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

GOE 419 AIRFOIL (goe419-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: GOE 419 AIRFOIL (goe419-il)
Reynolds number: 100,000
Max Cl/Cd: 59.99 at α=3.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe419-il-100000.txt
Download as CSV file: xf-goe419-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 419 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.750  -0.3702   0.09513   0.09040  -0.0233   1.0000   0.0472
  -7.500  -0.3720   0.09294   0.08830  -0.0232   1.0000   0.0483
  -7.250  -0.3721   0.09062   0.08605  -0.0240   1.0000   0.0495
  -7.000  -0.3694   0.08843   0.08393  -0.0263   1.0000   0.0507
  -6.750  -0.3631   0.08692   0.08246  -0.0317   1.0000   0.0517
  -6.500  -0.3523   0.08514   0.08066  -0.0367   1.0000   0.0521
  -6.250  -0.3447   0.08165   0.07720  -0.0383   1.0000   0.0525
  -6.000  -0.3492   0.07625   0.07193  -0.0321   1.0000   0.0539
  -5.750  -0.3465   0.07315   0.06888  -0.0300   1.0000   0.0559
  -5.500  -0.3409   0.07042   0.06616  -0.0298   1.0000   0.0583
  -5.250  -0.3306   0.06779   0.06343  -0.0315   1.0000   0.0624
  -5.000  -0.2990   0.06635   0.06166  -0.0402   1.0000   0.0656
  -4.750  -0.3085   0.06154   0.05712  -0.0341   1.0000   0.0686
  -4.500  -0.2962   0.05891   0.05444  -0.0344   1.0000   0.0752
  -4.250  -0.2722   0.05561   0.05093  -0.0384   1.0000   0.0803
  -4.000  -0.2639   0.05259   0.04795  -0.0368   1.0000   0.0844
  -3.750  -0.2356   0.04979   0.04485  -0.0404   1.0000   0.0935
  -3.500  -0.2250   0.04680   0.04191  -0.0391   1.0000   0.0991
  -3.250  -0.2019   0.04395   0.03887  -0.0407   1.0000   0.1090
  -3.000  -0.1792   0.04155   0.03626  -0.0418   1.0000   0.1213
  -2.750  -0.1588   0.03916   0.03374  -0.0421   1.0000   0.1351
  -2.500  -0.1384   0.03689   0.03136  -0.0423   1.0000   0.1500
  -2.250  -0.1169   0.03493   0.02917  -0.0426   1.0000   0.1756
  -1.750  -0.0183   0.02776   0.02053  -0.0489   0.9931   0.1024
  -1.500   0.0327   0.02424   0.01593  -0.0512   0.9870   0.0776
  -1.250   0.0784   0.02228   0.01364  -0.0546   0.9797   0.0866
  -1.000   0.1230   0.02078   0.01179  -0.0574   0.9719   0.0980
  -0.750   0.1638   0.01985   0.01072  -0.0599   0.9631   0.1156
  -0.500   0.2082   0.01893   0.00979  -0.0631   0.9560   0.1344
  -0.250   0.2477   0.01796   0.00901  -0.0653   0.9469   0.1944
   0.000   0.3069   0.01571   0.00840  -0.0715   0.9439   1.0000
   0.250   0.3482   0.01576   0.00821  -0.0743   0.9329   1.0000
   0.500   0.3888   0.01579   0.00810  -0.0770   0.9219   1.0000
   0.750   0.4312   0.01576   0.00796  -0.0799   0.9116   1.0000
   1.000   0.4837   0.01552   0.00766  -0.0846   0.9051   1.0000
   1.250   0.5245   0.01543   0.00755  -0.0870   0.8937   1.0000
   1.500   0.5649   0.01531   0.00745  -0.0892   0.8821   1.0000
   1.750   0.6044   0.01519   0.00743  -0.0912   0.8701   1.0000
   2.000   0.6422   0.01508   0.00739  -0.0927   0.8571   1.0000
   2.250   0.6773   0.01501   0.00740  -0.0937   0.8429   1.0000
   2.500   0.7103   0.01497   0.00747  -0.0941   0.8277   1.0000
   2.750   0.7424   0.01492   0.00755  -0.0942   0.8119   1.0000
   3.000   0.7632   0.01371   0.00632  -0.0890   0.7521   1.0000
   3.250   0.7844   0.01338   0.00595  -0.0859   0.7038   1.0000
   3.500   0.8020   0.01337   0.00562  -0.0820   0.6253   1.0000
   3.750   0.8108   0.01408   0.00554  -0.0770   0.4715   1.0000
   4.000   0.8160   0.01593   0.00603  -0.0727   0.2470   1.0000
   4.250   0.8229   0.01853   0.00737  -0.0695   0.0566   1.0000
   4.500   0.8439   0.01940   0.00845  -0.0679   0.0516   1.0000
   4.750   0.8630   0.02045   0.00972  -0.0659   0.0490   1.0000
   5.000   0.8795   0.02169   0.01116  -0.0636   0.0476   1.0000
   5.250   0.8902   0.02359   0.01320  -0.0605   0.0446   1.0000
   5.500   0.9095   0.02498   0.01466  -0.0587   0.0422   1.0000
   5.750   0.9345   0.02707   0.01670  -0.0576   0.0432   1.0000
   6.000   0.9687   0.03018   0.01973  -0.0580   0.0451   1.0000
   6.250   0.9981   0.03106   0.02105  -0.0564   0.0500   1.0000
   6.500   1.0357   0.03600   0.02606  -0.0571   0.0568   1.0000
   6.750   1.0677   0.03949   0.03005  -0.0557   0.0688   1.0000
   7.000   1.0580   0.02983   0.02236  -0.0465   0.1031   1.0000
   9.750   0.9234   0.11123   0.10673  -0.0579   0.2287   1.0000
  11.750   0.7723   0.14366   0.13939  -0.0590   0.1650   1.0000
  12.000   0.7963   0.14818   0.14394  -0.0568   0.1574   1.0000
<< Back to GOE 419 AIRFOIL (goe419-il)

Polar data table (+)

Polar graphs


<< Back to GOE 419 AIRFOIL (goe419-il)