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MH 38 9.68% (mh38-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: MH 38 9.68% (mh38-il)
Reynolds number: 200,000
Max Cl/Cd: 86.83 at α=6.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-mh38-il-200000.txt
Download as CSV file: xf-mh38-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: MH 38  9.68%                                    
  
 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
  ------ -------- --------- --------- -------- -------- --------
   0.250   0.4650   0.01058   0.00466  -0.0976   0.7596   1.0000
   0.500   0.4921   0.01068   0.00468  -0.0974   0.7533   1.0000
   0.750   0.5194   0.01080   0.00471  -0.0973   0.7476   1.0000
   1.000   0.5463   0.01092   0.00479  -0.0971   0.7407   1.0000
   1.250   0.5740   0.01102   0.00480  -0.0969   0.7355   1.0000
   1.500   0.6005   0.01116   0.00493  -0.0968   0.7279   1.0000
   1.750   0.6284   0.01124   0.00493  -0.0966   0.7227   1.0000
   2.000   0.6548   0.01139   0.00510  -0.0964   0.7148   1.0000
   2.250   0.6827   0.01147   0.00513  -0.0962   0.7093   1.0000
   2.500   0.7090   0.01162   0.00531  -0.0960   0.7010   1.0000
   3.000   0.7631   0.01182   0.00551  -0.0956   0.6864   1.0000
   3.250   0.7904   0.01191   0.00561  -0.0953   0.6794   1.0000
   3.500   0.8172   0.01200   0.00573  -0.0950   0.6710   1.0000
   3.750   0.8438   0.01211   0.00587  -0.0947   0.6624   1.0000
   4.000   0.8715   0.01216   0.00590  -0.0945   0.6548   1.0000
   4.250   0.8974   0.01226   0.00613  -0.0941   0.6447   1.0000
   4.500   0.9239   0.01236   0.00627  -0.0937   0.6350   1.0000
   4.750   0.9508   0.01242   0.00636  -0.0934   0.6256   1.0000
   5.000   0.9772   0.01248   0.00647  -0.0929   0.6146   1.0000
   5.250   1.0029   0.01255   0.00663  -0.0924   0.6022   1.0000
   5.500   1.0285   0.01261   0.00675  -0.0918   0.5886   1.0000
   5.750   1.0539   0.01265   0.00689  -0.0911   0.5732   1.0000
   6.000   1.0779   0.01270   0.00704  -0.0903   0.5525   1.0000
   6.250   1.1011   0.01277   0.00712  -0.0892   0.5261   1.0000
   6.500   1.1236   0.01294   0.00730  -0.0880   0.4964   1.0000
   6.750   1.1449   0.01325   0.00757  -0.0868   0.4612   1.0000
   7.000   1.1625   0.01382   0.00794  -0.0850   0.4086   1.0000
   7.250   1.1751   0.01479   0.00857  -0.0827   0.3350   1.0000
   7.500   1.1814   0.01632   0.00955  -0.0798   0.2430   1.0000
   7.750   1.1791   0.01859   0.01104  -0.0760   0.1280   1.0000
   8.000   1.1689   0.02152   0.01321  -0.0711   0.0442   1.0000
   8.250   1.1624   0.02373   0.01545  -0.0662   0.0288   1.0000
   8.500   1.1688   0.02502   0.01683  -0.0634   0.0228   1.0000
   8.750   1.1605   0.02746   0.01932  -0.0593   0.0195   1.0000
   9.000   1.1677   0.02898   0.02098  -0.0570   0.0180   1.0000
   9.250   1.1737   0.03082   0.02292  -0.0547   0.0170   1.0000
   9.500   1.1822   0.03276   0.02496  -0.0526   0.0160   1.0000
   9.750   1.1948   0.03474   0.02704  -0.0509   0.0153   1.0000
  10.000   1.2124   0.03688   0.02931  -0.0495   0.0149   1.0000
  10.250   1.2346   0.03934   0.03197  -0.0485   0.0147   1.0000
  10.500   1.2576   0.04252   0.03546  -0.0476   0.0149   1.0000
  10.750   1.2737   0.04641   0.03974  -0.0463   0.0155   1.0000
  11.000   1.2785   0.05041   0.04412  -0.0443   0.0162   1.0000
  11.250   1.2754   0.05450   0.04857  -0.0421   0.0168   1.0000
  11.500   1.2673   0.05851   0.05289  -0.0401   0.0172   1.0000
  11.750   1.2566   0.06254   0.05718  -0.0385   0.0175   1.0000
  12.000   1.2433   0.06677   0.06167  -0.0374   0.0177   1.0000
  12.250   1.2273   0.07162   0.06675  -0.0369   0.0181   1.0000
  12.500   1.2098   0.07671   0.07207  -0.0371   0.0183   1.0000
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