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NASA SC(2)-0503 AIRFOIL (sc20503-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9


Details Polar file
Airfoil: NASA SC(2)-0503 AIRFOIL (sc20503-il)
Reynolds number: 1,000,000
Max Cl/Cd: 71.34 at α=2°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-sc20503-il-1000000.txt
Download as CSV file: xf-sc20503-il-1000000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NASA SC(2)-0503 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
  ------ -------- --------- --------- -------- -------- --------
  -6.000  -0.4730   0.02913   0.02712  -0.0297   1.0000   0.0048
  -5.750  -0.4535   0.02418   0.02198  -0.0318   1.0000   0.0050
  -5.500  -0.4317   0.02019   0.01781  -0.0335   1.0000   0.0052
  -5.250  -0.4081   0.01685   0.01427  -0.0348   1.0000   0.0055
  -5.000  -0.3823   0.01387   0.01105  -0.0360   1.0000   0.0059
  -4.750  -0.3548   0.01126   0.00819  -0.0370   1.0000   0.0065
  -4.500  -0.3261   0.00906   0.00572  -0.0378   1.0000   0.0072
  -4.250  -0.2969   0.00735   0.00372  -0.0383   1.0000   0.0082
  -4.000  -0.2689   0.00659   0.00278  -0.0380   1.0000   0.0099
  -2.500  -0.0843   0.00657   0.00110  -0.0429   1.0000   0.0096
  -2.250  -0.0566   0.00631   0.00086  -0.0429   1.0000   0.0288
  -2.000  -0.0218   0.00464   0.00066  -0.0459   1.0000   0.4844
  -1.750   0.0073   0.00413   0.00065  -0.0465   1.0000   0.6508
  -1.500   0.0346   0.00394   0.00069  -0.0464   1.0000   0.7288
  -1.250   0.0613   0.00381   0.00075  -0.0461   1.0000   0.7884
  -1.000   0.0880   0.00376   0.00078  -0.0458   1.0000   0.8156
  -0.750   0.1147   0.00373   0.00082  -0.0456   1.0000   0.8340
  -0.500   0.1415   0.00371   0.00087  -0.0453   1.0000   0.8491
  -0.250   0.1682   0.00369   0.00093  -0.0451   1.0000   0.8637
   0.000   0.2189   0.00407   0.00081  -0.0498   0.7891   0.8735
   0.250   0.2391   0.00571   0.00099  -0.0487   0.3684   0.8869
   0.500   0.2629   0.00704   0.00121  -0.0485   0.0213   0.9012
   1.000   0.3120   0.00740   0.00172  -0.0466   0.0057   0.9421
   1.250   0.3351   0.00872   0.00334  -0.0451   0.0041   1.0000
   1.500   0.3628   0.00987   0.00457  -0.0448   0.0043   1.0000
   1.750   0.3918   0.01182   0.00674  -0.0438   0.0072   1.0000
   2.000   0.4045   0.00567   0.00128  -0.0395   0.0102   1.0000
   2.250   0.4306   0.00689   0.00278  -0.0387   0.0101   1.0000
   2.500   0.4592   0.00709   0.00317  -0.0381   0.0089   1.0000
   2.750   0.4857   0.00820   0.00450  -0.0372   0.0075   1.0000
   3.000   0.5107   0.00997   0.00654  -0.0362   0.0066   1.0000
   3.250   0.5347   0.01219   0.00902  -0.0352   0.0060   1.0000
   3.500   0.5577   0.01477   0.01185  -0.0344   0.0056   1.0000
   3.750   0.5796   0.01776   0.01505  -0.0338   0.0052   1.0000
   4.000   0.6002   0.02112   0.01861  -0.0334   0.0049   1.0000
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