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=9


Details Polar file
Airfoil: GOE 134 (MVA H.12) AIRFOIL (goe134-il)
Reynolds number: 100,000
Max Cl/Cd: 50.73 at α=9°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe134-il-100000.txt
Download as CSV file: xf-goe134-il-100000.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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.3911   0.10585   0.10092  -0.0242   1.0000   0.0630
  -8.250  -0.4006   0.10406   0.09924  -0.0280   1.0000   0.0635
  -8.000  -0.4054   0.10153   0.09676  -0.0342   1.0000   0.0638
  -7.750  -0.3945   0.09548   0.09075  -0.0285   1.0000   0.0654
  -7.500  -0.3834   0.09196   0.08725  -0.0254   1.0000   0.0679
  -7.250  -0.3806   0.08898   0.08432  -0.0259   1.0000   0.0705
  -7.000  -0.3799   0.08598   0.08137  -0.0279   1.0000   0.0736
  -6.750  -0.3800   0.08322   0.07860  -0.0350   1.0000   0.0768
  -6.500  -0.3778   0.07984   0.07507  -0.0409   1.0000   0.0783
  -6.250  -0.3724   0.07557   0.07104  -0.0341   1.0000   0.0809
  -6.000  -0.3662   0.07278   0.06828  -0.0329   1.0000   0.0845
  -5.750  -0.3564   0.06976   0.06515  -0.0364   1.0000   0.0905
  -5.500  -0.3470   0.06573   0.06098  -0.0394   1.0000   0.0941
  -5.250  -0.3409   0.06285   0.05822  -0.0365   1.0000   0.0973
  -5.000  -0.3215   0.06073   0.05564  -0.0414   1.0000   0.1082
  -4.750  -0.3167   0.05674   0.05193  -0.0384   1.0000   0.1109
  -4.250  -0.2946   0.05174   0.04686  -0.0375   1.0000   0.1283
  -4.000  -0.2822   0.04945   0.04447  -0.0378   1.0000   0.1419
  -3.750  -0.2703   0.04746   0.04239  -0.0376   1.0000   0.1566
  -3.500  -0.2586   0.04558   0.04044  -0.0372   1.0000   0.1719
  -3.250  -0.2352   0.04322   0.03811  -0.0384   0.9972   0.1925
  -3.000  -0.1951   0.04034   0.03518  -0.0428   0.9902   0.2367
  -2.750  -0.1591   0.03805   0.03296  -0.0456   0.9835   0.3100
  -2.000   0.0161   0.02905   0.02225  -0.0653   0.9594   0.2475
  -1.750   0.0835   0.02691   0.01853  -0.0679   0.9510   0.1227
  -1.500   0.1350   0.02476   0.01601  -0.0714   0.9435   0.1067
  -1.250   0.1804   0.02288   0.01401  -0.0744   0.9332   0.1006
  -1.000   0.2336   0.02143   0.01232  -0.0783   0.9270   0.0965
  -0.750   0.2742   0.02035   0.01123  -0.0803   0.9141   0.0987
  -0.500   0.3131   0.01934   0.01023  -0.0816   0.9010   0.0984
  -0.250   0.3493   0.01846   0.00938  -0.0825   0.8870   0.0989
   0.000   0.3831   0.01775   0.00865  -0.0828   0.8719   0.1009
   0.250   0.4152   0.01721   0.00802  -0.0827   0.8557   0.1046
   0.500   0.4457   0.01672   0.00745  -0.0823   0.8390   0.1131
   0.750   0.4747   0.01630   0.00705  -0.0818   0.8208   0.1396
   1.000   0.5011   0.01397   0.00684  -0.0802   0.8025   1.0000
   1.250   0.5283   0.01403   0.00659  -0.0792   0.7833   1.0000
   1.500   0.5554   0.01414   0.00643  -0.0783   0.7648   1.0000
   1.750   0.5816   0.01435   0.00641  -0.0775   0.7452   1.0000
   2.000   0.6076   0.01460   0.00648  -0.0767   0.7253   1.0000
   2.250   0.6338   0.01487   0.00657  -0.0760   0.7064   1.0000
   2.500   0.6600   0.01517   0.00668  -0.0753   0.6882   1.0000
   2.750   0.6859   0.01547   0.00687  -0.0747   0.6689   1.0000
   3.000   0.7117   0.01577   0.00708  -0.0742   0.6497   1.0000
   3.250   0.7378   0.01606   0.00725  -0.0736   0.6316   1.0000
   3.500   0.7640   0.01635   0.00743  -0.0731   0.6141   1.0000
   3.750   0.7902   0.01665   0.00764  -0.0726   0.5973   1.0000
   4.000   0.8163   0.01697   0.00788  -0.0721   0.5801   1.0000
   4.250   0.8423   0.01732   0.00819  -0.0717   0.5632   1.0000
   4.500   0.8685   0.01771   0.00852  -0.0714   0.5474   1.0000
   4.750   0.8946   0.01812   0.00891  -0.0710   0.5322   1.0000
   5.000   0.9207   0.01855   0.00930  -0.0707   0.5177   1.0000
   5.250   0.9471   0.01901   0.00972  -0.0704   0.5045   1.0000
   5.500   0.9735   0.01949   0.01016  -0.0702   0.4919   1.0000
   5.750   0.9991   0.02000   0.01075  -0.0699   0.4787   1.0000
   6.000   1.0246   0.02051   0.01132  -0.0695   0.4659   1.0000
   6.250   1.0498   0.02096   0.01179  -0.0691   0.4522   1.0000
   6.500   1.0747   0.02137   0.01225  -0.0686   0.4383   1.0000
   6.750   1.0998   0.02184   0.01276  -0.0682   0.4258   1.0000
   7.000   1.1253   0.02234   0.01327  -0.0678   0.4143   1.0000
   7.250   1.1506   0.02279   0.01371  -0.0673   0.4020   1.0000
   7.500   1.1742   0.02322   0.01427  -0.0667   0.3883   1.0000
   7.750   1.1976   0.02366   0.01482  -0.0660   0.3746   1.0000
   8.000   1.2208   0.02414   0.01541  -0.0653   0.3609   1.0000
   8.500   1.2657   0.02508   0.01658  -0.0636   0.3320   1.0000
   8.750   1.2867   0.02540   0.01693  -0.0625   0.3147   1.0000
   9.000   1.3049   0.02572   0.01736  -0.0610   0.2951   1.0000
   9.250   1.3215   0.02605   0.01782  -0.0593   0.2741   1.0000
   9.500   1.3371   0.02645   0.01829  -0.0575   0.2530   1.0000
   9.750   1.3507   0.02698   0.01897  -0.0555   0.2297   1.0000
  10.000   1.3631   0.02769   0.01977  -0.0534   0.2063   1.0000
  10.250   1.3726   0.02855   0.02066  -0.0511   0.1796   1.0000
  10.500   1.3790   0.02978   0.02187  -0.0484   0.1522   1.0000
  10.750   1.3773   0.03168   0.02365  -0.0449   0.1247   1.0000
  11.000   1.3643   0.03420   0.02599  -0.0400   0.1029   1.0000
  11.250   1.3514   0.03707   0.02883  -0.0357   0.0839   1.0000
  11.500   1.3421   0.03994   0.03167  -0.0326   0.0732   1.0000
  11.750   1.3344   0.04288   0.03454  -0.0302   0.0668   1.0000
  12.000   1.3330   0.04553   0.03734  -0.0284   0.0608   1.0000
  12.250   1.3305   0.04841   0.04016  -0.0268   0.0571   1.0000
  12.500   1.3349   0.05100   0.04289  -0.0252   0.0540   1.0000
  12.750   1.3403   0.05355   0.04559  -0.0238   0.0512   1.0000
  13.000   1.3457   0.05610   0.04821  -0.0227   0.0488   1.0000
  13.250   1.3581   0.05875   0.05076  -0.0211   0.0460   1.0000
  13.500   1.3566   0.06193   0.05426  -0.0204   0.0448   1.0000
  13.750   1.3555   0.06534   0.05795  -0.0198   0.0435   1.0000
  14.000   1.3538   0.06904   0.06191  -0.0194   0.0426   1.0000
  14.250   1.3490   0.07315   0.06628  -0.0192   0.0420   1.0000
  14.500   1.3404   0.07773   0.07112  -0.0196   0.0417   1.0000
  14.750   1.3279   0.08285   0.07651  -0.0205   0.0415   1.0000
  15.000   1.3116   0.08860   0.08254  -0.0222   0.0415   1.0000
  15.250   1.2916   0.09511   0.08932  -0.0247   0.0416   1.0000
  15.500   1.2684   0.10245   0.09692  -0.0282   0.0419   1.0000
  15.750   1.2426   0.11081   0.10552  -0.0329   0.0423   1.0000
  16.000   1.2143   0.12035   0.11530  -0.0388   0.0429   1.0000
  16.250   1.1837   0.13129   0.12643  -0.0460   0.0437   1.0000
  16.500   1.1522   0.14355   0.13883  -0.0541   0.0447   1.0000
  16.750   1.1241   0.15611   0.15143  -0.0621   0.0456   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)