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NASA/AMES A-02 AIRFOIL (ames02-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: NASA/AMES A-02 AIRFOIL (ames02-il)
Reynolds number: 100,000
Max Cl/Cd: 31.2 at α=6°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-ames02-il-100000-n5.txt
Download as CSV file: xf-ames02-il-100000-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 =     0.100 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.7407   0.07140   0.06621  -0.0019   1.0000   0.0262
  -8.000  -0.7455   0.06488   0.05946  -0.0049   1.0000   0.0262
  -7.750  -0.7385   0.06148   0.05596  -0.0059   1.0000   0.0269
  -7.500  -0.7319   0.05695   0.05120  -0.0071   1.0000   0.0275
  -7.250  -0.7242   0.05169   0.04557  -0.0081   1.0000   0.0278
  -7.000  -0.7135   0.04632   0.03971  -0.0087   1.0000   0.0277
  -6.750  -0.6988   0.04147   0.03432  -0.0088   1.0000   0.0277
  -6.500  -0.6804   0.03739   0.02969  -0.0087   1.0000   0.0279
  -6.250  -0.6591   0.03390   0.02564  -0.0084   1.0000   0.0283
  -6.000  -0.6357   0.03080   0.02197  -0.0081   1.0000   0.0291
  -5.750  -0.6109   0.02878   0.01964  -0.0079   1.0000   0.0306
  -5.500  -0.5851   0.02706   0.01763  -0.0078   1.0000   0.0326
  -5.250  -0.5585   0.02509   0.01529  -0.0074   1.0000   0.0338
  -5.000  -0.5320   0.02366   0.01380  -0.0073   1.0000   0.0351
  -4.750  -0.5050   0.02227   0.01221  -0.0070   1.0000   0.0372
  -4.500  -0.4784   0.02121   0.01115  -0.0069   1.0000   0.0401
  -4.250  -0.4517   0.02009   0.00997  -0.0066   1.0000   0.0429
  -4.000  -0.4252   0.01909   0.00895  -0.0064   1.0000   0.0461
  -3.750  -0.3986   0.01825   0.00812  -0.0062   1.0000   0.0510
  -3.500  -0.3720   0.01743   0.00731  -0.0061   1.0000   0.0568
  -3.250  -0.3453   0.01668   0.00660  -0.0059   1.0000   0.0654
  -3.000  -0.3184   0.01597   0.00594  -0.0058   1.0000   0.0792
  -2.750  -0.2918   0.01516   0.00532  -0.0058   1.0000   0.1063
  -2.500  -0.2661   0.01410   0.00479  -0.0058   1.0000   0.1943
  -2.250  -0.2440   0.01254   0.00445  -0.0055   1.0000   0.4479
  -2.000  -0.2230   0.01179   0.00447  -0.0037   1.0000   0.6323
  -1.750  -0.2014   0.01148   0.00447  -0.0016   1.0000   0.7343
  -1.500  -0.1804   0.01125   0.00446   0.0008   1.0000   0.8103
  -1.250  -0.1554   0.01111   0.00447   0.0026   1.0000   0.8833
  -1.000  -0.1096   0.01110   0.00447  -0.0002   1.0000   0.9577
  -0.750  -0.0585   0.01108   0.00436  -0.0050   1.0000   0.9942
  -0.500  -0.0347   0.01104   0.00423  -0.0049   1.0000   1.0000
  -0.250  -0.0184   0.01106   0.00418  -0.0033   1.0000   1.0000
   0.000  -0.0006   0.01114   0.00420  -0.0020   1.0000   1.0000
   0.250   0.0453   0.01118   0.00418  -0.0059   0.9856   1.0000
   0.500   0.0940   0.01120   0.00416  -0.0100   0.9625   1.0000
   0.750   0.1318   0.01124   0.00417  -0.0118   0.9372   1.0000
   1.000   0.1614   0.01131   0.00421  -0.0117   0.9113   1.0000
   1.250   0.1866   0.01139   0.00424  -0.0105   0.8842   1.0000
   1.500   0.2099   0.01144   0.00427  -0.0089   0.8536   1.0000
   1.750   0.2328   0.01149   0.00427  -0.0072   0.8196   1.0000
   2.000   0.2554   0.01154   0.00426  -0.0053   0.7787   1.0000
   2.250   0.2779   0.01163   0.00423  -0.0034   0.7250   1.0000
   2.500   0.3002   0.01182   0.00416  -0.0014   0.6473   1.0000
   2.750   0.3224   0.01227   0.00413   0.0004   0.5373   1.0000
   3.000   0.3463   0.01300   0.00426   0.0013   0.4167   1.0000
   3.250   0.3718   0.01376   0.00453   0.0015   0.3213   1.0000
   3.500   0.3981   0.01444   0.00488   0.0016   0.2557   1.0000
   3.750   0.4248   0.01505   0.00526   0.0016   0.2097   1.0000
   4.000   0.4516   0.01564   0.00568   0.0016   0.1756   1.0000
   4.250   0.4783   0.01626   0.00614   0.0016   0.1500   1.0000
   4.500   0.5052   0.01685   0.00668   0.0017   0.1296   1.0000
   4.750   0.5318   0.01751   0.00724   0.0017   0.1139   1.0000
   5.000   0.5582   0.01819   0.00785   0.0018   0.1013   1.0000
   5.250   0.5847   0.01889   0.00855   0.0019   0.0906   1.0000
   5.500   0.6109   0.01963   0.00928   0.0020   0.0820   1.0000
   5.750   0.6369   0.02043   0.01006   0.0021   0.0748   1.0000
   6.000   0.6629   0.02125   0.01091   0.0023   0.0685   1.0000
   6.250   0.6887   0.02216   0.01191   0.0025   0.0630   1.0000
   6.500   0.7143   0.02316   0.01299   0.0027   0.0584   1.0000
   6.750   0.7396   0.02418   0.01405   0.0029   0.0546   1.0000
   7.000   0.7648   0.02527   0.01524   0.0032   0.0512   1.0000
   7.250   0.7896   0.02663   0.01678   0.0035   0.0482   1.0000
   7.500   0.8140   0.02772   0.01790   0.0036   0.0457   1.0000
   7.750   0.8379   0.02948   0.02001   0.0040   0.0431   1.0000
   8.000   0.8613   0.03103   0.02170   0.0043   0.0415   1.0000
   8.250   0.8831   0.03318   0.02422   0.0046   0.0397   1.0000
   8.500   0.9041   0.03511   0.02644   0.0049   0.0379   1.0000
   8.750   0.9244   0.03688   0.02838   0.0052   0.0365   1.0000
   9.000   0.9377   0.04085   0.03305   0.0058   0.0354   1.0000
   9.250   0.9478   0.04487   0.03760   0.0062   0.0346   1.0000
   9.500   0.9545   0.04887   0.04206   0.0065   0.0340   1.0000
   9.750   0.9622   0.05199   0.04543   0.0068   0.0334   1.0000
  10.000   0.9665   0.05527   0.04892   0.0070   0.0328   1.0000
  10.250   0.9439   0.06223   0.05640   0.0060   0.0326   1.0000
  10.500   0.9186   0.06829   0.06271   0.0042   0.0327   1.0000
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