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NASA SC(2)-0610 AIRFOIL (sc20610-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: NASA SC(2)-0610 AIRFOIL (sc20610-il)
Reynolds number: 200,000
Max Cl/Cd: 41.91 at α=2°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-sc20610-il-200000.txt
Download as CSV file: xf-sc20610-il-200000.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 =     0.200 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.750  -0.4752   0.09058   0.08682  -0.0339   1.0000   0.0773
 -10.500  -0.4782   0.08642   0.08268  -0.0346   1.0000   0.0787
 -10.250  -0.4864   0.08139   0.07769  -0.0359   1.0000   0.0803
 -10.000  -0.5028   0.07493   0.07126  -0.0385   1.0000   0.0819
  -9.750  -0.5389   0.06533   0.06169  -0.0441   1.0000   0.0826
  -9.500  -0.7726   0.05017   0.04511  -0.0516   1.0000   0.0519
  -9.250  -0.7558   0.04890   0.04401  -0.0506   1.0000   0.0533
  -9.000  -0.7592   0.03921   0.03347  -0.0514   1.0000   0.0451
  -8.750  -0.7457   0.03288   0.02624  -0.0524   1.0000   0.0436
  -8.500  -0.7250   0.03052   0.02364  -0.0524   1.0000   0.0442
  -8.250  -0.7031   0.02883   0.02178  -0.0523   1.0000   0.0453
  -8.000  -0.6800   0.02735   0.02012  -0.0522   1.0000   0.0471
  -7.750  -0.6557   0.02572   0.01820  -0.0521   1.0000   0.0489
  -7.500  -0.6306   0.02431   0.01649  -0.0520   1.0000   0.0503
  -7.250  -0.6066   0.02222   0.01427  -0.0518   1.0000   0.0524
  -7.000  -0.5822   0.02121   0.01327  -0.0516   1.0000   0.0550
  -6.750  -0.5568   0.02046   0.01244  -0.0514   1.0000   0.0588
  -6.500  -0.5311   0.01936   0.01126  -0.0513   1.0000   0.0632
  -6.250  -0.5052   0.01860   0.01055  -0.0514   1.0000   0.0680
  -6.000  -0.4785   0.01818   0.01001  -0.0514   1.0000   0.0732
  -5.750  -0.4508   0.01715   0.00912  -0.0520   1.0000   0.0805
  -5.500  -0.4221   0.01656   0.00851  -0.0525   1.0000   0.0881
  -5.250  -0.3923   0.01595   0.00796  -0.0535   1.0000   0.0976
  -5.000  -0.3605   0.01527   0.00736  -0.0549   1.0000   0.1098
  -4.750  -0.3273   0.01464   0.00683  -0.0567   1.0000   0.1280
  -4.500  -0.2876   0.01366   0.00621  -0.0602   1.0000   0.1815
  -4.250  -0.2290   0.01160   0.00613  -0.0686   1.0000   0.5755
  -4.000  -0.2057   0.01196   0.00655  -0.0673   1.0000   0.6324
  -3.750  -0.1834   0.01239   0.00696  -0.0657   1.0000   0.6604
  -3.500  -0.1631   0.01287   0.00743  -0.0637   1.0000   0.6783
  -3.250  -0.1430   0.01334   0.00790  -0.0616   1.0000   0.6928
  -3.000  -0.1207   0.01376   0.00827  -0.0602   1.0000   0.7062
  -2.750  -0.1050   0.01426   0.00882  -0.0571   1.0000   0.7141
  -2.500  -0.0842   0.01466   0.00921  -0.0554   1.0000   0.7250
  -2.250  -0.0674   0.01521   0.00978  -0.0526   1.0000   0.7373
  -2.000  -0.0559   0.01575   0.01038  -0.0484   1.0000   0.7462
  -1.750  -0.0365   0.01616   0.01079  -0.0465   1.0000   0.7578
  -1.500  -0.0221   0.01656   0.01123  -0.0433   1.0000   0.7667
  -1.250  -0.0036   0.01686   0.01155  -0.0412   1.0000   0.7757
  -1.000   0.0178   0.01712   0.01183  -0.0400   1.0000   0.7844
  -0.750   0.0400   0.01728   0.01201  -0.0391   1.0000   0.7902
  -0.500   0.0632   0.01744   0.01220  -0.0384   0.9995   0.7951
  -0.250   0.1141   0.01758   0.01233  -0.0435   0.9941   0.8003
   0.000   0.1618   0.01752   0.01230  -0.0476   0.9857   0.8041
   0.250   0.2127   0.01729   0.01213  -0.0518   0.9743   0.8071
   0.500   0.2661   0.01684   0.01173  -0.0565   0.9614   0.8095
   0.750   0.3260   0.01604   0.01101  -0.0620   0.9460   0.8116
   1.000   0.3779   0.01537   0.01041  -0.0661   0.9330   0.8143
   1.250   0.4299   0.01465   0.00979  -0.0703   0.9194   0.8168
   1.500   0.4739   0.01385   0.00907  -0.0723   0.8993   0.8187
   1.750   0.5044   0.01332   0.00860  -0.0716   0.8644   0.8209
   2.000   0.5398   0.01288   0.00800  -0.0716   0.7723   0.8231
   2.250   0.5532   0.01432   0.00787  -0.0677   0.4591   0.8257
   2.500   0.5667   0.01619   0.00844  -0.0658   0.2113   0.8288
   2.750   0.5924   0.01726   0.00900  -0.0661   0.1377   0.8317
   3.000   0.6179   0.01789   0.00949  -0.0658   0.1158   0.8341
   3.250   0.6411   0.01839   0.00997  -0.0649   0.1031   0.8364
   3.500   0.6657   0.01901   0.01055  -0.0643   0.0931   0.8389
   3.750   0.6917   0.01971   0.01118  -0.0642   0.0845   0.8413
   4.000   0.7199   0.02030   0.01182  -0.0643   0.0772   0.8440
   4.250   0.7479   0.02148   0.01287  -0.0648   0.0701   0.8470
   4.500   0.7757   0.02195   0.01347  -0.0648   0.0650   0.8494
   4.750   0.8005   0.02269   0.01420  -0.0642   0.0607   0.8518
   5.000   0.8262   0.02399   0.01557  -0.0638   0.0567   0.8546
   5.250   0.8536   0.02474   0.01644  -0.0637   0.0529   0.8575
   5.500   0.8819   0.02582   0.01755  -0.0640   0.0503   0.8601
   5.750   0.9110   0.02835   0.02018  -0.0646   0.0482   0.8626
   6.000   0.9359   0.02969   0.02185  -0.0638   0.0470   0.8649
   6.250   0.9591   0.03150   0.02401  -0.0627   0.0459   0.8674
   6.500   0.9814   0.03338   0.02625  -0.0617   0.0444   0.8703
   6.750   1.0037   0.03525   0.02837  -0.0609   0.0431   0.8736
   7.000   1.0232   0.03848   0.03205  -0.0598   0.0429   0.8767
   7.250   1.0325   0.04366   0.03792  -0.0571   0.0441   0.8792
   7.500   1.0374   0.04889   0.04366  -0.0543   0.0460   0.8817
   7.750   1.0437   0.05395   0.04894  -0.0526   0.0476   0.8845
   8.000   1.0043   0.07098   0.06732  -0.0465   0.0724   0.8864
   8.250   1.0203   0.07494   0.07119  -0.0466   0.0713   0.8893
   8.500   1.0101   0.08415   0.08049  -0.0466   0.0701   0.8916
   8.750   0.9878   0.08723   0.08393  -0.0443   0.0697   0.8944
   9.000   0.9489   0.09128   0.08826  -0.0424   0.0693   0.8977
   9.250   0.9125   0.09813   0.09529  -0.0458   0.0690   0.9008
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