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GOE 114 (MVA MK.1) AIRFOIL (goe114-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: GOE 114 (MVA MK.1) AIRFOIL (goe114-il)
Reynolds number: 100,000
Max Cl/Cd: 44.42 at α=2.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe114-il-100000-n5.txt
Download as CSV file: xf-goe114-il-100000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 114 (MVA MK.1) 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
  ------ -------- --------- --------- -------- -------- --------
  -9.250  -0.3121   0.09332   0.08916  -0.0465   1.0000   0.0759
  -9.000  -0.2943   0.08978   0.08562  -0.0427   1.0000   0.0829
  -8.500  -0.3825   0.09484   0.09037  -0.0437   1.0000   0.0734
  -5.500  -0.2965   0.03558   0.02923  -0.0873   0.9486   0.0267
  -5.250  -0.2610   0.03296   0.02585  -0.0884   0.9442   0.0198
  -5.000  -0.2331   0.02983   0.02229  -0.0892   0.9388   0.0181
  -4.750  -0.2026   0.02744   0.01947  -0.0898   0.9334   0.0167
  -4.500  -0.1677   0.02534   0.01697  -0.0909   0.9300   0.0157
  -4.250  -0.1401   0.02376   0.01517  -0.0907   0.9236   0.0162
  -4.000  -0.1088   0.02231   0.01357  -0.0912   0.9188   0.0178
  -3.750  -0.0777   0.02105   0.01217  -0.0916   0.9142   0.0183
  -3.500  -0.0513   0.02012   0.01109  -0.0913   0.9070   0.0186
  -3.250  -0.0176   0.01915   0.00980  -0.0923   0.9027   0.0196
  -3.000   0.0088   0.01830   0.00861  -0.0921   0.8951   0.0232
  -2.750   0.0417   0.01752   0.00745  -0.0929   0.8901   0.0287
  -2.500   0.0677   0.01594   0.00685  -0.0933   0.8835   0.2623
  -2.250   0.0975   0.01641   0.00727  -0.0934   0.8772   0.3626
  -2.000   0.1250   0.01656   0.00730  -0.0933   0.8705   0.3958
  -1.750   0.1547   0.01641   0.00691  -0.0936   0.8641   0.4046
  -1.500   0.1842   0.01626   0.00663  -0.0940   0.8579   0.4123
  -1.250   0.2127   0.01611   0.00638  -0.0942   0.8510   0.4195
  -1.000   0.2421   0.01596   0.00612  -0.0945   0.8447   0.4253
  -0.750   0.2704   0.01586   0.00593  -0.0947   0.8374   0.4306
  -0.500   0.2994   0.01570   0.00578  -0.0949   0.8308   0.4356
  -0.250   0.3274   0.01560   0.00567  -0.0950   0.8234   0.4420
   0.000   0.3552   0.01550   0.00560  -0.0950   0.8162   0.4484
   0.250   0.3836   0.01541   0.00554  -0.0951   0.8089   0.4555
   0.500   0.4111   0.01535   0.00555  -0.0951   0.8012   0.4627
   0.750   0.4399   0.01526   0.00556  -0.0952   0.7939   0.4714
   1.000   0.4660   0.01522   0.00566  -0.0950   0.7854   0.4816
   1.250   0.4951   0.01510   0.00569  -0.0951   0.7785   0.4958
   1.500   0.5200   0.01506   0.00589  -0.0946   0.7691   0.5149
   1.750   0.5472   0.01491   0.00609  -0.0944   0.7615   0.5484
   2.000   0.5682   0.01419   0.00577  -0.0920   0.7303   0.6355
   2.500   0.6148   0.01384   0.00481  -0.0873   0.4851   1.0000
   2.750   0.6051   0.01727   0.00596  -0.0816   0.0362   1.0000
   3.000   0.6269   0.01793   0.00666  -0.0803   0.0242   1.0000
   3.250   0.6483   0.01861   0.00753  -0.0789   0.0203   1.0000
   3.500   0.6691   0.01932   0.00847  -0.0775   0.0192   1.0000
   3.750   0.6881   0.02017   0.00955  -0.0757   0.0186   1.0000
   4.000   0.7050   0.02114   0.01070  -0.0736   0.0177   1.0000
   4.250   0.7191   0.02224   0.01198  -0.0713   0.0162   1.0000
   4.500   0.7304   0.02360   0.01340  -0.0685   0.0150   1.0000
   4.750   0.7463   0.02500   0.01480  -0.0663   0.0152   1.0000
   5.000   0.7698   0.02654   0.01634  -0.0650   0.0157   1.0000
   5.250   0.7997   0.02826   0.01813  -0.0645   0.0168   1.0000
   5.500   0.8347   0.03057   0.02050  -0.0647   0.0188   1.0000
   5.750   0.8746   0.03413   0.02400  -0.0660   0.0220   1.0000
   6.000   0.9035   0.03516   0.02549  -0.0641   0.0281   1.0000
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