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GRUMMAN K-2 AIRFOIL (k2-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: GRUMMAN K-2 AIRFOIL (k2-il)
Reynolds number: 50,000
Max Cl/Cd: 21.1 at α=6.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-k2-il-50000-n5.txt
Download as CSV file: xf-k2-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GRUMMAN K-2 AIRFOIL                             
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.750  -0.6414   0.09633   0.08796  -0.0333   1.0000   0.0557
  -9.500  -0.6582   0.08940   0.08103  -0.0371   1.0000   0.0549
  -9.250  -0.6811   0.08301   0.07461  -0.0401   1.0000   0.0542
  -9.000  -0.7072   0.07765   0.06918  -0.0413   1.0000   0.0535
  -8.750  -0.7342   0.07341   0.06488  -0.0404   1.0000   0.0530
  -8.500  -0.7573   0.06956   0.06090  -0.0388   1.0000   0.0526
  -8.250  -0.7730   0.06574   0.05687  -0.0374   1.0000   0.0524
  -8.000  -0.7822   0.06212   0.05301  -0.0357   1.0000   0.0523
  -7.750  -0.7864   0.05859   0.04917  -0.0341   1.0000   0.0522
  -7.500  -0.7859   0.05520   0.04547  -0.0325   1.0000   0.0522
  -7.250  -0.7811   0.05201   0.04194  -0.0310   1.0000   0.0524
  -7.000  -0.7724   0.04902   0.03861  -0.0296   1.0000   0.0527
  -6.750  -0.7604   0.04622   0.03547  -0.0283   1.0000   0.0532
  -6.500  -0.7458   0.04368   0.03260  -0.0271   1.0000   0.0541
  -6.250  -0.7293   0.04138   0.02991  -0.0260   1.0000   0.0559
  -6.000  -0.7108   0.03922   0.02726  -0.0248   1.0000   0.0579
  -5.750  -0.6910   0.03731   0.02530  -0.0239   1.0000   0.0597
  -5.500  -0.6699   0.03564   0.02352  -0.0228   1.0000   0.0616
  -5.250  -0.6478   0.03408   0.02179  -0.0215   1.0000   0.0638
  -5.000  -0.6252   0.03276   0.02023  -0.0201   1.0000   0.0673
  -4.750  -0.6036   0.03150   0.01905  -0.0188   1.0000   0.0712
  -4.500  -0.5814   0.03045   0.01794  -0.0172   1.0000   0.0755
  -4.250  -0.5595   0.02945   0.01684  -0.0156   1.0000   0.0808
  -4.000  -0.5394   0.02848   0.01589  -0.0141   1.0000   0.0880
  -3.750  -0.5199   0.02745   0.01488  -0.0126   1.0000   0.0962
  -3.500  -0.5006   0.02640   0.01389  -0.0113   1.0000   0.1089
  -3.250  -0.4815   0.02519   0.01285  -0.0103   1.0000   0.1300
  -3.000  -0.4632   0.02346   0.01175  -0.0098   1.0000   0.1924
  -2.750  -0.4526   0.02105   0.01118  -0.0079   1.0000   0.4449
  -2.500  -0.4506   0.02227   0.01332   0.0024   1.0000   0.6205
  -2.000  -0.4224   0.02459   0.01543   0.0125   1.0000   0.7594
  -1.750  -0.4126   0.02588   0.01665   0.0194   1.0000   0.8047
  -1.500  -0.3972   0.02633   0.01696   0.0239   1.0000   0.8339
  -1.250  -0.3784   0.02619   0.01667   0.0263   1.0000   0.8509
  -1.000  -0.3586   0.02595   0.01627   0.0283   1.0000   0.8665
  -0.750  -0.3373   0.02563   0.01583   0.0299   1.0000   0.8813
  -0.500  -0.3154   0.02526   0.01534   0.0312   1.0000   0.8950
  -0.250  -0.2911   0.02484   0.01482   0.0319   1.0000   0.9065
   0.000  -0.2665   0.02442   0.01430   0.0322   1.0000   0.9160
   0.250  -0.2438   0.02402   0.01383   0.0327   1.0000   0.9246
   0.500  -0.2222   0.02364   0.01341   0.0334   1.0000   0.9327
   0.750  -0.1954   0.02330   0.01302   0.0330   1.0000   0.9390
   1.000  -0.1736   0.02299   0.01269   0.0334   1.0000   0.9458
   1.250  -0.1464   0.02272   0.01242   0.0329   1.0000   0.9514
   1.500  -0.1223   0.02248   0.01220   0.0328   1.0000   0.9573
   1.750  -0.0946   0.02230   0.01205   0.0320   1.0000   0.9623
   2.000  -0.0683   0.02216   0.01197   0.0315   1.0000   0.9673
   2.250  -0.0395   0.02207   0.01195   0.0304   1.0000   0.9719
   2.500  -0.0126   0.02203   0.01200   0.0296   1.0000   0.9766
   2.750   0.0158   0.02204   0.01214   0.0284   1.0000   0.9812
   3.000   0.0424   0.02209   0.01233   0.0275   1.0000   0.9860
   3.250   0.0922   0.02244   0.01288   0.0223   0.9885   0.9872
   3.500   0.1699   0.02267   0.01339   0.0127   0.9539   0.9848
   3.750   0.2369   0.02211   0.01315   0.0062   0.9053   0.9831
   4.000   0.2805   0.02137   0.01268   0.0045   0.8467   0.9843
   4.250   0.3952   0.02051   0.01118  -0.0067   0.5447   0.9763
   4.500   0.4188   0.02182   0.01145  -0.0057   0.3690   0.9786
   4.750   0.4422   0.02291   0.01204  -0.0054   0.2834   0.9817
   5.000   0.4676   0.02388   0.01270  -0.0054   0.2307   0.9848
   5.250   0.4972   0.02476   0.01339  -0.0064   0.1836   0.9873
   5.500   0.5268   0.02571   0.01417  -0.0073   0.1466   0.9901
   5.750   0.5570   0.02679   0.01520  -0.0081   0.1158   0.9930
   6.000   0.5874   0.02797   0.01641  -0.0089   0.0924   0.9957
   6.250   0.6157   0.02918   0.01764  -0.0092   0.0782   0.9988
   6.500   0.6374   0.03041   0.01898  -0.0082   0.0693   1.0000
   6.750   0.6549   0.03165   0.02030  -0.0066   0.0640   1.0000
   7.000   0.6716   0.03292   0.02176  -0.0047   0.0591   1.0000
   7.250   0.6865   0.03427   0.02324  -0.0027   0.0555   1.0000
   7.500   0.7010   0.03592   0.02518  -0.0005   0.0525   1.0000
   7.750   0.7145   0.03756   0.02703   0.0016   0.0508   1.0000
   8.000   0.7278   0.03931   0.02891   0.0036   0.0493   1.0000
   8.250   0.7386   0.04199   0.03215   0.0059   0.0477   1.0000
   8.500   0.7486   0.04460   0.03519   0.0079   0.0460   1.0000
   8.750   0.7596   0.04693   0.03781   0.0094   0.0444   1.0000
   9.000   0.7732   0.04888   0.03988   0.0103   0.0430   1.0000
   9.250   0.7774   0.05250   0.04396   0.0119   0.0424   1.0000
   9.500   0.7777   0.05654   0.04847   0.0133   0.0422   1.0000
   9.750   0.7739   0.06084   0.05318   0.0144   0.0422   1.0000
  10.000   0.7663   0.06531   0.05799   0.0150   0.0422   1.0000
  10.250   0.7546   0.06977   0.06270   0.0152   0.0424   1.0000
  10.500   0.7408   0.07457   0.06771   0.0144   0.0425   1.0000
  10.750   0.7272   0.07989   0.07317   0.0121   0.0428   1.0000
  11.000   0.6824   0.09510   0.08862  -0.0014   0.0448   1.0000
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