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

S6063 7.05% (s6063-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: S6063 7.05% (s6063-il)
Reynolds number: 200,000
Max Cl/Cd: 61.5 at α=3.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-s6063-il-200000.txt
Download as CSV file: xf-s6063-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: S6063 7.05%                                     
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.5351   0.08514   0.08179  -0.0081   1.0000   0.0445
  -7.750  -0.5381   0.08101   0.07772  -0.0117   1.0000   0.0458
  -7.500  -0.5383   0.07522   0.07194  -0.0208   1.0000   0.0472
  -7.250  -0.4961   0.05516   0.05195  -0.0307   1.0000   0.0490
  -7.000  -0.4828   0.05188   0.04880  -0.0277   1.0000   0.0506
  -6.750  -0.4753   0.04816   0.04507  -0.0281   1.0000   0.0521
  -6.500  -0.4687   0.04367   0.04049  -0.0300   1.0000   0.0541
  -6.250  -0.4607   0.03866   0.03533  -0.0326   1.0000   0.0571
  -6.000  -0.4562   0.03175   0.02785  -0.0365   1.0000   0.0622
  -5.750  -0.4418   0.02838   0.02458  -0.0361   1.0000   0.0640
  -5.500  -0.4268   0.02546   0.02158  -0.0359   1.0000   0.0672
  -5.250  -0.4138   0.02161   0.01725  -0.0363   1.0000   0.0760
  -5.000  -0.3987   0.01920   0.01488  -0.0354   1.0000   0.0790
  -4.750  -0.3922   0.02556   0.01948  -0.0337   1.0000   0.0389
  -4.500  -0.3702   0.02218   0.01568  -0.0320   1.0000   0.0306
  -4.250  -0.3471   0.01949   0.01234  -0.0301   1.0000   0.0268
  -4.000  -0.3248   0.01812   0.01074  -0.0286   1.0000   0.0262
  -3.750  -0.3030   0.01650   0.00895  -0.0273   1.0000   0.0263
  -3.500  -0.2821   0.01516   0.00758  -0.0262   1.0000   0.0299
  -3.250  -0.2603   0.01422   0.00658  -0.0250   1.0000   0.0317
  -3.000  -0.2385   0.01333   0.00564  -0.0238   1.0000   0.0329
  -2.750  -0.2164   0.01263   0.00487  -0.0227   1.0000   0.0352
  -2.500  -0.1936   0.01187   0.00405  -0.0218   1.0000   0.0436
  -2.250  -0.1707   0.01029   0.00343  -0.0217   1.0000   0.2513
  -2.000  -0.1513   0.00915   0.00343  -0.0208   1.0000   0.5214
  -1.750  -0.1364   0.00845   0.00360  -0.0180   0.9995   0.7393
  -1.500  -0.0861   0.00805   0.00360  -0.0211   0.9984   1.0000
  -1.250  -0.0406   0.00816   0.00350  -0.0251   0.9914   1.0000
  -1.000   0.0062   0.00826   0.00343  -0.0293   0.9847   1.0000
  -0.750   0.0503   0.00831   0.00336  -0.0328   0.9764   1.0000
  -0.500   0.0997   0.00833   0.00327  -0.0374   0.9709   1.0000
  -0.250   0.1432   0.00830   0.00318  -0.0407   0.9618   1.0000
   0.000   0.1871   0.00825   0.00308  -0.0439   0.9530   1.0000
   0.250   0.2287   0.00818   0.00298  -0.0465   0.9433   1.0000
   0.500   0.2614   0.00813   0.00291  -0.0472   0.9295   1.0000
   0.750   0.2879   0.00812   0.00287  -0.0465   0.9132   1.0000
   1.000   0.3119   0.00811   0.00284  -0.0452   0.8957   1.0000
   1.250   0.3351   0.00812   0.00283  -0.0437   0.8779   1.0000
   1.500   0.3581   0.00812   0.00280  -0.0421   0.8600   1.0000
   1.750   0.3810   0.00815   0.00281  -0.0406   0.8391   1.0000
   2.000   0.4045   0.00818   0.00280  -0.0392   0.8184   1.0000
   2.250   0.4285   0.00823   0.00283  -0.0380   0.7946   1.0000
   2.500   0.4527   0.00830   0.00289  -0.0368   0.7695   1.0000
   2.750   0.4772   0.00840   0.00294  -0.0357   0.7417   1.0000
   3.000   0.5017   0.00852   0.00300  -0.0346   0.7105   1.0000
   3.250   0.5261   0.00868   0.00308  -0.0335   0.6733   1.0000
   3.500   0.5492   0.00893   0.00316  -0.0321   0.6169   1.0000
   3.750   0.5716   0.00933   0.00325  -0.0307   0.5414   1.0000
   4.000   0.5945   0.00982   0.00345  -0.0297   0.4674   1.0000
   4.250   0.6175   0.01040   0.00373  -0.0289   0.3884   1.0000
   4.500   0.6379   0.01144   0.00413  -0.0280   0.2567   1.0000
   4.750   0.6543   0.01361   0.00513  -0.0270   0.0541   1.0000
   5.000   0.6770   0.01483   0.00630  -0.0261   0.0355   1.0000
   5.250   0.7012   0.01564   0.00726  -0.0253   0.0327   1.0000
   5.500   0.7243   0.01665   0.00837  -0.0244   0.0308   1.0000
   5.750   0.7469   0.01782   0.00962  -0.0235   0.0298   1.0000
   6.000   0.7697   0.01921   0.01111  -0.0225   0.0293   1.0000
   6.250   0.7931   0.02064   0.01260  -0.0216   0.0281   1.0000
   6.500   0.8156   0.02277   0.01478  -0.0209   0.0262   1.0000
   6.750   0.8396   0.02501   0.01722  -0.0201   0.0261   1.0000
   7.000   0.8637   0.02703   0.01950  -0.0192   0.0264   1.0000
   7.250   0.8873   0.02946   0.02240  -0.0179   0.0279   1.0000
   7.500   0.9045   0.03464   0.02838  -0.0158   0.0323   1.0000
  11.500   0.8056   0.13110   0.12769  -0.0471   0.0519   1.0000
  11.750   0.8090   0.13531   0.13190  -0.0486   0.0504   1.0000
  12.000   0.8161   0.13879   0.13540  -0.0489   0.0490   1.0000
  14.250   0.6688   0.16714   0.16408  -0.0605   0.0428   1.0000
  14.500   0.6719   0.17033   0.16728  -0.0622   0.0403   1.0000
<< Back to S6063 7.05% (s6063-il)

Polar data table (+)

Polar graphs


<< Back to S6063 7.05% (s6063-il)