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

GOE 374 AIRFOIL (goe374-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: GOE 374 AIRFOIL (goe374-il)
Reynolds number: 100,000
Max Cl/Cd: 51.76 at α=5.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe374-il-100000.txt
Download as CSV file: xf-goe374-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 374 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
  ------ -------- --------- --------- -------- -------- --------
  -7.750  -0.3441   0.09554   0.09079  -0.0162   1.0000   0.0360
  -7.500  -0.3420   0.09301   0.08832  -0.0167   1.0000   0.0367
  -7.250  -0.3413   0.09060   0.08599  -0.0172   1.0000   0.0375
  -7.000  -0.3378   0.08799   0.08344  -0.0186   1.0000   0.0382
  -6.750  -0.3325   0.08545   0.08096  -0.0208   1.0000   0.0391
  -6.500  -0.3247   0.08314   0.07868  -0.0240   1.0000   0.0400
  -6.250  -0.3120   0.08138   0.07691  -0.0292   1.0000   0.0407
  -6.000  -0.2969   0.07970   0.07514  -0.0335   1.0000   0.0410
  -5.750  -0.2909   0.07518   0.07069  -0.0333   1.0000   0.0416
  -5.500  -0.2916   0.07056   0.06620  -0.0295   1.0000   0.0428
  -5.250  -0.2871   0.06749   0.06318  -0.0278   1.0000   0.0448
  -5.000  -0.2782   0.06480   0.06050  -0.0279   1.0000   0.0469
  -4.750  -0.2653   0.06224   0.05790  -0.0291   1.0000   0.0503
  -4.250  -0.2335   0.05666   0.05209  -0.0321   1.0000   0.0551
  -4.000  -0.2271   0.05374   0.04922  -0.0306   1.0000   0.0575
  -3.750  -0.2133   0.05129   0.04670  -0.0306   1.0000   0.0619
  -3.500  -0.1835   0.05022   0.04518  -0.0331   1.0000   0.0674
  -3.250  -0.1791   0.04621   0.04138  -0.0315   1.0000   0.0703
  -3.000  -0.1643   0.04415   0.03925  -0.0312   1.0000   0.0762
  -2.750  -0.1445   0.04201   0.03688  -0.0318   1.0000   0.0829
  -2.500  -0.1038   0.03898   0.03365  -0.0363   0.9936   0.0967
  -2.250  -0.0582   0.03582   0.03029  -0.0419   0.9841   0.1251
  -2.000  -0.0178   0.03300   0.02742  -0.0463   0.9732   0.1710
  -1.500   0.0485   0.02728   0.02196  -0.0509   0.9505   0.3542
  -1.250   0.0844   0.02436   0.01924  -0.0518   0.9409   0.4471
  -1.000   0.1433   0.02238   0.01681  -0.0575   0.9284   0.4345
  -0.750   0.2371   0.02231   0.01472  -0.0646   0.9172   0.1628
  -0.500   0.2866   0.02055   0.01253  -0.0665   0.9043   0.1282
  -0.250   0.3323   0.01940   0.01097  -0.0679   0.8892   0.1097
   0.000   0.3732   0.01816   0.00963  -0.0689   0.8721   0.1026
   0.250   0.4075   0.01722   0.00860  -0.0688   0.8507   0.1006
   0.500   0.4392   0.01643   0.00775  -0.0683   0.8279   0.1038
   0.750   0.4663   0.01584   0.00709  -0.0669   0.8014   0.1031
   1.000   0.4930   0.01546   0.00656  -0.0657   0.7739   0.1037
   1.250   0.5194   0.01524   0.00613  -0.0644   0.7458   0.1059
   1.500   0.5457   0.01509   0.00578  -0.0633   0.7184   0.1109
   1.750   0.5718   0.01503   0.00555  -0.0623   0.6924   0.1254
   2.000   0.6180   0.01328   0.00546  -0.0657   0.6659   1.0000
   2.250   0.6430   0.01363   0.00547  -0.0646   0.6438   1.0000
   2.500   0.6677   0.01402   0.00560  -0.0637   0.6231   1.0000
   2.750   0.6924   0.01442   0.00582  -0.0628   0.6038   1.0000
   3.000   0.7173   0.01482   0.00604  -0.0620   0.5864   1.0000
   3.250   0.7422   0.01521   0.00626  -0.0613   0.5698   1.0000
   3.500   0.7663   0.01556   0.00654  -0.0604   0.5518   1.0000
   3.750   0.7903   0.01587   0.00680  -0.0595   0.5336   1.0000
   4.000   0.8142   0.01613   0.00695  -0.0585   0.5154   1.0000
   4.250   0.8383   0.01643   0.00712  -0.0576   0.4984   1.0000
   4.500   0.8623   0.01678   0.00744  -0.0568   0.4818   1.0000
   4.750   0.8865   0.01717   0.00781  -0.0560   0.4664   1.0000
   5.000   0.9106   0.01760   0.00826  -0.0553   0.4514   1.0000
   5.250   0.9347   0.01806   0.00872  -0.0546   0.4365   1.0000
   5.500   0.9586   0.01854   0.00921  -0.0538   0.4210   1.0000
   5.750   0.9821   0.01902   0.00970  -0.0530   0.4048   1.0000
   6.000   1.0054   0.01950   0.01023  -0.0521   0.3881   1.0000
   6.250   1.0286   0.01997   0.01070  -0.0512   0.3712   1.0000
   6.500   1.0510   0.02047   0.01128  -0.0502   0.3532   1.0000
   6.750   1.0729   0.02098   0.01188  -0.0492   0.3340   1.0000
   7.000   1.0949   0.02144   0.01232  -0.0481   0.3149   1.0000
   7.250   1.1146   0.02175   0.01276  -0.0467   0.2924   1.0000
   7.500   1.1342   0.02202   0.01295  -0.0454   0.2716   1.0000
   7.750   1.1529   0.02237   0.01344  -0.0440   0.2498   1.0000
   8.000   1.1720   0.02281   0.01390  -0.0428   0.2315   1.0000
   8.250   1.1919   0.02341   0.01461  -0.0416   0.2171   1.0000
   8.500   1.2103   0.02389   0.01527  -0.0404   0.2000   1.0000
   8.750   1.2283   0.02429   0.01586  -0.0391   0.1786   1.0000
   9.000   1.2447   0.02490   0.01652  -0.0377   0.1494   1.0000
   9.250   1.2583   0.02613   0.01768  -0.0360   0.1095   1.0000
   9.500   1.2650   0.02825   0.01945  -0.0336   0.0601   1.0000
   9.750   1.2683   0.03048   0.02155  -0.0308   0.0487   1.0000
  10.000   1.2742   0.03226   0.02352  -0.0282   0.0433   1.0000
  10.250   1.2741   0.03430   0.02570  -0.0252   0.0400   1.0000
  10.500   1.2679   0.03654   0.02802  -0.0216   0.0384   1.0000
  10.750   1.2671   0.03854   0.03022  -0.0190   0.0376   1.0000
  11.000   1.2660   0.04077   0.03262  -0.0167   0.0368   1.0000
  11.250   1.2657   0.04320   0.03521  -0.0148   0.0360   1.0000
  11.500   1.2675   0.04573   0.03789  -0.0132   0.0354   1.0000
  11.750   1.2714   0.04837   0.04069  -0.0118   0.0348   1.0000
  12.000   1.2770   0.05123   0.04372  -0.0104   0.0343   1.0000
  12.250   1.2822   0.05424   0.04694  -0.0093   0.0341   1.0000
  12.500   1.2817   0.05764   0.05057  -0.0087   0.0333   1.0000
  12.750   1.2784   0.06127   0.05441  -0.0085   0.0327   1.0000
  13.000   1.2730   0.06533   0.05872  -0.0087   0.0326   1.0000
  13.250   1.2645   0.06968   0.06329  -0.0095   0.0318   1.0000
  13.500   1.2531   0.07472   0.06860  -0.0110   0.0320   1.0000
  13.750   1.2404   0.08000   0.07408  -0.0131   0.0316   1.0000
  14.000   1.2217   0.08665   0.08102  -0.0164   0.0322   1.0000
  14.250   1.1979   0.09467   0.08931  -0.0209   0.0334   1.0000
  14.500   1.1748   0.10287   0.09772  -0.0261   0.0337   1.0000
  14.750   1.1454   0.11321   0.10828  -0.0329   0.0351   1.0000
  15.000   1.1179   0.12389   0.11911  -0.0398   0.0361   1.0000
  15.250   1.0923   0.13502   0.13032  -0.0469   0.0372   1.0000
  15.500   1.0694   0.14627   0.14160  -0.0536   0.0380   1.0000
  15.750   1.0529   0.15656   0.15191  -0.0594   0.0387   1.0000
<< Back to GOE 374 AIRFOIL (goe374-il)

Polar data table (+)

Polar graphs


<< Back to GOE 374 AIRFOIL (goe374-il)