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

GOE 11K AIRFOIL (goe11k-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: GOE 11K AIRFOIL (goe11k-il)
Reynolds number: 100,000
Max Cl/Cd: 49.02 at α=4.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe11k-il-100000-n5.txt
Download as CSV file: xf-goe11k-il-100000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 11K 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
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.5057   0.09235   0.08763  -0.0301   1.0000   0.0247
  -8.000  -0.5187   0.08972   0.08507  -0.0286   1.0000   0.0245
  -7.750  -0.5301   0.08697   0.08239  -0.0275   1.0000   0.0243
  -7.500  -0.5384   0.08346   0.07892  -0.0276   1.0000   0.0241
  -7.250  -0.5445   0.07972   0.07520  -0.0278   1.0000   0.0240
  -7.000  -0.5486   0.07578   0.07126  -0.0283   1.0000   0.0238
  -6.750  -0.5510   0.07118   0.06662  -0.0291   1.0000   0.0239
  -6.500  -0.5488   0.06622   0.06158  -0.0305   0.9996   0.0239
  -6.250  -0.5312   0.05924   0.05436  -0.0357   0.9959   0.0243
  -6.000  -0.5134   0.05186   0.04661  -0.0394   0.9922   0.0249
  -5.750  -0.4967   0.04492   0.03921  -0.0417   0.9887   0.0253
  -5.500  -0.4793   0.03971   0.03350  -0.0427   0.9859   0.0260
  -5.250  -0.4600   0.03682   0.03024  -0.0427   0.9828   0.0272
  -5.000  -0.4365   0.03490   0.02798  -0.0430   0.9800   0.0299
  -4.750  -0.4126   0.03130   0.02371  -0.0429   0.9779   0.0326
  -4.500  -0.3896   0.02790   0.01943  -0.0418   0.9758   0.0362
  -4.250  -0.3662   0.02694   0.01830  -0.0414   0.9730   0.0400
  -4.000  -0.3398   0.02560   0.01638  -0.0409   0.9707   0.0474
  -3.750  -0.3128   0.02467   0.01533  -0.0411   0.9685   0.0541
  -3.500  -0.2837   0.02388   0.01426  -0.0415   0.9667   0.0638
  -3.250  -0.2605   0.02318   0.01333  -0.0407   0.9635   0.0720
  -3.000  -0.2355   0.02275   0.01274  -0.0404   0.9607   0.0804
  -2.750  -0.2090   0.02221   0.01220  -0.0404   0.9583   0.0875
  -2.500  -0.1803   0.02179   0.01158  -0.0407   0.9561   0.0933
  -2.250  -0.1510   0.02127   0.01104  -0.0412   0.9543   0.0973
  -2.000  -0.1299   0.02091   0.01065  -0.0400   0.9506   0.1011
  -1.750  -0.1058   0.02068   0.01032  -0.0394   0.9475   0.1062
  -1.500  -0.0796   0.02046   0.01011  -0.0393   0.9448   0.1139
  -1.250  -0.0509   0.02031   0.00990  -0.0396   0.9424   0.1217
  -1.000  -0.0205   0.02013   0.00973  -0.0404   0.9405   0.1347
  -0.750   0.0010   0.01994   0.00961  -0.0393   0.9369   0.1563
  -0.500   0.0951   0.01775   0.01005  -0.0543   0.9445   1.0000
  -0.250   0.1244   0.01796   0.01006  -0.0549   0.9421   1.0000
   0.000   0.1546   0.01818   0.01013  -0.0556   0.9400   1.0000
   0.250   0.1732   0.01832   0.01017  -0.0540   0.9347   1.0000
   0.500   0.1992   0.01851   0.01027  -0.0539   0.9312   1.0000
   0.750   0.2292   0.01871   0.01040  -0.0546   0.9281   1.0000
   1.000   0.2646   0.01888   0.01051  -0.0564   0.9251   1.0000
   1.250   0.2848   0.01897   0.01057  -0.0550   0.9174   1.0000
   1.500   0.3193   0.01905   0.01065  -0.0564   0.9128   1.0000
   1.750   0.3491   0.01915   0.01075  -0.0569   0.9075   1.0000
   2.000   0.3756   0.01923   0.01087  -0.0567   0.9006   1.0000
   2.250   0.4122   0.01926   0.01095  -0.0585   0.8965   1.0000
   2.500   0.4354   0.01934   0.01109  -0.0576   0.8882   1.0000
   2.750   0.4723   0.01924   0.01110  -0.0592   0.8824   1.0000
   3.000   0.5031   0.01899   0.01095  -0.0594   0.8707   1.0000
   3.250   0.5517   0.01788   0.01001  -0.0619   0.8532   1.0000
   3.500   0.5928   0.01640   0.00864  -0.0623   0.8224   1.0000
   4.000   0.6373   0.01556   0.00801  -0.0579   0.7530   1.0000
   4.250   0.7318   0.01493   0.00638  -0.0690   0.4831   1.0000
   4.500   0.7414   0.01613   0.00695  -0.0654   0.3607   1.0000
   4.750   0.7506   0.01753   0.00753  -0.0621   0.2107   1.0000
   5.000   0.7671   0.01859   0.00822  -0.0602   0.1512   1.0000
   5.250   0.7854   0.01936   0.00895  -0.0584   0.1298   1.0000
   5.500   0.8031   0.02012   0.00968  -0.0565   0.1137   1.0000
   5.750   0.8205   0.02089   0.01041  -0.0547   0.0968   1.0000
   6.000   0.8377   0.02169   0.01120  -0.0529   0.0824   1.0000
   6.250   0.8552   0.02251   0.01211  -0.0511   0.0683   1.0000
   6.500   0.8720   0.02347   0.01313  -0.0491   0.0550   1.0000
   6.750   0.8896   0.02450   0.01424  -0.0472   0.0441   1.0000
   7.000   0.9079   0.02565   0.01545  -0.0455   0.0376   1.0000
   7.250   0.9274   0.02672   0.01664  -0.0440   0.0323   1.0000
   7.500   0.9483   0.02843   0.01843  -0.0430   0.0293   1.0000
   7.750   0.9731   0.03023   0.02053  -0.0424   0.0259   1.0000
   8.000   0.9971   0.03238   0.02296  -0.0417   0.0239   1.0000
   8.250   1.0169   0.03448   0.02531  -0.0405   0.0224   1.0000
   8.500   1.0318   0.03686   0.02794  -0.0387   0.0208   1.0000
   8.750   1.0437   0.04006   0.03159  -0.0363   0.0197   1.0000
   9.000   1.0511   0.04317   0.03521  -0.0330   0.0191   1.0000
   9.250   1.0531   0.04642   0.03892  -0.0292   0.0188   1.0000
   9.500   1.0495   0.04983   0.04275  -0.0248   0.0186   1.0000
   9.750   1.0409   0.05314   0.04643  -0.0200   0.0185   1.0000
  10.000   1.0278   0.05635   0.04995  -0.0149   0.0184   1.0000
  10.250   1.0120   0.05966   0.05354  -0.0102   0.0184   1.0000
  10.500   0.9946   0.06312   0.05724  -0.0061   0.0185   1.0000
  10.750   0.9750   0.06690   0.06124  -0.0029   0.0185   1.0000
  11.000   0.9532   0.07114   0.06569  -0.0008   0.0184   1.0000
  11.250   0.9312   0.07587   0.07058  -0.0001   0.0186   1.0000
  11.500   0.9103   0.08108   0.07592  -0.0009   0.0189   1.0000
  11.750   0.8876   0.08769   0.08265  -0.0038   0.0190   1.0000
<< Back to GOE 11K AIRFOIL (goe11k-il)

Polar data table (+)

Polar graphs


<< Back to GOE 11K AIRFOIL (goe11k-il)