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MA409 (original) (modified line 7) (ma409-il) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: MA409 (original) (modified line 7) (ma409-il)
Reynolds number: 500,000
Max Cl/Cd: 89.3 at α=1.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-ma409-il-500000-n5.txt
Download as CSV file: xf-ma409-il-500000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: MA409 (original)                                
  
 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
  ------ -------- --------- --------- -------- -------- --------
 -10.500  -0.3136   0.08532   0.08221  -0.0386   0.7422   0.0057
 -10.250  -0.3139   0.08099   0.07790  -0.0397   0.7419   0.0056
 -10.000  -0.3153   0.07638   0.07330  -0.0410   0.7415   0.0054
  -9.750  -0.3172   0.07164   0.06859  -0.0423   0.7410   0.0053
  -9.500  -0.3201   0.06673   0.06369  -0.0439   0.7406   0.0052
  -9.250  -0.3247   0.06137   0.05836  -0.0459   0.7403   0.0051
  -9.000  -0.3305   0.05578   0.05281  -0.0483   0.7398   0.0050
  -8.500  -0.3955   0.03089   0.02712  -0.0947   0.7388   0.0041
  -8.250  -0.3717   0.02500   0.02059  -0.0995   0.7382   0.0041
  -8.000  -0.3472   0.02165   0.01675  -0.1014   0.7377   0.0042
  -7.750  -0.3216   0.01935   0.01404  -0.1023   0.7372   0.0043
  -7.500  -0.2955   0.01765   0.01204  -0.1028   0.7366   0.0044
  -7.250  -0.2689   0.01632   0.01044  -0.1030   0.7357   0.0047
  -7.000  -0.2420   0.01518   0.00906  -0.1032   0.7346   0.0049
  -6.750  -0.2149   0.01422   0.00788  -0.1032   0.7333   0.0054
  -6.500  -0.1875   0.01341   0.00689  -0.1032   0.7321   0.0058
  -6.000  -0.1326   0.01196   0.00510  -0.1033   0.7296   0.0075
  -5.750  -0.1047   0.01141   0.00441  -0.1033   0.7283   0.0084
  -5.500  -0.0768   0.01097   0.00384  -0.1033   0.7272   0.0095
  -5.250  -0.0488   0.01064   0.00338  -0.1032   0.7263   0.0107
  -5.000  -0.0206   0.01025   0.00288  -0.1032   0.7253   0.0164
  -4.750   0.0070   0.00981   0.00257  -0.1032   0.7240   0.0448
  -4.500   0.0349   0.00969   0.00248  -0.1032   0.7224   0.0579
  -4.250   0.0629   0.00965   0.00238  -0.1032   0.7206   0.0627
  -4.000   0.0907   0.00951   0.00223  -0.1032   0.7191   0.0693
  -3.750   0.1186   0.00946   0.00213  -0.1032   0.7174   0.0747
  -3.500   0.1467   0.00935   0.00198  -0.1031   0.7159   0.0769
  -3.250   0.1746   0.00918   0.00181  -0.1031   0.7138   0.0811
  -3.000   0.2026   0.00904   0.00167  -0.1031   0.7113   0.0854
  -2.750   0.2306   0.00894   0.00157  -0.1031   0.7090   0.0904
  -2.500   0.2585   0.00881   0.00148  -0.1032   0.7067   0.1027
  -2.250   0.2862   0.00868   0.00142  -0.1032   0.7043   0.1289
  -2.000   0.3141   0.00856   0.00137  -0.1032   0.7017   0.1552
  -1.750   0.3420   0.00838   0.00132  -0.1033   0.6980   0.1869
  -1.500   0.3698   0.00818   0.00132  -0.1033   0.6942   0.2461
  -1.250   0.3974   0.00806   0.00131  -0.1033   0.6893   0.2911
  -1.000   0.4254   0.00788   0.00129  -0.1034   0.6756   0.3235
  -0.750   0.4484   0.00793   0.00107  -0.1023   0.5809   0.3495
  -0.500   0.4741   0.00805   0.00120  -0.1020   0.5683   0.3881
  -0.250   0.5007   0.00807   0.00132  -0.1019   0.5601   0.4423
   0.000   0.5273   0.00805   0.00145  -0.1017   0.5505   0.4969
   0.250   0.5515   0.00758   0.00163  -0.1012   0.5410   0.7240
   0.500   0.5780   0.00710   0.00168  -0.1005   0.5322   1.0000
   0.750   0.6056   0.00718   0.00176  -0.1005   0.5229   1.0000
   1.000   0.6332   0.00725   0.00185  -0.1004   0.5066   1.0000
   1.250   0.6599   0.00739   0.00192  -0.1002   0.4798   1.0000
   1.500   0.6847   0.00772   0.00204  -0.0997   0.4356   1.0000
   1.750   0.7066   0.00838   0.00228  -0.0988   0.3449   1.0000
   2.000   0.7305   0.00887   0.00253  -0.0982   0.2972   1.0000
   2.250   0.7516   0.00972   0.00287  -0.0973   0.1856   1.0000
   2.500   0.7725   0.01068   0.00332  -0.0964   0.0843   1.0000
   2.750   0.7947   0.01153   0.00386  -0.0955   0.0173   1.0000
   3.000   0.8203   0.01188   0.00427  -0.0950   0.0119   1.0000
   3.250   0.8459   0.01223   0.00470  -0.0946   0.0103   1.0000
   3.500   0.8708   0.01267   0.00521  -0.0940   0.0086   1.0000
   3.750   0.8940   0.01341   0.00605  -0.0932   0.0070   1.0000
   4.000   0.9182   0.01394   0.00667  -0.0925   0.0065   1.0000
   4.250   0.9415   0.01462   0.00742  -0.0917   0.0059   1.0000
   4.500   0.9639   0.01541   0.00829  -0.0907   0.0055   1.0000
   4.750   0.9858   0.01625   0.00920  -0.0896   0.0052   1.0000
   5.000   1.0082   0.01695   0.00993  -0.0888   0.0046   1.0000
   5.250   1.0283   0.01807   0.01111  -0.0876   0.0041   1.0000
   5.500   1.0496   0.01910   0.01230  -0.0864   0.0039   1.0000
   5.750   1.0702   0.02039   0.01372  -0.0851   0.0036   1.0000
   6.000   1.0908   0.02187   0.01535  -0.0837   0.0034   1.0000
   6.250   1.1113   0.02353   0.01718  -0.0824   0.0032   1.0000
   6.500   1.1314   0.02541   0.01926  -0.0810   0.0031   1.0000
   6.750   1.1504   0.02754   0.02164  -0.0794   0.0030   1.0000
   7.000   1.1674   0.03005   0.02444  -0.0776   0.0030   1.0000
   7.250   1.1815   0.03299   0.02774  -0.0753   0.0030   1.0000
   7.500   1.1919   0.03640   0.03153  -0.0726   0.0030   1.0000
   7.750   1.1981   0.04009   0.03560  -0.0695   0.0031   1.0000
   8.000   1.1996   0.04393   0.03982  -0.0661   0.0032   1.0000
   8.250   1.1968   0.04782   0.04405  -0.0626   0.0033   1.0000
   8.500   1.1899   0.05160   0.04812  -0.0590   0.0034   1.0000
   8.750   1.1777   0.05493   0.05170  -0.0550   0.0034   1.0000
   9.000   1.1611   0.05828   0.05525  -0.0512   0.0035   1.0000
   9.250   1.1435   0.06198   0.05914  -0.0486   0.0034   1.0000
   9.500   1.1246   0.06630   0.06364  -0.0475   0.0034   1.0000
   9.750   1.1048   0.07145   0.06895  -0.0482   0.0035   1.0000
  10.000   1.0849   0.07752   0.07518  -0.0507   0.0035   1.0000
  10.250   1.0652   0.08496   0.08276  -0.0555   0.0036   1.0000
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