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NASA/AMES A-02 AIRFOIL (ames02-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: NASA/AMES A-02 AIRFOIL (ames02-il)
Reynolds number: 100,000
Max Cl/Cd: 31.72 at α=3.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-ames02-il-100000.txt
Download as CSV file: xf-ames02-il-100000.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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.5610   0.08806   0.08352   0.0120   1.0000   0.1286
  -7.750  -0.5881   0.08240   0.07793   0.0055   1.0000   0.1329
  -7.500  -0.6270   0.07538   0.07082  -0.0038   1.0000   0.1340
  -7.250  -0.5828   0.07335   0.06895   0.0054   1.0000   0.1407
  -7.000  -0.6193   0.06595   0.06134  -0.0050   1.0000   0.1481
  -6.750  -0.6728   0.07124   0.06642   0.0024   1.0000   0.1429
  -6.500  -0.6674   0.06619   0.06124  -0.0010   1.0000   0.1526
  -6.250  -0.6586   0.06200   0.05684  -0.0038   1.0000   0.1664
  -6.000  -0.6401   0.05892   0.05383  -0.0027   1.0000   0.1749
  -5.750  -0.6254   0.05525   0.05009  -0.0033   1.0000   0.1890
  -5.500  -0.5924   0.03924   0.03191  -0.0106   1.0000   0.0859
  -5.250  -0.5655   0.03340   0.02487  -0.0099   1.0000   0.0708
  -5.000  -0.5391   0.03056   0.02153  -0.0096   1.0000   0.0710
  -4.750  -0.5117   0.02784   0.01849  -0.0093   1.0000   0.0706
  -4.500  -0.4833   0.02554   0.01582  -0.0089   1.0000   0.0708
  -4.250  -0.4557   0.02357   0.01384  -0.0088   1.0000   0.0737
  -4.000  -0.4272   0.02211   0.01215  -0.0084   1.0000   0.0783
  -3.750  -0.3994   0.02043   0.01048  -0.0081   1.0000   0.0822
  -3.500  -0.3718   0.01911   0.00920  -0.0078   1.0000   0.0899
  -3.250  -0.3449   0.01785   0.00804  -0.0074   1.0000   0.1000
  -3.000  -0.3187   0.01667   0.00706  -0.0070   1.0000   0.1169
  -2.750  -0.2932   0.01534   0.00601  -0.0065   1.0000   0.1526
  -2.500  -0.2809   0.01195   0.00530  -0.0036   1.0000   0.6328
  -2.250  -0.2666   0.01155   0.00539   0.0014   1.0000   0.7979
  -2.000  -0.2464   0.01144   0.00542   0.0053   1.0000   0.8853
  -1.750  -0.1916   0.01166   0.00548   0.0022   1.0000   0.9629
  -1.500  -0.1089   0.01184   0.00536  -0.0081   1.0000   1.0000
  -1.250  -0.0885   0.01151   0.00493  -0.0079   1.0000   1.0000
  -1.000  -0.0692   0.01125   0.00460  -0.0073   1.0000   1.0000
  -0.750  -0.0512   0.01110   0.00437  -0.0063   1.0000   1.0000
  -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   0.0188   0.01126   0.00427  -0.0009   1.0000   1.0000
   0.500   0.0394   0.01143   0.00440  -0.0001   1.0000   1.0000
   0.750   0.0608   0.01163   0.00458   0.0005   1.0000   1.0000
   1.000   0.0826   0.01186   0.00481   0.0010   1.0000   1.0000
   1.250   0.1048   0.01214   0.00510   0.0013   1.0000   1.0000
   1.500   0.1305   0.01244   0.00543   0.0009   0.9985   1.0000
   1.750   0.2068   0.01243   0.00557  -0.0087   0.9763   1.0000
   2.000   0.2755   0.01228   0.00559  -0.0161   0.9491   1.0000
   2.250   0.3137   0.01224   0.00565  -0.0167   0.9125   1.0000
   2.500   0.3354   0.01221   0.00567  -0.0136   0.8715   1.0000
   2.750   0.3529   0.01214   0.00560  -0.0098   0.8230   1.0000
   3.000   0.3699   0.01206   0.00543  -0.0057   0.7578   1.0000
   3.250   0.3867   0.01219   0.00521  -0.0015   0.6396   1.0000
   3.500   0.4042   0.01332   0.00520   0.0015   0.4285   1.0000
   3.750   0.4274   0.01477   0.00577   0.0021   0.3003   1.0000
   4.000   0.4528   0.01593   0.00648   0.0024   0.2394   1.0000
   4.250   0.4788   0.01699   0.00724   0.0026   0.2015   1.0000
   4.500   0.5052   0.01802   0.00813   0.0029   0.1743   1.0000
   4.750   0.5317   0.01906   0.00901   0.0031   0.1541   1.0000
   5.000   0.5582   0.02020   0.01001   0.0034   0.1383   1.0000
   5.250   0.5850   0.02137   0.01115   0.0036   0.1254   1.0000
   5.500   0.6121   0.02258   0.01243   0.0038   0.1146   1.0000
   5.750   0.6391   0.02401   0.01398   0.0041   0.1059   1.0000
   6.000   0.6658   0.02560   0.01568   0.0043   0.0988   1.0000
   6.250   0.6921   0.02710   0.01719   0.0045   0.0926   1.0000
   6.500   0.7183   0.02929   0.01988   0.0048   0.0878   1.0000
   6.750   0.7437   0.03106   0.02163   0.0049   0.0837   1.0000
   7.000   0.7672   0.03406   0.02530   0.0053   0.0804   1.0000
   7.250   0.7887   0.03779   0.02966   0.0056   0.0794   1.0000
   7.500   0.8068   0.04224   0.03471   0.0058   0.0791   1.0000
   7.750   0.8205   0.04737   0.04044   0.0058   0.0793   1.0000
   8.000   0.8310   0.05307   0.04657   0.0055   0.0815   1.0000
   8.250   0.8496   0.05709   0.05052   0.0057   0.0831   1.0000
   8.500   0.7783   0.08743   0.08240  -0.0129   0.1531   1.0000
   8.750   0.7922   0.09233   0.08724  -0.0108   0.1508   1.0000
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