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MA409 (original) (modified line 7) (ma409-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: MA409 (original) (modified line 7) (ma409-il)
Reynolds number: 50,000
Max Cl/Cd: 42.94 at α=4.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-ma409-il-50000.txt
Download as CSV file: xf-ma409-il-50000.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 =     0.050 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.4382   0.10632   0.09973  -0.0199   1.0000   0.1838
  -8.750  -0.4207   0.10101   0.09441  -0.0181   1.0000   0.1943
  -8.500  -0.4411   0.10112   0.09471  -0.0194   1.0000   0.1979
  -8.250  -0.4234   0.09584   0.08941  -0.0168   1.0000   0.2081
  -8.000  -0.4318   0.09387   0.08760  -0.0178   1.0000   0.2137
  -7.750  -0.4286   0.09108   0.08481  -0.0172   1.0000   0.2244
  -7.500  -0.4218   0.08735   0.08113  -0.0156   1.0000   0.2328
  -7.250  -0.4246   0.08488   0.07876  -0.0168   1.0000   0.2424
  -7.000  -0.4260   0.08271   0.07667  -0.0189   1.0000   0.2545
  -6.750  -0.4225   0.07993   0.07395  -0.0194   1.0000   0.2682
  -6.500  -0.4168   0.07656   0.07062  -0.0180   1.0000   0.2826
  -6.250  -0.4090   0.07303   0.06713  -0.0129   1.0000   0.3031
  -6.000  -0.4062   0.07028   0.06445  -0.0122   1.0000   0.3270
  -5.750  -0.4020   0.06760   0.06181  -0.0114   1.0000   0.3533
  -5.500  -0.3964   0.06478   0.05902  -0.0100   1.0000   0.3807
  -5.250  -0.2822   0.04300   0.03553  -0.0589   1.0000   0.1536
  -5.000  -0.2531   0.03857   0.03071  -0.0616   1.0000   0.1643
  -4.750  -0.2252   0.03535   0.02712  -0.0632   1.0000   0.1895
  -4.500  -0.1945   0.03258   0.02382  -0.0650   1.0000   0.2134
  -4.250  -0.1643   0.03021   0.02101  -0.0661   1.0000   0.2228
  -4.000  -0.1327   0.02809   0.01834  -0.0674   1.0000   0.2293
  -3.750  -0.1045   0.02682   0.01676  -0.0679   1.0000   0.2428
  -3.500  -0.0752   0.02551   0.01510  -0.0684   1.0000   0.2512
  -3.250  -0.0463   0.02439   0.01375  -0.0687   1.0000   0.2608
  -3.000  -0.0177   0.02350   0.01265  -0.0690   1.0000   0.2757
  -2.750   0.0097   0.02280   0.01183  -0.0690   1.0000   0.3006
  -2.500   0.0391   0.02204   0.01100  -0.0692   1.0000   0.3346
  -2.250   0.0694   0.02107   0.01026  -0.0694   1.0000   0.4051
  -2.000   0.0961   0.01848   0.00936  -0.0679   1.0000   0.6916
  -1.750   0.1094   0.01800   0.00873  -0.0651   1.0000   1.0000
  -1.500   0.1356   0.01827   0.00854  -0.0652   1.0000   1.0000
  -1.250   0.1608   0.01857   0.00852  -0.0652   1.0000   1.0000
  -1.000   0.1856   0.01888   0.00860  -0.0652   1.0000   1.0000
  -0.750   0.2100   0.01922   0.00875  -0.0651   1.0000   1.0000
  -0.500   0.2341   0.01959   0.00895  -0.0650   1.0000   1.0000
  -0.250   0.2580   0.01998   0.00923  -0.0649   1.0000   1.0000
   0.000   0.2816   0.02040   0.00956  -0.0648   1.0000   1.0000
   0.250   0.3049   0.02086   0.00996  -0.0648   1.0000   1.0000
   0.500   0.3278   0.02136   0.01044  -0.0647   1.0000   1.0000
   0.750   0.3503   0.02192   0.01101  -0.0647   1.0000   1.0000
   1.000   0.3721   0.02256   0.01167  -0.0647   1.0000   1.0000
   1.250   0.3931   0.02329   0.01246  -0.0648   1.0000   1.0000
   1.500   0.4128   0.02419   0.01342  -0.0651   1.0000   1.0000
   1.750   0.4305   0.02530   0.01463  -0.0654   1.0000   1.0000
   2.000   0.4808   0.02677   0.01631  -0.0721   0.9840   1.0000
   2.250   0.5430   0.02800   0.01778  -0.0802   0.9577   1.0000
   2.500   0.5993   0.02890   0.01895  -0.0866   0.9308   1.0000
   2.750   0.6593   0.02924   0.01968  -0.0926   0.9015   1.0000
   3.000   0.7102   0.02894   0.01974  -0.0956   0.8663   1.0000
   3.250   0.7713   0.02778   0.01905  -0.0988   0.8344   1.0000
   3.500   0.8231   0.02659   0.01838  -0.1000   0.8050   1.0000
   3.750   0.8686   0.02542   0.01765  -0.0997   0.7715   1.0000
   4.000   0.9117   0.02290   0.01536  -0.0950   0.7111   1.0000
   4.250   0.9161   0.02235   0.01495  -0.0883   0.6316   1.0000
   4.500   0.9409   0.02191   0.01373  -0.0827   0.5276   1.0000
   4.750   0.9411   0.02290   0.01432  -0.0765   0.3843   1.0000
   5.000   0.9343   0.02734   0.01645  -0.0711   0.1513   1.0000
   5.250   0.9528   0.02999   0.01872  -0.0691   0.1127   1.0000
   5.500   0.9789   0.03242   0.02103  -0.0678   0.0986   1.0000
   5.750   1.0113   0.03496   0.02377  -0.0671   0.0904   1.0000
   6.000   1.0410   0.03799   0.02689  -0.0667   0.0837   1.0000
   6.250   1.0669   0.04087   0.03017  -0.0657   0.0787   1.0000
   6.500   1.0913   0.04444   0.03428  -0.0644   0.0777   1.0000
   6.750   1.1104   0.04847   0.03891  -0.0627   0.0782   1.0000
   7.000   1.1240   0.05290   0.04395  -0.0605   0.0798   1.0000
   7.250   1.1335   0.05758   0.04916  -0.0584   0.0818   1.0000
   7.500   1.1406   0.06248   0.05446  -0.0564   0.0839   1.0000
   7.750   1.1549   0.06826   0.06035  -0.0556   0.0862   1.0000
   8.000   1.1260   0.07281   0.06597  -0.0516   0.0910   1.0000
   8.250   1.1113   0.07836   0.07185  -0.0500   0.0944   1.0000
   8.500   1.1253   0.08448   0.07798  -0.0496   0.0995   1.0000
   8.750   1.0828   0.08884   0.08272  -0.0481   0.1011   1.0000
   9.000   1.0456   0.09467   0.08869  -0.0494   0.1024   1.0000
   9.250   1.0107   0.10287   0.09695  -0.0548   0.1045   1.0000
   9.500   0.9951   0.11136   0.10542  -0.0598   0.1095   1.0000
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