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P-51D TIP (BL215) AIRFOIL (p51dtip-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: P-51D TIP (BL215) AIRFOIL (p51dtip-il)
Reynolds number: 50,000
Max Cl/Cd: 22.64 at α=6.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-p51dtip-il-50000.txt
Download as CSV file: xf-p51dtip-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: P-51D TIP (BL215) AIRFOIL                       
  
 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  -0.4622   0.11232   0.10595  -0.0148   1.0000   0.3004
  -9.750  -0.4622   0.10949   0.10318  -0.0152   1.0000   0.3187
  -9.500  -0.4677   0.10696   0.10075  -0.0156   1.0000   0.3356
  -9.250  -0.4445   0.10268   0.09644  -0.0145   1.0000   0.3607
  -9.000  -0.4343   0.09951   0.09329  -0.0135   1.0000   0.3896
  -8.750  -0.4194   0.09586   0.08966  -0.0127   1.0000   0.4153
  -8.250  -0.4147   0.09084   0.08472  -0.0108   1.0000   0.4645
  -8.000  -0.3912   0.08649   0.08035  -0.0109   1.0000   0.4831
  -7.750  -0.3778   0.08362   0.07749  -0.0097   1.0000   0.5118
  -6.750  -0.5346   0.07772   0.07236   0.0003   1.0000   0.4232
  -6.500  -0.5786   0.07672   0.07153   0.0052   1.0000   0.4219
  -6.250  -0.6439   0.07138   0.06630   0.0013   1.0000   0.3772
  -6.000  -0.6740   0.05597   0.04928  -0.0188   1.0000   0.1738
  -5.750  -0.6615   0.05095   0.04346  -0.0184   1.0000   0.1466
  -5.500  -0.6474   0.04732   0.03896  -0.0172   1.0000   0.1359
  -5.250  -0.6309   0.04407   0.03532  -0.0160   1.0000   0.1350
  -5.000  -0.6116   0.04093   0.03177  -0.0149   1.0000   0.1321
  -4.750  -0.5897   0.03811   0.02841  -0.0138   1.0000   0.1292
  -4.500  -0.5674   0.03592   0.02577  -0.0127   1.0000   0.1319
  -4.250  -0.5441   0.03422   0.02347  -0.0115   1.0000   0.1365
  -4.000  -0.5198   0.03215   0.02124  -0.0106   1.0000   0.1393
  -3.750  -0.4953   0.03065   0.01960  -0.0094   1.0000   0.1445
  -3.500  -0.4719   0.02951   0.01840  -0.0080   1.0000   0.1559
  -3.250  -0.1706   0.02963   0.02083  -0.0320   1.0000   1.0000
  -3.000  -0.1654   0.02950   0.02047  -0.0296   1.0000   1.0000
  -2.750  -0.1597   0.02938   0.02014  -0.0273   1.0000   1.0000
  -2.500  -0.1537   0.02927   0.01981  -0.0249   1.0000   1.0000
  -2.250  -0.1475   0.02915   0.01952  -0.0226   1.0000   1.0000
  -2.000  -0.1411   0.02905   0.01926  -0.0202   1.0000   1.0000
  -1.750  -0.1345   0.02894   0.01901  -0.0179   1.0000   1.0000
  -1.500  -0.1279   0.02884   0.01877  -0.0155   1.0000   1.0000
  -1.250  -0.1213   0.02873   0.01854  -0.0132   1.0000   1.0000
  -1.000  -0.1147   0.02862   0.01831  -0.0108   1.0000   1.0000
  -0.750  -0.1082   0.02850   0.01806  -0.0084   1.0000   1.0000
  -0.500  -0.1019   0.02837   0.01784  -0.0060   1.0000   1.0000
  -0.250  -0.0957   0.02824   0.01762  -0.0035   1.0000   1.0000
   0.000  -0.0899   0.02808   0.01740  -0.0010   1.0000   1.0000
   0.250  -0.0843   0.02791   0.01716   0.0016   1.0000   1.0000
   0.500  -0.0790   0.02772   0.01691   0.0042   1.0000   1.0000
   0.750  -0.0738   0.02751   0.01663   0.0069   1.0000   1.0000
   1.000  -0.0684   0.02729   0.01637   0.0094   1.0000   1.0000
   1.250  -0.0622   0.02709   0.01612   0.0119   1.0000   1.0000
   1.500  -0.0539   0.02695   0.01594   0.0139   1.0000   1.0000
   1.750  -0.0420   0.02695   0.01589   0.0153   1.0000   1.0000
   2.000  -0.0269   0.02707   0.01597   0.0161   1.0000   1.0000
   2.250  -0.0095   0.02730   0.01617   0.0164   1.0000   1.0000
   2.500   0.0093   0.02762   0.01647   0.0165   1.0000   1.0000
   2.750   0.0291   0.02800   0.01684   0.0164   1.0000   1.0000
   3.000   0.0494   0.02845   0.01730   0.0162   1.0000   1.0000
   3.250   0.0701   0.02896   0.01783   0.0159   1.0000   1.0000
   3.500   0.0909   0.02952   0.01844   0.0156   1.0000   1.0000
   3.750   0.1117   0.03013   0.01911   0.0152   1.0000   1.0000
   4.000   0.1323   0.03080   0.01984   0.0148   1.0000   1.0000
   4.250   0.1527   0.03153   0.02065   0.0144   1.0000   1.0000
   4.500   0.1727   0.03231   0.02156   0.0139   1.0000   1.0000
   4.750   0.1923   0.03316   0.02252   0.0135   1.0000   1.0000
   5.000   0.2114   0.03409   0.02359   0.0130   1.0000   1.0000
   5.250   0.2297   0.03512   0.02476   0.0125   1.0000   1.0000
   5.500   0.2472   0.03625   0.02606   0.0119   1.0000   1.0000
   5.750   0.2636   0.03753   0.02756   0.0112   1.0000   1.0000
   6.000   0.4576   0.04316   0.03446  -0.0159   0.8504   1.0000
   6.250   0.6294   0.02970   0.02297  -0.0110   0.5732   1.0000
   6.500   0.6668   0.03000   0.01940  -0.0008   0.1747   1.0000
   6.750   0.7411   0.03273   0.02192  -0.0044   0.1340   1.0000
   7.000   0.7875   0.03596   0.02510  -0.0061   0.1175   1.0000
   7.250   0.8168   0.03870   0.02843  -0.0055   0.1107   1.0000
   7.500   0.8440   0.04215   0.03195  -0.0055   0.1035   1.0000
   7.750   0.8661   0.04574   0.03606  -0.0044   0.1029   1.0000
   8.000   0.8859   0.04982   0.04053  -0.0034   0.1036   1.0000
   8.250   0.9005   0.05393   0.04505  -0.0021   0.1039   1.0000
   8.500   0.9113   0.05809   0.04959  -0.0007   0.1041   1.0000
   8.750   0.9003   0.06147   0.05397   0.0027   0.1108   1.0000
   9.000   0.8998   0.06601   0.05883   0.0041   0.1140   1.0000
   9.250   0.8980   0.07066   0.06381   0.0054   0.1205   1.0000
   9.500   0.8807   0.07586   0.06931   0.0065   0.1289   1.0000
   9.750   0.8454   0.08042   0.07413   0.0075   0.1391   1.0000
  10.000   0.8090   0.08672   0.08054   0.0056   0.1528   1.0000
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