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

GOE 122 (MVA H.2) AIRFOIL (goe122-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: GOE 122 (MVA H.2) AIRFOIL (goe122-il)
Reynolds number: 500,000
Max Cl/Cd: 95.7 at α=5.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe122-il-500000.txt
Download as CSV file: xf-goe122-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 122 (MVA H.2) AIRFOIL                       
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.4306   0.09743   0.09528   0.0004   1.0000   0.0168
  -8.000  -0.4245   0.09396   0.09183  -0.0019   1.0000   0.0174
  -7.750  -0.4195   0.09039   0.08829  -0.0061   1.0000   0.0179
  -7.500  -0.4104   0.08615   0.08407  -0.0130   1.0000   0.0181
  -7.250  -0.3954   0.08134   0.07927  -0.0199   1.0000   0.0182
  -7.000  -0.3785   0.07648   0.07439  -0.0260   1.0000   0.0182
  -6.750  -0.3661   0.07077   0.06867  -0.0305   1.0000   0.0184
  -6.500  -0.3567   0.06758   0.06549  -0.0304   1.0000   0.0187
  -6.250  -0.3427   0.06465   0.06256  -0.0317   1.0000   0.0190
  -6.000  -0.3275   0.06149   0.05940  -0.0339   1.0000   0.0195
  -5.750  -0.2997   0.05756   0.05541  -0.0392   0.9933   0.0205
  -5.500  -0.2486   0.05107   0.04869  -0.0502   0.9810   0.0230
  -5.250  -0.2085   0.04518   0.04252  -0.0565   0.9685   0.0233
  -4.250  -0.1164   0.03165   0.02838  -0.0618   0.9132   0.0268
  -4.000  -0.0829   0.02981   0.02600  -0.0613   0.9012   0.0299
  -3.750  -0.0615   0.02430   0.02014  -0.0621   0.8899   0.0312
  -3.500  -0.0378   0.02282   0.01858  -0.0620   0.8778   0.0323
  -3.250  -0.0125   0.02147   0.01705  -0.0617   0.8655   0.0341
  -3.000   0.0173   0.02225   0.01748  -0.0606   0.8524   0.0387
  -2.750   0.0411   0.01816   0.01315  -0.0611   0.8397   0.0417
  -2.500   0.0676   0.01725   0.01209  -0.0609   0.8251   0.0449
  -1.250   0.2064   0.01208   0.00570  -0.0587   0.7343   0.0473
  -1.000   0.2341   0.01112   0.00454  -0.0583   0.7124   0.0445
  -0.750   0.2615   0.01046   0.00369  -0.0578   0.6908   0.0424
  -0.500   0.2890   0.01004   0.00316  -0.0575   0.6701   0.0418
  -0.250   0.3164   0.00975   0.00276  -0.0572   0.6503   0.0419
   0.000   0.3439   0.00953   0.00245  -0.0570   0.6312   0.0422
   0.250   0.3716   0.00937   0.00221  -0.0569   0.6120   0.0429
   0.500   0.3993   0.00926   0.00203  -0.0568   0.5916   0.0445
   0.750   0.4271   0.00919   0.00187  -0.0567   0.5704   0.0453
   1.000   0.4549   0.00916   0.00174  -0.0566   0.5472   0.0465
   1.250   0.4826   0.00918   0.00165  -0.0565   0.5230   0.0483
   1.500   0.5102   0.00925   0.00160  -0.0564   0.4989   0.0515
   1.750   0.5379   0.00927   0.00162  -0.0563   0.4787   0.0782
   2.000   0.5629   0.00726   0.00178  -0.0562   0.4634   1.0000
   2.250   0.5904   0.00744   0.00184  -0.0561   0.4498   1.0000
   2.500   0.6178   0.00763   0.00192  -0.0560   0.4354   1.0000
   2.750   0.6451   0.00782   0.00201  -0.0559   0.4213   1.0000
   3.000   0.6725   0.00800   0.00211  -0.0559   0.4097   1.0000
   3.250   0.7000   0.00815   0.00221  -0.0558   0.3996   1.0000
   3.500   0.7275   0.00830   0.00234  -0.0557   0.3901   1.0000
   3.750   0.7548   0.00848   0.00247  -0.0557   0.3802   1.0000
   4.000   0.7821   0.00866   0.00259  -0.0556   0.3669   1.0000
   4.250   0.8096   0.00880   0.00273  -0.0555   0.3550   1.0000
   4.500   0.8369   0.00896   0.00287  -0.0555   0.3434   1.0000
   4.750   0.8641   0.00913   0.00302  -0.0554   0.3301   1.0000
   5.000   0.8911   0.00934   0.00319  -0.0553   0.3138   1.0000
   5.250   0.9178   0.00959   0.00338  -0.0552   0.2923   1.0000
   5.500   0.9434   0.01002   0.00362  -0.0551   0.2467   1.0000
   5.750   0.9666   0.01089   0.00410  -0.0547   0.1796   1.0000
   6.000   0.9911   0.01152   0.00458  -0.0544   0.1487   1.0000
   6.250   1.0110   0.01297   0.00547  -0.0538   0.0527   1.0000
   6.500   1.0357   0.01355   0.00605  -0.0533   0.0438   1.0000
   6.750   1.0609   0.01401   0.00658  -0.0530   0.0411   1.0000
   7.000   1.0852   0.01460   0.00723  -0.0526   0.0381   1.0000
   7.250   1.1078   0.01544   0.00817  -0.0520   0.0348   1.0000
   7.500   1.1307   0.01617   0.00899  -0.0514   0.0327   1.0000
   7.750   1.1541   0.01678   0.00969  -0.0508   0.0308   1.0000
   8.000   1.1761   0.01755   0.01052  -0.0502   0.0287   1.0000
   8.250   1.1958   0.01859   0.01162  -0.0493   0.0264   1.0000
   8.500   1.2107   0.02020   0.01335  -0.0478   0.0244   1.0000
   8.750   1.2330   0.02077   0.01400  -0.0471   0.0231   1.0000
   9.000   1.2531   0.02159   0.01490  -0.0462   0.0216   1.0000
   9.250   1.2723   0.02245   0.01581  -0.0453   0.0203   1.0000
   9.500   1.2887   0.02358   0.01697  -0.0441   0.0191   1.0000
   9.750   1.2981   0.02562   0.01910  -0.0421   0.0180   1.0000
  10.000   1.3171   0.02629   0.01989  -0.0411   0.0171   1.0000
  10.250   1.3325   0.02739   0.02108  -0.0397   0.0163   1.0000
  10.500   1.3466   0.02851   0.02230  -0.0383   0.0155   1.0000
  10.750   1.3584   0.02961   0.02349  -0.0366   0.0149   1.0000
  11.000   1.3671   0.03081   0.02475  -0.0347   0.0144   1.0000
  11.250   1.3729   0.03248   0.02648  -0.0326   0.0140   1.0000
  11.500   1.3741   0.03563   0.02976  -0.0303   0.0134   1.0000
  11.750   1.3805   0.03705   0.03134  -0.0288   0.0131   1.0000
  12.000   1.3858   0.03880   0.03326  -0.0275   0.0126   1.0000
  12.250   1.3894   0.04094   0.03556  -0.0263   0.0122   1.0000
  12.500   1.3914   0.04338   0.03815  -0.0252   0.0118   1.0000
  12.750   1.3919   0.04602   0.04096  -0.0245   0.0116   1.0000
  13.000   1.3914   0.04883   0.04392  -0.0240   0.0113   1.0000
  13.250   1.3893   0.05191   0.04716  -0.0239   0.0111   1.0000
  13.500   1.3855   0.05532   0.05072  -0.0242   0.0109   1.0000
  13.750   1.3807   0.05896   0.05450  -0.0249   0.0108   1.0000
  14.000   1.3740   0.06302   0.05872  -0.0258   0.0106   1.0000
  14.250   1.3661   0.06736   0.06319  -0.0272   0.0105   1.0000
  14.500   1.3561   0.07219   0.06818  -0.0288   0.0104   1.0000
  14.750   1.3441   0.07743   0.07358  -0.0307   0.0104   1.0000
  15.000   1.3301   0.08323   0.07954  -0.0331   0.0103   1.0000
  15.250   1.3142   0.08966   0.08615  -0.0360   0.0103   1.0000
  15.500   1.2964   0.09678   0.09344  -0.0395   0.0103   1.0000
  15.750   1.2776   0.10443   0.10127  -0.0434   0.0103   1.0000
  16.000   1.2542   0.11339   0.11044  -0.0483   0.0104   1.0000
  16.250   1.2106   0.12777   0.12515  -0.0564   0.0107   1.0000
  16.500   1.1242   0.15682   0.15461  -0.0745   0.0119   1.0000
  16.750   1.0800   0.17793   0.17581  -0.0867   0.0127   1.0000
  17.000   1.0636   0.18954   0.18739  -0.0926   0.0133   1.0000
<< Back to GOE 122 (MVA H.2) AIRFOIL (goe122-il)

Polar data table (+)

Polar graphs


<< Back to GOE 122 (MVA H.2) AIRFOIL (goe122-il)