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MA409 (original) (modified line 7) (ma409-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9


Details Polar file
Airfoil: MA409 (original) (modified line 7) (ma409-il)
Reynolds number: 1,000,000
Max Cl/Cd: 110.64 at α=1.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-ma409-il-1000000.txt
Download as CSV file: xf-ma409-il-1000000.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 =     1.000 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.4269   0.02887   0.02563  -0.0954   0.7470   0.0070
  -8.500  -0.4053   0.02355   0.01978  -0.0991   0.7465   0.0071
  -8.250  -0.3820   0.02006   0.01584  -0.1007   0.7457   0.0072
  -8.000  -0.3567   0.01793   0.01338  -0.1013   0.7445   0.0071
  -7.750  -0.3312   0.01591   0.01104  -0.1016   0.7430   0.0071
  -7.500  -0.3057   0.01383   0.00863  -0.1019   0.7417   0.0074
  -7.250  -0.2791   0.01266   0.00727  -0.1020   0.7409   0.0080
  -7.000  -0.2519   0.01199   0.00648  -0.1020   0.7404   0.0087
  -6.750  -0.2244   0.01139   0.00576  -0.1020   0.7397   0.0094
  -6.500  -0.1968   0.01085   0.00510  -0.1019   0.7387   0.0100
  -6.250  -0.1691   0.01039   0.00453  -0.1019   0.7378   0.0106
  -5.750  -0.1134   0.00931   0.00319  -0.1018   0.7363   0.0132
  -5.500  -0.0854   0.00891   0.00272  -0.1017   0.7353   0.0155
  -5.250  -0.0573   0.00864   0.00237  -0.1017   0.7341   0.0176
  -5.000  -0.0295   0.00826   0.00216  -0.1016   0.7325   0.0476
  -4.750  -0.0015   0.00820   0.00212  -0.1016   0.7307   0.0575
  -4.500   0.0267   0.00825   0.00216  -0.1016   0.7292   0.0601
  -4.250   0.0546   0.00805   0.00194  -0.1016   0.7278   0.0651
  -4.000   0.0827   0.00796   0.00185  -0.1016   0.7264   0.0687
  -3.750   0.1108   0.00794   0.00179  -0.1015   0.7248   0.0721
  -3.500   0.1389   0.00787   0.00169  -0.1015   0.7232   0.0751
  -3.250   0.1669   0.00769   0.00154  -0.1015   0.7216   0.0812
  -3.000   0.1951   0.00758   0.00143  -0.1015   0.7200   0.0849
  -2.750   0.2232   0.00748   0.00133  -0.1015   0.7181   0.0888
  -2.500   0.2513   0.00735   0.00123  -0.1015   0.7161   0.0998
  -2.250   0.2792   0.00717   0.00118  -0.1016   0.7137   0.1373
  -2.000   0.3069   0.00702   0.00115  -0.1016   0.7102   0.1836
  -1.750   0.3347   0.00674   0.00111  -0.1016   0.7057   0.2547
  -1.500   0.3625   0.00661   0.00108  -0.1016   0.7005   0.2971
  -1.250   0.3903   0.00650   0.00106  -0.1016   0.6959   0.3298
  -1.000   0.4184   0.00635   0.00103  -0.1016   0.6888   0.3652
  -0.750   0.4464   0.00615   0.00102  -0.1017   0.6783   0.4257
  -0.500   0.4708   0.00623   0.00089  -0.1010   0.5843   0.4700
  -0.250   0.4965   0.00622   0.00104  -0.1008   0.5696   0.5573
   0.000   0.5164   0.00549   0.00121  -0.0990   0.5628   0.8926
   0.250   0.5480   0.00547   0.00127  -0.0998   0.5546   1.0000
   0.500   0.5755   0.00557   0.00136  -0.0997   0.5460   1.0000
   0.750   0.6029   0.00568   0.00143  -0.0996   0.5357   1.0000
   1.000   0.6306   0.00575   0.00147  -0.0995   0.5123   1.0000
   1.250   0.6572   0.00594   0.00149  -0.0993   0.4681   1.0000
   1.500   0.6822   0.00629   0.00162  -0.0988   0.4162   1.0000
   1.750   0.7058   0.00681   0.00182  -0.0982   0.3438   1.0000
   2.000   0.7314   0.00712   0.00199  -0.0978   0.3120   1.0000
   2.250   0.7533   0.00789   0.00227  -0.0970   0.2037   1.0000
   2.500   0.7739   0.00889   0.00270  -0.0959   0.0892   1.0000
   2.750   0.7965   0.00972   0.00318  -0.0950   0.0172   1.0000
   3.000   0.8227   0.00999   0.00351  -0.0947   0.0154   1.0000
   3.250   0.8485   0.01033   0.00390  -0.0943   0.0136   1.0000
   3.500   0.8731   0.01088   0.00454  -0.0936   0.0113   1.0000
   3.750   0.8972   0.01151   0.00525  -0.0928   0.0103   1.0000
   4.000   0.9223   0.01194   0.00571  -0.0923   0.0099   1.0000
   4.250   0.9469   0.01244   0.00629  -0.0917   0.0094   1.0000
   4.500   0.9709   0.01302   0.00692  -0.0910   0.0087   1.0000
   4.750   0.9943   0.01367   0.00762  -0.0902   0.0080   1.0000
   5.000   1.0169   0.01444   0.00844  -0.0892   0.0075   1.0000
   5.250   1.0383   0.01543   0.00949  -0.0881   0.0072   1.0000
   5.500   1.0592   0.01659   0.01073  -0.0867   0.0071   1.0000
   5.750   1.0801   0.01780   0.01200  -0.0856   0.0068   1.0000
   6.000   1.0991   0.01994   0.01421  -0.0842   0.0064   1.0000
   6.250   1.1207   0.02168   0.01607  -0.0831   0.0064   1.0000
   6.500   1.1421   0.02350   0.01803  -0.0819   0.0064   1.0000
   6.750   1.1623   0.02566   0.02037  -0.0807   0.0064   1.0000
   7.000   1.1791   0.02885   0.02382  -0.0789   0.0063   1.0000
   7.250   1.1873   0.03421   0.02963  -0.0758   0.0062   1.0000
   7.500   1.1976   0.03611   0.03188  -0.0731   0.0061   1.0000
   7.750   1.2123   0.03725   0.03325  -0.0711   0.0058   1.0000
   9.500   1.0451   0.05725   0.05486  -0.0355   0.0062   1.0000
   9.750   1.0218   0.06167   0.05942  -0.0343   0.0062   1.0000
  10.000   0.9978   0.06683   0.06470  -0.0343   0.0062   1.0000
  10.250   0.9735   0.07256   0.07054  -0.0351   0.0062   1.0000
  10.500   0.9489   0.07913   0.07723  -0.0374   0.0062   1.0000
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