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

GOE 7K AIRFOIL (goe07k-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: GOE 7K AIRFOIL (goe07k-il)
Reynolds number: 200,000
Max Cl/Cd: 69.69 at α=6.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe07k-il-200000.txt
Download as CSV file: xf-goe07k-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 7K AIRFOIL                                  
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.500  -0.3871   0.09242   0.08914  -0.0613   0.9548   0.0414
  -7.250  -0.3838   0.08866   0.08536  -0.0646   0.9515   0.0424
  -7.000  -0.3928   0.08577   0.08248  -0.0640   0.9443   0.0429
  -6.750  -0.3860   0.08147   0.07813  -0.0675   0.9404   0.0440
  -6.500  -0.3853   0.07762   0.07422  -0.0690   0.9353   0.0458
  -6.250  -0.3775   0.07298   0.06922  -0.0738   0.9282   0.0487
  -5.500  -0.3669   0.06063   0.05662  -0.0711   0.9167   0.0513
  -5.250  -0.3511   0.05782   0.05372  -0.0713   0.9145   0.0527
  -5.000  -0.3304   0.05468   0.05040  -0.0723   0.9128   0.0553
  -4.750  -0.3322   0.05276   0.04833  -0.0682   0.9067   0.0572
  -4.500  -0.3139   0.05245   0.04721  -0.0657   0.9025   0.0621
  -4.250  -0.3041   0.04613   0.04092  -0.0655   0.9011   0.0644
  -4.000  -0.2818   0.04426   0.03904  -0.0658   0.8997   0.0681
  -3.500  -0.2657   0.04108   0.03541  -0.0596   0.8917   0.0808
  -3.250  -0.2473   0.03994   0.03402  -0.0581   0.8892   0.0871
  -3.000  -0.2237   0.03403   0.02664  -0.0527   0.8876   0.0499
  -2.750  -0.1996   0.03241   0.02499  -0.0527   0.8863   0.0537
  -2.500  -0.1703   0.03119   0.02347  -0.0526   0.8851   0.0557
  -2.250  -0.1656   0.03054   0.02262  -0.0481   0.8800   0.0565
  -2.000  -0.1478   0.02991   0.02177  -0.0461   0.8772   0.0587
  -1.750  -0.1239   0.02936   0.02096  -0.0452   0.8750   0.0633
  -1.500  -0.0949   0.02838   0.01994  -0.0455   0.8733   0.0672
  -1.250  -0.0630   0.02808   0.01956  -0.0463   0.8718   0.0743
  -1.000  -0.0312   0.02747   0.01894  -0.0471   0.8707   0.0796
  -0.750  -0.0354   0.02753   0.01901  -0.0414   0.8645   0.0825
  -0.500  -0.0127   0.02747   0.01891  -0.0406   0.8615   0.0895
  -0.250   0.0139   0.02714   0.01864  -0.0405   0.8594   0.0974
   0.000   0.0473   0.02702   0.01850  -0.0417   0.8576   0.1071
   0.250   0.2355   0.02548   0.01978  -0.0789   0.8655   1.0000
   0.500   0.2595   0.02568   0.01985  -0.0784   0.8614   1.0000
   0.750   0.2943   0.02572   0.01978  -0.0799   0.8588   1.0000
   1.000   0.3350   0.02567   0.01963  -0.0824   0.8570   1.0000
   1.250   0.3374   0.02613   0.02004  -0.0780   0.8484   1.0000
   1.500   0.3698   0.02618   0.02003  -0.0790   0.8456   1.0000
   1.750   0.4066   0.02617   0.01997  -0.0809   0.8437   1.0000
   2.000   0.4457   0.02614   0.01991  -0.0831   0.8424   1.0000
   2.250   0.4427   0.02677   0.02053  -0.0778   0.8326   1.0000
   2.500   0.4781   0.02673   0.02048  -0.0794   0.8305   1.0000
   2.750   0.5171   0.02661   0.02036  -0.0815   0.8290   1.0000
   3.000   0.5187   0.02725   0.02100  -0.0772   0.8194   1.0000
   3.250   0.5539   0.02715   0.02092  -0.0786   0.8170   1.0000
   3.500   0.5919   0.02698   0.02079  -0.0805   0.8154   1.0000
   3.750   0.5981   0.02752   0.02136  -0.0769   0.8057   1.0000
   4.000   0.6340   0.02733   0.02123  -0.0784   0.8032   1.0000
   4.250   0.6803   0.02659   0.02057  -0.0813   0.8014   1.0000
   4.500   0.6947   0.02667   0.02069  -0.0788   0.7910   1.0000
   4.750   0.7674   0.02400   0.01815  -0.0851   0.7884   1.0000
   5.000   0.8306   0.02166   0.01595  -0.0900   0.7862   1.0000
   5.500   0.9131   0.01840   0.01294  -0.0923   0.7693   1.0000
   5.750   0.9536   0.01653   0.01117  -0.0929   0.7520   1.0000
   6.000   0.9849   0.01533   0.01004  -0.0922   0.7244   1.0000
   6.250   1.0109   0.01473   0.00943  -0.0908   0.6852   1.0000
   6.500   1.0244   0.01470   0.00864  -0.0867   0.5361   1.0000
   6.750   1.0053   0.01615   0.00943  -0.0776   0.4392   1.0000
   7.000   0.9855   0.01787   0.01056  -0.0691   0.3349   1.0000
   7.250   0.9657   0.01997   0.01193  -0.0613   0.2177   1.0000
   7.500   0.9545   0.02204   0.01330  -0.0554   0.1137   1.0000
   7.750   0.9549   0.02361   0.01462  -0.0512   0.0809   1.0000
   8.000   0.9597   0.02497   0.01594  -0.0477   0.0663   1.0000
   8.250   0.9643   0.02640   0.01738  -0.0443   0.0584   1.0000
   8.500   0.9758   0.02742   0.01843  -0.0420   0.0522   1.0000
   8.750   0.9848   0.02888   0.01987  -0.0393   0.0471   1.0000
   9.000   1.0008   0.02983   0.02091  -0.0376   0.0438   1.0000
   9.250   1.0183   0.03089   0.02199  -0.0362   0.0408   1.0000
   9.500   1.0508   0.03262   0.02369  -0.0369   0.0371   1.0000
   9.750   1.0707   0.03350   0.02473  -0.0357   0.0349   1.0000
  10.000   1.1059   0.03503   0.02642  -0.0367   0.0330   1.0000
  10.250   1.1446   0.03701   0.02859  -0.0383   0.0317   1.0000
  10.500   1.1686   0.03875   0.03042  -0.0382   0.0298   1.0000
  10.750   1.2172   0.04457   0.03669  -0.0422   0.0282   1.0000
  11.000   1.2406   0.04958   0.04213  -0.0422   0.0284   1.0000
  11.250   1.2441   0.05211   0.04494  -0.0386   0.0285   1.0000
  11.500   1.2441   0.05357   0.04665  -0.0343   0.0289   1.0000
  11.750   1.2082   0.05913   0.05316  -0.0240   0.0342   1.0000
  12.000   1.1003   0.05139   0.04582  -0.0082   0.0326   1.0000
  12.250   1.0787   0.05603   0.05076  -0.0035   0.0340   1.0000
  12.500   1.0561   0.06072   0.05567   0.0006   0.0351   1.0000
<< Back to GOE 7K AIRFOIL (goe07k-il)

Polar data table (+)

Polar graphs


<< Back to GOE 7K AIRFOIL (goe07k-il)