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

GOE 9K AIRFOIL (goe09k-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: GOE 9K AIRFOIL (goe09k-il)
Reynolds number: 100,000
Max Cl/Cd: 23.83 at α=2°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe09k-il-100000.txt
Download as CSV file: xf-goe09k-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 9K 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
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.5970   0.10392   0.09901   0.0057   1.0000   0.0462
  -8.000  -0.5974   0.10151   0.09667   0.0026   1.0000   0.0468
  -7.750  -0.5968   0.09906   0.09428  -0.0029   1.0000   0.0472
  -7.500  -0.5893   0.09598   0.09120  -0.0095   1.0000   0.0475
  -7.250  -0.5848   0.08938   0.08465  -0.0003   1.0000   0.0520
  -7.000  -0.5781   0.08567   0.08096  -0.0032   1.0000   0.0548
  -6.750  -0.5680   0.08188   0.07716  -0.0090   1.0000   0.0572
  -6.500  -0.5516   0.07834   0.07349  -0.0176   1.0000   0.0582
  -6.250  -0.5478   0.07340   0.06867  -0.0132   1.0000   0.0625
  -6.000  -0.5276   0.06964   0.06473  -0.0203   1.0000   0.0668
  -5.750  -0.5190   0.06506   0.06022  -0.0192   1.0000   0.0710
  -5.500  -0.4973   0.06105   0.05594  -0.0243   1.0000   0.0771
  -5.250  -0.4743   0.05862   0.05312  -0.0276   1.0000   0.0866
  -5.000  -0.4668   0.05312   0.04792  -0.0251   1.0000   0.0924
  -4.750  -0.4470   0.04925   0.04387  -0.0267   1.0000   0.1036
  -4.500  -0.4275   0.04569   0.04015  -0.0275   1.0000   0.1175
  -4.250  -0.4055   0.04277   0.03692  -0.0287   1.0000   0.1409
  -4.000  -0.3888   0.03911   0.03328  -0.0278   1.0000   0.1599
  -3.250  -0.3407   0.03074   0.02495  -0.0234   1.0000   0.3060
  -3.000  -0.2681   0.02658   0.01853  -0.0275   1.0000   0.1001
  -2.750  -0.2355   0.02311   0.01431  -0.0256   1.0000   0.0578
  -2.500  -0.2088   0.02035   0.01128  -0.0245   1.0000   0.0516
  -2.250  -0.1808   0.01848   0.00891  -0.0230   1.0000   0.0463
  -2.000  -0.1546   0.01653   0.00680  -0.0217   1.0000   0.0456
  -1.750  -0.1301   0.01488   0.00515  -0.0205   1.0000   0.0530
  -1.500  -0.1063   0.01354   0.00381  -0.0192   1.0000   0.0702
  -1.250  -0.0608   0.00956   0.00231  -0.0219   1.0000   1.0000
  -1.000  -0.0369   0.00955   0.00185  -0.0210   1.0000   1.0000
  -0.750  -0.0129   0.00955   0.00158  -0.0202   1.0000   1.0000
  -0.500   0.0109   0.00956   0.00140  -0.0194   1.0000   1.0000
  -0.250   0.0347   0.00959   0.00128  -0.0187   1.0000   1.0000
   0.000   0.0585   0.00962   0.00121  -0.0180   1.0000   1.0000
   0.250   0.0822   0.00967   0.00121  -0.0173   1.0000   1.0000
   0.500   0.1059   0.00972   0.00127  -0.0165   1.0000   1.0000
   0.750   0.1294   0.00979   0.00139  -0.0158   1.0000   1.0000
   1.000   0.1530   0.00987   0.00156  -0.0151   1.0000   1.0000
   1.250   0.1764   0.00996   0.00178  -0.0143   1.0000   1.0000
   1.500   0.1998   0.01007   0.00207  -0.0136   1.0000   1.0000
   1.750   0.2232   0.01019   0.00253  -0.0128   1.0000   1.0000
   2.000   0.2464   0.01034   0.00304  -0.0120   1.0000   1.0000
   2.250   0.3528   0.01573   0.00558  -0.0255   0.0474   1.0000
   2.500   0.3770   0.01731   0.00726  -0.0239   0.0435   1.0000
   2.750   0.4039   0.01928   0.00935  -0.0227   0.0433   1.0000
   3.000   0.4319   0.02145   0.01178  -0.0215   0.0443   1.0000
   3.250   0.4611   0.02387   0.01466  -0.0200   0.0521   1.0000
   3.500   0.4915   0.02722   0.01853  -0.0182   0.0752   1.0000
   4.750   0.6313   0.04437   0.03859  -0.0130   0.1614   1.0000
   5.250   0.6617   0.05148   0.04590  -0.0117   0.1170   1.0000
   5.500   0.6755   0.05558   0.05006  -0.0113   0.1021   1.0000
   5.750   0.6899   0.05932   0.05409  -0.0119   0.0899   1.0000
   6.000   0.7030   0.06355   0.05853  -0.0130   0.0810   1.0000
   6.250   0.7135   0.06818   0.06322  -0.0133   0.0753   1.0000
   6.500   0.7251   0.07220   0.06733  -0.0137   0.0686   1.0000
   6.750   0.7335   0.07685   0.07215  -0.0168   0.0661   1.0000
   7.000   0.7411   0.08114   0.07654  -0.0196   0.0619   1.0000
   7.250   0.7483   0.08675   0.08204  -0.0167   0.0586   1.0000
   7.500   0.7514   0.09053   0.08595  -0.0219   0.0581   1.0000
   7.750   0.7525   0.09461   0.09005  -0.0271   0.0572   1.0000
   8.000   0.7493   0.09858   0.09398  -0.0310   0.0559   1.0000
   8.250   0.7479   0.10254   0.09790  -0.0339   0.0544   1.0000
<< Back to GOE 9K AIRFOIL (goe09k-il)

Polar data table (+)

Polar graphs


<< Back to GOE 9K AIRFOIL (goe09k-il)