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GOE 113 AIRFOIL (goe113-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9


Details Polar file
Airfoil: GOE 113 AIRFOIL (goe113-il)
Reynolds number: 1,000,000
Max Cl/Cd: 97.12 at α=2.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe113-il-1000000.txt
Download as CSV file: xf-goe113-il-1000000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 113 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.4568   0.09343   0.09187  -0.0069   1.0000   0.0076
  -7.750  -0.4531   0.08987   0.08833  -0.0090   1.0000   0.0076
  -7.500  -0.4498   0.08625   0.08473  -0.0115   1.0000   0.0076
  -7.250  -0.4412   0.08200   0.08049  -0.0160   1.0000   0.0077
  -7.000  -0.4311   0.07762   0.07611  -0.0203   1.0000   0.0077
  -6.750  -0.4205   0.07321   0.07169  -0.0242   1.0000   0.0077
  -6.500  -0.3760   0.05338   0.05200  -0.0261   1.0000   0.0081
  -6.250  -0.3723   0.05011   0.04873  -0.0268   1.0000   0.0083
  -6.000  -0.3673   0.04711   0.04572  -0.0272   1.0000   0.0086
  -5.750  -0.3463   0.04260   0.04114  -0.0319   0.9986   0.0093
  -5.500  -0.3129   0.03710   0.03552  -0.0387   0.9966   0.0116
  -5.250  -0.2819   0.03216   0.03044  -0.0434   0.9949   0.0118
  -4.000  -0.1657   0.02301   0.01989  -0.0549   0.9880   0.0102
  -3.750  -0.1289   0.01534   0.01139  -0.0558   0.9857   0.0112
  -3.500  -0.0942   0.01122   0.00666  -0.0570   0.9837   0.0126
  -3.250  -0.0620   0.01161   0.00710  -0.0581   0.9772   0.0136
  -3.000  -0.0282   0.01155   0.00700  -0.0594   0.9689   0.0155
  -2.750   0.0015   0.01094   0.00625  -0.0595   0.9545   0.0172
  -2.500   0.0289   0.00983   0.00495  -0.0592   0.9330   0.0190
  -2.250   0.0544   0.00952   0.00455  -0.0585   0.8971   0.0208
  -2.000   0.0788   0.00957   0.00441  -0.0576   0.8446   0.0230
  -1.750   0.1041   0.00937   0.00395  -0.0568   0.7945   0.0246
  -1.500   0.1303   0.00925   0.00359  -0.0563   0.7509   0.0259
  -1.250   0.1569   0.00891   0.00307  -0.0561   0.7177   0.0284
  -1.000   0.1837   0.00842   0.00251  -0.0559   0.6942   0.0311
  -0.750   0.2111   0.00820   0.00219  -0.0559   0.6716   0.0320
  -0.500   0.2386   0.00803   0.00192  -0.0558   0.6508   0.0323
  -0.250   0.2664   0.00788   0.00167  -0.0557   0.6319   0.0323
   0.000   0.2940   0.00779   0.00148  -0.0557   0.6116   0.0330
   0.250   0.3218   0.00771   0.00130  -0.0557   0.5913   0.0328
   0.500   0.3497   0.00766   0.00116  -0.0557   0.5741   0.0332
   0.750   0.3776   0.00762   0.00105  -0.0557   0.5582   0.0350
   1.250   0.4333   0.00764   0.00094  -0.0557   0.5270   0.0389
   1.500   0.4612   0.00767   0.00092  -0.0557   0.5137   0.0415
   1.750   0.4891   0.00771   0.00094  -0.0558   0.4999   0.0477
   2.000   0.5111   0.00554   0.00111  -0.0552   0.4868   1.0000
   2.250   0.5389   0.00565   0.00114  -0.0552   0.4708   1.0000
   2.500   0.5662   0.00583   0.00120  -0.0552   0.4418   1.0000
   2.750   0.5921   0.00626   0.00130  -0.0550   0.3679   1.0000
   3.000   0.6173   0.00684   0.00151  -0.0548   0.2828   1.0000
   3.250   0.6435   0.00725   0.00171  -0.0547   0.2368   1.0000
   3.500   0.6701   0.00757   0.00188  -0.0546   0.2065   1.0000
   3.750   0.6969   0.00786   0.00204  -0.0546   0.1783   1.0000
   4.000   0.7230   0.00825   0.00223  -0.0545   0.1390   1.0000
   4.250   0.7489   0.00870   0.00250  -0.0543   0.1040   1.0000
   4.500   0.7723   0.00963   0.00304  -0.0538   0.0174   1.0000
   4.750   0.7991   0.00990   0.00335  -0.0536   0.0147   1.0000
   5.000   0.8256   0.01022   0.00373  -0.0534   0.0131   1.0000
   5.250   0.8516   0.01067   0.00426  -0.0531   0.0112   1.0000
   5.500   0.8752   0.01162   0.00540  -0.0523   0.0096   1.0000
   5.750   0.9017   0.01188   0.00567  -0.0521   0.0091   1.0000
   6.000   0.9269   0.01241   0.00627  -0.0517   0.0087   1.0000
   6.250   0.9515   0.01302   0.00694  -0.0512   0.0082   1.0000
   6.500   0.9756   0.01369   0.00769  -0.0506   0.0076   1.0000
   6.750   0.9988   0.01448   0.00856  -0.0499   0.0069   1.0000
   7.000   1.0209   0.01544   0.00960  -0.0490   0.0065   1.0000
   7.250   1.0414   0.01670   0.01095  -0.0478   0.0064   1.0000
   7.500   1.0603   0.01861   0.01301  -0.0460   0.0070   1.0000
   7.750   1.0817   0.01989   0.01436  -0.0450   0.0067   1.0000
   8.000   1.1029   0.02089   0.01539  -0.0444   0.0061   1.0000
   8.750   1.1592   0.02828   0.02354  -0.0395   0.0071   1.0000
   9.000   1.1777   0.02951   0.02483  -0.0387   0.0065   1.0000
  16.750   0.7312   0.17989   0.17848  -0.0627   0.0057   1.0000
  17.000   0.7225   0.18583   0.18439  -0.0646   0.0054   1.0000
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