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

GOE 142 (MVA H.19) AIRFOIL (goe142-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: GOE 142 (MVA H.19) AIRFOIL (goe142-il)
Reynolds number: 50,000
Max Cl/Cd: 41.01 at α=5.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe142-il-50000-n5.txt
Download as CSV file: xf-goe142-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 142 (MVA H.19) 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
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.3911   0.10615   0.09929  -0.0112   1.0000   0.0780
  -7.750  -0.3892   0.10388   0.09711  -0.0133   1.0000   0.0802
  -7.500  -0.3884   0.10223   0.09556  -0.0174   1.0000   0.0817
  -7.250  -0.3825   0.10026   0.09365  -0.0235   1.0000   0.0823
  -7.000  -0.3737   0.09645   0.08991  -0.0259   1.0000   0.0829
  -6.750  -0.3646   0.09122   0.08472  -0.0218   1.0000   0.0848
  -6.500  -0.3555   0.08772   0.08126  -0.0219   1.0000   0.0872
  -6.250  -0.3462   0.08456   0.07812  -0.0235   1.0000   0.0900
  -6.000  -0.3357   0.08165   0.07523  -0.0270   1.0000   0.0939
  -5.750  -0.3174   0.07978   0.07325  -0.0366   1.0000   0.0967
  -5.500  -0.3100   0.07524   0.06883  -0.0346   1.0000   0.0979
  -5.250  -0.3024   0.07171   0.06538  -0.0324   1.0000   0.1012
  -5.000  -0.2786   0.06973   0.06317  -0.0400   1.0000   0.1108
  -4.750  -0.2746   0.06552   0.05917  -0.0362   1.0000   0.1145
  -4.250  -0.2421   0.05939   0.05294  -0.0395   1.0000   0.1284
  -4.000  -0.1981   0.05292   0.04589  -0.0467   1.0000   0.0696
  -3.750  -0.1815   0.04966   0.04259  -0.0465   1.0000   0.0633
  -3.500  -0.1566   0.04604   0.03874  -0.0484   1.0000   0.0596
  -3.250  -0.1272   0.04229   0.03455  -0.0506   1.0000   0.0569
  -3.000  -0.1052   0.03998   0.03209  -0.0511   1.0000   0.0596
  -2.750  -0.0795   0.03753   0.02933  -0.0521   1.0000   0.0622
  -2.500  -0.0518   0.03493   0.02635  -0.0530   1.0000   0.0625
  -2.250  -0.0235   0.03252   0.02352  -0.0538   1.0000   0.0628
  -2.000   0.0053   0.03044   0.02095  -0.0544   1.0000   0.0649
  -1.750   0.0332   0.02876   0.01884  -0.0548   1.0000   0.0702
  -1.500   0.0732   0.02704   0.01678  -0.0575   0.9942   0.0744
  -1.250   0.1187   0.02550   0.01463  -0.0607   0.9852   0.0820
  -1.000   0.1608   0.02427   0.01325  -0.0636   0.9749   0.0922
  -0.750   0.2038   0.02322   0.01188  -0.0662   0.9646   0.1032
  -0.500   0.2475   0.02239   0.01087  -0.0690   0.9538   0.1207
  -0.250   0.2911   0.02160   0.01002  -0.0719   0.9423   0.1435
   0.000   0.3328   0.02087   0.00933  -0.0745   0.9289   0.1770
   0.250   0.3729   0.01946   0.00881  -0.0773   0.9153   0.3430
   0.750   0.4406   0.01811   0.00809  -0.0778   0.8799   1.0000
   1.000   0.4746   0.01813   0.00789  -0.0784   0.8611   1.0000
   1.250   0.5078   0.01814   0.00771  -0.0787   0.8421   1.0000
   1.500   0.5370   0.01819   0.00764  -0.0784   0.8200   1.0000
   1.750   0.5671   0.01821   0.00754  -0.0780   0.7987   1.0000
   2.000   0.5949   0.01828   0.00750  -0.0773   0.7752   1.0000
   2.250   0.6222   0.01837   0.00749  -0.0764   0.7511   1.0000
   2.500   0.6492   0.01848   0.00748  -0.0754   0.7268   1.0000
   2.750   0.6755   0.01862   0.00755  -0.0744   0.7018   1.0000
   3.000   0.7010   0.01883   0.00766  -0.0733   0.6755   1.0000
   3.250   0.7262   0.01908   0.00782  -0.0722   0.6492   1.0000
   3.500   0.7512   0.01937   0.00802  -0.0712   0.6232   1.0000
   3.750   0.7762   0.01970   0.00826  -0.0702   0.5977   1.0000
   4.000   0.8013   0.02007   0.00858  -0.0692   0.5729   1.0000
   4.250   0.8260   0.02049   0.00896  -0.0683   0.5476   1.0000
   4.500   0.8506   0.02095   0.00939  -0.0675   0.5233   1.0000
   4.750   0.8752   0.02144   0.00983  -0.0667   0.5005   1.0000
   5.000   0.8996   0.02197   0.01042  -0.0659   0.4777   1.0000
   5.250   0.9240   0.02253   0.01098  -0.0652   0.4564   1.0000
   5.500   0.9482   0.02313   0.01161  -0.0644   0.4355   1.0000
   5.750   0.9724   0.02375   0.01223  -0.0637   0.4167   1.0000
   6.000   0.9963   0.02442   0.01307  -0.0630   0.3972   1.0000
   6.250   1.0201   0.02513   0.01386  -0.0623   0.3795   1.0000
   6.500   1.0437   0.02585   0.01466  -0.0615   0.3626   1.0000
   6.750   1.0672   0.02662   0.01552  -0.0608   0.3467   1.0000
   7.000   1.0901   0.02744   0.01651  -0.0600   0.3302   1.0000
   7.250   1.1126   0.02830   0.01761  -0.0592   0.3141   1.0000
   7.500   1.1345   0.02917   0.01868  -0.0583   0.2978   1.0000
   7.750   1.1555   0.03004   0.01973  -0.0573   0.2808   1.0000
   8.000   1.1750   0.03083   0.02062  -0.0561   0.2623   1.0000
   8.250   1.1910   0.03163   0.02165  -0.0546   0.2383   1.0000
   8.500   1.2047   0.03243   0.02264  -0.0530   0.2114   1.0000
   8.750   1.2167   0.03342   0.02373  -0.0513   0.1822   1.0000
   9.000   1.2269   0.03476   0.02512  -0.0496   0.1502   1.0000
   9.250   1.2346   0.03665   0.02700  -0.0478   0.1186   1.0000
   9.500   1.2390   0.03902   0.02927  -0.0457   0.0950   1.0000
   9.750   1.2417   0.04153   0.03175  -0.0437   0.0795   1.0000
  10.000   1.2432   0.04417   0.03444  -0.0415   0.0710   1.0000
  10.250   1.2417   0.04674   0.03707  -0.0392   0.0650   1.0000
  10.500   1.2414   0.04949   0.04000  -0.0373   0.0596   1.0000
  10.750   1.2395   0.05236   0.04301  -0.0358   0.0551   1.0000
  11.000   1.2358   0.05552   0.04624  -0.0346   0.0522   1.0000
  11.250   1.2374   0.05874   0.04979  -0.0334   0.0498   1.0000
  11.500   1.2370   0.06225   0.05361  -0.0325   0.0478   1.0000
  11.750   1.2336   0.06606   0.05768  -0.0322   0.0459   1.0000
  12.000   1.2283   0.07009   0.06191  -0.0324   0.0443   1.0000
  12.250   1.2223   0.07426   0.06620  -0.0330   0.0427   1.0000
  12.500   1.2176   0.07843   0.07041  -0.0333   0.0411   1.0000
  12.750   1.2050   0.08413   0.07638  -0.0354   0.0406   1.0000
  13.000   1.1907   0.09041   0.08294  -0.0382   0.0403   1.0000
  13.250   1.1753   0.09726   0.09003  -0.0416   0.0402   1.0000
  13.500   1.1587   0.10469   0.09768  -0.0455   0.0402   1.0000
  13.750   1.1413   0.11271   0.10588  -0.0500   0.0404   1.0000
  14.000   1.1238   0.12118   0.11451  -0.0548   0.0407   1.0000
  14.250   1.1065   0.13017   0.12361  -0.0600   0.0410   1.0000
<< Back to GOE 142 (MVA H.19) AIRFOIL (goe142-il)

Polar data table (+)

Polar graphs


<< Back to GOE 142 (MVA H.19) AIRFOIL (goe142-il)