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=1,000,000 Ncrit=5


Details Polar file
Airfoil: GOE 134 (MVA H.12) AIRFOIL (goe134-il)
Reynolds number: 1,000,000
Max Cl/Cd: 109.66 at α=5.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe134-il-1000000-n5.txt
Download as CSV file: xf-goe134-il-1000000-n5.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 =     1.000 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.4100   0.10261   0.10098  -0.0193   1.0000   0.0056
  -9.250  -0.4122   0.09743   0.09581  -0.0216   1.0000   0.0059
  -9.000  -0.4197   0.09122   0.08962  -0.0245   1.0000   0.0063
  -8.750  -0.4168   0.08799   0.08641  -0.0262   1.0000   0.0064
  -8.500  -0.4159   0.08493   0.08337  -0.0274   1.0000   0.0064
  -8.250  -0.4115   0.08139   0.07985  -0.0305   0.9959   0.0065
  -8.000  -0.3982   0.07677   0.07522  -0.0370   0.9887   0.0067
  -7.750  -0.3823   0.07178   0.07021  -0.0441   0.9809   0.0070
  -7.500  -0.3665   0.06610   0.06449  -0.0515   0.9718   0.0074
  -7.250  -0.3582   0.05321   0.05140  -0.0642   0.9601   0.0085
  -7.000  -0.3385   0.04982   0.04792  -0.0668   0.9509   0.0087
  -6.750  -0.3184   0.04663   0.04462  -0.0688   0.9412   0.0089
  -6.500  -0.2980   0.04302   0.04086  -0.0706   0.9308   0.0093
  -6.250  -0.2816   0.03217   0.02946  -0.0740   0.9194   0.0111
  -6.000  -0.2575   0.03094   0.02810  -0.0741   0.9060   0.0113
  -5.500  -0.2122   0.02090   0.01709  -0.0744   0.8712   0.0142
  -5.250  -0.1867   0.02074   0.01679  -0.0741   0.8370   0.0144
  -5.000  -0.1615   0.02043   0.01622  -0.0737   0.7891   0.0146
  -4.750  -0.1354   0.01978   0.01531  -0.0737   0.7534   0.0151
  -4.500  -0.1083   0.01606   0.01091  -0.0734   0.7319   0.0175
  -4.250  -0.0810   0.01531   0.00996  -0.0736   0.7129   0.0179
  -4.000  -0.0532   0.01496   0.00949  -0.0738   0.6976   0.0181
  -3.750  -0.0252   0.01459   0.00900  -0.0739   0.6849   0.0184
  -3.500   0.0030   0.01424   0.00855  -0.0741   0.6736   0.0188
  -3.250   0.0313   0.01386   0.00806  -0.0743   0.6624   0.0193
  -3.000   0.0597   0.01323   0.00727  -0.0744   0.6505   0.0198
  -2.750   0.0882   0.01257   0.00646  -0.0745   0.6384   0.0203
  -2.500   0.1168   0.01203   0.00577  -0.0747   0.6258   0.0208
  -2.250   0.1453   0.01173   0.00535  -0.0748   0.6121   0.0214
  -2.000   0.1739   0.01150   0.00500  -0.0749   0.5973   0.0217
  -1.750   0.2023   0.01116   0.00455  -0.0750   0.5819   0.0219
  -1.500   0.2303   0.01059   0.00390  -0.0752   0.5656   0.0224
  -1.250   0.2586   0.01033   0.00358  -0.0753   0.5500   0.0227
  -1.000   0.2869   0.01013   0.00332  -0.0755   0.5347   0.0229
  -0.750   0.3153   0.00996   0.00309  -0.0757   0.5203   0.0231
  -0.500   0.3437   0.00981   0.00289  -0.0758   0.5063   0.0233
  -0.250   0.3722   0.00968   0.00270  -0.0760   0.4941   0.0236
   0.000   0.4008   0.00960   0.00258  -0.0762   0.4821   0.0241
   0.250   0.4293   0.00950   0.00243  -0.0764   0.4692   0.0244
   0.500   0.4579   0.00942   0.00230  -0.0766   0.4556   0.0246
   0.750   0.4864   0.00937   0.00220  -0.0768   0.4423   0.0250
   1.000   0.5149   0.00935   0.00212  -0.0770   0.4301   0.0254
   1.250   0.5434   0.00934   0.00207  -0.0772   0.4183   0.0260
   1.500   0.5718   0.00936   0.00204  -0.0774   0.4063   0.0265
   1.750   0.6003   0.00939   0.00204  -0.0776   0.3944   0.0269
   2.000   0.6286   0.00943   0.00204  -0.0778   0.3824   0.0271
   2.250   0.6569   0.00947   0.00204  -0.0780   0.3704   0.0277
   2.500   0.6852   0.00953   0.00205  -0.0781   0.3592   0.0284
   2.750   0.7134   0.00959   0.00208  -0.0783   0.3492   0.0297
   3.000   0.7417   0.00965   0.00212  -0.0785   0.3406   0.0306
   3.250   0.7697   0.00975   0.00219  -0.0786   0.3315   0.0316
   3.500   0.7978   0.00985   0.00226  -0.0788   0.3209   0.0332
   3.750   0.8257   0.00995   0.00235  -0.0789   0.3112   0.0381
   4.250   0.8748   0.00830   0.00272  -0.0783   0.2933   1.0000
   4.500   0.9026   0.00848   0.00285  -0.0784   0.2842   1.0000
   4.750   0.9300   0.00869   0.00300  -0.0785   0.2714   1.0000
   5.000   0.9576   0.00888   0.00314  -0.0786   0.2604   1.0000
   5.250   0.9851   0.00907   0.00329  -0.0787   0.2503   1.0000
   5.500   1.0124   0.00928   0.00346  -0.0788   0.2408   1.0000
   5.750   1.0396   0.00948   0.00363  -0.0789   0.2310   1.0000
   6.000   1.0664   0.00975   0.00383  -0.0789   0.2162   1.0000
   6.250   1.0928   0.01004   0.00406  -0.0788   0.2010   1.0000
   6.500   1.1189   0.01037   0.00431  -0.0788   0.1833   1.0000
   6.750   1.1436   0.01086   0.00464  -0.0785   0.1531   1.0000
   7.000   1.1665   0.01159   0.00515  -0.0780   0.1151   1.0000
   7.250   1.1909   0.01208   0.00555  -0.0777   0.0957   1.0000
   7.500   1.2145   0.01265   0.00599  -0.0773   0.0745   1.0000
   7.750   1.2377   0.01325   0.00647  -0.0768   0.0554   1.0000
   8.000   1.2614   0.01376   0.00692  -0.0764   0.0427   1.0000
   8.250   1.2815   0.01467   0.00766  -0.0754   0.0166   1.0000
   8.500   1.3051   0.01513   0.00814  -0.0749   0.0131   1.0000
   8.750   1.3287   0.01557   0.00860  -0.0744   0.0111   1.0000
   9.000   1.3522   0.01598   0.00906  -0.0740   0.0100   1.0000
   9.250   1.3748   0.01646   0.00957  -0.0733   0.0089   1.0000
   9.500   1.3967   0.01699   0.01014  -0.0726   0.0080   1.0000
   9.750   1.4190   0.01744   0.01063  -0.0720   0.0076   1.0000
  10.000   1.4405   0.01794   0.01118  -0.0712   0.0071   1.0000
  10.250   1.4613   0.01847   0.01175  -0.0703   0.0065   1.0000
  10.500   1.4806   0.01909   0.01241  -0.0692   0.0060   1.0000
  10.750   1.4991   0.01974   0.01311  -0.0680   0.0056   1.0000
  11.000   1.5176   0.02032   0.01375  -0.0668   0.0054   1.0000
  11.250   1.5347   0.02095   0.01445  -0.0654   0.0051   1.0000
  11.500   1.5504   0.02162   0.01518  -0.0638   0.0049   1.0000
  11.750   1.5644   0.02233   0.01595  -0.0619   0.0046   1.0000
  12.000   1.5746   0.02309   0.01677  -0.0593   0.0045   1.0000
  12.250   1.5831   0.02398   0.01772  -0.0567   0.0043   1.0000
  12.500   1.5905   0.02502   0.01884  -0.0540   0.0041   1.0000
  12.750   1.5953   0.02631   0.02023  -0.0514   0.0039   1.0000
  13.000   1.6007   0.02765   0.02166  -0.0490   0.0038   1.0000
  13.250   1.6073   0.02899   0.02309  -0.0471   0.0038   1.0000
  13.500   1.6129   0.03049   0.02469  -0.0453   0.0037   1.0000
  13.750   1.6175   0.03219   0.02648  -0.0438   0.0036   1.0000
  14.000   1.6212   0.03410   0.02849  -0.0425   0.0035   1.0000
  14.250   1.6239   0.03625   0.03074  -0.0415   0.0034   1.0000
  14.500   1.6258   0.03864   0.03324  -0.0408   0.0033   1.0000
  14.750   1.6263   0.04137   0.03607  -0.0405   0.0032   1.0000
  15.000   1.6255   0.04440   0.03920  -0.0404   0.0031   1.0000
  15.250   1.6232   0.04769   0.04260  -0.0405   0.0031   1.0000
  15.500   1.6190   0.05135   0.04638  -0.0409   0.0030   1.0000
  15.750   1.6129   0.05536   0.05050  -0.0415   0.0030   1.0000
  16.000   1.6042   0.05979   0.05505  -0.0423   0.0029   1.0000
  16.250   1.5945   0.06454   0.05992  -0.0434   0.0029   1.0000
  16.500   1.5831   0.06967   0.06517  -0.0447   0.0028   1.0000
  16.750   1.5693   0.07523   0.07085  -0.0463   0.0028   1.0000
  17.000   1.5545   0.08112   0.07687  -0.0481   0.0028   1.0000
  17.250   1.5393   0.08721   0.08308  -0.0501   0.0028   1.0000
  17.500   1.5222   0.09369   0.08969  -0.0523   0.0028   1.0000
  17.750   1.5055   0.10022   0.09633  -0.0546   0.0028   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)