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NACA 16-015 (naca16015-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: NACA 16-015 (naca16015-il)
Reynolds number: 100,000
Max Cl/Cd: 29.9 at α=4.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-naca16015-il-100000.txt
Download as CSV file: xf-naca16015-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 16-015                                     
  
 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
  ------ -------- --------- --------- -------- -------- --------
 -11.000  -0.6344   0.10508   0.10012  -0.0361   1.0000   0.1733
 -10.750  -0.6632   0.09945   0.09456  -0.0375   1.0000   0.1778
  -9.250  -1.0134   0.06385   0.05699   0.0071   1.0000   0.0795
  -9.000  -1.0127   0.05949   0.05251   0.0105   1.0000   0.0780
  -8.750  -1.0157   0.05581   0.04856   0.0151   1.0000   0.0763
  -8.500  -1.0189   0.05225   0.04463   0.0202   1.0000   0.0745
  -8.250  -1.0193   0.04883   0.04077   0.0252   1.0000   0.0731
  -8.000  -1.0147   0.04621   0.03776   0.0293   1.0000   0.0736
  -7.750  -1.0070   0.04402   0.03520   0.0330   1.0000   0.0753
  -7.500  -0.9959   0.04176   0.03253   0.0362   1.0000   0.0766
  -7.250  -0.9804   0.03955   0.02990   0.0388   1.0000   0.0776
  -7.000  -0.9654   0.03803   0.02794   0.0415   1.0000   0.0796
  -6.750  -0.9404   0.03560   0.02539   0.0416   1.0000   0.0832
  -6.500  -0.9161   0.03411   0.02378   0.0422   1.0000   0.0865
  -6.250  -0.8918   0.03288   0.02236   0.0429   1.0000   0.0908
  -6.000  -0.8635   0.03144   0.02080   0.0427   1.0000   0.0958
  -5.750  -0.8357   0.03021   0.01962   0.0424   1.0000   0.1013
  -5.500  -0.8137   0.02945   0.01870   0.0435   1.0000   0.1080
  -5.250  -0.7904   0.02821   0.01767   0.0440   1.0000   0.1171
  -5.000  -0.7732   0.02721   0.01674   0.0459   1.0000   0.1267
  -4.750  -0.7611   0.02628   0.01591   0.0487   1.0000   0.1422
  -4.500  -0.7526   0.02518   0.01505   0.0523   1.0000   0.1711
  -4.250  -0.6659   0.02582   0.01943   0.0464   1.0000   0.8225
  -4.000  -0.5786   0.03037   0.02358   0.0393   1.0000   0.8639
  -3.750  -0.4742   0.03445   0.02725   0.0280   1.0000   0.9002
  -3.500  -0.3649   0.03670   0.02916   0.0137   1.0000   0.9263
  -3.250  -0.2754   0.03760   0.02982   0.0017   1.0000   0.9532
  -3.000  -0.2256   0.03759   0.02963  -0.0037   1.0000   0.9697
  -2.750  -0.1742   0.03736   0.02925  -0.0099   1.0000   0.9837
  -2.500  -0.1168   0.03684   0.02858  -0.0175   1.0000   0.9954
  -2.250  -0.0860   0.03647   0.02813  -0.0199   1.0000   1.0000
  -2.000  -0.0768   0.03621   0.02783  -0.0177   1.0000   1.0000
  -1.750  -0.0675   0.03599   0.02757  -0.0155   1.0000   1.0000
  -1.500  -0.0581   0.03580   0.02734  -0.0133   1.0000   1.0000
  -1.250  -0.0486   0.03564   0.02716  -0.0111   1.0000   1.0000
  -1.000  -0.0389   0.03551   0.02701  -0.0089   1.0000   1.0000
  -0.750  -0.0293   0.03541   0.02689  -0.0067   1.0000   1.0000
  -0.500  -0.0195   0.03534   0.02680  -0.0044   1.0000   1.0000
  -0.250  -0.0098   0.03530   0.02675  -0.0022   1.0000   1.0000
   0.000   0.0000   0.03528   0.02674   0.0000   1.0000   1.0000
   0.250   0.0098   0.03529   0.02675   0.0022   1.0000   1.0000
   0.500   0.0195   0.03533   0.02680   0.0044   1.0000   1.0000
   0.750   0.0293   0.03540   0.02688   0.0067   1.0000   1.0000
   1.000   0.0390   0.03550   0.02699   0.0089   1.0000   1.0000
   1.250   0.0486   0.03563   0.02714   0.0111   1.0000   1.0000
   1.500   0.0581   0.03578   0.02733   0.0133   1.0000   1.0000
   1.750   0.0675   0.03597   0.02755   0.0155   1.0000   1.0000
   2.000   0.0768   0.03619   0.02781   0.0177   1.0000   1.0000
   2.250   0.0860   0.03645   0.02811   0.0199   1.0000   1.0000
   2.500   0.1166   0.03681   0.02855   0.0176   0.9955   1.0000
   2.750   0.1739   0.03733   0.02922   0.0100   0.9838   1.0000
   3.000   0.2256   0.03757   0.02961   0.0038   0.9698   1.0000
   3.250   0.2753   0.03758   0.02979  -0.0017   0.9533   1.0000
   3.500   0.3650   0.03667   0.02913  -0.0137   0.9264   1.0000
   3.750   0.4750   0.03441   0.02722  -0.0281   0.9003   1.0000
   4.000   0.5788   0.03035   0.02355  -0.0394   0.8639   1.0000
   4.250   0.6661   0.02580   0.01941  -0.0465   0.8226   1.0000
   4.500   0.7525   0.02517   0.01504  -0.0523   0.1711   1.0000
   4.750   0.7610   0.02627   0.01590  -0.0487   0.1421   1.0000
   5.000   0.7731   0.02720   0.01673  -0.0458   0.1267   1.0000
   5.250   0.7903   0.02820   0.01766  -0.0439   0.1170   1.0000
   5.500   0.8135   0.02943   0.01868  -0.0435   0.1081   1.0000
   5.750   0.8356   0.03020   0.01961  -0.0424   0.1013   1.0000
   6.000   0.8634   0.03143   0.02079  -0.0427   0.0958   1.0000
   6.250   0.8918   0.03287   0.02235  -0.0429   0.0908   1.0000
   6.500   0.9159   0.03410   0.02377  -0.0421   0.0865   1.0000
   6.750   0.9402   0.03560   0.02538  -0.0416   0.0832   1.0000
   7.000   0.9652   0.03802   0.02793  -0.0415   0.0796   1.0000
   7.250   0.9802   0.03953   0.02989  -0.0388   0.0777   1.0000
   7.500   0.9957   0.04174   0.03252  -0.0362   0.0766   1.0000
   7.750   1.0069   0.04400   0.03519  -0.0330   0.0754   1.0000
   8.000   1.0146   0.04620   0.03776  -0.0293   0.0737   1.0000
   8.250   1.0192   0.04882   0.04076  -0.0252   0.0731   1.0000
   8.500   1.0188   0.05223   0.04461  -0.0202   0.0745   1.0000
   8.750   1.0157   0.05579   0.04854  -0.0151   0.0763   1.0000
   9.000   1.0128   0.05947   0.05249  -0.0105   0.0780   1.0000
   9.250   1.0136   0.06387   0.05701  -0.0071   0.0795   1.0000
   9.500   0.9075   0.06106   0.05538   0.0100   0.0924   1.0000
  10.250   0.6676   0.08842   0.08397   0.0359   0.1970   1.0000
  10.500   0.6152   0.09445   0.08998   0.0369   0.1969   1.0000
  10.750   0.5186   0.10441   0.09976   0.0315   0.1892   1.0000
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