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Cambered plate C=6% T=5% R=2.11 (cp-060-050-gn) Xfoil prediction polar at RE=50,000 Ncrit=1


Details Polar file
Airfoil: Cambered plate C=6% T=5% R=2.11 (cp-060-050-gn)
Reynolds number: 50,000
Max Cl/Cd: 21.16 at α=11°
Description: Mach=0 Ncrit=1
Source: Xfoil prediction
Download polar: xf-cp-060-050-gn-50000-n1.txt
Download as CSV file: xf-cp-060-050-gn-50000-n1.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: Cambered plate C=6% T=5% R=2.11                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   1.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -5.750  -0.2537   0.10399   0.09490  -0.0468   0.9462   0.0478
  -5.500  -0.2477   0.10132   0.09226  -0.0479   0.9406   0.0484
  -5.250  -0.2425   0.09842   0.08938  -0.0493   0.9351   0.0492
  -5.000  -0.2377   0.09604   0.08706  -0.0499   0.9283   0.0498
  -4.750  -0.2310   0.09355   0.08461  -0.0509   0.9224   0.0502
  -4.500  -0.2256   0.09108   0.08219  -0.0517   0.9159   0.0506
  -4.250  -0.2219   0.08864   0.07980  -0.0523   0.9086   0.0508
  -4.000  -0.2163   0.08606   0.07725  -0.0535   0.9023   0.0510
  -3.750  -0.2162   0.08376   0.07500  -0.0535   0.8936   0.0513
  -3.500  -0.2151   0.08124   0.07252  -0.0543   0.8856   0.0519
  -3.250  -0.2157   0.07862   0.06994  -0.0550   0.8765   0.0526
  -3.000  -0.2133   0.07552   0.06684  -0.0567   0.8677   0.0533
  -2.750  -0.2072   0.07284   0.06418  -0.0580   0.8604   0.0538
  -2.500  -0.2023   0.07037   0.06173  -0.0585   0.8518   0.0541
  -2.250  -0.1921   0.06761   0.05895  -0.0600   0.8453   0.0544
  -2.000  -0.1852   0.06514   0.05646  -0.0604   0.8372   0.0546
  -1.750  -0.1750   0.06258   0.05387  -0.0612   0.8296   0.0550
  -1.250  -0.1510   0.05718   0.04832  -0.0630   0.8154   0.0569
  -1.000  -0.1361   0.05433   0.04535  -0.0640   0.8085   0.0577
  -0.500  -0.1031   0.05014   0.04102  -0.0642   0.7956   0.0583
  -0.250  -0.0848   0.04817   0.03897  -0.0642   0.7891   0.0585
   0.000  -0.0629   0.04618   0.03687  -0.0647   0.7840   0.0592
   0.250  -0.0449   0.04434   0.03491  -0.0643   0.7771   0.0603
   0.500  -0.0247   0.04238   0.03279  -0.0640   0.7703   0.0614
   0.750  -0.0006   0.04043   0.03065  -0.0642   0.7637   0.0620
   1.000   0.0193   0.03889   0.02897  -0.0632   0.7492   0.0623
   1.250   0.0448   0.03741   0.02734  -0.0627   0.7228   0.0626
   1.500   0.1092   0.03659   0.02379  -0.0674   0.2437   0.0639
   1.750   0.1179   0.03705   0.02328  -0.0649   0.0898   0.0647
   2.000   0.1388   0.03616   0.02214  -0.0637   0.0826   0.0656
   2.250   0.1601   0.03531   0.02107  -0.0625   0.0786   0.0662
   2.500   0.1818   0.03457   0.02016  -0.0613   0.0757   0.0665
   2.750   0.2037   0.03392   0.01933  -0.0601   0.0731   0.0669
   3.000   0.2254   0.03349   0.01879  -0.0588   0.0712   0.0675
   3.250   0.2471   0.03317   0.01839  -0.0575   0.0692   0.0683
   3.500   0.2688   0.03286   0.01799  -0.0561   0.0676   0.0693
   3.750   0.2906   0.03257   0.01759  -0.0548   0.0664   0.0701
   4.000   0.3127   0.03233   0.01724  -0.0534   0.0655   0.0707
   4.250   0.3346   0.03216   0.01697  -0.0520   0.0647   0.0710
   4.500   0.3559   0.03203   0.01679  -0.0505   0.0637   0.0713
   4.750   0.3779   0.03199   0.01668  -0.0490   0.0626   0.0716
   5.000   0.4005   0.03204   0.01665  -0.0477   0.0614   0.0723
   5.250   0.4221   0.03219   0.01678  -0.0463   0.0604   0.0731
   5.500   0.4435   0.03237   0.01693  -0.0449   0.0596   0.0741
   5.750   0.4648   0.03257   0.01709  -0.0435   0.0589   0.0751
   6.000   0.4860   0.03279   0.01730  -0.0420   0.0586   0.0759
   6.250   0.5070   0.03306   0.01757  -0.0405   0.0582   0.0763
   6.500   0.5281   0.03337   0.01788  -0.0390   0.0579   0.0767
   6.750   0.5493   0.03373   0.01823  -0.0375   0.0576   0.0771
   7.000   0.5709   0.03412   0.01863  -0.0361   0.0572   0.0775
   7.250   0.5929   0.03454   0.01909  -0.0349   0.0567   0.0782
   7.500   0.6156   0.03499   0.01957  -0.0337   0.0560   0.0793
   7.750   0.6389   0.03547   0.02007  -0.0326   0.0554   0.0807
   8.000   0.6629   0.03596   0.02057  -0.0316   0.0546   0.0821
   8.250   0.6874   0.03648   0.02108  -0.0307   0.0540   0.0834
   8.500   0.7138   0.03705   0.02170  -0.0300   0.0534   0.0843
   8.750   0.7411   0.03769   0.02242  -0.0295   0.0529   0.0856
   9.000   0.7687   0.03842   0.02320  -0.0290   0.0526   0.0875
   9.250   0.7961   0.03922   0.02407  -0.0285   0.0523   0.0899
   9.500   0.8231   0.04010   0.02506  -0.0280   0.0520   0.0922
   9.750   0.8491   0.04107   0.02612  -0.0274   0.0517   0.0952
  10.000   0.8738   0.04211   0.02730  -0.0267   0.0513   0.0986
  10.250   0.8965   0.04321   0.02853  -0.0257   0.0508   0.1030
  10.500   0.9171   0.04434   0.02981  -0.0245   0.0502   0.1095
  11.000   0.9586   0.04531   0.03270  -0.0230   0.0491   1.0000
  11.250   0.9730   0.04667   0.03410  -0.0211   0.0487   1.0000
  11.500   0.9869   0.04841   0.03601  -0.0192   0.0483   1.0000
  11.750   0.9978   0.05024   0.03800  -0.0171   0.0480   1.0000
  12.000   1.0056   0.05217   0.04009  -0.0148   0.0478   1.0000
  12.250   1.0104   0.05424   0.04235  -0.0124   0.0476   1.0000
  12.500   1.0114   0.05653   0.04485  -0.0099   0.0473   1.0000
  12.750   1.0088   0.05903   0.04756  -0.0073   0.0470   1.0000
  13.000   1.0023   0.06180   0.05057  -0.0050   0.0466   1.0000
  13.250   0.9919   0.06492   0.05391  -0.0029   0.0463   1.0000
  13.500   0.9775   0.06850   0.05773  -0.0014   0.0459   1.0000
  13.750   0.9581   0.07278   0.06225  -0.0007   0.0456   1.0000
  14.000   0.9327   0.07816   0.06789  -0.0013   0.0453   1.0000
  14.250   0.8978   0.08561   0.07564  -0.0041   0.0451   1.0000
  14.500   0.8541   0.09633   0.08666  -0.0106   0.0447   1.0000
  14.750   0.8193   0.10821   0.09876  -0.0186   0.0443   1.0000
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