Airfoil Tools
Search 1636 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

GOE 723 AIRFOIL (goe723-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: GOE 723 AIRFOIL (goe723-il)
Reynolds number: 100,000
Max Cl/Cd: 58.52 at α=6.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe723-il-100000-n5.txt
Download as CSV file: xf-goe723-il-100000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 723 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
  ------ -------- --------- --------- -------- -------- --------
 -10.000  -0.2293   0.10595   0.10104  -0.0416   1.0000   0.0840
  -9.750  -0.2518   0.10377   0.09896  -0.0439   1.0000   0.0865
  -9.500  -0.2684   0.10138   0.09667  -0.0448   1.0000   0.0869
  -9.250  -0.2606   0.09796   0.09332  -0.0423   1.0000   0.0879
  -9.000  -0.2406   0.09505   0.09042  -0.0411   0.9978   0.0898
  -8.750  -0.2247   0.09100   0.08636  -0.0439   0.9915   0.0916
  -8.500  -0.2133   0.08655   0.08190  -0.0474   0.9843   0.0923
  -8.000  -0.3474   0.08009   0.07520  -0.0526   0.9869   0.0541
  -7.750  -0.3282   0.07660   0.07169  -0.0559   0.9763   0.0533
  -7.500  -0.3113   0.07151   0.06658  -0.0621   0.9630   0.0522
  -7.250  -0.2965   0.06457   0.05959  -0.0712   0.9489   0.0508
  -7.000  -0.2894   0.04794   0.04240  -0.0928   0.9323   0.0486
  -6.750  -0.2701   0.04224   0.03622  -0.0988   0.9201   0.0490
  -6.500  -0.2468   0.03792   0.03139  -0.1026   0.9101   0.0492
  -6.250  -0.2201   0.03417   0.02708  -0.1056   0.9015   0.0489
  -6.000  -0.1948   0.03153   0.02397  -0.1068   0.8913   0.0488
  -5.500  -0.1382   0.02761   0.01926  -0.1088   0.8735   0.0490
  -5.250  -0.1065   0.02604   0.01737  -0.1099   0.8663   0.0494
  -5.000  -0.0807   0.02486   0.01596  -0.1097   0.8557   0.0499
  -4.750  -0.0492   0.02377   0.01461  -0.1105   0.8485   0.0512
  -4.500  -0.0236   0.02293   0.01358  -0.1101   0.8380   0.0524
  -4.250   0.0066   0.02205   0.01249  -0.1105   0.8305   0.0533
  -4.000   0.0325   0.02131   0.01163  -0.1100   0.8208   0.0538
  -3.750   0.0617   0.02061   0.01080  -0.1101   0.8133   0.0543
  -3.500   0.0867   0.01994   0.01013  -0.1095   0.8040   0.0551
  -3.250   0.1145   0.01932   0.00947  -0.1094   0.7969   0.0561
  -3.000   0.1391   0.01889   0.00902  -0.1087   0.7878   0.0573
  -2.750   0.1675   0.01846   0.00849  -0.1086   0.7813   0.0590
  -2.500   0.1917   0.01817   0.00814  -0.1079   0.7722   0.0617
  -2.250   0.2196   0.01782   0.00769  -0.1077   0.7656   0.0652
  -2.000   0.2445   0.01753   0.00737  -0.1071   0.7575   0.0691
  -1.750   0.2718   0.01727   0.00700  -0.1067   0.7506   0.0743
  -1.500   0.2978   0.01694   0.00672  -0.1063   0.7434   0.0863
  -1.250   0.3230   0.01626   0.00658  -0.1060   0.7362   0.2038
  -1.000   0.3505   0.01594   0.00661  -0.1059   0.7303   0.3210
  -0.750   0.3750   0.01591   0.00673  -0.1052   0.7223   0.3818
  -0.500   0.4033   0.01587   0.00672  -0.1050   0.7161   0.4258
  -0.250   0.4293   0.01587   0.00680  -0.1045   0.7090   0.4676
   0.000   0.4555   0.01582   0.00688  -0.1039   0.7019   0.5202
   0.250   0.4834   0.01575   0.00687  -0.1037   0.6958   0.5670
   0.500   0.5091   0.01571   0.00692  -0.1031   0.6876   0.6031
   0.750   0.5379   0.01555   0.00684  -0.1029   0.6812   0.6501
   1.000   0.5655   0.01525   0.00688  -0.1026   0.6726   0.7405
   1.250   0.6274   0.01496   0.00672  -0.1095   0.6647   1.0000
   1.500   0.6522   0.01516   0.00683  -0.1088   0.6565   1.0000
   1.750   0.6788   0.01531   0.00687  -0.1084   0.6494   1.0000
   2.000   0.7043   0.01551   0.00699  -0.1078   0.6421   1.0000
   2.250   0.7298   0.01570   0.00711  -0.1073   0.6347   1.0000
   2.500   0.7563   0.01587   0.00720  -0.1068   0.6281   1.0000
   2.750   0.7804   0.01609   0.00741  -0.1060   0.6200   1.0000
   3.000   0.8076   0.01625   0.00749  -0.1057   0.6138   1.0000
   3.250   0.8308   0.01651   0.00778  -0.1048   0.6054   1.0000
   3.500   0.8571   0.01668   0.00791  -0.1043   0.5986   1.0000
   3.750   0.8808   0.01693   0.00819  -0.1034   0.5906   1.0000
   4.000   0.9062   0.01713   0.00837  -0.1028   0.5832   1.0000
   4.250   0.9303   0.01738   0.00866  -0.1020   0.5753   1.0000
   4.500   0.9549   0.01759   0.00887  -0.1012   0.5669   1.0000
   4.750   0.9777   0.01780   0.00910  -0.1001   0.5562   1.0000
   5.000   1.0010   0.01793   0.00922  -0.0990   0.5440   1.0000
   5.250   1.0236   0.01806   0.00932  -0.0977   0.5303   1.0000
   5.500   1.0450   0.01825   0.00952  -0.0963   0.5163   1.0000
   5.750   1.0660   0.01849   0.00982  -0.0949   0.5028   1.0000
   6.000   1.0867   0.01872   0.01007  -0.0934   0.4877   1.0000
   6.250   1.1062   0.01895   0.01029  -0.0917   0.4700   1.0000
   6.500   1.1253   0.01923   0.01057  -0.0900   0.4528   1.0000
   6.750   1.1442   0.01956   0.01092  -0.0883   0.4360   1.0000
   7.000   1.1619   0.01992   0.01129  -0.0864   0.4173   1.0000
   7.250   1.1780   0.02035   0.01170  -0.0843   0.3953   1.0000
   7.500   1.1922   0.02085   0.01213  -0.0819   0.3718   1.0000
   7.750   1.2045   0.02146   0.01269  -0.0793   0.3451   1.0000
   8.000   1.2130   0.02223   0.01333  -0.0761   0.3134   1.0000
   8.250   1.2169   0.02317   0.01411  -0.0723   0.2765   1.0000
   8.500   1.2126   0.02450   0.01513  -0.0676   0.2175   1.0000
   8.750   1.1971   0.02668   0.01675  -0.0622   0.1484   1.0000
   9.000   1.1920   0.02859   0.01842  -0.0585   0.1254   1.0000
   9.250   1.1907   0.03038   0.02014  -0.0555   0.1136   1.0000
   9.500   1.1936   0.03198   0.02181  -0.0531   0.1052   1.0000
   9.750   1.1936   0.03388   0.02371  -0.0509   0.0997   1.0000
  10.000   1.1976   0.03555   0.02548  -0.0490   0.0952   1.0000
  10.250   1.2014   0.03733   0.02735  -0.0474   0.0914   1.0000
  10.500   1.2036   0.03930   0.02939  -0.0458   0.0883   1.0000
  10.750   1.2041   0.04149   0.03158  -0.0443   0.0857   1.0000
  11.000   1.2102   0.04325   0.03347  -0.0431   0.0829   1.0000
  11.250   1.2156   0.04511   0.03544  -0.0419   0.0803   1.0000
  11.500   1.2208   0.04701   0.03740  -0.0409   0.0779   1.0000
  11.750   1.2258   0.04895   0.03937  -0.0398   0.0759   1.0000
  12.000   1.2316   0.05083   0.04123  -0.0386   0.0740   1.0000
  12.250   1.2409   0.05244   0.04296  -0.0376   0.0718   1.0000
  12.500   1.2499   0.05410   0.04473  -0.0367   0.0694   1.0000
  12.750   1.2585   0.05580   0.04648  -0.0358   0.0672   1.0000
  13.000   1.2672   0.05752   0.04820  -0.0349   0.0652   1.0000
  13.250   1.2788   0.05903   0.04969  -0.0338   0.0629   1.0000
  13.500   1.2878   0.06086   0.05170  -0.0330   0.0608   1.0000
  13.750   1.2960   0.06277   0.05376  -0.0323   0.0586   1.0000
  14.000   1.3034   0.06476   0.05581  -0.0317   0.0565   1.0000
  14.250   1.3140   0.06648   0.05750  -0.0310   0.0546   1.0000
  14.500   1.3235   0.06848   0.05960  -0.0302   0.0527   1.0000
  14.750   1.3261   0.07115   0.06253  -0.0299   0.0510   1.0000
  15.250   1.3322   0.07649   0.06821  -0.0294   0.0476   1.0000
  15.500   1.3363   0.07899   0.07077  -0.0292   0.0462   1.0000
  15.750   1.3448   0.08116   0.07289  -0.0288   0.0446   1.0000
  16.000   1.3365   0.08535   0.07743  -0.0294   0.0435   1.0000
  16.250   1.3292   0.08959   0.08196  -0.0302   0.0424   1.0000
  16.500   1.3223   0.09390   0.08651  -0.0311   0.0414   1.0000
<< Back to GOE 723 AIRFOIL (goe723-il)

Polar data table (+)

Polar graphs


<< Back to GOE 723 AIRFOIL (goe723-il)