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NASA/AMES A-03 AIRFOIL (ames03-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: NASA/AMES A-03 AIRFOIL (ames03-il)
Reynolds number: 50,000
Max Cl/Cd: 27.54 at α=4.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-ames03-il-50000.txt
Download as CSV file: xf-ames03-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NASA/AMES A-03 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.750  -0.4972   0.09259   0.08667   0.0165   1.0000   0.3879
  -7.250  -0.4958   0.08737   0.08155   0.0179   1.0000   0.4423
  -6.500  -0.5858   0.06036   0.05375  -0.0222   1.0000   0.2004
  -6.250  -0.5802   0.05328   0.04581  -0.0232   1.0000   0.1534
  -6.000  -0.5713   0.04886   0.04071  -0.0220   1.0000   0.1391
  -5.750  -0.5573   0.04523   0.03695  -0.0207   1.0000   0.1351
  -5.500  -0.5428   0.04186   0.03311  -0.0193   1.0000   0.1314
  -5.250  -0.5261   0.03911   0.02989  -0.0180   1.0000   0.1322
  -5.000  -0.5070   0.03658   0.02692  -0.0167   1.0000   0.1338
  -4.750  -0.4854   0.03414   0.02403  -0.0156   1.0000   0.1346
  -4.500  -0.4622   0.03203   0.02144  -0.0144   1.0000   0.1375
  -4.250  -0.4394   0.03012   0.01940  -0.0136   1.0000   0.1454
  -4.000  -0.4143   0.02856   0.01747  -0.0125   1.0000   0.1539
  -3.750  -0.3892   0.02697   0.01587  -0.0116   1.0000   0.1674
  -3.500  -0.3633   0.02550   0.01449  -0.0107   1.0000   0.1876
  -3.250  -0.1287   0.02212   0.01391  -0.0309   1.0000   1.0000
  -3.000  -0.1299   0.02182   0.01343  -0.0278   1.0000   1.0000
  -2.750  -0.1305   0.02156   0.01298  -0.0245   1.0000   1.0000
  -2.500  -0.1303   0.02132   0.01255  -0.0212   1.0000   1.0000
  -2.250  -0.1292   0.02109   0.01216  -0.0179   1.0000   1.0000
  -2.000  -0.1271   0.02089   0.01180  -0.0146   1.0000   1.0000
  -1.750  -0.1237   0.02071   0.01146  -0.0115   1.0000   1.0000
  -1.500  -0.1185   0.02057   0.01116  -0.0085   1.0000   1.0000
  -1.250  -0.1116   0.02046   0.01090  -0.0058   1.0000   1.0000
  -1.000  -0.1021   0.02040   0.01067  -0.0034   1.0000   1.0000
  -0.750  -0.0895   0.02041   0.01051  -0.0016   1.0000   1.0000
  -0.500  -0.0741   0.02048   0.01043  -0.0003   1.0000   1.0000
  -0.250  -0.0567   0.02060   0.01041   0.0007   1.0000   1.0000
   0.000  -0.0379   0.02077   0.01045   0.0015   1.0000   1.0000
   0.250  -0.0181   0.02099   0.01056   0.0020   1.0000   1.0000
   0.500   0.0024   0.02124   0.01072   0.0024   1.0000   1.0000
   0.750   0.0233   0.02154   0.01095   0.0028   1.0000   1.0000
   1.000   0.0446   0.02187   0.01122   0.0030   1.0000   1.0000
   1.250   0.0660   0.02225   0.01157   0.0031   1.0000   1.0000
   1.500   0.0876   0.02267   0.01197   0.0032   1.0000   1.0000
   1.750   0.1090   0.02314   0.01244   0.0031   1.0000   1.0000
   2.000   0.1304   0.02367   0.01300   0.0030   1.0000   1.0000
   2.250   0.1866   0.02480   0.01421  -0.0037   0.9858   1.0000
   2.500   0.2817   0.02642   0.01604  -0.0167   0.9490   1.0000
   2.750   0.3599   0.02724   0.01717  -0.0251   0.9054   1.0000
   3.000   0.4470   0.02690   0.01723  -0.0321   0.8527   1.0000
   3.250   0.4889   0.02535   0.01597  -0.0293   0.7946   1.0000
   3.500   0.5137   0.02298   0.01375  -0.0229   0.7208   1.0000
   3.750   0.5353   0.02103   0.01168  -0.0169   0.6248   1.0000
   4.000   0.5559   0.02031   0.01029  -0.0117   0.5564   1.0000
   4.250   0.5767   0.02094   0.01036  -0.0090   0.5106   1.0000
   4.500   0.6003   0.02189   0.01100  -0.0077   0.4778   1.0000
   4.750   0.6251   0.02288   0.01177  -0.0067   0.4534   1.0000
   5.000   0.6511   0.02392   0.01277  -0.0063   0.4327   1.0000
   5.250   0.6771   0.02501   0.01386  -0.0059   0.4142   1.0000
   5.500   0.7033   0.02618   0.01505  -0.0056   0.3980   1.0000
   5.750   0.7295   0.02743   0.01635  -0.0054   0.3826   1.0000
   6.000   0.7555   0.02880   0.01784  -0.0053   0.3680   1.0000
   6.250   0.7812   0.03029   0.01945  -0.0053   0.3540   1.0000
   6.500   0.8064   0.03190   0.02121  -0.0052   0.3401   1.0000
   6.750   0.8311   0.03365   0.02312  -0.0052   0.3268   1.0000
   7.000   0.8549   0.03549   0.02515  -0.0051   0.3131   1.0000
   7.250   0.8785   0.03746   0.02722  -0.0048   0.3001   1.0000
   7.500   0.9010   0.03949   0.02942  -0.0045   0.2860   1.0000
   7.750   0.9200   0.04205   0.03242  -0.0046   0.2729   1.0000
   8.000   0.9374   0.04470   0.03540  -0.0043   0.2590   1.0000
   8.250   0.9515   0.04796   0.03906  -0.0041   0.2471   1.0000
   8.500   0.9645   0.05106   0.04244  -0.0036   0.2346   1.0000
   8.750   0.9784   0.05398   0.04554  -0.0030   0.2222   1.0000
   9.000   0.9942   0.05654   0.04819  -0.0021   0.2089   1.0000
   9.250   0.9782   0.06257   0.05485  -0.0023   0.2035   1.0000
   9.500   0.9892   0.06553   0.05789  -0.0013   0.1926   1.0000
   9.750   0.9644   0.07233   0.06502  -0.0021   0.1916   1.0000
  10.000   0.9367   0.07944   0.07228  -0.0038   0.1924   1.0000
  10.250   0.9102   0.08657   0.07943  -0.0058   0.1936   1.0000
  10.500   0.8882   0.09446   0.08735  -0.0091   0.1949   1.0000
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