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=100,000 Ncrit=5


Details Polar file
Airfoil: GOE 134 (MVA H.12) AIRFOIL (goe134-il)
Reynolds number: 100,000
Max Cl/Cd: 55.18 at α=7°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe134-il-100000-n5.txt
Download as CSV file: xf-goe134-il-100000-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 =     0.100 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.3644   0.09728   0.09245  -0.0270   1.0000   0.0418
  -7.750  -0.3720   0.09485   0.09011  -0.0290   1.0000   0.0425
  -7.500  -0.3753   0.09181   0.08714  -0.0338   1.0000   0.0430
  -7.250  -0.3742   0.08841   0.08374  -0.0382   1.0000   0.0433
  -7.000  -0.3709   0.08461   0.07994  -0.0406   1.0000   0.0435
  -6.750  -0.3686   0.08074   0.07619  -0.0352   1.0000   0.0446
  -6.500  -0.3646   0.07791   0.07341  -0.0337   1.0000   0.0458
  -6.250  -0.3600   0.07501   0.07053  -0.0338   1.0000   0.0475
  -6.000  -0.3538   0.07191   0.06742  -0.0350   1.0000   0.0505
  -5.500  -0.3361   0.06446   0.05984  -0.0393   1.0000   0.0565
  -5.250  -0.3267   0.06187   0.05726  -0.0386   0.9996   0.0594
  -5.000  -0.2804   0.05638   0.05123  -0.0489   0.9929   0.0679
  -4.750  -0.2542   0.05248   0.04742  -0.0511   0.9873   0.0711
  -4.500  -0.2158   0.04850   0.04307  -0.0569   0.9804   0.0831
  -4.250  -0.1781   0.04531   0.03956  -0.0615   0.9743   0.0973
  -4.000  -0.1476   0.04187   0.03613  -0.0639   0.9668   0.1024
  -3.750  -0.0979   0.03597   0.02940  -0.0676   0.9610   0.0616
  -3.500  -0.0559   0.03230   0.02500  -0.0699   0.9534   0.0526
  -3.250  -0.0184   0.02951   0.02188  -0.0723   0.9461   0.0506
  -3.000   0.0188   0.02737   0.01932  -0.0742   0.9371   0.0506
  -2.750   0.0538   0.02569   0.01725  -0.0754   0.9262   0.0517
  -2.500   0.0893   0.02402   0.01527  -0.0767   0.9153   0.0509
  -2.250   0.1247   0.02259   0.01355  -0.0778   0.9038   0.0504
  -2.000   0.1578   0.02139   0.01212  -0.0784   0.8896   0.0503
  -1.750   0.1898   0.02049   0.01101  -0.0788   0.8735   0.0519
  -1.500   0.2211   0.01969   0.01003  -0.0790   0.8559   0.0531
  -1.250   0.2520   0.01885   0.00907  -0.0792   0.8378   0.0531
  -1.000   0.2823   0.01812   0.00824  -0.0792   0.8193   0.0533
  -0.750   0.3118   0.01749   0.00753  -0.0791   0.8006   0.0535
  -0.500   0.3402   0.01700   0.00695  -0.0788   0.7817   0.0540
  -0.250   0.3682   0.01659   0.00645  -0.0785   0.7636   0.0546
   0.000   0.3961   0.01624   0.00598  -0.0782   0.7462   0.0566
   0.250   0.4239   0.01607   0.00569  -0.0780   0.7293   0.0604
   0.500   0.4516   0.01598   0.00544  -0.0777   0.7130   0.0633
   0.750   0.4792   0.01594   0.00525  -0.0774   0.6969   0.0657
   1.000   0.5068   0.01591   0.00511  -0.0771   0.6809   0.0699
   1.500   0.5582   0.01363   0.00502  -0.0758   0.6497   1.0000
   1.750   0.5851   0.01381   0.00500  -0.0754   0.6338   1.0000
   2.000   0.6120   0.01400   0.00502  -0.0751   0.6174   1.0000
   2.250   0.6388   0.01419   0.00506  -0.0748   0.6011   1.0000
   2.500   0.6655   0.01440   0.00514  -0.0745   0.5851   1.0000
   2.750   0.6920   0.01461   0.00523  -0.0742   0.5690   1.0000
   3.000   0.7186   0.01484   0.00534  -0.0740   0.5533   1.0000
   3.250   0.7450   0.01508   0.00550  -0.0737   0.5378   1.0000
   3.500   0.7714   0.01534   0.00567  -0.0734   0.5226   1.0000
   3.750   0.7977   0.01562   0.00586  -0.0732   0.5080   1.0000
   4.000   0.8238   0.01591   0.00610  -0.0729   0.4937   1.0000
   4.250   0.8499   0.01621   0.00634  -0.0726   0.4796   1.0000
   4.500   0.8758   0.01653   0.00661  -0.0723   0.4662   1.0000
   4.750   0.9017   0.01686   0.00694  -0.0721   0.4534   1.0000
   5.000   0.9276   0.01720   0.00728  -0.0718   0.4408   1.0000
   5.250   0.9533   0.01756   0.00764  -0.0715   0.4287   1.0000
   5.500   0.9789   0.01794   0.00801  -0.0712   0.4174   1.0000
   5.750   1.0043   0.01834   0.00842  -0.0709   0.4073   1.0000
   6.000   1.0298   0.01873   0.00888  -0.0707   0.3965   1.0000
   6.250   1.0549   0.01915   0.00934  -0.0703   0.3858   1.0000
   6.500   1.0794   0.01957   0.00979  -0.0699   0.3741   1.0000
   6.750   1.1034   0.02000   0.01023  -0.0694   0.3613   1.0000
   7.000   1.1273   0.02043   0.01073  -0.0689   0.3485   1.0000
   7.250   1.1513   0.02088   0.01131  -0.0684   0.3375   1.0000
   7.500   1.1747   0.02135   0.01186  -0.0678   0.3261   1.0000
   7.750   1.1974   0.02182   0.01240  -0.0672   0.3136   1.0000
   8.000   1.2196   0.02232   0.01297  -0.0664   0.3012   1.0000
   8.250   1.2417   0.02283   0.01363  -0.0657   0.2889   1.0000
   8.500   1.2634   0.02335   0.01431  -0.0649   0.2763   1.0000
   8.750   1.2835   0.02390   0.01496  -0.0639   0.2607   1.0000
   9.000   1.3017   0.02450   0.01561  -0.0627   0.2409   1.0000
   9.250   1.3190   0.02517   0.01635  -0.0613   0.2168   1.0000
   9.500   1.3348   0.02599   0.01720  -0.0599   0.1900   1.0000
   9.750   1.3495   0.02693   0.01815  -0.0583   0.1652   1.0000
  10.000   1.3611   0.02810   0.01926  -0.0564   0.1411   1.0000
  10.250   1.3691   0.02952   0.02060  -0.0541   0.1189   1.0000
  10.500   1.3746   0.03109   0.02212  -0.0516   0.0985   1.0000
  10.750   1.3760   0.03277   0.02378  -0.0485   0.0781   1.0000
  11.000   1.3749   0.03463   0.02563  -0.0453   0.0601   1.0000
  11.250   1.3719   0.03678   0.02778  -0.0423   0.0462   1.0000
  11.500   1.3680   0.03915   0.03021  -0.0398   0.0372   1.0000
  11.750   1.3632   0.04178   0.03293  -0.0377   0.0324   1.0000
  12.000   1.3573   0.04473   0.03599  -0.0362   0.0293   1.0000
  12.250   1.3499   0.04810   0.03946  -0.0352   0.0274   1.0000
  12.500   1.3431   0.05171   0.04325  -0.0349   0.0259   1.0000
  12.750   1.3349   0.05572   0.04745  -0.0351   0.0247   1.0000
  13.000   1.3252   0.06017   0.05207  -0.0357   0.0238   1.0000
  13.250   1.3137   0.06507   0.05713  -0.0368   0.0231   1.0000
  13.500   1.3012   0.07032   0.06254  -0.0381   0.0226   1.0000
  13.750   1.2877   0.07588   0.06825  -0.0398   0.0222   1.0000
  14.000   1.2740   0.08160   0.07413  -0.0416   0.0219   1.0000
  14.250   1.2613   0.08730   0.07998  -0.0434   0.0217   1.0000
  14.500   1.2509   0.09271   0.08555  -0.0452   0.0214   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)