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NACA 0012-34 (naca001234-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: NACA 0012-34 (naca001234-il)
Reynolds number: 50,000
Max Cl/Cd: 25.93 at α=5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-naca001234-il-50000.txt
Download as CSV file: xf-naca001234-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 0012-34                                    
  
 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  -0.6304   0.09725   0.09044  -0.0247   1.0000   0.1825
  -9.750  -0.6668   0.08674   0.08006  -0.0332   1.0000   0.1577
  -9.500  -0.7097   0.08125   0.07459  -0.0345   1.0000   0.1546
  -9.250  -0.7325   0.07621   0.06949  -0.0333   1.0000   0.1441
  -9.000  -0.7750   0.07221   0.06525  -0.0300   1.0000   0.1397
  -8.750  -0.7807   0.06716   0.05999  -0.0280   1.0000   0.1288
  -8.500  -0.7946   0.06312   0.05557  -0.0243   1.0000   0.1209
  -8.250  -0.8039   0.05893   0.05100  -0.0205   1.0000   0.1145
  -8.000  -0.8079   0.05554   0.04707  -0.0164   1.0000   0.1086
  -7.750  -0.8022   0.05173   0.04297  -0.0134   1.0000   0.1051
  -7.500  -0.7984   0.04826   0.03897  -0.0097   1.0000   0.1013
  -7.250  -0.7942   0.04577   0.03564  -0.0052   1.0000   0.0979
  -7.000  -0.7797   0.04309   0.03251  -0.0025   1.0000   0.0974
  -6.750  -0.7589   0.03961   0.02896  -0.0016   1.0000   0.1012
  -6.500  -0.7381   0.03715   0.02622   0.0001   1.0000   0.1048
  -6.250  -0.7114   0.03470   0.02339   0.0012   1.0000   0.1074
  -6.000  -0.6766   0.03223   0.02070   0.0012   1.0000   0.1124
  -5.750  -0.6427   0.03011   0.01866   0.0008   1.0000   0.1271
  -5.500  -0.6139   0.02820   0.01686   0.0015   1.0000   0.1452
  -5.250  -0.5997   0.02636   0.01522   0.0042   1.0000   0.1780
  -5.000  -0.6032   0.02327   0.01359   0.0092   1.0000   0.2994
  -4.750  -0.4085   0.03433   0.02537   0.0087   1.0000   0.8524
  -4.500  -0.3529   0.03390   0.02445   0.0035   1.0000   0.8795
  -4.250  -0.3002   0.03321   0.02334  -0.0018   1.0000   0.9051
  -4.000  -0.2508   0.03233   0.02209  -0.0071   1.0000   0.9298
  -3.750  -0.1962   0.03124   0.02068  -0.0137   1.0000   0.9529
  -3.500  -0.1381   0.02989   0.01902  -0.0215   1.0000   0.9744
  -3.250  -0.0769   0.02839   0.01726  -0.0303   1.0000   0.9951
  -3.000  -0.0523   0.02745   0.01621  -0.0321   1.0000   1.0000
  -2.750  -0.0437   0.02685   0.01556  -0.0306   1.0000   1.0000
  -2.500  -0.0359   0.02633   0.01498  -0.0288   1.0000   1.0000
  -2.250  -0.0288   0.02588   0.01449  -0.0268   1.0000   1.0000
  -2.000  -0.0224   0.02550   0.01408  -0.0246   1.0000   1.0000
  -1.750  -0.0170   0.02517   0.01374  -0.0222   1.0000   1.0000
  -1.500  -0.0125   0.02491   0.01346  -0.0195   1.0000   1.0000
  -1.250  -0.0089   0.02470   0.01324  -0.0166   1.0000   1.0000
  -1.000  -0.0061   0.02453   0.01307  -0.0135   1.0000   1.0000
  -0.750  -0.0040   0.02441   0.01293  -0.0102   1.0000   1.0000
  -0.500  -0.0024   0.02432   0.01284  -0.0069   1.0000   1.0000
  -0.250  -0.0011   0.02427   0.01279  -0.0035   1.0000   1.0000
   0.000   0.0000   0.02426   0.01277   0.0000   1.0000   1.0000
   0.250   0.0011   0.02427   0.01278   0.0035   1.0000   1.0000
   0.500   0.0024   0.02432   0.01284   0.0069   1.0000   1.0000
   0.750   0.0040   0.02441   0.01292   0.0102   1.0000   1.0000
   1.000   0.0061   0.02453   0.01306   0.0135   1.0000   1.0000
   1.250   0.0089   0.02469   0.01323   0.0166   1.0000   1.0000
   1.500   0.0125   0.02490   0.01345   0.0195   1.0000   1.0000
   1.750   0.0170   0.02516   0.01373   0.0222   1.0000   1.0000
   2.000   0.0225   0.02549   0.01407   0.0246   1.0000   1.0000
   2.250   0.0289   0.02587   0.01448   0.0268   1.0000   1.0000
   2.500   0.0359   0.02632   0.01496   0.0288   1.0000   1.0000
   2.750   0.0438   0.02683   0.01554   0.0306   1.0000   1.0000
   3.000   0.0524   0.02743   0.01619   0.0321   1.0000   1.0000
   3.250   0.0768   0.02837   0.01724   0.0303   0.9952   1.0000
   3.500   0.1380   0.02987   0.01900   0.0215   0.9745   1.0000
   3.750   0.1963   0.03122   0.02066   0.0137   0.9530   1.0000
   4.000   0.2507   0.03231   0.02207   0.0071   0.9299   1.0000
   4.250   0.3003   0.03319   0.02332   0.0018   0.9052   1.0000
   4.500   0.3530   0.03389   0.02443  -0.0035   0.8795   1.0000
   4.750   0.4087   0.03431   0.02535  -0.0087   0.8524   1.0000
   5.000   0.6032   0.02326   0.01358  -0.0092   0.3001   1.0000
   5.250   0.5996   0.02635   0.01521  -0.0042   0.1783   1.0000
   5.500   0.6139   0.02819   0.01686  -0.0015   0.1453   1.0000
   5.750   0.6426   0.03010   0.01865  -0.0008   0.1272   1.0000
   6.000   0.6765   0.03223   0.02069  -0.0011   0.1124   1.0000
   6.250   0.7113   0.03469   0.02338  -0.0012   0.1074   1.0000
   6.500   0.7381   0.03715   0.02621  -0.0001   0.1048   1.0000
   6.750   0.7588   0.03960   0.02895   0.0016   0.1012   1.0000
   7.000   0.7797   0.04308   0.03249   0.0025   0.0975   1.0000
   7.250   0.7942   0.04576   0.03563   0.0052   0.0979   1.0000
   7.500   0.7984   0.04826   0.03896   0.0097   0.1013   1.0000
   7.750   0.8022   0.05173   0.04296   0.0134   0.1051   1.0000
   8.000   0.8079   0.05554   0.04706   0.0164   0.1086   1.0000
   8.250   0.8039   0.05893   0.05100   0.0205   0.1145   1.0000
   8.500   0.7947   0.06312   0.05558   0.0243   0.1209   1.0000
   8.750   0.7807   0.06716   0.06000   0.0279   0.1288   1.0000
   9.000   0.7749   0.07222   0.06526   0.0300   0.1397   1.0000
   9.250   0.7325   0.07624   0.06951   0.0332   0.1442   1.0000
   9.500   0.7095   0.08129   0.07463   0.0344   0.1546   1.0000
   9.750   0.6668   0.08683   0.08015   0.0331   0.1579   1.0000
  10.000   0.6296   0.09745   0.09063   0.0243   0.1828   1.0000
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