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MA409 (original) (modified line 7) (ma409-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: MA409 (original) (modified line 7) (ma409-il)
Reynolds number: 500,000
Max Cl/Cd: 99.3 at α=1°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-ma409-il-500000.txt
Download as CSV file: xf-ma409-il-500000.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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.250  -0.4392   0.09967   0.09754  -0.0265   1.0000   0.0186
 -10.000  -0.4324   0.09747   0.09535  -0.0262   1.0000   0.0193
  -9.750  -0.4287   0.09482   0.09272  -0.0265   1.0000   0.0198
  -9.500  -0.4272   0.09211   0.09004  -0.0267   1.0000   0.0203
  -9.250  -0.4204   0.08868   0.08662  -0.0290   0.9976   0.0210
  -9.000  -0.4094   0.08456   0.08251  -0.0331   0.9930   0.0220
  -8.750  -0.3994   0.08017   0.07813  -0.0378   0.9860   0.0231
  -7.750  -0.3229   0.01991   0.01586  -0.1036   0.9405   0.0124
  -7.500  -0.2968   0.01714   0.01257  -0.1043   0.9332   0.0124
  -7.250  -0.2692   0.01517   0.01026  -0.1048   0.9270   0.0130
  -7.000  -0.2420   0.01383   0.00867  -0.1049   0.9191   0.0138
  -6.750  -0.2142   0.01273   0.00735  -0.1050   0.9103   0.0147
  -6.500  -0.1865   0.01186   0.00627  -0.1049   0.8989   0.0158
  -6.250  -0.1590   0.01127   0.00551  -0.1048   0.8853   0.0172
  -6.000  -0.1316   0.01043   0.00443  -0.1046   0.8705   0.0184
  -5.750  -0.1041   0.00975   0.00357  -0.1043   0.8552   0.0221
  -5.500  -0.0768   0.00926   0.00303  -0.1040   0.8412   0.0341
  -5.250  -0.0492   0.00942   0.00321  -0.1037   0.8288   0.0595
  -5.000  -0.0217   0.00937   0.00311  -0.1036   0.8205   0.0670
  -4.750   0.0061   0.00943   0.00305  -0.1035   0.8149   0.0721
  -4.500   0.0335   0.00923   0.00277  -0.1033   0.8094   0.0758
  -4.250   0.0609   0.00904   0.00255  -0.1032   0.8045   0.0810
  -4.000   0.0886   0.00895   0.00239  -0.1032   0.8004   0.0856
  -3.750   0.1162   0.00873   0.00211  -0.1031   0.7963   0.0903
  -3.250   0.1716   0.00842   0.00171  -0.1029   0.7889   0.0999
  -3.000   0.1994   0.00825   0.00156  -0.1029   0.7853   0.1085
  -2.750   0.2272   0.00811   0.00145  -0.1028   0.7820   0.1246
  -2.500   0.2549   0.00787   0.00134  -0.1028   0.7792   0.1699
  -2.250   0.2825   0.00756   0.00133  -0.1030   0.7767   0.2642
  -2.000   0.3102   0.00741   0.00133  -0.1030   0.7736   0.3201
  -1.750   0.3377   0.00732   0.00135  -0.1029   0.7696   0.3648
  -1.500   0.3652   0.00725   0.00139  -0.1029   0.7656   0.4213
  -1.250   0.3925   0.00711   0.00142  -0.1028   0.7611   0.4791
  -1.000   0.4184   0.00679   0.00150  -0.1025   0.7542   0.6155
  -0.750   0.4445   0.00601   0.00157  -0.1016   0.7469   1.0000
  -0.500   0.4719   0.00611   0.00161  -0.1015   0.7404   1.0000
  -0.250   0.4994   0.00622   0.00167  -0.1013   0.7343   1.0000
   0.000   0.5268   0.00630   0.00172  -0.1011   0.7270   1.0000
   0.250   0.5538   0.00638   0.00178  -0.1009   0.7175   1.0000
   0.500   0.5808   0.00644   0.00183  -0.1006   0.7071   1.0000
   0.750   0.6079   0.00647   0.00186  -0.1004   0.6963   1.0000
   1.000   0.6355   0.00640   0.00188  -0.1003   0.6785   1.0000
   1.250   0.6555   0.00671   0.00172  -0.0985   0.5784   1.0000
   1.500   0.6791   0.00724   0.00204  -0.0977   0.5501   1.0000
   1.750   0.7057   0.00741   0.00223  -0.0975   0.5302   1.0000
   2.000   0.7323   0.00753   0.00232  -0.0972   0.4894   1.0000
   2.250   0.7557   0.00798   0.00246  -0.0965   0.4240   1.0000
   2.500   0.7770   0.00871   0.00273  -0.0954   0.3339   1.0000
   2.750   0.7985   0.00949   0.00307  -0.0945   0.2298   1.0000
   3.000   0.8149   0.01106   0.00376  -0.0929   0.0718   1.0000
   3.250   0.8374   0.01190   0.00436  -0.0920   0.0223   1.0000
   3.500   0.8625   0.01239   0.00495  -0.0913   0.0186   1.0000
   3.750   0.8866   0.01302   0.00571  -0.0905   0.0172   1.0000
   4.000   0.9107   0.01362   0.00638  -0.0897   0.0164   1.0000
   4.250   0.9343   0.01426   0.00711  -0.0889   0.0150   1.0000
   4.500   0.9567   0.01508   0.00801  -0.0878   0.0142   1.0000
   4.750   0.9783   0.01603   0.00904  -0.0866   0.0134   1.0000
   5.000   0.9991   0.01715   0.01024  -0.0853   0.0128   1.0000
   5.250   1.0198   0.01848   0.01165  -0.0838   0.0126   1.0000
   5.500   1.0413   0.02014   0.01346  -0.0823   0.0132   1.0000
   5.750   1.0643   0.02299   0.01654  -0.0806   0.0156   1.0000
  13.750   0.7464   0.16685   0.16492  -0.0750   0.0135   1.0000
  14.000   0.7456   0.17026   0.16834  -0.0763   0.0131   1.0000
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