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

GOE 143 (MVA H.20) AIRFOIL (goe143-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: GOE 143 (MVA H.20) AIRFOIL (goe143-il)
Reynolds number: 50,000
Max Cl/Cd: 35.33 at α=5.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe143-il-50000.txt
Download as CSV file: xf-goe143-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 143 (MVA H.20) 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.4180   0.11330   0.10633  -0.0040   1.0000   0.1411
  -8.500  -0.4208   0.11172   0.10484  -0.0066   1.0000   0.1462
  -8.250  -0.4348   0.11204   0.10532  -0.0117   1.0000   0.1478
  -8.000  -0.4054   0.10425   0.09744  -0.0072   1.0000   0.1583
  -7.750  -0.4169   0.10402   0.09735  -0.0122   1.0000   0.1625
  -7.500  -0.3972   0.09817   0.09149  -0.0086   1.0000   0.1722
  -7.250  -0.4064   0.09800   0.09144  -0.0172   1.0000   0.1781
  -7.000  -0.3874   0.09210   0.08556  -0.0110   1.0000   0.1877
  -6.500  -0.3786   0.08685   0.08043  -0.0172   1.0000   0.2077
  -6.250  -0.3673   0.08277   0.07641  -0.0147   1.0000   0.2183
  -6.000  -0.3600   0.07973   0.07340  -0.0151   1.0000   0.2328
  -5.750  -0.3534   0.07693   0.07067  -0.0158   1.0000   0.2504
  -5.250  -0.3406   0.07119   0.06507  -0.0157   1.0000   0.2878
  -5.000  -0.3333   0.06841   0.06235  -0.0133   1.0000   0.3097
  -4.750  -0.3281   0.06573   0.05975  -0.0115   1.0000   0.3356
  -4.250  -0.3230   0.06101   0.05522  -0.0038   1.0000   0.4098
  -4.000   0.0140   0.03836   0.03154  -0.0159   1.0000   1.0000
  -3.750   0.0270   0.03651   0.02974  -0.0172   1.0000   1.0000
  -3.500   0.0399   0.03474   0.02803  -0.0184   1.0000   1.0000
  -3.250   0.0058   0.03549   0.02898  -0.0088   1.0000   0.9769
  -3.000  -0.0489   0.03699   0.03074   0.0041   1.0000   0.9358
  -2.750  -0.1001   0.03787   0.03190   0.0149   1.0000   0.8983
  -2.500  -0.1499   0.03832   0.03267   0.0249   1.0000   0.8692
  -2.250  -0.2038   0.03865   0.03327   0.0359   1.0000   0.8526
  -2.000  -0.2567   0.03863   0.03349   0.0461   1.0000   0.8379
  -1.750  -0.0036   0.03205   0.02300  -0.0530   1.0000   0.2083
  -1.500   0.0228   0.03052   0.02129  -0.0536   1.0000   0.2052
  -1.250   0.0515   0.02926   0.01970  -0.0545   1.0000   0.1981
  -1.000   0.0801   0.02850   0.01852  -0.0551   1.0000   0.1921
  -0.750   0.1056   0.02776   0.01760  -0.0556   1.0000   0.1902
  -0.500   0.1298   0.02737   0.01701  -0.0560   1.0000   0.1911
  -0.250   0.1521   0.02715   0.01669  -0.0565   1.0000   0.1960
   0.000   0.1892   0.02696   0.01640  -0.0594   0.9945   0.2018
   0.250   0.2575   0.02647   0.01570  -0.0670   0.9744   0.2111
   0.500   0.3240   0.02592   0.01524  -0.0743   0.9537   0.2402
   0.750   0.3841   0.02507   0.01482  -0.0805   0.9303   0.2962
   1.000   0.4388   0.02313   0.01420  -0.0838   0.9089   1.0000
   1.250   0.4915   0.02329   0.01397  -0.0876   0.8839   1.0000
   1.500   0.5356   0.02341   0.01391  -0.0897   0.8590   1.0000
   1.750   0.5790   0.02338   0.01375  -0.0912   0.8371   1.0000
   2.000   0.6107   0.02364   0.01392  -0.0910   0.8127   1.0000
   2.250   0.6442   0.02375   0.01396  -0.0906   0.7918   1.0000
   2.500   0.6707   0.02416   0.01431  -0.0896   0.7692   1.0000
   2.750   0.6984   0.02447   0.01458  -0.0886   0.7487   1.0000
   3.000   0.7263   0.02471   0.01476  -0.0873   0.7302   1.0000
   3.250   0.7502   0.02529   0.01536  -0.0862   0.7094   1.0000
   3.500   0.7755   0.02572   0.01577  -0.0849   0.6902   1.0000
   3.750   0.8016   0.02594   0.01596  -0.0833   0.6716   1.0000
   4.000   0.8252   0.02617   0.01620  -0.0813   0.6486   1.0000
   4.250   0.8506   0.02593   0.01583  -0.0787   0.6253   1.0000
   4.500   0.8733   0.02610   0.01598  -0.0765   0.5993   1.0000
   4.750   0.8974   0.02629   0.01609  -0.0745   0.5757   1.0000
   5.000   0.9215   0.02663   0.01643  -0.0728   0.5524   1.0000
   5.250   0.9448   0.02713   0.01694  -0.0712   0.5281   1.0000
   5.500   0.9696   0.02747   0.01716  -0.0694   0.5037   1.0000
   5.750   0.9922   0.02808   0.01777  -0.0678   0.4757   1.0000
   6.000   1.0147   0.02885   0.01855  -0.0661   0.4464   1.0000
   6.250   1.0371   0.02981   0.01946  -0.0646   0.4166   1.0000
   6.500   1.0589   0.03098   0.02063  -0.0632   0.3886   1.0000
   6.750   1.0820   0.03216   0.02175  -0.0620   0.3661   1.0000
   7.000   1.1033   0.03364   0.02341  -0.0610   0.3470   1.0000
   7.250   1.1256   0.03523   0.02513  -0.0602   0.3330   1.0000
   7.500   1.1500   0.03674   0.02669  -0.0595   0.3232   1.0000
   7.750   1.1677   0.03913   0.02951  -0.0589   0.3141   1.0000
   8.000   1.1888   0.04102   0.03164  -0.0582   0.3066   1.0000
   8.250   1.2126   0.04155   0.03217  -0.0571   0.2940   1.0000
   8.500   1.2307   0.04193   0.03272  -0.0556   0.2775   1.0000
   8.750   1.2509   0.04223   0.03313  -0.0543   0.2627   1.0000
   9.000   1.2692   0.04241   0.03350  -0.0528   0.2464   1.0000
   9.250   1.2867   0.04248   0.03371  -0.0511   0.2288   1.0000
   9.500   1.3024   0.04178   0.03292  -0.0489   0.2043   1.0000
   9.750   1.3104   0.04230   0.03341  -0.0462   0.1759   1.0000
  10.000   1.3127   0.04463   0.03565  -0.0432   0.1455   1.0000
  10.250   1.3224   0.04720   0.03782  -0.0412   0.1241   1.0000
  10.500   1.3261   0.05027   0.04119  -0.0389   0.1123   1.0000
  10.750   1.3290   0.05413   0.04546  -0.0369   0.1053   1.0000
  11.000   1.3355   0.05744   0.04895  -0.0352   0.0995   1.0000
  11.250   1.3378   0.06133   0.05302  -0.0335   0.0958   1.0000
  11.500   1.3226   0.06565   0.05780  -0.0312   0.0944   1.0000
  11.750   1.3024   0.07007   0.06256  -0.0293   0.0938   1.0000
  12.000   1.2788   0.07519   0.06797  -0.0287   0.0938   1.0000
  12.250   1.2524   0.08123   0.07425  -0.0298   0.0944   1.0000
  12.500   1.2246   0.08827   0.08148  -0.0324   0.0953   1.0000
  12.750   1.1974   0.09619   0.08953  -0.0361   0.0963   1.0000
<< Back to GOE 143 (MVA H.20) AIRFOIL (goe143-il)

Polar data table (+)

Polar graphs


<< Back to GOE 143 (MVA H.20) AIRFOIL (goe143-il)