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GOE 587 AIRFOIL (goe587-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: GOE 587 AIRFOIL (goe587-il)
Reynolds number: 100,000
Max Cl/Cd: 50.13 at α=4°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe587-il-100000.txt
Download as CSV file: xf-goe587-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 587 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
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.4588   0.09610   0.09123  -0.0129   1.0000   0.0526
  -7.750  -0.4615   0.09366   0.08887  -0.0153   1.0000   0.0536
  -7.500  -0.4624   0.09153   0.08679  -0.0195   1.0000   0.0545
  -7.250  -0.4578   0.08947   0.08469  -0.0241   1.0000   0.0550
  -7.000  -0.4502   0.08720   0.08234  -0.0274   1.0000   0.0553
  -6.750  -0.4468   0.07951   0.07485  -0.0215   1.0000   0.0581
  -6.500  -0.4395   0.07619   0.07155  -0.0213   1.0000   0.0615
  -6.250  -0.4301   0.07319   0.06846  -0.0235   1.0000   0.0661
  -6.000  -0.4125   0.07276   0.06772  -0.0297   1.0000   0.0687
  -5.750  -0.4095   0.06610   0.06124  -0.0269   1.0000   0.0707
  -5.500  -0.4005   0.06264   0.05781  -0.0257   1.0000   0.0738
  -5.250  -0.3849   0.05986   0.05490  -0.0268   1.0000   0.0800
  -5.000  -0.3699   0.05653   0.05138  -0.0280   1.0000   0.0835
  -4.750  -0.3603   0.05299   0.04791  -0.0263   1.0000   0.0875
  -4.500  -0.3400   0.05093   0.04547  -0.0273   1.0000   0.0966
  -4.250  -0.3305   0.04712   0.04180  -0.0255   1.0000   0.1005
  -4.000  -0.3133   0.04469   0.03911  -0.0251   1.0000   0.1111
  -3.500  -0.2859   0.03941   0.03372  -0.0224   1.0000   0.1296
  -3.250  -0.2704   0.03779   0.03177  -0.0212   1.0000   0.1512
  -3.000  -0.2601   0.03523   0.02932  -0.0192   1.0000   0.1727
  -2.250  -0.2269   0.02919   0.02322  -0.0132   1.0000   0.2774
  -2.000  -0.1620   0.02483   0.01678  -0.0131   1.0000   0.0805
  -1.750  -0.1411   0.02274   0.01443  -0.0119   1.0000   0.0760
  -1.500  -0.1202   0.02180   0.01304  -0.0103   1.0000   0.0813
  -1.250  -0.0868   0.01994   0.01091  -0.0114   0.9964   0.0848
  -1.000  -0.0393   0.01859   0.00932  -0.0151   0.9871   0.0967
  -0.750   0.0064   0.01731   0.00806  -0.0187   0.9772   0.1159
  -0.500   0.0518   0.01631   0.00720  -0.0224   0.9664   0.1467
  -0.250   0.1325   0.01316   0.00633  -0.0315   0.9738   1.0000
   0.000   0.1873   0.01322   0.00606  -0.0373   0.9602   1.0000
   0.250   0.2426   0.01319   0.00584  -0.0431   0.9470   1.0000
   0.500   0.2990   0.01305   0.00560  -0.0489   0.9338   1.0000
   0.750   0.3523   0.01286   0.00536  -0.0540   0.9191   1.0000
   1.000   0.4025   0.01266   0.00512  -0.0583   0.9026   1.0000
   1.250   0.4396   0.01262   0.00506  -0.0599   0.8808   1.0000
   1.500   0.4734   0.01263   0.00504  -0.0607   0.8604   1.0000
   1.750   0.4991   0.01277   0.00521  -0.0599   0.8384   1.0000
   2.000   0.5241   0.01291   0.00534  -0.0589   0.8188   1.0000
   2.250   0.5452   0.01295   0.00532  -0.0566   0.7929   1.0000
   2.500   0.5645   0.01300   0.00530  -0.0539   0.7640   1.0000
   2.750   0.5857   0.01317   0.00548  -0.0521   0.7423   1.0000
   3.000   0.6068   0.01334   0.00566  -0.0503   0.7203   1.0000
   3.250   0.6273   0.01346   0.00585  -0.0481   0.6955   1.0000
   3.500   0.6479   0.01360   0.00605  -0.0461   0.6717   1.0000
   3.750   0.6642   0.01353   0.00594  -0.0428   0.6277   1.0000
   4.000   0.6788   0.01354   0.00583  -0.0390   0.5657   1.0000
   4.250   0.6896   0.01393   0.00583  -0.0346   0.4468   1.0000
   4.500   0.6860   0.01690   0.00673  -0.0296   0.1109   1.0000
   4.750   0.7014   0.01821   0.00799  -0.0272   0.0856   1.0000
   5.000   0.7161   0.01949   0.00941  -0.0247   0.0766   1.0000
   5.250   0.7322   0.02064   0.01063  -0.0223   0.0688   1.0000
   5.500   0.7480   0.02226   0.01219  -0.0200   0.0622   1.0000
   5.750   0.7705   0.02403   0.01396  -0.0185   0.0596   1.0000
   6.000   0.7973   0.02619   0.01619  -0.0174   0.0588   1.0000
   6.250   0.8234   0.02851   0.01862  -0.0165   0.0563   1.0000
   6.500   0.8477   0.03162   0.02184  -0.0157   0.0537   1.0000
   6.750   0.8723   0.03405   0.02505  -0.0133   0.0596   1.0000
   7.000   0.8969   0.03737   0.02895  -0.0110   0.0696   1.0000
  10.250   0.8432   0.10453   0.09989  -0.0113   0.1559   1.0000
  10.500   0.8087   0.11116   0.10631  -0.0193   0.1530   1.0000
  10.750   0.6926   0.11342   0.10888  -0.0164   0.1684   1.0000
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