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NASA/AMES A-02 AIRFOIL (ames02-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: NASA/AMES A-02 AIRFOIL (ames02-il)
Reynolds number: 50,000
Max Cl/Cd: 22.8 at α=5.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-ames02-il-50000-n5.txt
Download as CSV file: xf-ames02-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NASA/AMES A-02 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
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.6867   0.11430   0.10710   0.0227   1.0000   0.0572
  -9.250  -0.6828   0.10986   0.10269   0.0214   1.0000   0.0553
  -9.000  -0.6925   0.10213   0.09500   0.0147   1.0000   0.0519
  -8.750  -0.6905   0.09746   0.09036   0.0125   1.0000   0.0512
  -8.500  -0.6922   0.09185   0.08478   0.0084   1.0000   0.0505
  -8.250  -0.6978   0.08580   0.07873   0.0035   1.0000   0.0497
  -8.000  -0.7026   0.07966   0.07251  -0.0007   1.0000   0.0489
  -7.750  -0.7061   0.07361   0.06631  -0.0042   1.0000   0.0484
  -7.500  -0.7056   0.06805   0.06052  -0.0067   1.0000   0.0483
  -7.250  -0.7004   0.06304   0.05523  -0.0083   1.0000   0.0485
  -7.000  -0.6918   0.05823   0.05008  -0.0096   1.0000   0.0489
  -6.750  -0.6799   0.05358   0.04500  -0.0104   1.0000   0.0491
  -6.500  -0.6648   0.04912   0.04006  -0.0109   1.0000   0.0490
  -6.250  -0.6465   0.04512   0.03555  -0.0111   1.0000   0.0492
  -6.000  -0.6257   0.04152   0.03143  -0.0111   1.0000   0.0500
  -5.750  -0.6029   0.03813   0.02732  -0.0109   1.0000   0.0519
  -5.500  -0.5789   0.03597   0.02501  -0.0109   1.0000   0.0544
  -5.250  -0.5531   0.03345   0.02204  -0.0106   1.0000   0.0565
  -5.000  -0.5266   0.03121   0.01947  -0.0102   1.0000   0.0586
  -4.750  -0.5001   0.02949   0.01757  -0.0099   1.0000   0.0628
  -4.500  -0.4735   0.02783   0.01579  -0.0095   1.0000   0.0668
  -4.250  -0.4465   0.02632   0.01414  -0.0089   1.0000   0.0721
  -3.750  -0.3941   0.02382   0.01152  -0.0078   1.0000   0.0880
  -3.500  -0.3683   0.02263   0.01039  -0.0074   1.0000   0.1000
  -3.250  -0.3420   0.02148   0.00927  -0.0072   1.0000   0.1200
  -3.000  -0.3171   0.02010   0.00820  -0.0069   1.0000   0.1596
  -2.750  -0.2963   0.01785   0.00723  -0.0066   1.0000   0.3456
  -2.500  -0.2833   0.01648   0.00728  -0.0023   1.0000   0.6476
  -2.250  -0.2654   0.01611   0.00727   0.0020   1.0000   0.7871
  -2.000  -0.2229   0.01604   0.00728   0.0024   1.0000   0.9163
  -1.750  -0.1332   0.01609   0.00689  -0.0084   1.0000   0.9964
  -1.500  -0.1095   0.01579   0.00642  -0.0087   1.0000   1.0000
  -1.250  -0.0904   0.01554   0.00605  -0.0079   1.0000   1.0000
  -1.000  -0.0714   0.01537   0.00574  -0.0070   1.0000   1.0000
  -0.750  -0.0525   0.01526   0.00552  -0.0060   1.0000   1.0000
  -0.500  -0.0336   0.01521   0.00538  -0.0049   1.0000   1.0000
  -0.250  -0.0148   0.01521   0.00529  -0.0038   1.0000   1.0000
   0.000   0.0041   0.01525   0.00527  -0.0026   1.0000   1.0000
   0.250   0.0231   0.01534   0.00530  -0.0015   1.0000   1.0000
   0.500   0.0426   0.01546   0.00539  -0.0004   1.0000   1.0000
   0.750   0.0625   0.01562   0.00553   0.0005   1.0000   1.0000
   1.000   0.0827   0.01581   0.00573   0.0013   1.0000   1.0000
   1.250   0.1034   0.01605   0.00598   0.0020   1.0000   1.0000
   1.500   0.1365   0.01632   0.00630   0.0003   0.9926   1.0000
   1.750   0.2047   0.01648   0.00661  -0.0076   0.9631   1.0000
   2.000   0.2587   0.01657   0.00686  -0.0122   0.9275   1.0000
   2.250   0.2979   0.01664   0.00705  -0.0133   0.8852   1.0000
   2.500   0.3250   0.01669   0.00717  -0.0118   0.8353   1.0000
   2.750   0.3480   0.01672   0.00721  -0.0093   0.7759   1.0000
   3.000   0.3688   0.01679   0.00717  -0.0061   0.6957   1.0000
   3.250   0.3876   0.01710   0.00706  -0.0025   0.5790   1.0000
   3.500   0.4064   0.01795   0.00718   0.0002   0.4427   1.0000
   3.750   0.4282   0.01909   0.00766   0.0013   0.3378   1.0000
   4.000   0.4518   0.02020   0.00833   0.0017   0.2714   1.0000
   4.250   0.4764   0.02122   0.00910   0.0020   0.2266   1.0000
   4.500   0.5019   0.02225   0.00998   0.0022   0.1942   1.0000
   4.750   0.5275   0.02326   0.01090   0.0025   0.1693   1.0000
   5.000   0.5531   0.02431   0.01185   0.0029   0.1505   1.0000
   5.250   0.5790   0.02539   0.01295   0.0033   0.1346   1.0000
   5.500   0.6047   0.02655   0.01415   0.0037   0.1218   1.0000
   5.750   0.6303   0.02778   0.01539   0.0040   0.1113   1.0000
   6.000   0.6558   0.02908   0.01675   0.0044   0.1022   1.0000
   6.250   0.6816   0.03060   0.01851   0.0047   0.0943   1.0000
   6.500   0.7068   0.03231   0.02042   0.0051   0.0878   1.0000
   6.750   0.7312   0.03394   0.02210   0.0053   0.0826   1.0000
   7.000   0.7553   0.03611   0.02468   0.0056   0.0775   1.0000
   7.250   0.7780   0.03839   0.02728   0.0058   0.0734   1.0000
   7.500   0.7994   0.04102   0.03030   0.0061   0.0699   1.0000
   7.750   0.8191   0.04373   0.03334   0.0062   0.0669   1.0000
   8.000   0.8347   0.04757   0.03777   0.0063   0.0644   1.0000
   8.250   0.8501   0.05079   0.04132   0.0064   0.0625   1.0000
   8.500   0.8603   0.05484   0.04579   0.0063   0.0609   1.0000
   8.750   0.8597   0.06063   0.05217   0.0056   0.0596   1.0000
   9.000   0.8548   0.06639   0.05833   0.0044   0.0588   1.0000
   9.250   0.8435   0.07262   0.06485   0.0024   0.0586   1.0000
   9.500   0.8249   0.07953   0.07195  -0.0010   0.0590   1.0000
   9.750   0.8028   0.08731   0.07977  -0.0063   0.0596   1.0000
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