Airfoil Tools
Search 1636 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

GOE 10K AIRFOIL (goe10k-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9


Details Polar file
Airfoil: GOE 10K AIRFOIL (goe10k-il)
Reynolds number: 1,000,000
Max Cl/Cd: 111.64 at α=4°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe10k-il-1000000.txt
Download as CSV file: xf-goe10k-il-1000000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 10K 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
  ------ -------- --------- --------- -------- -------- --------
  -7.750  -0.5396   0.08743   0.08586  -0.0087   1.0000   0.0037
  -7.500  -0.5450   0.08481   0.08326  -0.0085   1.0000   0.0036
  -7.250  -0.5451   0.08125   0.07973  -0.0103   1.0000   0.0038
  -7.000  -0.5417   0.07758   0.07606  -0.0127   1.0000   0.0037
  -6.750  -0.5366   0.07362   0.07210  -0.0154   1.0000   0.0039
  -6.500  -0.5295   0.06957   0.06803  -0.0179   1.0000   0.0039
  -6.250  -0.5207   0.06539   0.06382  -0.0203   1.0000   0.0041
  -6.000  -0.5093   0.06084   0.05922  -0.0226   1.0000   0.0043
  -5.750  -0.4959   0.05641   0.05472  -0.0244   1.0000   0.0044
  -5.500  -0.4809   0.05190   0.05011  -0.0256   1.0000   0.0045
  -5.250  -0.4648   0.04740   0.04549  -0.0261   1.0000   0.0045
  -5.000  -0.4349   0.04211   0.04002  -0.0293   0.9988   0.0046
  -4.750  -0.4066   0.03621   0.03390  -0.0330   0.9971   0.0048
  -4.500  -0.3767   0.03249   0.02998  -0.0354   0.9957   0.0051
  -4.250  -0.3466   0.02841   0.02564  -0.0368   0.9942   0.0056
  -4.000  -0.3144   0.02373   0.02059  -0.0364   0.9922   0.0065
  -3.750  -0.2868   0.01890   0.01527  -0.0367   0.9904   0.0069
  -3.500  -0.2575   0.01705   0.01319  -0.0375   0.9891   0.0078
  -3.250  -0.2271   0.01491   0.01069  -0.0378   0.9881   0.0109
  -3.000  -0.1964   0.01312   0.00857  -0.0382   0.9873   0.0139
  -2.750  -0.1685   0.01215   0.00749  -0.0381   0.9853   0.0158
  -2.500  -0.1426   0.00887   0.00367  -0.0357   0.9839   0.0062
  -2.250  -0.1139   0.00804   0.00273  -0.0358   0.9821   0.0075
  -2.000  -0.0831   0.00758   0.00222  -0.0365   0.9806   0.0107
  -1.750  -0.0508   0.00710   0.00178  -0.0374   0.9793   0.0379
  -1.500  -0.0166   0.00684   0.00163  -0.0390   0.9782   0.0801
  -1.250   0.0114   0.00641   0.00150  -0.0393   0.9746   0.1749
  -1.000   0.0382   0.00527   0.00139  -0.0397   0.9708   0.5172
  -0.750   0.0638   0.00419   0.00132  -0.0393   0.9679   0.8310
  -0.500   0.1479   0.00408   0.00151  -0.0523   0.9777   0.9975
  -0.250   0.1811   0.00401   0.00141  -0.0536   0.9732   1.0000
   0.000   0.2169   0.00390   0.00129  -0.0554   0.9694   1.0000
   0.250   0.2528   0.00373   0.00112  -0.0571   0.9601   1.0000
   0.500   0.2983   0.00356   0.00095  -0.0611   0.9443   1.0000
   0.750   0.3501   0.00346   0.00081  -0.0665   0.9173   1.0000
   1.000   0.3829   0.00354   0.00076  -0.0675   0.8721   1.0000
   1.250   0.4011   0.00385   0.00075  -0.0650   0.7831   1.0000
   1.500   0.4024   0.00515   0.00090  -0.0593   0.4879   1.0000
   1.750   0.4124   0.00650   0.00117  -0.0560   0.1911   1.0000
   2.000   0.4318   0.00715   0.00139  -0.0544   0.0754   1.0000
   2.250   0.4536   0.00759   0.00162  -0.0532   0.0170   1.0000
   2.500   0.4760   0.00805   0.00212  -0.0519   0.0090   1.0000
   2.750   0.4974   0.00866   0.00280  -0.0503   0.0069   1.0000
   3.000   0.5197   0.00918   0.00339  -0.0490   0.0062   1.0000
   3.250   0.5412   0.00982   0.00407  -0.0475   0.0052   1.0000
   3.750   2.1728   0.59188   0.59038  -0.5010   0.0069   1.0000
   4.000   0.5783   0.00518   0.00044  -0.0361   0.0098   1.0000
   4.250   0.6023   0.00544   0.00104  -0.0344   0.0082   1.0000
   4.500   0.6216   0.00662   0.00239  -0.0329   0.0075   1.0000
   4.750   0.6338   0.00965   0.00575  -0.0303   0.0070   1.0000
   5.000   0.6548   0.01183   0.00838  -0.0272   0.0062   1.0000
   5.250   0.6704   0.01461   0.01144  -0.0248   0.0056   1.0000
   5.500   0.6847   0.01702   0.01403  -0.0233   0.0052   1.0000
   5.750   0.6898   0.02160   0.01887  -0.0211   0.0049   1.0000
   6.000   0.7052   0.02607   0.02363  -0.0183   0.0048   1.0000
   6.250   0.7168   0.03119   0.02898  -0.0160   0.0042   1.0000
   6.500   0.7246   0.03602   0.03398  -0.0145   0.0039   1.0000
   6.750   0.7303   0.04083   0.03894  -0.0133   0.0037   1.0000
   7.000   0.7337   0.04553   0.04377  -0.0125   0.0036   1.0000
   7.250   0.7342   0.05030   0.04864  -0.0119   0.0035   1.0000
   7.500   0.7317   0.05522   0.05365  -0.0116   0.0035   1.0000
   7.750   0.7234   0.05965   0.05813  -0.0113   0.0033   1.0000
   8.000   0.7118   0.06414   0.06268  -0.0109   0.0034   1.0000
   8.250   0.6895   0.07025   0.06883  -0.0141   0.0036   1.0000
   8.500   0.6745   0.07621   0.07477  -0.0185   0.0037   1.0000
   8.750   0.6659   0.08171   0.08026  -0.0212   0.0037   1.0000
<< Back to GOE 10K AIRFOIL (goe10k-il)

Polar data table (+)

Polar graphs


<< Back to GOE 10K AIRFOIL (goe10k-il)