0763   0.02617   0.01552  -0.0298   1.0000   0.1769
  -1.750  -0.0453   0.02457   0.01373  -0.0299   1.0000   0.1788
  -1.500  -0.0158   0.02305   0.01219  -0.0298   1.0000   0.1818
  -1.250   0.0135   0.02185   0.01096  -0.0295   1.0000   0.1891
  -1.000   0.0420   0.02072   0.00987  -0.0289   1.0000   0.2017
  -0.750   0.0715   0.01964   0.00885  -0.0285   1.0000   0.2163
  -0.500   0.1003   0.01855   0.00800  -0.0281   1.0000   0.2455
  -0.250   0.1295   0.01713   0.00721  -0.0281   1.0000   0.3109
   0.000   0.1516   0.01418   0.00624  -0.0252   1.0000   1.0000
   0.250   0.1845   0.01433   0.00641  -0.0272   1.0000   1.0000
   0.500   0.2525   0.01532   0.00668  -0.0344   0.7913   1.0000
   0.750   0.2759   0.01616   0.00694  -0.0323   0.7338   1.0000
   1.000   0.3023   0.01690   0.00730  -0.0314   0.6973   1.0000
   1.250   0.3296   0.01759   0.00772  -0.0308   0.6704   1.0000
   1.500   0.3582   0.01824   0.00823  -0.0307   0.6467   1.0000
   1.750   0.3863   0.01890   0.00874  -0.0304   0.6258   1.0000
   2.000   0.4136   0.01957   0.00927  -0.0299   0.6055   1.0000
   2.250   0.4412   0.02024   0.00989  -0.0296   0.5835   1.0000
   2.500   0.4679   0.02094   0.01050  -0.0289   0.5624   1.0000
   2.750   0.4943   0.02171   0.01116  -0.0282   0.5427   1.0000
   3.000   0.5219   0.02254   0.01202  -0.0280   0.5216   1.0000
   3.250   0.5489   0.02344   0.01292  -0.0276   0.5006   1.0000
   3.500   0.5754   0.02442   0.01386  -0.0269   0.4803   1.0000
   3.750   0.6024   0.02557   0.01507  -0.0267   0.4589   1.0000
   4.000   0.6293   0.02681   0.01642  -0.0264   0.4368   1.0000
   4.250   0.6551   0.02806   0.01766  -0.0257   0.4162   1.0000
   4.500   0.6821   0.02967   0.01944  -0.0257   0.3956   1.0000
   4.750   0.7078   0.03101   0.02094  -0.0253   0.3723   1.0000
   5.000   0.7317   0.03181   0.02175  -0.0241   0.3452   1.0000
   5.250   0.7543   0.03160   0.02139  -0.0222   0.3118   1.0000
   5.500   0.7779   0.03135   0.02105  -0.0206   0.2792   1.0000
   5.750   0.8024   0.03121   0.02083  -0.0190   0.2447   1.0000
   6.000   0.8275   0.03147   0.02100  -0.0176   0.2048   1.0000
   6.250   0.8526   0.03260   0.02200  -0.0167   0.1800   1.0000
   6.500   0.8776   0.03498   0.02454  -0.0164   0.1662   1.0000
   6.750   0.9023   0.03727   0.02684  -0.0159   0.1575   1.0000
   7.000   0.9236   0.04189   0.03223  -0.0167   0.1514   1.0000
   7.250   0.9450   0.04495   0.03553  -0.0167   0.1453   1.0000
   7.500   0.9637   0.04885   0.03968  -0.0170   0.1412   1.0000
   7.750   0.9737   0.05570   0.04723  -0.0191   0.1406   1.0000
   8.000   0.9785   0.06311   0.05512  -0.0217   0.1413   1.0000
   8.250   0.9838   0.06981   0.06212  -0.0238   0.1423   1.0000
   8.500   0.9006   0.09812   0.09086  -0.0523   0.1671   1.0000
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