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=100,000 Ncrit=9


Details Polar file
Airfoil: GOE 142 (MVA H.19) AIRFOIL (goe142-il)
Reynolds number: 100,000
Max Cl/Cd: 55.73 at α=4.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe142-il-100000.txt
Download as CSV file: xf-goe142-il-100000.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 =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.4035   0.11082   0.10589  -0.0099   1.0000   0.0535
  -8.250  -0.4034   0.10963   0.10477  -0.0136   1.0000   0.0540
  -8.000  -0.4019   0.10813   0.10335  -0.0182   1.0000   0.0543
  -7.750  -0.3934   0.10585   0.10110  -0.0245   1.0000   0.0545
  -7.500  -0.3868   0.09924   0.09455  -0.0215   1.0000   0.0553
  -7.250  -0.3768   0.09390   0.08923  -0.0166   1.0000   0.0569
  -7.000  -0.3683   0.09047   0.08583  -0.0169   1.0000   0.0585
  -6.750  -0.3600   0.08733   0.08272  -0.0184   1.0000   0.0604
  -6.500  -0.3513   0.08425   0.07968  -0.0207   1.0000   0.0625
  -6.250  -0.3412   0.08127   0.07672  -0.0240   1.0000   0.0650
  -6.000  -0.3204   0.07974   0.07511  -0.0351   1.0000   0.0679
  -5.750  -0.3074   0.07592   0.07127  -0.0390   1.0000   0.0688
  -5.500  -0.3091   0.07138   0.06689  -0.0325   1.0000   0.0710
  -5.250  -0.3011   0.06849   0.06404  -0.0314   1.0000   0.0742
  -5.000  -0.2875   0.06561   0.06114  -0.0334   1.0000   0.0784
  -4.750  -0.2544   0.06316   0.05837  -0.0430   1.0000   0.0830
  -4.500  -0.2521   0.05897   0.05435  -0.0396   1.0000   0.0847
  -4.250  -0.2426   0.05613   0.05155  -0.0383   1.0000   0.0878
  -3.750  -0.1998   0.05014   0.04529  -0.0430   1.0000   0.0991
  -3.500  -0.1861   0.04753   0.04266  -0.0423   1.0000   0.1039
  -3.250  -0.1594   0.04476   0.03965  -0.0451   1.0000   0.1132
  -3.000  -0.1363   0.04262   0.03732  -0.0464   1.0000   0.1264
  -2.750  -0.1163   0.04029   0.03490  -0.0470   1.0000   0.1406
  -2.500  -0.0950   0.03797   0.03249  -0.0476   1.0000   0.1550
  -2.250  -0.0741   0.03564   0.03012  -0.0480   1.0000   0.1705
  -2.000  -0.0504   0.03389   0.02821  -0.0489   1.0000   0.1974
  -1.250   0.1019   0.02429   0.01639  -0.0592   0.9913   0.0953
  -1.000   0.1542   0.02224   0.01406  -0.0640   0.9845   0.1026
  -0.750   0.2044   0.02065   0.01205  -0.0678   0.9752   0.1070
  -0.500   0.2542   0.01950   0.01074  -0.0718   0.9654   0.1226
  -0.250   0.3062   0.01829   0.00952  -0.0762   0.9566   0.1413
   0.000   0.3518   0.01729   0.00864  -0.0794   0.9445   0.1687
   0.250   0.3983   0.01625   0.00788  -0.0827   0.9325   0.2325
   0.500   0.4396   0.01398   0.00720  -0.0841   0.9207   1.0000
   0.750   0.4799   0.01382   0.00682  -0.0857   0.9045   1.0000
   1.000   0.5164   0.01363   0.00648  -0.0864   0.8873   1.0000
   1.250   0.5471   0.01349   0.00623  -0.0860   0.8674   1.0000
   1.500   0.5756   0.01336   0.00600  -0.0850   0.8462   1.0000
   1.750   0.6021   0.01328   0.00582  -0.0837   0.8234   1.0000
   2.000   0.6284   0.01321   0.00563  -0.0822   0.8007   1.0000
   2.250   0.6531   0.01324   0.00555  -0.0807   0.7745   1.0000
   2.500   0.6779   0.01332   0.00551  -0.0793   0.7476   1.0000
   2.750   0.7028   0.01345   0.00553  -0.0779   0.7199   1.0000
   3.000   0.7276   0.01363   0.00558  -0.0766   0.6917   1.0000
   3.250   0.7524   0.01387   0.00567  -0.0755   0.6629   1.0000
   3.500   0.7772   0.01416   0.00582  -0.0744   0.6343   1.0000
   3.750   0.8019   0.01449   0.00605  -0.0733   0.6059   1.0000
   4.000   0.8267   0.01487   0.00630  -0.0724   0.5780   1.0000
   4.250   0.8515   0.01528   0.00660  -0.0715   0.5509   1.0000
   4.500   0.8762   0.01573   0.00694  -0.0707   0.5252   1.0000
   4.750   0.9010   0.01622   0.00735  -0.0699   0.5011   1.0000
   5.000   0.9257   0.01672   0.00780  -0.0692   0.4770   1.0000
   5.250   0.9505   0.01727   0.00830  -0.0686   0.4545   1.0000
   5.500   0.9752   0.01784   0.00880  -0.0679   0.4336   1.0000
   5.750   0.9998   0.01844   0.00939  -0.0673   0.4134   1.0000
   6.000   1.0244   0.01905   0.01006  -0.0666   0.3940   1.0000
   6.250   1.0491   0.01971   0.01068  -0.0660   0.3764   1.0000
   6.500   1.0735   0.02039   0.01139  -0.0654   0.3591   1.0000
   6.750   1.0974   0.02106   0.01217  -0.0647   0.3415   1.0000
   7.000   1.1211   0.02172   0.01295  -0.0640   0.3242   1.0000
   7.250   1.1442   0.02230   0.01352  -0.0632   0.3060   1.0000
   7.500   1.1654   0.02274   0.01409  -0.0622   0.2846   1.0000
   7.750   1.1858   0.02308   0.01439  -0.0611   0.2623   1.0000
   8.000   1.2054   0.02345   0.01497  -0.0599   0.2385   1.0000
   8.250   1.2240   0.02388   0.01559  -0.0586   0.2123   1.0000
   8.500   1.2396   0.02444   0.01624  -0.0570   0.1712   1.0000
   8.750   1.2471   0.02656   0.01792  -0.0546   0.1069   1.0000
   9.000   1.2563   0.02891   0.02007  -0.0522   0.0813   1.0000
   9.250   1.2672   0.03116   0.02231  -0.0499   0.0699   1.0000
   9.500   1.2786   0.03347   0.02457  -0.0479   0.0617   1.0000
   9.750   1.2932   0.03556   0.02679  -0.0462   0.0558   1.0000
  10.000   1.3110   0.03904   0.03017  -0.0450   0.0519   1.0000
  10.250   1.3260   0.04142   0.03301  -0.0432   0.0491   1.0000
  10.500   1.3379   0.04387   0.03574  -0.0415   0.0457   1.0000
  10.750   1.3496   0.04677   0.03883  -0.0399   0.0435   1.0000
  11.000   1.3600   0.05065   0.04294  -0.0385   0.0423   1.0000
  11.250   1.3638   0.05513   0.04776  -0.0366   0.0417   1.0000
  11.500   1.3607   0.05944   0.05245  -0.0344   0.0415   1.0000
  11.750   1.3516   0.06325   0.05661  -0.0317   0.0416   1.0000
  12.000   1.3374   0.06666   0.06033  -0.0289   0.0417   1.0000
  12.250   1.3202   0.07018   0.06415  -0.0270   0.0419   1.0000
  12.500   1.3000   0.07421   0.06847  -0.0262   0.0421   1.0000
  12.750   1.2769   0.07909   0.07366  -0.0270   0.0425   1.0000
  13.000   1.2493   0.08530   0.08016  -0.0297   0.0431   1.0000
  13.250   1.2124   0.09397   0.08915  -0.0357   0.0440   1.0000
  13.500   1.1656   0.10668   0.10214  -0.0457   0.0455   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)