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-02 AIRFOIL (ames02-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: NASA/AMES A-02 AIRFOIL (ames02-il)
Reynolds number: 200,000
Max Cl/Cd: 37.34 at α=5.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-ames02-il-200000.txt
Download as CSV file: xf-ames02-il-200000.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 =     0.200 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.6981   0.09587   0.09243   0.0223   1.0000   0.0497
  -8.250  -0.7058   0.08920   0.08580   0.0155   1.0000   0.0510
  -8.000  -0.7408   0.07609   0.07235  -0.0025   1.0000   0.0520
  -7.750  -0.7255   0.07315   0.06956  -0.0009   1.0000   0.0525
  -7.500  -0.7143   0.07041   0.06682  -0.0005   1.0000   0.0535
  -7.250  -0.7072   0.06596   0.06226  -0.0025   1.0000   0.0555
  -7.000  -0.7054   0.05887   0.05479  -0.0063   1.0000   0.0586
  -6.750  -0.6895   0.05603   0.05194  -0.0064   1.0000   0.0602
  -6.500  -0.6784   0.05073   0.04619  -0.0080   1.0000   0.0651
  -6.250  -0.6592   0.04802   0.04348  -0.0081   1.0000   0.0673
  -6.000  -0.6418   0.04404   0.03914  -0.0088   1.0000   0.0730
  -5.750  -0.6207   0.03115   0.02429  -0.0071   1.0000   0.0419
  -5.500  -0.5960   0.02831   0.02134  -0.0072   1.0000   0.0426
  -5.250  -0.5692   0.02530   0.01796  -0.0068   1.0000   0.0404
  -5.000  -0.5417   0.02259   0.01482  -0.0063   1.0000   0.0395
  -4.750  -0.5137   0.02067   0.01262  -0.0060   1.0000   0.0399
  -4.500  -0.4854   0.01923   0.01089  -0.0057   1.0000   0.0414
  -4.250  -0.4577   0.01804   0.00972  -0.0057   1.0000   0.0441
  -4.000  -0.4294   0.01693   0.00848  -0.0055   1.0000   0.0463
  -3.750  -0.4019   0.01579   0.00742  -0.0054   1.0000   0.0489
  -3.500  -0.3742   0.01490   0.00656  -0.0053   1.0000   0.0535
  -3.250  -0.3466   0.01406   0.00579  -0.0053   1.0000   0.0594
  -3.000  -0.3189   0.01327   0.00510  -0.0053   1.0000   0.0676
  -2.750  -0.2911   0.01250   0.00444  -0.0054   1.0000   0.0828
  -2.500  -0.2637   0.01144   0.00375  -0.0056   1.0000   0.1387
  -2.250  -0.2430   0.00915   0.00334  -0.0054   1.0000   0.5400
  -2.000  -0.2207   0.00858   0.00338  -0.0039   1.0000   0.6930
  -1.750  -0.1980   0.00834   0.00337  -0.0022   1.0000   0.7708
  -1.500  -0.1771   0.00816   0.00337  -0.0001   1.0000   0.8297
  -1.250  -0.1573   0.00807   0.00342   0.0024   1.0000   0.8856
  -1.000  -0.1267   0.00809   0.00349   0.0026   1.0000   0.9451
  -0.750  -0.0702   0.00817   0.00352  -0.0029   1.0000   0.9894
  -0.500  -0.0384   0.00816   0.00345  -0.0044   1.0000   1.0000
  -0.250  -0.0252   0.00823   0.00344  -0.0025   1.0000   1.0000
   0.000  -0.0018   0.00833   0.00349  -0.0022   0.9986   1.0000
   0.250   0.0514   0.00827   0.00340  -0.0075   0.9883   1.0000
   0.500   0.1008   0.00823   0.00333  -0.0119   0.9747   1.0000
   0.750   0.1397   0.00823   0.00332  -0.0138   0.9559   1.0000
   1.000   0.1666   0.00830   0.00337  -0.0130   0.9342   1.0000
   1.250   0.1874   0.00837   0.00341  -0.0109   0.9107   1.0000
   1.500   0.2071   0.00842   0.00342  -0.0084   0.8869   1.0000
   1.750   0.2282   0.00844   0.00340  -0.0062   0.8605   1.0000
   2.000   0.2497   0.00845   0.00335  -0.0041   0.8305   1.0000
   2.250   0.2727   0.00846   0.00329  -0.0023   0.7918   1.0000
   2.500   0.2960   0.00853   0.00321  -0.0006   0.7386   1.0000
   2.750   0.3196   0.00876   0.00314   0.0010   0.6541   1.0000
   3.000   0.3435   0.00939   0.00313   0.0022   0.5162   1.0000
   3.250   0.3693   0.01033   0.00334   0.0022   0.3660   1.0000
   3.500   0.3962   0.01117   0.00368   0.0020   0.2666   1.0000
   3.750   0.4237   0.01185   0.00405   0.0017   0.2073   1.0000
   4.000   0.4512   0.01251   0.00445   0.0015   0.1689   1.0000
   4.250   0.4787   0.01311   0.00492   0.0014   0.1419   1.0000
   4.500   0.5060   0.01378   0.00545   0.0014   0.1224   1.0000
   4.750   0.5334   0.01439   0.00599   0.0014   0.1072   1.0000
   5.000   0.5606   0.01504   0.00656   0.0014   0.0954   1.0000
   5.250   0.5877   0.01574   0.00722   0.0014   0.0859   1.0000
   5.500   0.6146   0.01647   0.00792   0.0015   0.0779   1.0000
   5.750   0.6413   0.01727   0.00869   0.0016   0.0713   1.0000
   6.000   0.6679   0.01816   0.00956   0.0018   0.0659   1.0000
   6.250   0.6942   0.01908   0.01041   0.0019   0.0612   1.0000
   6.500   0.7210   0.01996   0.01140   0.0021   0.0571   1.0000
   6.750   0.7473   0.02120   0.01273   0.0023   0.0540   1.0000
   7.000   0.7735   0.02219   0.01372   0.0025   0.0510   1.0000
   7.250   0.7993   0.02365   0.01545   0.0028   0.0483   1.0000
   7.500   0.8249   0.02499   0.01688   0.0030   0.0464   1.0000
   7.750   0.8495   0.02693   0.01901   0.0032   0.0451   1.0000
   8.000   0.8723   0.02930   0.02194   0.0037   0.0434   1.0000
   8.250   0.8960   0.03072   0.02345   0.0039   0.0417   1.0000
   8.500   0.9175   0.03311   0.02602   0.0042   0.0407   1.0000
   8.750   0.9328   0.03719   0.03078   0.0049   0.0405   1.0000
   9.000   0.9441   0.04185   0.03599   0.0055   0.0406   1.0000
   9.250   0.9529   0.04665   0.04122   0.0060   0.0411   1.0000
   9.500   0.7785   0.09488   0.09124  -0.0202   0.0735   1.0000
   9.750   0.7689   0.10154   0.09785  -0.0255   0.0713   1.0000
<< Back to NASA/AMES A-02 AIRFOIL (ames02-il)

Polar data table (+)

Polar graphs


<< Back to NASA/AMES A-02 AIRFOIL (ames02-il)