Airfoil Tools
Search 1636 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

NASA/AMES A-01 AIRFOIL (ames01-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: NASA/AMES A-01 AIRFOIL (ames01-il)
Reynolds number: 50,000
Max Cl/Cd: 27.77 at α=6.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-ames01-il-50000-n5.txt
Download as CSV file: xf-ames01-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NASA/AMES A-01 AIRFOIL                          
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.500  -0.5952   0.11733   0.11032  -0.0009   1.0000   0.0676
 -10.250  -0.5867   0.11321   0.10620  -0.0010   1.0000   0.0668
 -10.000  -0.5853   0.10839   0.10141  -0.0032   1.0000   0.0665
  -9.750  -0.5856   0.10332   0.09638  -0.0061   1.0000   0.0662
  -9.500  -0.5872   0.09795   0.09104  -0.0095   1.0000   0.0659
  -9.250  -0.5902   0.09222   0.08533  -0.0136   1.0000   0.0653
  -9.000  -0.5969   0.08631   0.07944  -0.0181   1.0000   0.0647
  -8.750  -0.6067   0.08115   0.07424  -0.0206   1.0000   0.0640
  -8.500  -0.6146   0.07604   0.06905  -0.0227   1.0000   0.0634
  -8.250  -0.6203   0.07105   0.06392  -0.0241   1.0000   0.0628
  -8.000  -0.6234   0.06624   0.05890  -0.0249   1.0000   0.0623
  -7.750  -0.6233   0.06166   0.05406  -0.0252   1.0000   0.0620
  -7.500  -0.6199   0.05740   0.04947  -0.0250   1.0000   0.0620
  -7.250  -0.6134   0.05358   0.04528  -0.0244   1.0000   0.0631
  -7.000  -0.6048   0.05005   0.04130  -0.0234   1.0000   0.0647
  -6.750  -0.5938   0.04679   0.03754  -0.0222   1.0000   0.0661
  -6.500  -0.5807   0.04381   0.03407  -0.0207   1.0000   0.0670
  -6.250  -0.5660   0.04117   0.03090  -0.0191   1.0000   0.0677
  -6.000  -0.5505   0.03873   0.02834  -0.0178   1.0000   0.0697
  -5.750  -0.5340   0.03703   0.02655  -0.0164   1.0000   0.0726
  -5.500  -0.5171   0.03529   0.02455  -0.0147   1.0000   0.0754
  -5.250  -0.4987   0.03348   0.02241  -0.0131   1.0000   0.0774
  -5.000  -0.4790   0.03185   0.02044  -0.0115   1.0000   0.0797
  -4.750  -0.4589   0.03043   0.01874  -0.0101   1.0000   0.0834
  -4.500  -0.4387   0.02918   0.01749  -0.0089   1.0000   0.0876
  -4.250  -0.4168   0.02801   0.01619  -0.0077   1.0000   0.0912
  -4.000  -0.3942   0.02704   0.01501  -0.0065   1.0000   0.0959
  -3.750  -0.3727   0.02605   0.01408  -0.0054   1.0000   0.1021
  -3.500  -0.3505   0.02527   0.01319  -0.0043   1.0000   0.1084
  -3.250  -0.3284   0.02448   0.01235  -0.0033   1.0000   0.1146
  -3.000  -0.3068   0.02385   0.01164  -0.0023   1.0000   0.1256
  -2.750  -0.2859   0.02314   0.01097  -0.0013   1.0000   0.1380
  -2.500  -0.2654   0.02246   0.01034  -0.0003   1.0000   0.1550
  -2.250  -0.2450   0.02144   0.00978   0.0005   1.0000   0.2079
  -2.000  -0.2323   0.01923   0.00962   0.0031   1.0000   0.5759
  -1.750  -0.0846   0.01954   0.01049  -0.0148   1.0000   1.0000
  -1.500  -0.0792   0.01951   0.01029  -0.0117   1.0000   1.0000
  -1.250  -0.0718   0.01953   0.01013  -0.0089   1.0000   1.0000
  -1.000  -0.0625   0.01960   0.01004  -0.0063   1.0000   1.0000
  -0.750  -0.0520   0.01970   0.00999  -0.0039   1.0000   1.0000
  -0.500   0.0080   0.01993   0.00999  -0.0108   0.9848   1.0000
  -0.250   0.0909   0.01999   0.00987  -0.0215   0.9614   1.0000
   0.000   0.1593   0.01988   0.00964  -0.0290   0.9380   1.0000
   0.250   0.2040   0.01984   0.00953  -0.0320   0.9192   1.0000
   0.500   0.2365   0.01984   0.00948  -0.0326   0.9013   1.0000
   0.750   0.2633   0.01983   0.00944  -0.0320   0.8814   1.0000
   1.000   0.2898   0.01976   0.00935  -0.0311   0.8607   1.0000
   1.250   0.3127   0.01967   0.00925  -0.0295   0.8376   1.0000
   1.500   0.3357   0.01952   0.00909  -0.0277   0.8130   1.0000
   1.750   0.3567   0.01937   0.00894  -0.0256   0.7829   1.0000
   2.000   0.3771   0.01923   0.00880  -0.0233   0.7454   1.0000
   2.250   0.3985   0.01904   0.00856  -0.0210   0.6988   1.0000
   2.500   0.4203   0.01886   0.00817  -0.0184   0.6359   1.0000
   2.750   0.4409   0.01892   0.00780  -0.0155   0.5589   1.0000
   3.000   0.4604   0.01939   0.00774  -0.0132   0.4924   1.0000
   3.250   0.4805   0.02002   0.00796  -0.0116   0.4466   1.0000
   3.500   0.5018   0.02065   0.00833  -0.0104   0.4155   1.0000
   3.750   0.5243   0.02127   0.00876  -0.0095   0.3924   1.0000
   4.000   0.5476   0.02188   0.00922  -0.0087   0.3749   1.0000
   4.250   0.5715   0.02247   0.00973  -0.0080   0.3600   1.0000
   4.500   0.5959   0.02307   0.01027  -0.0075   0.3472   1.0000
   4.750   0.6208   0.02368   0.01083  -0.0069   0.3365   1.0000
   5.000   0.6460   0.02430   0.01143  -0.0064   0.3262   1.0000
   5.250   0.6713   0.02494   0.01214  -0.0060   0.3164   1.0000
   5.500   0.6968   0.02563   0.01278  -0.0056   0.3084   1.0000
   5.750   0.7219   0.02632   0.01362  -0.0052   0.2995   1.0000
   6.000   0.7472   0.02708   0.01439  -0.0048   0.2919   1.0000
   6.250   0.7719   0.02786   0.01532  -0.0044   0.2835   1.0000
   6.500   0.7969   0.02870   0.01619  -0.0040   0.2767   1.0000
   6.750   0.8209   0.02958   0.01729  -0.0036   0.2684   1.0000
   7.000   0.8453   0.03047   0.01823  -0.0032   0.2613   1.0000
   7.250   0.8684   0.03146   0.01947  -0.0027   0.2531   1.0000
   7.500   0.8919   0.03240   0.02047  -0.0022   0.2456   1.0000
   7.750   0.9137   0.03345   0.02181  -0.0016   0.2370   1.0000
   8.000   0.9358   0.03447   0.02297  -0.0010   0.2289   1.0000
   8.250   0.9569   0.03548   0.02418  -0.0004   0.2200   1.0000
   8.500   0.9764   0.03664   0.02560   0.0004   0.2108   1.0000
   8.750   0.9983   0.03736   0.02637   0.0011   0.2023   1.0000
   9.000   1.0141   0.03864   0.02803   0.0022   0.1914   1.0000
   9.250   1.0311   0.03971   0.02932   0.0032   0.1816   1.0000
   9.500   1.0505   0.04016   0.02977   0.0042   0.1722   1.0000
   9.750   1.0628   0.04152   0.03151   0.0055   0.1612   1.0000
  10.000   1.0745   0.04274   0.03303   0.0069   0.1506   1.0000
  10.250   1.0867   0.04374   0.03419   0.0083   0.1409   1.0000
  10.500   1.1004   0.04418   0.03462   0.0098   0.1317   1.0000
  10.750   1.1037   0.04625   0.03703   0.0115   0.1220   1.0000
  11.000   1.1080   0.04798   0.03892   0.0132   0.1141   1.0000
  11.250   1.1128   0.04926   0.04024   0.0149   0.1070   1.0000
  11.500   1.1057   0.05208   0.04330   0.0169   0.1016   1.0000
  11.750   1.0983   0.05474   0.04616   0.0186   0.0971   1.0000
  12.000   1.0987   0.05653   0.04788   0.0198   0.0928   1.0000
  12.250   1.0768   0.06161   0.05331   0.0199   0.0904   1.0000
  12.500   1.0513   0.06773   0.05970   0.0184   0.0889   1.0000
  12.750   1.0139   0.07678   0.06896   0.0140   0.0892   1.0000
  13.000   0.9565   0.09181   0.08415   0.0049   0.0912   1.0000
<< Back to NASA/AMES A-01 AIRFOIL (ames01-il)

Polar data table (+)

Polar graphs


<< Back to NASA/AMES A-01 AIRFOIL (ames01-il)