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NACA 0010-65 (naca001065-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: NACA 0010-65 (naca001065-il)
Reynolds number: 50,000
Max Cl/Cd: 20.4 at α=4.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-naca001065-il-50000.txt
Download as CSV file: xf-naca001065-il-50000.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 =     0.050 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.000  -0.5652   0.11425   0.10659  -0.0044   1.0000   0.3264
  -9.750  -0.5636   0.11101   0.10337  -0.0035   1.0000   0.3420
  -9.500  -0.5706   0.10836   0.10077  -0.0024   1.0000   0.3595
  -9.250  -0.5645   0.10479   0.09723  -0.0010   1.0000   0.3782
  -9.000  -0.5590   0.10213   0.09457   0.0013   1.0000   0.4049
  -8.750  -0.5390   0.09839   0.09080   0.0032   1.0000   0.4306
  -8.000  -0.5152   0.08923   0.08168   0.0088   1.0000   0.5043
  -7.750  -0.5045   0.08630   0.07874   0.0109   1.0000   0.5334
  -7.250  -0.7585   0.06427   0.05667  -0.0056   1.0000   0.2033
  -7.000  -0.7670   0.05849   0.05030  -0.0029   1.0000   0.1724
  -6.750  -0.7685   0.05421   0.04546   0.0007   1.0000   0.1598
  -6.500  -0.7625   0.05073   0.04165   0.0037   1.0000   0.1570
  -6.250  -0.7554   0.04747   0.03795   0.0068   1.0000   0.1549
  -6.000  -0.7467   0.04427   0.03411   0.0101   1.0000   0.1514
  -5.750  -0.7341   0.04176   0.03085   0.0134   1.0000   0.1488
  -5.500  -0.7164   0.03916   0.02782   0.0155   1.0000   0.1490
  -5.250  -0.6957   0.03662   0.02529   0.0165   1.0000   0.1540
  -5.000  -0.6748   0.03466   0.02300   0.0181   1.0000   0.1589
  -4.750  -0.6509   0.03272   0.02061   0.0195   1.0000   0.1624
  -4.500  -0.6235   0.03072   0.01846   0.0199   1.0000   0.1680
  -4.250  -0.5961   0.02923   0.01676   0.0205   1.0000   0.1800
  -4.000  -0.1965   0.02624   0.01609  -0.0360   1.0000   1.0000
  -3.750  -0.1840   0.02578   0.01542  -0.0342   1.0000   1.0000
  -3.500  -0.1714   0.02538   0.01482  -0.0324   1.0000   1.0000
  -3.250  -0.1587   0.02502   0.01430  -0.0304   1.0000   1.0000
  -3.000  -0.1461   0.02471   0.01384  -0.0284   1.0000   1.0000
  -2.750  -0.1335   0.02443   0.01342  -0.0263   1.0000   1.0000
  -2.500  -0.1210   0.02418   0.01306  -0.0241   1.0000   1.0000
  -2.250  -0.1086   0.02397   0.01275  -0.0219   1.0000   1.0000
  -2.000  -0.0963   0.02378   0.01247  -0.0196   1.0000   1.0000
  -1.750  -0.0840   0.02362   0.01223  -0.0172   1.0000   1.0000
  -1.500  -0.0719   0.02349   0.01202  -0.0148   1.0000   1.0000
  -1.250  -0.0598   0.02337   0.01185  -0.0124   1.0000   1.0000
  -1.000  -0.0477   0.02328   0.01172  -0.0100   1.0000   1.0000
  -0.750  -0.0358   0.02321   0.01162  -0.0075   1.0000   1.0000
  -0.500  -0.0238   0.02316   0.01154  -0.0050   1.0000   1.0000
  -0.250  -0.0119   0.02313   0.01149  -0.0025   1.0000   1.0000
   0.000   0.0000   0.02312   0.01148   0.0000   1.0000   1.0000
   0.250   0.0119   0.02313   0.01149   0.0025   1.0000   1.0000
   0.500   0.0238   0.02316   0.01154   0.0050   1.0000   1.0000
   0.750   0.0358   0.02321   0.01161   0.0075   1.0000   1.0000
   1.000   0.0478   0.02328   0.01172   0.0100   1.0000   1.0000
   1.250   0.0598   0.02337   0.01185   0.0124   1.0000   1.0000
   1.500   0.0719   0.02348   0.01202   0.0148   1.0000   1.0000
   1.750   0.0840   0.02362   0.01223   0.0172   1.0000   1.0000
   2.000   0.0963   0.02378   0.01246   0.0196   1.0000   1.0000
   2.250   0.1086   0.02396   0.01274   0.0219   1.0000   1.0000
   2.500   0.1210   0.02417   0.01305   0.0241   1.0000   1.0000
   2.750   0.1335   0.02442   0.01341   0.0263   1.0000   1.0000
   3.000   0.1461   0.02469   0.01383   0.0284   1.0000   1.0000
   3.250   0.1588   0.02501   0.01429   0.0304   1.0000   1.0000
   3.500   0.1714   0.02536   0.01481   0.0324   1.0000   1.0000
   3.750   0.1841   0.02576   0.01540   0.0342   1.0000   1.0000
   4.000   0.1966   0.02622   0.01607   0.0360   1.0000   1.0000
   4.250   0.5960   0.02922   0.01676  -0.0205   0.1800   1.0000
   4.500   0.6234   0.03072   0.01845  -0.0199   0.1680   1.0000
   4.750   0.6509   0.03272   0.02060  -0.0195   0.1624   1.0000
   5.000   0.6747   0.03465   0.02300  -0.0181   0.1589   1.0000
   5.250   0.6956   0.03662   0.02529  -0.0165   0.1540   1.0000
   5.500   0.7164   0.03915   0.02781  -0.0155   0.1490   1.0000
   5.750   0.7340   0.04175   0.03085  -0.0134   0.1488   1.0000
   6.000   0.7466   0.04426   0.03410  -0.0101   0.1514   1.0000
   6.250   0.7553   0.04747   0.03795  -0.0068   0.1549   1.0000
   6.500   0.7625   0.05073   0.04164  -0.0037   0.1570   1.0000
   6.750   0.7685   0.05420   0.04545  -0.0007   0.1598   1.0000
   7.000   0.7670   0.05848   0.05029   0.0029   0.1724   1.0000
   8.000   0.5158   0.08922   0.08167  -0.0089   0.5043   1.0000
   8.250   0.5388   0.09317   0.08567  -0.0081   0.4834   1.0000
   8.750   0.5396   0.09838   0.09079  -0.0032   0.4305   1.0000
   9.000   0.5587   0.10207   0.09451  -0.0014   0.4048   1.0000
   9.250   0.5681   0.10494   0.09739   0.0009   0.3777   1.0000
   9.500   0.5703   0.10830   0.10070   0.0023   0.3593   1.0000
   9.750   0.5639   0.11098   0.10334   0.0034   0.3417   1.0000
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