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NACA 16-012 (naca16012-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: NACA 16-012 (naca16012-il)
Reynolds number: 200,000
Max Cl/Cd: 44.87 at α=2.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-naca16012-il-200000.txt
Download as CSV file: xf-naca16012-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 16-012                                     
  
 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
  ------ -------- --------- --------- -------- -------- --------
 -11.000  -0.5324   0.09132   0.08798  -0.0454   1.0000   0.0622
 -10.750  -0.5524   0.08477   0.08144  -0.0477   1.0000   0.0619
 -10.500  -0.5768   0.07887   0.07554  -0.0488   1.0000   0.0614
 -10.250  -0.6013   0.07389   0.07053  -0.0484   1.0000   0.0607
 -10.000  -0.6296   0.06950   0.06610  -0.0468   1.0000   0.0601
  -9.750  -0.6592   0.06582   0.06237  -0.0436   1.0000   0.0595
  -9.500  -0.6895   0.06285   0.05934  -0.0388   1.0000   0.0592
  -9.250  -0.7199   0.06054   0.05697  -0.0327   1.0000   0.0590
  -9.000  -0.7494   0.05862   0.05498  -0.0256   1.0000   0.0589
  -8.750  -0.7737   0.05614   0.05238  -0.0194   1.0000   0.0598
  -8.500  -0.7985   0.05412   0.05014  -0.0125   1.0000   0.0617
  -7.000  -0.8886   0.03808   0.03128   0.0259   1.0000   0.0387
  -6.750  -0.8787   0.03449   0.02730   0.0292   1.0000   0.0379
  -6.500  -0.8640   0.03184   0.02425   0.0319   1.0000   0.0379
  -6.250  -0.8467   0.03003   0.02205   0.0343   1.0000   0.0387
  -6.000  -0.8273   0.02764   0.01930   0.0359   1.0000   0.0409
  -5.750  -0.8050   0.02587   0.01743   0.0368   1.0000   0.0429
  -5.500  -0.7829   0.02465   0.01609   0.0379   1.0000   0.0449
  -5.250  -0.7612   0.02368   0.01500   0.0390   1.0000   0.0478
  -5.000  -0.7410   0.02318   0.01430   0.0405   1.0000   0.0508
  -4.750  -0.7185   0.02150   0.01264   0.0414   1.0000   0.0543
  -4.500  -0.6997   0.02081   0.01196   0.0429   1.0000   0.0583
  -4.250  -0.6807   0.02036   0.01142   0.0445   1.0000   0.0628
  -4.000  -0.6643   0.01942   0.01052   0.0465   1.0000   0.0674
  -3.750  -0.6467   0.01895   0.01002   0.0483   1.0000   0.0737
  -3.500  -0.6298   0.01836   0.00945   0.0502   1.0000   0.0815
  -3.250  -0.6125   0.01783   0.00894   0.0520   1.0000   0.0945
  -3.000  -0.5996   0.01661   0.00835   0.0545   1.0000   0.1891
  -2.750  -0.6067   0.01353   0.00808   0.0614   0.9997   0.7181
  -2.500  -0.5631   0.01483   0.00979   0.0604   0.9992   0.8817
  -2.250  -0.5220   0.01600   0.01083   0.0586   0.9973   0.9121
  -2.000  -0.4633   0.01764   0.01231   0.0533   0.9976   0.9318
  -1.750  -0.2575   0.02189   0.01624   0.0199   1.0000   0.9649
  -1.500  -0.2218   0.02204   0.01631   0.0175   1.0000   0.9724
  -1.250  -0.1811   0.02208   0.01629   0.0139   1.0000   0.9769
  -1.000  -0.1488   0.02221   0.01635   0.0121   1.0000   0.9837
  -0.750  -0.1092   0.02219   0.01629   0.0086   1.0000   0.9874
  -0.500  -0.0734   0.02225   0.01632   0.0059   1.0000   0.9924
  -0.250  -0.0366   0.02231   0.01636   0.0029   1.0000   0.9964
   0.000   0.0000   0.02239   0.01643   0.0000   1.0000   1.0000
   0.250   0.0365   0.02231   0.01636  -0.0029   0.9964   1.0000
   0.500   0.0733   0.02225   0.01632  -0.0059   0.9924   1.0000
   0.750   0.1091   0.02218   0.01628  -0.0086   0.9874   1.0000
   1.000   0.1486   0.02221   0.01635  -0.0120   0.9838   1.0000
   1.250   0.1811   0.02208   0.01628  -0.0139   0.9770   1.0000
   1.500   0.2217   0.02203   0.01630  -0.0174   0.9724   1.0000
   1.750   0.2573   0.02189   0.01623  -0.0198   0.9650   1.0000
   2.000   0.4633   0.01763   0.01231  -0.0533   0.9318   0.9977
   2.250   0.5221   0.01599   0.01082  -0.0586   0.9121   0.9974
   2.500   0.5631   0.01482   0.00978  -0.0605   0.8819   0.9992
   2.750   0.6067   0.01352   0.00808  -0.0614   0.7195   0.9997
   3.000   0.5995   0.01660   0.00834  -0.0544   0.1903   1.0000
   3.250   0.6123   0.01782   0.00894  -0.0520   0.0945   1.0000
   3.500   0.6297   0.01835   0.00944  -0.0501   0.0816   1.0000
   3.750   0.6465   0.01894   0.01001  -0.0482   0.0737   1.0000
   4.000   0.6642   0.01942   0.01051  -0.0465   0.0674   1.0000
   4.250   0.6805   0.02036   0.01142  -0.0445   0.0629   1.0000
   4.500   0.6995   0.02081   0.01196  -0.0429   0.0583   1.0000
   4.750   0.7184   0.02149   0.01263  -0.0414   0.0543   1.0000
   5.000   0.7408   0.02316   0.01429  -0.0405   0.0508   1.0000
   5.250   0.7610   0.02367   0.01499  -0.0390   0.0478   1.0000
   5.500   0.7827   0.02464   0.01608  -0.0378   0.0449   1.0000
   5.750   0.8048   0.02586   0.01742  -0.0367   0.0429   1.0000
   6.000   0.8271   0.02762   0.01928  -0.0359   0.0409   1.0000
   6.250   0.8466   0.03003   0.02205  -0.0343   0.0387   1.0000
   6.500   0.8638   0.03183   0.02423  -0.0319   0.0379   1.0000
   6.750   0.8786   0.03448   0.02728  -0.0292   0.0379   1.0000
   7.000   0.8885   0.03805   0.03125  -0.0259   0.0387   1.0000
   9.000   0.7493   0.05849   0.05486   0.0256   0.0589   1.0000
   9.250   0.7195   0.06040   0.05683   0.0327   0.0590   1.0000
   9.500   0.6891   0.06271   0.05921   0.0389   0.0592   1.0000
   9.750   0.6592   0.06568   0.06224   0.0436   0.0596   1.0000
  10.000   0.6295   0.06935   0.06596   0.0468   0.0601   1.0000
  10.250   0.6017   0.07372   0.07037   0.0485   0.0608   1.0000
  10.500   0.5753   0.07883   0.07549   0.0487   0.0613   1.0000
  10.750   0.5530   0.08451   0.08118   0.0478   0.0619   1.0000
  11.000   0.5317   0.09122   0.08788   0.0453   0.0622   1.0000
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