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RAE 5215 AIRFOIL (rae5215-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: RAE 5215 AIRFOIL (rae5215-il)
Reynolds number: 100,000
Max Cl/Cd: 37.97 at α=2.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-rae5215-il-100000.txt
Download as CSV file: xf-rae5215-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: RAE 5215 AIRFOIL                                
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.5424   0.09648   0.09135  -0.0145   1.0000   0.1264
  -8.500  -0.5347   0.09342   0.08831  -0.0145   1.0000   0.1321
  -8.250  -0.5757   0.08654   0.08157  -0.0285   1.0000   0.1374
  -8.000  -0.5471   0.08445   0.07950  -0.0204   1.0000   0.1423
  -7.750  -0.5434   0.08127   0.07635  -0.0210   1.0000   0.1487
  -7.500  -0.5718   0.07306   0.06815  -0.0338   1.0000   0.1565
  -7.250  -0.5489   0.07163   0.06680  -0.0288   1.0000   0.1632
  -7.000  -0.5585   0.06622   0.06134  -0.0340   1.0000   0.1746
  -6.750  -0.5623   0.06232   0.05733  -0.0369   1.0000   0.1896
  -6.500  -0.5434   0.06031   0.05547  -0.0332   1.0000   0.1978
  -6.250  -0.5374   0.05752   0.05271  -0.0325   1.0000   0.2147
  -6.000  -0.5194   0.03840   0.03057  -0.0456   1.0000   0.0834
  -5.750  -0.4995   0.03495   0.02708  -0.0449   1.0000   0.0812
  -5.500  -0.4783   0.03164   0.02345  -0.0442   1.0000   0.0774
  -5.250  -0.4541   0.02881   0.01985  -0.0432   1.0000   0.0737
  -5.000  -0.4318   0.02684   0.01776  -0.0423   1.0000   0.0748
  -4.750  -0.4095   0.02540   0.01624  -0.0414   1.0000   0.0779
  -4.500  -0.3857   0.02395   0.01458  -0.0404   1.0000   0.0799
  -4.250  -0.3616   0.02286   0.01324  -0.0394   1.0000   0.0828
  -4.000  -0.3382   0.02135   0.01186  -0.0387   1.0000   0.0869
  -3.750  -0.3141   0.02033   0.01084  -0.0378   1.0000   0.0907
  -3.500  -0.2898   0.01936   0.00990  -0.0371   1.0000   0.0968
  -3.250  -0.2652   0.01850   0.00918  -0.0366   1.0000   0.1049
  -3.000  -0.2396   0.01760   0.00841  -0.0363   1.0000   0.1169
  -2.750  -0.2106   0.01648   0.00752  -0.0369   1.0000   0.1550
  -2.500  -0.1871   0.01493   0.00828  -0.0356   1.0000   0.6497
  -2.250  -0.1693   0.01544   0.00878  -0.0325   1.0000   0.6947
  -2.000  -0.1549   0.01584   0.00926  -0.0285   1.0000   0.7316
  -1.750  -0.1442   0.01616   0.00967  -0.0236   1.0000   0.7684
  -1.500  -0.1365   0.01631   0.00993  -0.0181   1.0000   0.8047
  -1.250  -0.1282   0.01629   0.00995  -0.0129   1.0000   0.8364
  -1.000  -0.1141   0.01617   0.00982  -0.0098   1.0000   0.8584
  -0.750  -0.0920   0.01605   0.00965  -0.0090   1.0000   0.8708
  -0.500  -0.0666   0.01600   0.00954  -0.0093   1.0000   0.8822
  -0.250  -0.0441   0.01592   0.00944  -0.0087   1.0000   0.8935
   0.000  -0.0204   0.01589   0.00939  -0.0086   1.0000   0.9054
   0.250   0.0035   0.01589   0.00939  -0.0085   1.0000   0.9188
   0.500   0.0272   0.01591   0.00942  -0.0084   1.0000   0.9344
   0.750   0.0526   0.01596   0.00951  -0.0089   1.0000   0.9544
   1.000   0.0923   0.01613   0.00972  -0.0126   0.9945   0.9881
   1.250   0.2033   0.01612   0.00971  -0.0275   0.9565   0.9794
   1.500   0.2635   0.01585   0.00951  -0.0333   0.9369   1.0000
   1.750   0.3265   0.01540   0.00916  -0.0393   0.9182   1.0000
   2.000   0.3863   0.01462   0.00849  -0.0436   0.8957   1.0000
   2.250   0.4239   0.01393   0.00788  -0.0439   0.8643   1.0000
   2.500   0.4553   0.01317   0.00717  -0.0424   0.8163   1.0000
   2.750   0.4780   0.01259   0.00606  -0.0381   0.6222   1.0000
   3.000   0.4887   0.01508   0.00637  -0.0351   0.2971   1.0000
   3.250   0.5156   0.01620   0.00701  -0.0356   0.2556   1.0000
   3.500   0.5449   0.01705   0.00766  -0.0364   0.2327   1.0000
   3.750   0.5745   0.01794   0.00834  -0.0371   0.2162   1.0000
   4.000   0.6039   0.01892   0.00909  -0.0378   0.2024   1.0000
   4.250   0.6337   0.01970   0.00984  -0.0383   0.1902   1.0000
   4.500   0.6631   0.02073   0.01083  -0.0388   0.1798   1.0000
   4.750   0.6920   0.02187   0.01178  -0.0392   0.1698   1.0000
   5.000   0.7208   0.02288   0.01292  -0.0394   0.1606   1.0000
   5.250   0.7490   0.02432   0.01421  -0.0397   0.1518   1.0000
   5.500   0.7765   0.02527   0.01536  -0.0396   0.1429   1.0000
   5.750   0.8033   0.02685   0.01689  -0.0397   0.1352   1.0000
   6.000   0.8296   0.02796   0.01822  -0.0394   0.1270   1.0000
   6.250   0.8551   0.02981   0.02007  -0.0393   0.1204   1.0000
   6.500   0.8796   0.03130   0.02193  -0.0387   0.1138   1.0000
   6.750   0.9047   0.03301   0.02352  -0.0387   0.1083   1.0000
   7.000   0.9255   0.03547   0.02656  -0.0376   0.1039   1.0000
   7.250   0.9464   0.03769   0.02913  -0.0367   0.0998   1.0000
   7.500   0.9706   0.03927   0.03056  -0.0367   0.0960   1.0000
   7.750   0.9859   0.04281   0.03457  -0.0355   0.0938   1.0000
   8.000   0.9978   0.04681   0.03914  -0.0339   0.0930   1.0000
   8.250   1.0060   0.05129   0.04412  -0.0324   0.0927   1.0000
   8.500   1.0100   0.05612   0.04941  -0.0309   0.0928   1.0000
   8.750   1.0111   0.06115   0.05480  -0.0297   0.0931   1.0000
   9.000   1.0131   0.06610   0.05998  -0.0288   0.0939   1.0000
   9.250   0.8294   0.10314   0.09817  -0.0502   0.1632   1.0000
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