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

GOE 134 (MVA H.12) AIRFOIL (goe134-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: GOE 134 (MVA H.12) AIRFOIL (goe134-il)
Reynolds number: 200,000
Max Cl/Cd: 70.4 at α=8°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe134-il-200000.txt
Download as CSV file: xf-goe134-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 134 (MVA H.12) AIRFOIL                      
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.3803   0.09417   0.09070  -0.0263   1.0000   0.0325
  -7.750  -0.3878   0.09125   0.08785  -0.0307   1.0000   0.0329
  -7.500  -0.3876   0.08752   0.08414  -0.0365   1.0000   0.0331
  -7.250  -0.3852   0.08368   0.08024  -0.0403   1.0000   0.0332
  -7.000  -0.3847   0.07924   0.07590  -0.0367   1.0000   0.0337
  -6.750  -0.3817   0.07643   0.07314  -0.0342   1.0000   0.0342
  -6.500  -0.3781   0.07364   0.07037  -0.0333   1.0000   0.0348
  -6.250  -0.3737   0.07072   0.06745  -0.0333   1.0000   0.0355
  -6.000  -0.3679   0.06763   0.06434  -0.0338   1.0000   0.0365
  -5.750  -0.3599   0.06435   0.06103  -0.0347   1.0000   0.0377
  -5.500  -0.3470   0.06075   0.05731  -0.0367   1.0000   0.0397
  -5.250  -0.3268   0.05596   0.05207  -0.0407   1.0000   0.0418
  -5.000  -0.3224   0.05284   0.04910  -0.0391   1.0000   0.0427
  -4.750  -0.3013   0.05016   0.04642  -0.0405   0.9979   0.0447
  -4.500  -0.2428   0.04817   0.04359  -0.0478   0.9930   0.0517
  -4.250  -0.2162   0.04133   0.03696  -0.0518   0.9886   0.0540
  -4.000  -0.1770   0.03850   0.03403  -0.0561   0.9845   0.0588
  -3.750  -0.1357   0.03507   0.03023  -0.0602   0.9785   0.0673
  -3.500  -0.0915   0.03286   0.02757  -0.0643   0.9734   0.0795
  -3.250  -0.0526   0.03007   0.02485  -0.0680   0.9691   0.0864
  -3.000  -0.0132   0.02779   0.02235  -0.0712   0.9617   0.1000
  -2.750   0.0310   0.02587   0.02015  -0.0754   0.9577   0.1258
  -1.250   0.2491   0.01611   0.00863  -0.0792   0.8802   0.0739
  -1.000   0.2775   0.01510   0.00742  -0.0784   0.8594   0.0680
  -0.750   0.3054   0.01464   0.00677  -0.0775   0.8388   0.0652
  -0.500   0.3325   0.01425   0.00627  -0.0767   0.8167   0.0643
  -0.250   0.3590   0.01358   0.00552  -0.0760   0.7958   0.0651
   0.000   0.3855   0.01310   0.00498  -0.0754   0.7759   0.0665
   0.250   0.4124   0.01283   0.00463  -0.0749   0.7563   0.0666
   0.500   0.4395   0.01265   0.00435  -0.0745   0.7380   0.0673
   0.750   0.4667   0.01257   0.00414  -0.0742   0.7205   0.0689
   1.000   0.4942   0.01255   0.00399  -0.0739   0.7037   0.0719
   1.250   0.5217   0.01258   0.00389  -0.0736   0.6872   0.0767
   1.500   0.5483   0.01180   0.00390  -0.0738   0.6709   0.3839
   1.750   0.5710   0.01034   0.00384  -0.0719   0.6549   1.0000
   2.000   0.5982   0.01050   0.00384  -0.0716   0.6378   1.0000
   2.250   0.6254   0.01067   0.00386  -0.0714   0.6209   1.0000
   2.500   0.6525   0.01084   0.00390  -0.0712   0.6039   1.0000
   2.750   0.6795   0.01102   0.00395  -0.0710   0.5867   1.0000
   3.000   0.7063   0.01122   0.00403  -0.0707   0.5694   1.0000
   3.250   0.7331   0.01145   0.00413  -0.0705   0.5525   1.0000
   3.500   0.7599   0.01169   0.00427  -0.0703   0.5359   1.0000
   3.750   0.7867   0.01194   0.00445  -0.0701   0.5190   1.0000
   4.000   0.8134   0.01220   0.00463  -0.0700   0.5023   1.0000
   4.250   0.8400   0.01248   0.00484  -0.0698   0.4862   1.0000
   4.500   0.8665   0.01277   0.00506  -0.0696   0.4711   1.0000
   4.750   0.8931   0.01308   0.00533  -0.0695   0.4571   1.0000
   5.000   0.9195   0.01341   0.00560  -0.0693   0.4440   1.0000
   5.250   0.9458   0.01375   0.00587  -0.0691   0.4308   1.0000
   5.500   0.9720   0.01406   0.00617  -0.0689   0.4174   1.0000
   5.750   0.9979   0.01436   0.00647  -0.0687   0.4037   1.0000
   6.000   1.0238   0.01467   0.00679  -0.0685   0.3907   1.0000
   6.250   1.0497   0.01501   0.00713  -0.0683   0.3790   1.0000
   6.500   1.0751   0.01535   0.00748  -0.0680   0.3668   1.0000
   6.750   1.1003   0.01570   0.00782  -0.0677   0.3548   1.0000
   7.000   1.1255   0.01605   0.00818  -0.0674   0.3436   1.0000
   7.250   1.1507   0.01637   0.00862  -0.0671   0.3327   1.0000
   7.500   1.1755   0.01673   0.00902  -0.0667   0.3215   1.0000
   7.750   1.1995   0.01706   0.00939  -0.0662   0.3085   1.0000
   8.000   1.2229   0.01737   0.00976  -0.0657   0.2937   1.0000
   8.250   1.2461   0.01770   0.01016  -0.0651   0.2783   1.0000
   8.500   1.2691   0.01805   0.01059  -0.0645   0.2628   1.0000
   8.750   1.2912   0.01844   0.01102  -0.0638   0.2443   1.0000
   9.000   1.3125   0.01890   0.01150  -0.0629   0.2224   1.0000
   9.250   1.3329   0.01947   0.01209  -0.0620   0.1990   1.0000
   9.500   1.3515   0.02021   0.01277  -0.0609   0.1683   1.0000
   9.750   1.3661   0.02132   0.01368  -0.0593   0.1329   1.0000
  10.000   1.3775   0.02270   0.01488  -0.0573   0.1042   1.0000
  10.250   1.3857   0.02433   0.01633  -0.0548   0.0690   1.0000
  10.500   1.3871   0.02633   0.01817  -0.0514   0.0464   1.0000
  10.750   1.3911   0.02786   0.01974  -0.0481   0.0387   1.0000
  11.000   1.3920   0.02938   0.02136  -0.0445   0.0351   1.0000
  11.250   1.3947   0.03093   0.02301  -0.0416   0.0323   1.0000
  11.500   1.3923   0.03295   0.02511  -0.0386   0.0303   1.0000
  11.750   1.3884   0.03529   0.02756  -0.0359   0.0292   1.0000
  12.000   1.3880   0.03754   0.02997  -0.0340   0.0283   1.0000
  12.250   1.3865   0.04008   0.03266  -0.0325   0.0275   1.0000
  12.500   1.3846   0.04286   0.03558  -0.0314   0.0267   1.0000
  12.750   1.3823   0.04584   0.03869  -0.0306   0.0261   1.0000
  13.000   1.3800   0.04896   0.04192  -0.0301   0.0255   1.0000
  13.250   1.3775   0.05217   0.04523  -0.0297   0.0250   1.0000
  13.500   1.3746   0.05547   0.04861  -0.0293   0.0244   1.0000
  13.750   1.3714   0.05876   0.05196  -0.0285   0.0237   1.0000
  14.000   1.3702   0.06188   0.05516  -0.0273   0.0231   1.0000
  14.250   1.3700   0.06508   0.05854  -0.0273   0.0227   1.0000
  14.500   1.3690   0.06839   0.06202  -0.0273   0.0223   1.0000
  14.750   1.3678   0.07180   0.06561  -0.0271   0.0220   1.0000
  15.000   1.3653   0.07546   0.06945  -0.0272   0.0218   1.0000
  15.250   1.3613   0.07943   0.07361  -0.0275   0.0216   1.0000
  15.500   1.3555   0.08376   0.07814  -0.0282   0.0215   1.0000
  15.750   1.3479   0.08849   0.08308  -0.0293   0.0214   1.0000
  16.000   1.3381   0.09371   0.08850  -0.0309   0.0214   1.0000
  16.250   1.3263   0.09942   0.09443  -0.0330   0.0214   1.0000
  16.500   1.3127   0.10565   0.10089  -0.0357   0.0214   1.0000
  16.750   1.2976   0.11245   0.10790  -0.0390   0.0215   1.0000
  17.000   1.2810   0.11981   0.11547  -0.0430   0.0216   1.0000
  17.250   1.2632   0.12775   0.12361  -0.0476   0.0218   1.0000
  17.500   1.2442   0.13636   0.13242  -0.0529   0.0219   1.0000
  17.750   1.2240   0.14574   0.14199  -0.0589   0.0221   1.0000
  18.000   1.2027   0.15593   0.15235  -0.0657   0.0224   1.0000
  18.250   1.1794   0.16733   0.16389  -0.0733   0.0227   1.0000
  18.500   1.1553   0.17965   0.17630  -0.0812   0.0232   1.0000
<< Back to GOE 134 (MVA H.12) AIRFOIL (goe134-il)

Polar data table (+)

Polar graphs


<< Back to GOE 134 (MVA H.12) AIRFOIL (goe134-il)