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

Cambered plate C=6% T=5% R=2.11 (cp-060-050-gn) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: Cambered plate C=6% T=5% R=2.11 (cp-060-050-gn)
Reynolds number: 500,000
Max Cl/Cd: 45.25 at α=12.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-cp-060-050-gn-500000-n5.txt
Download as CSV file: xf-cp-060-050-gn-500000-n5.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 =     0.500 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.3153   0.11377   0.11078  -0.0473   0.9930   0.0429
  -7.750  -0.3042   0.11088   0.10790  -0.0492   0.9915   0.0434
  -7.250  -0.3036   0.10015   0.09713  -0.0562   0.9872   0.0455
  -7.000  -0.2887   0.09789   0.09488  -0.0581   0.9855   0.0457
  -6.750  -0.2724   0.09578   0.09278  -0.0600   0.9839   0.0459
  -6.500  -0.2562   0.09337   0.09039  -0.0623   0.9824   0.0461
  -6.250  -0.2338   0.09063   0.08766  -0.0661   0.9804   0.0466
  -6.000  -0.2244   0.08805   0.08509  -0.0675   0.9755   0.0470
  -5.750  -0.2114   0.08475   0.08179  -0.0706   0.9717   0.0476
  -5.500  -0.2249   0.07631   0.07330  -0.0780   0.9689   0.0495
  -5.250  -0.2046   0.07439   0.07139  -0.0802   0.9641   0.0496
  -5.000  -0.1855   0.07234   0.06935  -0.0825   0.9566   0.0499
  -4.750  -0.1633   0.06970   0.06672  -0.0865   0.9530   0.0502
  -4.500  -0.1494   0.06714   0.06416  -0.0893   0.9450   0.0505
  -4.000  -0.1463   0.05149   0.04829  -0.1103   0.9268   0.0536
  -3.750  -0.1186   0.04948   0.04623  -0.1136   0.9201   0.0538
  -3.500  -0.1010   0.04774   0.04445  -0.1145   0.9090   0.0540
  -3.250  -0.0800   0.04551   0.04215  -0.1165   0.8987   0.0543
  -2.750  -0.0780   0.03318   0.02920  -0.1190   0.8787   0.0578
  -2.500  -0.0567   0.03204   0.02797  -0.1186   0.8660   0.0580
  -1.500   0.0000   0.02034   0.01497  -0.1133   0.8011   0.0624
  -1.250   0.0204   0.01999   0.01443  -0.1116   0.7630   0.0626
  -1.000   0.0351   0.01987   0.01398  -0.1087   0.6943   0.0628
  -0.750   0.0336   0.02049   0.01378  -0.1026   0.5138   0.0629
  -0.500   0.0245   0.02218   0.01401  -0.0958   0.0883   0.0630
  -0.250   0.0471   0.02185   0.01358  -0.0946   0.0768   0.0632
   0.000   0.0698   0.02142   0.01306  -0.0935   0.0723   0.0636
   0.250   0.0927   0.02090   0.01244  -0.0923   0.0698   0.0640
   0.500   0.1155   0.02034   0.01178  -0.0911   0.0677   0.0646
   0.750   0.1377   0.01951   0.01075  -0.0898   0.0663   0.0655
   1.000   0.1592   0.01844   0.00940  -0.0883   0.0652   0.0665
   1.250   0.1823   0.01779   0.00856  -0.0870   0.0644   0.0671
   1.500   0.2061   0.01753   0.00827  -0.0858   0.0636   0.0674
   1.750   0.2299   0.01737   0.00809  -0.0847   0.0628   0.0676
   2.000   0.2537   0.01723   0.00793  -0.0834   0.0621   0.0679
   2.250   0.2773   0.01716   0.00783  -0.0822   0.0614   0.0683
   2.500   0.3009   0.01708   0.00773  -0.0809   0.0609   0.0687
   2.750   0.3241   0.01702   0.00763  -0.0796   0.0603   0.0692
   3.000   0.3469   0.01695   0.00751  -0.0781   0.0598   0.0697
   3.250   0.3691   0.01689   0.00740  -0.0766   0.0593   0.0701
   3.500   0.3918   0.01680   0.00726  -0.0751   0.0590   0.0706
   3.750   0.4146   0.01673   0.00715  -0.0736   0.0587   0.0710
   4.000   0.4372   0.01668   0.00706  -0.0721   0.0582   0.0715
   4.250   0.4596   0.01668   0.00702  -0.0705   0.0577   0.0718
   4.500   0.4818   0.01671   0.00701  -0.0689   0.0573   0.0722
   4.750   0.5034   0.01668   0.00701  -0.0672   0.0569   0.0725
   5.000   0.5248   0.01675   0.00710  -0.0655   0.0565   0.0730
   5.250   0.5458   0.01685   0.00721  -0.0637   0.0562   0.0734
   5.500   0.5662   0.01697   0.00733  -0.0617   0.0559   0.0739
   5.750   0.5862   0.01708   0.00745  -0.0597   0.0556   0.0743
   6.000   0.6062   0.01721   0.00757  -0.0577   0.0553   0.0747
   6.250   0.6263   0.01736   0.00772  -0.0557   0.0550   0.0751
   6.500   0.6463   0.01753   0.00789  -0.0537   0.0547   0.0755
   6.750   0.6664   0.01775   0.00809  -0.0518   0.0545   0.0760
   7.000   0.6866   0.01801   0.00834  -0.0499   0.0541   0.0764
   7.250   0.7073   0.01835   0.00867  -0.0481   0.0538   0.0768
   7.500   0.7287   0.01858   0.00892  -0.0464   0.0535   0.0774
   7.750   0.7500   0.01881   0.00917  -0.0447   0.0533   0.0779
   8.000   0.7715   0.01907   0.00948  -0.0430   0.0530   0.0785
   8.250   0.7931   0.01936   0.00980  -0.0414   0.0528   0.0789
   8.500   0.8149   0.01969   0.01016  -0.0399   0.0525   0.0795
   8.750   0.8368   0.02002   0.01053  -0.0383   0.0522   0.0801
   9.000   0.8592   0.02040   0.01094  -0.0369   0.0519   0.0807
   9.250   0.8814   0.02078   0.01134  -0.0354   0.0516   0.0814
   9.500   0.9036   0.02117   0.01176  -0.0339   0.0514   0.0821
   9.750   0.9259   0.02158   0.01220  -0.0325   0.0511   0.0829
  10.000   0.9478   0.02199   0.01264  -0.0310   0.0508   0.0837
  10.250   0.9697   0.02241   0.01310  -0.0295   0.0505   0.0848
  10.500   0.9908   0.02282   0.01354  -0.0279   0.0503   0.0864
  10.750   1.0113   0.02322   0.01396  -0.0263   0.0500   0.0882
  11.000   1.0322   0.02367   0.01443  -0.0247   0.0497   0.0900
  11.250   1.0528   0.02413   0.01491  -0.0231   0.0495   0.0924
  11.500   1.0737   0.02463   0.01543  -0.0215   0.0493   0.0962
  11.750   1.0958   0.02522   0.01605  -0.0202   0.0490   0.1040
  12.000   1.1099   0.02494   0.01691  -0.0178   0.0488   0.7037
  12.500   1.1963   0.02644   0.01913  -0.0240   0.0482   1.0000
  12.750   1.2145   0.02721   0.01996  -0.0221   0.0480   1.0000
  13.000   1.2322   0.02803   0.02086  -0.0202   0.0477   1.0000
  13.250   1.2491   0.02888   0.02179  -0.0182   0.0474   1.0000
  13.500   1.2653   0.02977   0.02275  -0.0162   0.0472   1.0000
  13.750   1.2804   0.03068   0.02374  -0.0140   0.0469   1.0000
  14.000   1.2941   0.03150   0.02463  -0.0117   0.0466   1.0000
  14.250   1.3070   0.03234   0.02553  -0.0094   0.0463   1.0000
  14.500   1.3193   0.03325   0.02651  -0.0070   0.0461   1.0000
  14.750   1.3307   0.03417   0.02749  -0.0046   0.0458   1.0000
  15.000   1.3411   0.03507   0.02847  -0.0022   0.0456   1.0000
  15.250   1.3510   0.03607   0.02953   0.0002   0.0454   1.0000
  15.500   1.3601   0.03709   0.03062   0.0026   0.0453   1.0000
  15.750   1.3684   0.03822   0.03182   0.0050   0.0452   1.0000
  16.000   1.3763   0.03927   0.03293   0.0073   0.0450   1.0000
  16.250   1.3830   0.04051   0.03424   0.0095   0.0449   1.0000
  16.500   1.3891   0.04180   0.03560   0.0117   0.0448   1.0000
  16.750   1.3946   0.04313   0.03700   0.0139   0.0447   1.0000
  17.000   1.3992   0.04455   0.03848   0.0159   0.0446   1.0000
  17.250   1.4025   0.04617   0.04019   0.0178   0.0445   1.0000
  17.500   1.4042   0.04798   0.04207   0.0196   0.0444   1.0000
  17.750   1.4047   0.04993   0.04411   0.0213   0.0443   1.0000
  18.000   1.4041   0.05207   0.04633   0.0228   0.0441   1.0000
  18.250   1.3907   0.05545   0.04993   0.0242   0.0440   1.0000
  18.500   1.3596   0.06090   0.05574   0.0248   0.0437   1.0000
  18.750   1.2949   0.07114   0.06653   0.0224   0.0431   1.0000
<< Back to Cambered plate C=6% T=5% R=2.11 (cp-060-050-gn)

Polar data table (+)

Polar graphs


<< Back to Cambered plate C=6% T=5% R=2.11 (cp-060-050-gn)