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

S2060 8% (s2060-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: S2060 8% (s2060-il)
Reynolds number: 500,000
Max Cl/Cd: 90.99 at α=3°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-s2060-il-500000.txt
Download as CSV file: xf-s2060-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: S2060 8%                                        
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.6355   0.03789   0.03504  -0.0512   1.0000   0.0128
  -8.000  -0.6355   0.03381   0.03060  -0.0500   1.0000   0.0127
  -7.750  -0.6250   0.03171   0.02825  -0.0484   1.0000   0.0124
  -7.500  -0.6245   0.02664   0.02274  -0.0466   1.0000   0.0120
  -7.250  -0.6154   0.02297   0.01858  -0.0449   1.0000   0.0118
  -7.000  -0.6009   0.02033   0.01553  -0.0433   1.0000   0.0118
  -6.750  -0.5832   0.01840   0.01329  -0.0420   1.0000   0.0119
  -6.500  -0.5637   0.01699   0.01163  -0.0408   1.0000   0.0122
  -6.250  -0.5399   0.01529   0.00964  -0.0404   0.9995   0.0126
  -6.000  -0.5077   0.01380   0.00800  -0.0418   0.9976   0.0138
  -5.750  -0.4733   0.01311   0.00722  -0.0434   0.9955   0.0152
  -5.500  -0.4389   0.01247   0.00648  -0.0449   0.9936   0.0168
  -5.250  -0.4061   0.01176   0.00572  -0.0463   0.9908   0.0208
  -5.000  -0.3717   0.01117   0.00512  -0.0477   0.9881   0.0293
  -4.750  -0.3360   0.01079   0.00474  -0.0495   0.9858   0.0388
  -4.500  -0.2993   0.01053   0.00444  -0.0515   0.9840   0.0461
  -4.250  -0.2666   0.01031   0.00422  -0.0527   0.9804   0.0537
  -4.000  -0.2328   0.00997   0.00390  -0.0541   0.9768   0.0630
  -3.750  -0.1969   0.00963   0.00355  -0.0558   0.9741   0.0731
  -3.500  -0.1597   0.00929   0.00325  -0.0579   0.9721   0.0875
  -3.250  -0.1215   0.00895   0.00297  -0.0602   0.9706   0.1113
  -3.000  -0.0913   0.00859   0.00275  -0.0608   0.9651   0.1444
  -2.750  -0.0563   0.00821   0.00255  -0.0625   0.9616   0.1952
  -2.500  -0.0210   0.00776   0.00235  -0.0642   0.9585   0.2668
  -2.250   0.0106   0.00733   0.00219  -0.0651   0.9536   0.3486
  -2.000   0.0397   0.00686   0.00206  -0.0655   0.9471   0.4484
  -1.750   0.0705   0.00636   0.00194  -0.0661   0.9423   0.5638
  -1.500   0.0956   0.00606   0.00191  -0.0653   0.9331   0.6526
  -1.250   0.1232   0.00585   0.00185  -0.0649   0.9255   0.7126
  -1.000   0.1491   0.00568   0.00181  -0.0641   0.9161   0.7585
  -0.750   0.1744   0.00555   0.00177  -0.0632   0.9057   0.7987
  -0.500   0.1995   0.00544   0.00174  -0.0621   0.8953   0.8343
  -0.250   0.2241   0.00536   0.00170  -0.0609   0.8844   0.8688
   0.000   0.2476   0.00529   0.00167  -0.0594   0.8722   0.9023
   0.250   0.2728   0.00523   0.00164  -0.0583   0.8592   0.9360
   0.500   0.3065   0.00520   0.00159  -0.0592   0.8459   0.9656
   0.750   0.3461   0.00519   0.00153  -0.0616   0.8314   0.9869
   1.000   0.3824   0.00520   0.00148  -0.0633   0.8149   1.0000
   1.250   0.4066   0.00526   0.00147  -0.0624   0.7959   1.0000
   1.500   0.4315   0.00534   0.00147  -0.0617   0.7747   1.0000
   1.750   0.4566   0.00544   0.00149  -0.0610   0.7515   1.0000
   2.000   0.4818   0.00557   0.00152  -0.0603   0.7250   1.0000
   2.250   0.5068   0.00573   0.00157  -0.0596   0.6945   1.0000
   2.500   0.5317   0.00592   0.00163  -0.0588   0.6607   1.0000
   2.750   0.5567   0.00614   0.00171  -0.0582   0.6261   1.0000
   3.000   0.5814   0.00639   0.00184  -0.0575   0.5878   1.0000
   3.250   0.6052   0.00673   0.00197  -0.0566   0.5346   1.0000
   3.500   0.6280   0.00720   0.00213  -0.0557   0.4659   1.0000
   3.750   0.6510   0.00770   0.00233  -0.0549   0.3983   1.0000
   4.000   0.6752   0.00813   0.00257  -0.0543   0.3467   1.0000
   4.250   0.6994   0.00858   0.00281  -0.0538   0.2967   1.0000
   4.500   0.7231   0.00912   0.00309  -0.0532   0.2401   1.0000
   4.750   0.7465   0.00971   0.00341  -0.0526   0.1819   1.0000
   5.000   0.7698   0.01033   0.00379  -0.0520   0.1302   1.0000
   5.250   0.7915   0.01120   0.00430  -0.0512   0.0636   1.0000
   5.500   0.8134   0.01210   0.00497  -0.0502   0.0273   1.0000
   5.750   0.8376   0.01270   0.00562  -0.0494   0.0232   1.0000
   6.000   0.8622   0.01319   0.00618  -0.0488   0.0214   1.0000
   6.250   0.8861   0.01375   0.00679  -0.0482   0.0197   1.0000
   6.500   0.9089   0.01448   0.00758  -0.0473   0.0187   1.0000
   6.750   0.9298   0.01548   0.00866  -0.0461   0.0178   1.0000
   7.000   0.9477   0.01700   0.01032  -0.0445   0.0171   1.0000
   7.250   0.9698   0.01789   0.01130  -0.0435   0.0168   1.0000
   7.500   0.9914   0.01897   0.01249  -0.0425   0.0166   1.0000
   7.750   1.0134   0.02001   0.01364  -0.0415   0.0161   1.0000
   8.000   1.0353   0.02108   0.01482  -0.0406   0.0154   1.0000
   8.250   1.0563   0.02249   0.01637  -0.0395   0.0150   1.0000
   8.500   1.0768   0.02412   0.01816  -0.0384   0.0147   1.0000
   8.750   1.0964   0.02600   0.02024  -0.0372   0.0145   1.0000
   9.000   1.1141   0.02825   0.02277  -0.0358   0.0144   1.0000
   9.250   1.1281   0.03128   0.02615  -0.0340   0.0145   1.0000
   9.500   1.1383   0.03450   0.02973  -0.0318   0.0145   1.0000
   9.750   1.1455   0.03764   0.03321  -0.0295   0.0143   1.0000
  10.000   1.1428   0.04216   0.03816  -0.0265   0.0147   1.0000
<< Back to S2060 8% (s2060-il)

Polar data table (+)

Polar graphs


<< Back to S2060 8% (s2060-il)