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R140 (original) (r140-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9


Details Polar file
Airfoil: R140 (original) (r140-il)
Reynolds number: 1,000,000
Max Cl/Cd: 50.79 at α=5.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-r140-il-1000000.txt
Download as CSV file: xf-r140-il-1000000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: R140 (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
  ------ -------- --------- --------- -------- -------- --------
 -11.500  -0.4617   0.09433   0.09284  -0.0315   1.0000   0.0065
 -11.250  -0.4805   0.08513   0.08366  -0.0348   1.0000   0.0061
 -11.000  -0.5001   0.07272   0.07126  -0.0398   1.0000   0.0056
 -10.750  -0.5082   0.06500   0.06349  -0.0435   1.0000   0.0054
 -10.500  -0.5129   0.06024   0.05868  -0.0458   1.0000   0.0053
 -10.250  -0.5220   0.05557   0.05395  -0.0479   1.0000   0.0052
 -10.000  -0.5367   0.05086   0.04916  -0.0496   1.0000   0.0052
  -9.750  -0.5551   0.04655   0.04475  -0.0506   1.0000   0.0052
  -9.500  -0.5769   0.04255   0.04065  -0.0503   1.0000   0.0052
  -9.250  -0.5983   0.03929   0.03729  -0.0486   1.0000   0.0052
  -9.000  -0.6190   0.03642   0.03430  -0.0453   1.0000   0.0052
  -8.750  -0.6389   0.03405   0.03183  -0.0403   1.0000   0.0052
  -8.500  -0.6557   0.03193   0.02961  -0.0348   1.0000   0.0052
  -8.250  -0.6622   0.02852   0.02600  -0.0316   0.9982   0.0052
  -6.000  -0.5144   0.01283   0.00759  -0.0280   0.9510   0.0065
  -5.750  -0.4844   0.01238   0.00705  -0.0284   0.9431   0.0054
  -5.500  -0.4680   0.01129   0.00580  -0.0261   0.9301   0.0047
  -5.250  -0.4495   0.01052   0.00486  -0.0241   0.9164   0.0043
  -5.000  -0.4276   0.01005   0.00421  -0.0228   0.9047   0.0041
  -4.750  -0.4050   0.00968   0.00370  -0.0217   0.8935   0.0041
  -4.500  -0.3821   0.00940   0.00330  -0.0206   0.8815   0.0042
  -4.250  -0.3594   0.00912   0.00290  -0.0195   0.8658   0.0049
  -4.000  -0.3366   0.00895   0.00261  -0.0184   0.8386   0.0055
  -3.750  -0.3137   0.00887   0.00238  -0.0174   0.8195   0.0072
  -3.500  -0.2911   0.00868   0.00221  -0.0164   0.8063   0.0244
  -3.250  -0.2671   0.00852   0.00207  -0.0157   0.7919   0.0407
  -2.750  -0.2182   0.00837   0.00179  -0.0145   0.7526   0.0486
  -2.500  -0.1942   0.00827   0.00167  -0.0138   0.7393   0.0615
  -2.250  -0.1712   0.00808   0.00156  -0.0130   0.7281   0.1015
  -1.750  -0.1471   0.00610   0.00101  -0.0076   0.7050   0.5067
  -1.500  -0.1250   0.00589   0.00098  -0.0066   0.6940   0.5640
  -1.250  -0.1001   0.00585   0.00095  -0.0060   0.6825   0.5908
  -1.000  -0.0750   0.00583   0.00094  -0.0056   0.6724   0.6088
  -0.750  -0.0493   0.00583   0.00093  -0.0052   0.6631   0.6270
  -0.500  -0.0228   0.00582   0.00092  -0.0050   0.6545   0.6379
  -0.250   0.0035   0.00583   0.00091  -0.0048   0.6467   0.6484
   0.000   0.0303   0.00582   0.00090  -0.0047   0.6377   0.6570
   0.250   0.0565   0.00581   0.00089  -0.0044   0.6258   0.6648
   0.750   0.1089   0.00583   0.00088  -0.0040   0.6069   0.6797
   1.000   0.1355   0.00584   0.00089  -0.0038   0.6001   0.6871
   1.250   0.1611   0.00583   0.00091  -0.0035   0.5932   0.6970
   1.500   0.1865   0.00579   0.00094  -0.0031   0.5861   0.7103
   1.750   0.2108   0.00576   0.00100  -0.0024   0.5764   0.7314
   2.000   0.2339   0.00572   0.00106  -0.0014   0.5629   0.7692
   2.250   0.2579   0.00572   0.00113  -0.0007   0.5477   0.7994
   2.500   0.2826   0.00575   0.00118  -0.0001   0.5291   0.8190
   2.750   0.2972   0.00638   0.00131   0.0024   0.3965   0.8336
   3.000   0.2978   0.00810   0.00193   0.0071   0.0977   0.8506
   3.250   0.3157   0.00868   0.00219   0.0089   0.0074   0.8646
   3.500   0.3404   0.00880   0.00232   0.0094   0.0058   0.8772
   3.750   0.3658   0.00890   0.00253   0.0099   0.0056   0.8909
   4.000   0.3922   0.00904   0.00277   0.0101   0.0055   0.9059
   4.250   0.4203   0.00920   0.00305   0.0100   0.0056   0.9229
   4.500   0.4495   0.00942   0.00339   0.0095   0.0058   0.9366
   4.750   0.4816   0.00971   0.00376   0.0084   0.0061   0.9465
   5.000   0.5135   0.01013   0.00429   0.0074   0.0065   0.9553
   5.250   0.5485   0.01080   0.00510   0.0056   0.0070   0.9602
   5.500   0.5772   0.01195   0.00642   0.0050   0.0074   0.9673
   5.750   0.6095   0.01292   0.00748   0.0036   0.0073   0.9721
   6.000   0.6382   0.01374   0.00837   0.0031   0.0070   0.9784
   6.750   0.7241   0.01811   0.01310   0.0018   0.0067   0.9894
   7.000   0.7532   0.01902   0.01411   0.0011   0.0056   0.9925
   7.250   0.7810   0.02018   0.01537   0.0003   0.0046   0.9950
   7.500   0.8083   0.02226   0.01764  -0.0003   0.0042   0.9970
   8.000   0.8174   0.03264   0.02898   0.0059   0.0036   1.0000
   8.250   0.8144   0.03502   0.03158   0.0107   0.0036   1.0000
   8.500   0.8119   0.03725   0.03401   0.0151   0.0036   1.0000
   8.750   0.8049   0.03995   0.03692   0.0199   0.0036   1.0000
   9.000   0.7951   0.04271   0.03988   0.0246   0.0036   1.0000
   9.250   0.7825   0.04506   0.04238   0.0297   0.0036   1.0000
   9.500   0.7660   0.04737   0.04483   0.0347   0.0036   1.0000
   9.750   0.7512   0.05002   0.04761   0.0380   0.0036   1.0000
  10.000   0.7359   0.05312   0.05085   0.0402   0.0036   1.0000
  10.250   0.7196   0.05679   0.05465   0.0411   0.0036   1.0000
  10.500   0.7005   0.06140   0.05939   0.0407   0.0036   1.0000
  10.750   0.6859   0.06586   0.06395   0.0391   0.0036   1.0000
  11.000   0.6668   0.07203   0.07023   0.0361   0.0036   1.0000
  11.250   0.6501   0.07876   0.07705   0.0320   0.0036   1.0000
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