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NASA SC(2)-0610 AIRFOIL (sc20610-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9


Details Polar file
Airfoil: NASA SC(2)-0610 AIRFOIL (sc20610-il)
Reynolds number: 1,000,000
Max Cl/Cd: 74.25 at α=5.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-sc20610-il-1000000.txt
Download as CSV file: xf-sc20610-il-1000000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NASA SC(2)-0610 AIRFOIL                         
  
 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
  ------ -------- --------- --------- -------- -------- --------
   0.250   0.4575   0.00647   0.00203  -0.1046   0.6165   0.6993
   0.500   0.4804   0.00739   0.00229  -0.1037   0.4539   0.7025
   0.750   0.5052   0.00814   0.00251  -0.1033   0.3270   0.7052
   1.000   0.5309   0.00870   0.00269  -0.1030   0.2297   0.7080
   1.250   0.5573   0.00916   0.00288  -0.1028   0.1605   0.7099
   1.500   0.5842   0.00954   0.00307  -0.1027   0.1172   0.7119
   1.750   0.6117   0.00983   0.00324  -0.1026   0.0908   0.7141
   2.000   0.6393   0.01010   0.00342  -0.1025   0.0742   0.7163
   2.250   0.6673   0.01031   0.00359  -0.1024   0.0653   0.7186
   2.500   0.6951   0.01056   0.00379  -0.1024   0.0578   0.7206
   2.750   0.7232   0.01075   0.00395  -0.1023   0.0533   0.7224
   3.000   0.7506   0.01099   0.00418  -0.1022   0.0480   0.7246
   3.250   0.7785   0.01115   0.00436  -0.1021   0.0451   0.7266
   3.500   0.8057   0.01143   0.00461  -0.1019   0.0408   0.7287
   3.750   0.8332   0.01164   0.00484  -0.1018   0.0382   0.7311
   4.000   0.8605   0.01188   0.00506  -0.1016   0.0351   0.7334
   4.250   0.8876   0.01219   0.00537  -0.1014   0.0320   0.7356
   4.500   0.9150   0.01243   0.00560  -0.1012   0.0293   0.7376
   4.750   0.9416   0.01281   0.00597  -0.1009   0.0266   0.7395
   5.000   0.9686   0.01305   0.00624  -0.1007   0.0249   0.7418
   5.250   0.9950   0.01340   0.00658  -0.1003   0.0233   0.7439
   5.500   1.0205   0.01390   0.00712  -0.0998   0.0219   0.7459
   5.750   1.0469   0.01423   0.00750  -0.0994   0.0212   0.7482
   6.000   1.0730   0.01460   0.00790  -0.0990   0.0204   0.7506
   6.250   1.0989   0.01499   0.00832  -0.0986   0.0197   0.7530
   6.500   1.1244   0.01546   0.00880  -0.0981   0.0191   0.7551
   6.750   1.1483   0.01614   0.00954  -0.0973   0.0184   0.7573
   7.000   1.1715   0.01694   0.01042  -0.0964   0.0179   0.7595
   7.250   1.1964   0.01740   0.01095  -0.0958   0.0177   0.7616
   7.500   1.2211   0.01791   0.01153  -0.0952   0.0174   0.7639
   7.750   1.2456   0.01841   0.01210  -0.0945   0.0170   0.7662
   8.000   1.2697   0.01898   0.01272  -0.0938   0.0166   0.7686
   8.250   1.2933   0.01959   0.01339  -0.0930   0.0163   0.7707
   8.500   1.3166   0.02022   0.01409  -0.0922   0.0159   0.7729
   8.750   1.3393   0.02087   0.01482  -0.0913   0.0157   0.7752
   9.000   1.3613   0.02159   0.01561  -0.0903   0.0154   0.7773
   9.250   1.3821   0.02248   0.01658  -0.0892   0.0151   0.7796
   9.500   1.4006   0.02372   0.01793  -0.0877   0.0148   0.7819
   9.750   1.4139   0.02584   0.02026  -0.0855   0.0145   0.7842
  10.000   1.4306   0.02727   0.02185  -0.0837   0.0144   0.7863
  10.250   1.4488   0.02827   0.02299  -0.0823   0.0143   0.7885
  10.500   1.4658   0.02935   0.02421  -0.0806   0.0142   0.7907
  10.750   1.4811   0.03056   0.02558  -0.0787   0.0141   0.7929
  11.000   1.4943   0.03194   0.02712  -0.0766   0.0140   0.7952
  11.250   1.5055   0.03343   0.02878  -0.0743   0.0139   0.7977
  11.500   1.5150   0.03492   0.03043  -0.0718   0.0138   0.8001
  11.750   1.5224   0.03643   0.03209  -0.0690   0.0136   0.8022
  12.000   1.5252   0.03795   0.03377  -0.0656   0.0134   0.8049
  12.250   1.5215   0.03949   0.03547  -0.0612   0.0133   0.8074
  12.500   1.5177   0.04138   0.03753  -0.0573   0.0132   0.8099
  12.750   1.5120   0.04358   0.03989  -0.0538   0.0131   0.8124
  13.000   1.5017   0.04639   0.04288  -0.0506   0.0131   0.8148
  13.250   1.4906   0.04950   0.04617  -0.0480   0.0130   0.8168
  13.500   1.4758   0.05331   0.05017  -0.0463   0.0129   0.8191
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