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MH 121 8.76% (mh121-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: MH 121 8.76% (mh121-il)
Reynolds number: 200,000
Max Cl/Cd: 83.98 at α=4.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-mh121-il-200000.txt
Download as CSV file: xf-mh121-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: MH 121  8.76%                                   
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.3915   0.08989   0.08696  -0.0286   1.0000   0.0303
  -8.000  -0.3993   0.08712   0.08423  -0.0276   1.0000   0.0307
  -7.750  -0.4077   0.08433   0.08148  -0.0266   1.0000   0.0311
  -7.500  -0.4176   0.08145   0.07864  -0.0257   1.0000   0.0317
  -7.250  -0.4289   0.07862   0.07585  -0.0247   1.0000   0.0319
  -7.000  -0.4295   0.07380   0.07106  -0.0279   0.9972   0.0326
  -6.750  -0.5252   0.07811   0.07526  -0.0309   0.9998   0.0285
  -6.500  -0.5079   0.07283   0.06993  -0.0374   0.9963   0.0297
  -6.250  -0.4879   0.06641   0.06341  -0.0460   0.9922   0.0308
  -6.000  -0.4703   0.06060   0.05744  -0.0526   0.9880   0.0325
  -5.750  -0.4459   0.05466   0.05123  -0.0595   0.9843   0.0348
  -5.500  -0.4041   0.05154   0.04726  -0.0669   0.9808   0.0389
  -5.250  -0.3924   0.04347   0.03916  -0.0696   0.9778   0.0411
  -5.000  -0.3691   0.04020   0.03579  -0.0714   0.9751   0.0442
  -4.750  -0.3365   0.03640   0.03140  -0.0748   0.9727   0.0542
  -4.500  -0.3079   0.03423   0.02912  -0.0773   0.9708   0.0722
  -4.250  -0.2735   0.03158   0.02605  -0.0805   0.9694   0.0938
  -4.000  -0.2242   0.02551   0.01863  -0.0786   0.9694   0.0273
  -3.750  -0.1958   0.02308   0.01596  -0.0787   0.9676   0.0250
  -3.500  -0.1674   0.02119   0.01383  -0.0786   0.9652   0.0240
  -3.250  -0.1370   0.01970   0.01220  -0.0789   0.9631   0.0240
  -3.000  -0.1051   0.01853   0.01092  -0.0797   0.9611   0.0250
  -2.750  -0.0703   0.01744   0.00968  -0.0813   0.9595   0.0294
  -2.500  -0.0364   0.01493   0.00924  -0.0839   0.9588   0.5446
  -2.250  -0.0029   0.01511   0.00938  -0.0851   0.9569   0.5943
  -2.000   0.0190   0.01518   0.00941  -0.0840   0.9526   0.6210
  -1.750   0.0465   0.01529   0.00942  -0.0840   0.9491   0.6497
  -1.500   0.0764   0.01541   0.00959  -0.0844   0.9465   0.6838
  -1.250   0.1104   0.01552   0.00968  -0.0857   0.9444   0.7072
  -1.000   0.1322   0.01557   0.00973  -0.0847   0.9397   0.7232
  -0.750   0.1606   0.01562   0.00976  -0.0851   0.9359   0.7401
  -0.500   0.1936   0.01565   0.00975  -0.0863   0.9332   0.7583
  -0.250   0.2291   0.01568   0.00981  -0.0880   0.9312   0.7782
   0.000   0.2463   0.01573   0.00994  -0.0861   0.9249   0.8003
   0.250   0.2759   0.01570   0.00999  -0.0865   0.9213   0.8279
   0.500   0.3087   0.01561   0.01002  -0.0874   0.9188   0.8650
   0.750   0.3362   0.01545   0.01004  -0.0875   0.9132   0.9929
   1.000   0.3720   0.01556   0.01009  -0.0897   0.9091   1.0000
   1.250   0.4130   0.01559   0.01009  -0.0926   0.9064   1.0000
   1.500   0.4415   0.01575   0.01024  -0.0932   0.9005   1.0000
   1.750   0.4774   0.01575   0.01026  -0.0949   0.8957   1.0000
   2.000   0.5219   0.01557   0.01016  -0.0982   0.8929   1.0000
   2.250   0.5491   0.01562   0.01025  -0.0981   0.8848   1.0000
   2.500   0.5973   0.01516   0.00988  -0.1017   0.8811   1.0000
   2.750   0.6395   0.01465   0.00947  -0.1040   0.8739   1.0000
   3.000   0.7138   0.01288   0.00797  -0.1112   0.8667   1.0000
   3.250   0.7599   0.01186   0.00708  -0.1133   0.8534   1.0000
   3.500   0.7923   0.01126   0.00660  -0.1131   0.8355   1.0000
   3.750   0.8240   0.01061   0.00605  -0.1124   0.8095   1.0000
   4.000   0.8501   0.01021   0.00571  -0.1109   0.7728   1.0000
   4.250   0.8658   0.01031   0.00498  -0.1064   0.5801   1.0000
   4.500   0.8489   0.01290   0.00580  -0.0981   0.2898   1.0000
   4.750   0.8501   0.01513   0.00681  -0.0940   0.0929   1.0000
   5.000   0.8575   0.01750   0.00881  -0.0901   0.0149   1.0000
   5.250   0.8757   0.01868   0.01016  -0.0881   0.0129   1.0000
   5.500   0.8934   0.02017   0.01179  -0.0859   0.0118   1.0000
   5.750   0.9144   0.02201   0.01377  -0.0844   0.0113   1.0000
   6.000   0.9418   0.02437   0.01630  -0.0838   0.0113   1.0000
   6.250   0.9725   0.02729   0.01951  -0.0836   0.0116   1.0000
   6.500   1.0001   0.03073   0.02336  -0.0828   0.0122   1.0000
   6.750   1.0215   0.03446   0.02754  -0.0811   0.0130   1.0000
   7.000   1.0357   0.03911   0.03264  -0.0786   0.0139   1.0000
   9.250   1.0249   0.07810   0.07471  -0.0483   0.0313   1.0000
   9.500   1.0055   0.08198   0.07872  -0.0463   0.0311   1.0000
   9.750   0.9853   0.08645   0.08331  -0.0458   0.0311   1.0000
  10.000   0.9644   0.09176   0.08873  -0.0470   0.0311   1.0000
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