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

NACA 0010-65 (naca001065-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: NACA 0010-65 (naca001065-il)
Reynolds number: 100,000
Max Cl/Cd: 28.28 at α=4.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-naca001065-il-100000-n5.txt
Download as CSV file: xf-naca001065-il-100000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 0010-65                                    
  
 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
  ------ -------- --------- --------- -------- -------- --------
 -11.250  -0.6699   0.08957   0.08420  -0.0384   1.0000   0.0345
 -11.000  -0.6927   0.08061   0.07523  -0.0442   1.0000   0.0341
 -10.750  -0.7193   0.07409   0.06864  -0.0462   1.0000   0.0338
 -10.500  -0.7468   0.06921   0.06369  -0.0451   1.0000   0.0336
 -10.250  -0.7742   0.06542   0.05980  -0.0415   1.0000   0.0335
 -10.000  -0.8012   0.06234   0.05659  -0.0361   1.0000   0.0335
  -9.750  -0.8244   0.05879   0.05284  -0.0308   1.0000   0.0337
  -9.500  -0.8446   0.05513   0.04890  -0.0254   1.0000   0.0341
  -9.250  -0.8621   0.05142   0.04479  -0.0197   1.0000   0.0349
  -9.000  -0.8758   0.04793   0.04078  -0.0140   1.0000   0.0358
  -8.750  -0.8787   0.04511   0.03772  -0.0099   1.0000   0.0367
  -8.500  -0.8737   0.04329   0.03577  -0.0068   1.0000   0.0378
  -8.250  -0.8685   0.04126   0.03347  -0.0035   1.0000   0.0388
  -8.000  -0.8608   0.03943   0.03139  -0.0004   1.0000   0.0402
  -7.750  -0.8520   0.03763   0.02928   0.0026   1.0000   0.0423
  -7.500  -0.8423   0.03550   0.02669   0.0058   1.0000   0.0443
  -7.250  -0.8294   0.03342   0.02413   0.0085   1.0000   0.0457
  -7.000  -0.8140   0.03157   0.02204   0.0105   1.0000   0.0474
  -6.750  -0.7974   0.03052   0.02092   0.0122   1.0000   0.0496
  -6.500  -0.7793   0.02935   0.01954   0.0138   1.0000   0.0517
  -6.250  -0.7591   0.02801   0.01797   0.0153   1.0000   0.0534
  -6.000  -0.7382   0.02684   0.01658   0.0166   1.0000   0.0553
  -5.750  -0.7174   0.02600   0.01548   0.0179   1.0000   0.0578
  -5.500  -0.6964   0.02483   0.01432   0.0190   1.0000   0.0602
  -5.250  -0.6755   0.02395   0.01339   0.0202   1.0000   0.0621
  -5.000  -0.6549   0.02317   0.01254   0.0215   1.0000   0.0644
  -4.750  -0.6346   0.02244   0.01173   0.0229   1.0000   0.0670
  -4.500  -0.6147   0.02180   0.01099   0.0243   1.0000   0.0699
  -4.250  -0.5961   0.02111   0.01030   0.0259   1.0000   0.0739
  -4.000  -0.5769   0.02060   0.00978   0.0274   1.0000   0.0804
  -3.750  -0.5579   0.02004   0.00922   0.0290   1.0000   0.0877
  -3.500  -0.5386   0.01953   0.00869   0.0305   1.0000   0.0982
  -3.250  -0.5192   0.01901   0.00825   0.0320   1.0000   0.1144
  -3.000  -0.4994   0.01853   0.00790   0.0334   1.0000   0.1417
  -2.750  -0.4796   0.01802   0.00761   0.0346   1.0000   0.1865
  -2.500  -0.4607   0.01726   0.00735   0.0359   1.0000   0.2829
  -2.250  -0.3998   0.01650   0.00919   0.0321   1.0000   0.8756
  -2.000  -0.3602   0.01733   0.00988   0.0308   1.0000   0.9192
  -1.750  -0.2905   0.01827   0.01061   0.0230   1.0000   0.9437
  -1.500  -0.1889   0.01927   0.01138   0.0087   1.0000   0.9738
  -1.250  -0.1201   0.01941   0.01139   0.0000   1.0000   0.9867
  -1.000  -0.0843   0.01933   0.01124  -0.0023   1.0000   0.9919
  -0.750  -0.0524   0.01926   0.01113  -0.0039   1.0000   0.9962
  -0.500  -0.0212   0.01921   0.01104  -0.0054   1.0000   1.0000
  -0.250  -0.0106   0.01918   0.01100  -0.0027   1.0000   1.0000
   0.000   0.0000   0.01917   0.01099   0.0000   1.0000   1.0000
   0.250   0.0106   0.01918   0.01100   0.0027   1.0000   1.0000
   0.500   0.0212   0.01921   0.01104   0.0054   1.0000   1.0000
   0.750   0.0524   0.01926   0.01112   0.0040   0.9962   1.0000
   1.000   0.0843   0.01932   0.01123   0.0024   0.9919   1.0000
   1.250   0.1199   0.01940   0.01138   0.0000   0.9867   1.0000
   1.500   0.1889   0.01927   0.01138  -0.0087   0.9737   1.0000
   1.750   0.2904   0.01827   0.01061  -0.0230   0.9438   1.0000
   2.000   0.3602   0.01732   0.00988  -0.0308   0.9193   1.0000
   2.250   0.3999   0.01649   0.00919  -0.0321   0.8758   1.0000
   2.500   0.4606   0.01725   0.00735  -0.0359   0.2834   1.0000
   2.750   0.4795   0.01802   0.00761  -0.0346   0.1869   1.0000
   3.000   0.4993   0.01852   0.00789  -0.0333   0.1418   1.0000
   3.250   0.5191   0.01901   0.00824  -0.0320   0.1145   1.0000
   3.500   0.5385   0.01953   0.00869  -0.0305   0.0983   1.0000
   3.750   0.5578   0.02004   0.00921  -0.0290   0.0878   1.0000
   4.000   0.5767   0.02060   0.00978  -0.0274   0.0804   1.0000
   4.250   0.5959   0.02111   0.01029  -0.0259   0.0740   1.0000
   4.500   0.6146   0.02180   0.01098  -0.0243   0.0699   1.0000
   4.750   0.6345   0.02244   0.01173  -0.0228   0.0670   1.0000
   5.000   0.6548   0.02316   0.01253  -0.0214   0.0644   1.0000
   5.250   0.6754   0.02395   0.01338  -0.0201   0.0621   1.0000
   5.500   0.6963   0.02482   0.01431  -0.0189   0.0602   1.0000
   5.750   0.7173   0.02600   0.01547  -0.0179   0.0578   1.0000
   6.000   0.7381   0.02684   0.01657  -0.0165   0.0553   1.0000
   6.250   0.7590   0.02801   0.01797  -0.0152   0.0534   1.0000
   6.500   0.7791   0.02934   0.01953  -0.0138   0.0517   1.0000
   6.750   0.7972   0.03051   0.02092  -0.0121   0.0496   1.0000
   7.000   0.8139   0.03157   0.02203  -0.0105   0.0474   1.0000
   7.250   0.8293   0.03342   0.02412  -0.0085   0.0457   1.0000
   7.500   0.8422   0.03549   0.02668  -0.0057   0.0443   1.0000
   7.750   0.8519   0.03762   0.02928  -0.0026   0.0423   1.0000
   8.000   0.8607   0.03942   0.03139   0.0004   0.0402   1.0000
   8.250   0.8684   0.04125   0.03346   0.0035   0.0388   1.0000
   8.500   0.8737   0.04328   0.03576   0.0068   0.0378   1.0000
   8.750   0.8786   0.04512   0.03774   0.0099   0.0367   1.0000
   9.000   0.8759   0.04792   0.04077   0.0140   0.0358   1.0000
   9.250   0.8622   0.05140   0.04477   0.0197   0.0349   1.0000
   9.500   0.8447   0.05512   0.04889   0.0254   0.0341   1.0000
   9.750   0.8245   0.05879   0.05284   0.0308   0.0337   1.0000
  10.000   0.8013   0.06235   0.05660   0.0361   0.0335   1.0000
  10.250   0.7743   0.06544   0.05981   0.0415   0.0335   1.0000
  10.500   0.7469   0.06924   0.06372   0.0450   0.0336   1.0000
  10.750   0.7196   0.07414   0.06869   0.0461   0.0338   1.0000
  11.000   0.6931   0.08067   0.07529   0.0441   0.0341   1.0000
  11.250   0.6704   0.08969   0.08431   0.0382   0.0345   1.0000
<< Back to NACA 0010-65 (naca001065-il)

Polar data table (+)

Polar graphs


<< Back to NACA 0010-65 (naca001065-il)