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GOE 113 AIRFOIL (goe113-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: GOE 113 AIRFOIL (goe113-il)
Reynolds number: 50,000
Max Cl/Cd: 37.7 at α=6.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe113-il-50000.txt
Download as CSV file: xf-goe113-il-50000.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 =     0.050 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.4797   0.10744   0.10089  -0.0058   1.0000   0.1469
  -7.750  -0.4667   0.10254   0.09601  -0.0043   1.0000   0.1572
  -7.500  -0.4766   0.10134   0.09494  -0.0099   1.0000   0.1613
  -7.250  -0.4648   0.09666   0.09029  -0.0076   1.0000   0.1736
  -7.000  -0.4566   0.09268   0.08637  -0.0073   1.0000   0.1837
  -6.750  -0.4522   0.08926   0.08301  -0.0088   1.0000   0.1945
  -6.500  -0.4467   0.08587   0.07968  -0.0100   1.0000   0.2079
  -6.250  -0.4395   0.08248   0.07635  -0.0102   1.0000   0.2246
  -6.000  -0.4371   0.08000   0.07390  -0.0144   1.0000   0.2447
  -5.750  -0.4262   0.07591   0.06990  -0.0099   1.0000   0.2673
  -5.500  -0.4195   0.07274   0.06679  -0.0085   1.0000   0.2953
  -5.250  -0.4156   0.06987   0.06395  -0.0076   1.0000   0.3302
  -5.000  -0.4097   0.06687   0.06104  -0.0026   1.0000   0.3727
  -4.750  -0.0924   0.04378   0.03696  -0.0067   1.0000   1.0000
  -4.500  -0.0800   0.04143   0.03465  -0.0082   1.0000   1.0000
  -4.250  -0.0702   0.03933   0.03259  -0.0089   1.0000   0.9986
  -4.000  -0.1255   0.04156   0.03506   0.0055   1.0000   0.9579
  -3.750  -0.1743   0.04273   0.03647   0.0162   1.0000   0.9038
  -3.500  -0.2211   0.04324   0.03723   0.0242   1.0000   0.8477
  -3.250  -0.2686   0.04331   0.03761   0.0315   1.0000   0.8030
  -3.000  -0.1786   0.03529   0.02717  -0.0423   1.0000   0.2151
  -2.750  -0.1440   0.03230   0.02356  -0.0433   1.0000   0.1839
  -2.500  -0.1124   0.02989   0.02052  -0.0435   1.0000   0.1666
  -2.250  -0.0822   0.02787   0.01794  -0.0432   1.0000   0.1524
  -2.000  -0.0537   0.02628   0.01587  -0.0425   1.0000   0.1450
  -1.750  -0.0270   0.02479   0.01409  -0.0419   1.0000   0.1438
  -1.500  -0.0013   0.02362   0.01266  -0.0413   1.0000   0.1498
  -1.250   0.0237   0.02255   0.01145  -0.0403   1.0000   0.1513
  -1.000   0.0499   0.02172   0.01043  -0.0396   1.0000   0.1539
  -0.750   0.0747   0.02110   0.00969  -0.0389   1.0000   0.1588
  -0.500   0.0990   0.02057   0.00914  -0.0386   1.0000   0.1672
  -0.250   0.1237   0.02017   0.00878  -0.0387   1.0000   0.1875
   0.000   0.1466   0.01724   0.00863  -0.0380   1.0000   0.9111
   0.250   0.1630   0.01778   0.00859  -0.0365   1.0000   1.0000
   0.500   0.1827   0.01842   0.00875  -0.0361   1.0000   1.0000
   0.750   0.2017   0.01913   0.00917  -0.0361   1.0000   1.0000
   1.000   0.2203   0.01990   0.00973  -0.0363   1.0000   1.0000
   1.250   0.2388   0.02073   0.01039  -0.0366   1.0000   1.0000
   1.500   0.2709   0.02165   0.01116  -0.0395   0.9945   1.0000
   1.750   0.3235   0.02256   0.01192  -0.0460   0.9773   1.0000
   2.000   0.3765   0.02336   0.01266  -0.0522   0.9593   1.0000
   2.250   0.4330   0.02388   0.01318  -0.0583   0.9360   1.0000
   2.500   0.4946   0.02393   0.01333  -0.0643   0.9079   1.0000
   2.750   0.5521   0.02377   0.01329  -0.0690   0.8830   1.0000
   3.000   0.5999   0.02366   0.01332  -0.0717   0.8618   1.0000
   3.250   0.6380   0.02364   0.01351  -0.0725   0.8384   1.0000
   3.500   0.6756   0.02343   0.01346  -0.0726   0.8138   1.0000
   3.750   0.7086   0.02317   0.01333  -0.0712   0.7857   1.0000
   4.000   0.7374   0.02291   0.01319  -0.0689   0.7545   1.0000
   4.250   0.7639   0.02268   0.01312  -0.0659   0.7209   1.0000
   4.500   0.7870   0.02267   0.01316  -0.0627   0.6824   1.0000
   4.750   0.8095   0.02275   0.01325  -0.0595   0.6398   1.0000
   5.000   0.8305   0.02310   0.01359  -0.0565   0.5904   1.0000
   5.250   0.8531   0.02359   0.01398  -0.0539   0.5437   1.0000
   5.500   0.8761   0.02436   0.01481  -0.0520   0.5026   1.0000
   5.750   0.8974   0.02478   0.01522  -0.0499   0.4628   1.0000
   6.000   0.9145   0.02453   0.01483  -0.0469   0.4154   1.0000
   6.250   0.9365   0.02511   0.01553  -0.0454   0.3850   1.0000
   6.500   0.9539   0.02530   0.01576  -0.0432   0.3445   1.0000
   6.750   0.9645   0.02561   0.01586  -0.0404   0.2821   1.0000
   7.000   0.9737   0.02719   0.01715  -0.0380   0.1878   1.0000
   7.250   0.9846   0.02990   0.01935  -0.0356   0.1159   1.0000
   7.500   1.0006   0.03218   0.02151  -0.0335   0.0962   1.0000
   7.750   1.0204   0.03434   0.02376  -0.0317   0.0864   1.0000
   8.000   1.0410   0.03657   0.02603  -0.0302   0.0787   1.0000
   8.250   1.0623   0.03886   0.02843  -0.0289   0.0723   1.0000
   8.500   1.0871   0.04220   0.03186  -0.0279   0.0699   1.0000
   8.750   1.1089   0.04602   0.03617  -0.0267   0.0693   1.0000
   9.000   1.1252   0.05029   0.04095  -0.0253   0.0694   1.0000
   9.250   1.1366   0.05492   0.04604  -0.0238   0.0699   1.0000
   9.500   1.1432   0.05965   0.05136  -0.0222   0.0706   1.0000
   9.750   1.1310   0.06484   0.05735  -0.0201   0.0725   1.0000
  10.000   1.1110   0.07050   0.06347  -0.0187   0.0741   1.0000
  10.250   1.0883   0.07575   0.06904  -0.0179   0.0753   1.0000
  10.500   1.0629   0.08098   0.07444  -0.0181   0.0762   1.0000
  10.750   1.0361   0.08724   0.08084  -0.0209   0.0771   1.0000
  11.000   1.0098   0.09495   0.08861  -0.0261   0.0782   1.0000
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