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

GOE 500 AIRFOIL (goe500-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: GOE 500 AIRFOIL (goe500-il)
Reynolds number: 50,000
Max Cl/Cd: 36.26 at α=8.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe500-il-50000-n5.txt
Download as CSV file: xf-goe500-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 500 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.3309   0.12168   0.11514  -0.0319   1.0000   0.1126
  -8.500  -0.3435   0.12077   0.11432  -0.0308   1.0000   0.1132
  -8.250  -0.3548   0.11950   0.11315  -0.0295   1.0000   0.1134
  -8.000  -0.3655   0.11806   0.11179  -0.0279   1.0000   0.1135
  -7.750  -0.3744   0.11630   0.11012  -0.0265   0.9997   0.1135
  -7.500  -0.3649   0.10854   0.10228  -0.0316   0.9942   0.0845
  -7.250  -0.3508   0.10450   0.09821  -0.0339   0.9888   0.0838
  -7.000  -0.3370   0.10052   0.09421  -0.0375   0.9830   0.0832
  -6.750  -0.3232   0.09655   0.09022  -0.0407   0.9768   0.0816
  -6.500  -0.3104   0.09240   0.08605  -0.0448   0.9698   0.0796
  -6.250  -0.2941   0.08784   0.08146  -0.0506   0.9629   0.0787
  -6.000  -0.2808   0.08386   0.07745  -0.0555   0.9550   0.0800
  -5.750  -0.2583   0.07887   0.07239  -0.0632   0.9486   0.0811
  -5.500  -0.2412   0.07424   0.06769  -0.0690   0.9408   0.0808
  -5.250  -0.2128   0.06846   0.06176  -0.0778   0.9349   0.0805
  -5.000  -0.1773   0.06122   0.05426  -0.0899   0.9290   0.0825
  -4.750  -0.1258   0.05060   0.04286  -0.1077   0.9239   0.0848
  -4.500  -0.0885   0.04688   0.03883  -0.1133   0.9197   0.0865
  -4.250  -0.0503   0.04410   0.03573  -0.1181   0.9158   0.0895
  -4.000  -0.0159   0.04107   0.03215  -0.1225   0.9096   0.0953
  -3.750   0.0267   0.03788   0.02822  -0.1277   0.9056   0.1001
  -3.500   0.0655   0.03635   0.02642  -0.1307   0.9019   0.1049
  -3.250   0.0943   0.03501   0.02459  -0.1319   0.8950   0.1127
  -3.000   0.1276   0.03413   0.02359  -0.1336   0.8896   0.1216
  -2.750   0.1677   0.03297   0.02209  -0.1362   0.8859   0.1331
  -2.500   0.1912   0.03250   0.02141  -0.1359   0.8773   0.1461
  -2.250   0.2277   0.03192   0.02075  -0.1377   0.8723   0.1625
  -2.000   0.2534   0.03166   0.02045  -0.1377   0.8639   0.1788
  -1.750   0.2892   0.03133   0.02007  -0.1392   0.8580   0.1994
  -1.500   0.3168   0.03119   0.01982  -0.1393   0.8495   0.2191
  -1.250   0.3527   0.03085   0.01939  -0.1407   0.8433   0.2405
  -1.000   0.3798   0.03073   0.01927  -0.1406   0.8349   0.2620
  -0.750   0.4150   0.03040   0.01890  -0.1418   0.8286   0.2903
  -0.500   0.4418   0.03032   0.01889  -0.1418   0.8203   0.3247
  -0.250   0.4754   0.03007   0.01874  -0.1427   0.8140   0.3721
   0.000   0.5014   0.03006   0.01882  -0.1424   0.8056   0.4168
   0.250   0.5333   0.02987   0.01876  -0.1429   0.7992   0.4711
   0.500   0.5581   0.02987   0.01891  -0.1423   0.7907   0.5282
   0.750   0.5881   0.02948   0.01881  -0.1422   0.7842   0.6197
   1.000   0.6020   0.02879   0.01876  -0.1391   0.7748   1.0000
   1.250   0.6379   0.02882   0.01853  -0.1404   0.7687   1.0000
   1.500   0.6622   0.02923   0.01876  -0.1399   0.7592   1.0000
   1.750   0.6975   0.02920   0.01857  -0.1409   0.7534   1.0000
   2.000   0.7199   0.02965   0.01892  -0.1402   0.7435   1.0000
   2.250   0.7555   0.02959   0.01876  -0.1411   0.7379   1.0000
   2.500   0.7765   0.03010   0.01921  -0.1401   0.7276   1.0000
   2.750   0.8126   0.03001   0.01906  -0.1410   0.7223   1.0000
   3.000   0.8321   0.03060   0.01963  -0.1399   0.7116   1.0000
   3.250   0.8690   0.03047   0.01948  -0.1408   0.7065   1.0000
   3.500   0.8867   0.03116   0.02018  -0.1395   0.6955   1.0000
   3.750   0.9241   0.03100   0.02002  -0.1404   0.6906   1.0000
   4.000   0.9401   0.03181   0.02088  -0.1389   0.6793   1.0000
   4.250   0.9782   0.03159   0.02069  -0.1399   0.6747   1.0000
   4.500   0.9921   0.03256   0.02172  -0.1381   0.6631   1.0000
   4.750   1.0223   0.03274   0.02199  -0.1382   0.6564   1.0000
   5.000   1.0430   0.03338   0.02271  -0.1372   0.6469   1.0000
   5.250   1.0640   0.03403   0.02345  -0.1362   0.6379   1.0000
   5.500   1.0929   0.03427   0.02379  -0.1361   0.6306   1.0000
   5.750   1.1086   0.03518   0.02482  -0.1345   0.6203   1.0000
   6.000   1.1445   0.03498   0.02476  -0.1351   0.6139   1.0000
   6.250   1.1568   0.03594   0.02584  -0.1329   0.6022   1.0000
   6.500   1.1841   0.03606   0.02612  -0.1323   0.5929   1.0000
   6.750   1.2132   0.03605   0.02627  -0.1319   0.5835   1.0000
   7.000   1.2249   0.03703   0.02740  -0.1297   0.5719   1.0000
   7.250   1.2541   0.03710   0.02765  -0.1293   0.5632   1.0000
   7.500   1.2762   0.03753   0.02831  -0.1281   0.5531   1.0000
   7.750   1.2871   0.03859   0.02955  -0.1259   0.5416   1.0000
   8.000   1.3090   0.03871   0.02986  -0.1243   0.5288   1.0000
   8.250   1.3424   0.03702   0.02825  -0.1226   0.5058   1.0000
   8.500   1.3452   0.03728   0.02857  -0.1182   0.4795   1.0000
   8.750   1.3506   0.03742   0.02867  -0.1141   0.4489   1.0000
   9.000   1.3547   0.03818   0.02944  -0.1106   0.4209   1.0000
   9.250   1.3597   0.03919   0.03049  -0.1076   0.3948   1.0000
   9.500   1.3592   0.04084   0.03229  -0.1046   0.3689   1.0000
   9.750   1.3573   0.04268   0.03421  -0.1017   0.3365   1.0000
  10.000   1.3539   0.04472   0.03616  -0.0989   0.2912   1.0000
  10.250   1.3464   0.04730   0.03834  -0.0960   0.2392   1.0000
  10.500   1.3346   0.05079   0.04151  -0.0936   0.1967   1.0000
  10.750   1.3226   0.05470   0.04519  -0.0918   0.1624   1.0000
  11.000   1.3117   0.05877   0.04908  -0.0904   0.1393   1.0000
  11.250   1.3026   0.06282   0.05301  -0.0893   0.1234   1.0000
  11.500   1.2960   0.06672   0.05683  -0.0885   0.1117   1.0000
  11.750   1.2903   0.07059   0.06060  -0.0878   0.1031   1.0000
  12.000   1.2906   0.07382   0.06395  -0.0870   0.0945   1.0000
  12.250   1.2905   0.07711   0.06721  -0.0863   0.0886   1.0000
  12.500   1.2944   0.07998   0.07023  -0.0856   0.0820   1.0000
  12.750   1.2974   0.08288   0.07307  -0.0849   0.0770   1.0000
  13.000   1.3065   0.08520   0.07564  -0.0840   0.0719   1.0000
  13.250   1.3136   0.08769   0.07820  -0.0833   0.0673   1.0000
  13.500   1.3262   0.08962   0.08026  -0.0820   0.0632   1.0000
  13.750   1.3386   0.09180   0.08270  -0.0810   0.0593   1.0000
  14.000   1.3507   0.09387   0.08481  -0.0801   0.0557   1.0000
  14.250   1.3648   0.09621   0.08734  -0.0791   0.0528   1.0000
  14.500   1.3684   0.10001   0.09151  -0.0790   0.0509   1.0000
  14.750   1.3675   0.10427   0.09609  -0.0795   0.0492   1.0000
  15.000   1.3641   0.10882   0.10089  -0.0805   0.0477   1.0000
  15.250   1.3603   0.11339   0.10571  -0.0816   0.0463   1.0000
  15.500   1.3576   0.11787   0.11034  -0.0829   0.0451   1.0000
  15.750   1.3570   0.12229   0.11488  -0.0840   0.0441   1.0000
  16.000   1.3465   0.12845   0.12125  -0.0866   0.0438   1.0000
  16.250   1.3295   0.13587   0.12895  -0.0907   0.0437   1.0000
  16.500   1.3116   0.14394   0.13726  -0.0955   0.0437   1.0000
  16.750   1.2930   0.15273   0.14626  -0.1012   0.0438   1.0000
<< Back to GOE 500 AIRFOIL (goe500-il)

Polar data table (+)

Polar graphs


<< Back to GOE 500 AIRFOIL (goe500-il)