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GOE 654 AIRFOIL (goe654-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: GOE 654 AIRFOIL (goe654-il)
Reynolds number: 50,000
Max Cl/Cd: 28.45 at α=4°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe654-il-50000-n5.txt
Download as CSV file: xf-goe654-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 654 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
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.2283   0.11589   0.10958  -0.0374   1.0000   0.1432
  -9.250  -0.2534   0.11588   0.10978  -0.0372   1.0000   0.1446
  -9.000  -0.2779   0.11576   0.10986  -0.0354   1.0000   0.1449
  -8.750  -0.2508   0.11086   0.10499  -0.0319   1.0000   0.1481
  -8.500  -0.2255   0.10655   0.10066  -0.0360   0.9892   0.1522
  -8.250  -0.2091   0.10267   0.09677  -0.0412   0.9757   0.1558
  -7.750  -0.2178   0.08618   0.08013  -0.0643   0.9372   0.0839
  -7.500  -0.1897   0.08227   0.07619  -0.0668   0.9288   0.0820
  -7.250  -0.1731   0.07747   0.07135  -0.0723   0.9148   0.0807
  -7.000  -0.1597   0.07243   0.06622  -0.0782   0.9005   0.0796
  -6.750  -0.1480   0.06735   0.06102  -0.0837   0.8862   0.0787
  -6.500  -0.1369   0.06231   0.05578  -0.0884   0.8725   0.0774
  -6.250  -0.1253   0.05684   0.04996  -0.0929   0.8604   0.0754
  -6.000  -0.1179   0.05158   0.04412  -0.0957   0.8464   0.0733
  -5.750  -0.1052   0.04794   0.03984  -0.0965   0.8329   0.0722
  -5.500  -0.0845   0.04522   0.03678  -0.0970   0.8219   0.0721
  -5.250  -0.0612   0.04265   0.03381  -0.0976   0.8118   0.0720
  -5.000  -0.0420   0.04063   0.03144  -0.0971   0.7999   0.0726
  -4.750  -0.0158   0.03888   0.02954  -0.0975   0.7909   0.0742
  -4.500   0.0064   0.03747   0.02790  -0.0971   0.7799   0.0757
  -4.250   0.0305   0.03601   0.02613  -0.0967   0.7699   0.0769
  -4.000   0.0577   0.03452   0.02428  -0.0966   0.7611   0.0776
  -3.750   0.0814   0.03338   0.02285  -0.0960   0.7511   0.0785
  -3.500   0.1108   0.03217   0.02134  -0.0959   0.7432   0.0797
  -3.250   0.1343   0.03136   0.02030  -0.0950   0.7330   0.0810
  -3.000   0.1649   0.03046   0.01909  -0.0949   0.7255   0.0836
  -2.750   0.1870   0.02986   0.01849  -0.0938   0.7148   0.0867
  -2.500   0.2183   0.02905   0.01752  -0.0938   0.7072   0.0912
  -2.250   0.2403   0.02867   0.01701  -0.0926   0.6956   0.0949
  -2.000   0.2682   0.02809   0.01626  -0.0922   0.6866   0.0993
  -1.750   0.2930   0.02761   0.01571  -0.0914   0.6769   0.1060
  -1.500   0.3173   0.02720   0.01526  -0.0906   0.6676   0.1178
  -1.250   0.3438   0.02659   0.01476  -0.0902   0.6591   0.1467
  -1.000   0.3653   0.02564   0.01455  -0.0895   0.6503   0.2816
  -0.500   0.4584   0.02368   0.01436  -0.0941   0.6326   1.0000
  -0.250   0.4834   0.02380   0.01413  -0.0933   0.6248   1.0000
   0.000   0.5023   0.02418   0.01426  -0.0919   0.6155   1.0000
   0.250   0.5283   0.02432   0.01409  -0.0913   0.6081   1.0000
   0.500   0.5474   0.02474   0.01432  -0.0899   0.5988   1.0000
   0.750   0.5739   0.02488   0.01419  -0.0894   0.5914   1.0000
   1.000   0.5929   0.02534   0.01450  -0.0880   0.5820   1.0000
   1.250   0.6197   0.02551   0.01443  -0.0876   0.5747   1.0000
   1.500   0.6385   0.02602   0.01481  -0.0862   0.5654   1.0000
   1.750   0.6652   0.02622   0.01480  -0.0857   0.5581   1.0000
   2.000   0.6839   0.02676   0.01525  -0.0843   0.5489   1.0000
   2.250   0.7099   0.02702   0.01533  -0.0838   0.5415   1.0000
   2.500   0.7289   0.02759   0.01583  -0.0825   0.5327   1.0000
   2.750   0.7533   0.02794   0.01603  -0.0818   0.5250   1.0000
   3.000   0.7735   0.02849   0.01651  -0.0807   0.5169   1.0000
   3.250   0.7951   0.02899   0.01692  -0.0797   0.5089   1.0000
   3.500   0.8188   0.02942   0.01724  -0.0790   0.5020   1.0000
   3.750   0.8347   0.03020   0.01802  -0.0775   0.4933   1.0000
   4.000   0.8642   0.03038   0.01804  -0.0774   0.4877   1.0000
   4.250   0.8729   0.03150   0.01924  -0.0751   0.4785   1.0000
   4.500   0.8987   0.03185   0.01949  -0.0747   0.4725   1.0000
   4.750   0.9111   0.03286   0.02053  -0.0729   0.4649   1.0000
   5.000   0.9294   0.03358   0.02124  -0.0717   0.4583   1.0000
   5.250   0.9583   0.03379   0.02134  -0.0716   0.4536   1.0000
   5.500   0.9583   0.03543   0.02312  -0.0686   0.4452   1.0000
   5.750   0.9828   0.03582   0.02346  -0.0681   0.4401   1.0000
   6.000   0.9913   0.03709   0.02480  -0.0660   0.4337   1.0000
   6.250   0.9996   0.03836   0.02612  -0.0640   0.4273   1.0000
   6.500   1.0284   0.03853   0.02623  -0.0639   0.4232   1.0000
   6.750   1.0151   0.04095   0.02878  -0.0599   0.4158   1.0000
   7.000   1.0264   0.04200   0.02985  -0.0582   0.4104   1.0000
   7.250   1.0616   0.04183   0.02964  -0.0587   0.4070   1.0000
   7.500   1.0107   0.04674   0.03473  -0.0525   0.3980   1.0000
   7.750   1.0326   0.04723   0.03523  -0.0518   0.3938   1.0000
   8.000   1.0334   0.04931   0.03734  -0.0501   0.3884   1.0000
   8.250   0.9963   0.05494   0.04309  -0.0480   0.3791   1.0000
   8.500   1.0317   0.05414   0.04228  -0.0474   0.3766   1.0000
   9.000   0.9579   0.06808   0.05639  -0.0468   0.3566   1.0000
   9.500   0.9663   0.07340   0.06182  -0.0463   0.3451   1.0000
  10.500   0.9281   0.09221   0.08085  -0.0484   0.3144   1.0000
  10.750   0.9547   0.09185   0.08054  -0.0474   0.3122   1.0000
  11.250   0.9365   0.10142   0.09023  -0.0490   0.2970   1.0000
  11.500   0.9095   0.10900   0.09787  -0.0512   0.2874   1.0000
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