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=500,000 Ncrit=5


Details Polar file
Airfoil: NASA/AMES A-02 AIRFOIL (ames02-il)
Reynolds number: 500,000
Max Cl/Cd: 58.74 at α=7.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-ames02-il-500000-n5.txt
Download as CSV file: xf-ames02-il-500000-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 =     0.500 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -11.750  -0.7484   0.14342   0.14106   0.0520   1.0000   0.0080
 -11.500  -0.7471   0.13860   0.13624   0.0499   1.0000   0.0080
  -9.500  -0.9775   0.03867   0.03523  -0.0039   1.0000   0.0082
  -9.250  -0.9740   0.03302   0.02902  -0.0035   1.0000   0.0086
  -9.000  -0.9616   0.02907   0.02456  -0.0029   1.0000   0.0089
  -8.750  -0.9449   0.02586   0.02086  -0.0024   1.0000   0.0093
  -8.500  -0.9244   0.02353   0.01812  -0.0019   1.0000   0.0095
  -8.250  -0.9014   0.02190   0.01620  -0.0015   1.0000   0.0097
  -8.000  -0.8775   0.02061   0.01476  -0.0013   1.0000   0.0100
  -7.750  -0.8522   0.01971   0.01377  -0.0012   1.0000   0.0102
  -7.500  -0.8264   0.01890   0.01286  -0.0011   1.0000   0.0105
  -7.250  -0.8003   0.01806   0.01191  -0.0010   1.0000   0.0108
  -7.000  -0.7740   0.01723   0.01095  -0.0009   1.0000   0.0111
  -6.750  -0.7474   0.01641   0.01002  -0.0008   1.0000   0.0115
  -6.500  -0.7205   0.01566   0.00914  -0.0007   1.0000   0.0119
  -6.250  -0.6935   0.01494   0.00836  -0.0007   1.0000   0.0123
  -6.000  -0.6660   0.01445   0.00786  -0.0008   1.0000   0.0128
  -5.750  -0.6382   0.01398   0.00734  -0.0009   1.0000   0.0135
  -5.500  -0.6103   0.01347   0.00678  -0.0009   1.0000   0.0142
  -5.250  -0.5823   0.01294   0.00619  -0.0010   1.0000   0.0147
  -5.000  -0.5542   0.01246   0.00572  -0.0012   1.0000   0.0153
  -4.750  -0.5257   0.01205   0.00529  -0.0014   1.0000   0.0161
  -4.500  -0.4970   0.01166   0.00486  -0.0016   1.0000   0.0170
  -4.250  -0.4683   0.01126   0.00447  -0.0018   1.0000   0.0180
  -4.000  -0.4392   0.01094   0.00415  -0.0021   1.0000   0.0195
  -3.750  -0.4101   0.01062   0.00384  -0.0024   1.0000   0.0212
  -3.500  -0.3788   0.01037   0.00358  -0.0031   0.9769   0.0233
  -3.250  -0.3494   0.01015   0.00336  -0.0033   0.9639   0.0259
  -3.000  -0.3222   0.00996   0.00318  -0.0030   0.9522   0.0294
  -2.750  -0.2958   0.00978   0.00301  -0.0025   0.9417   0.0344
  -2.500  -0.2698   0.00962   0.00286  -0.0019   0.9319   0.0415
  -2.250  -0.2431   0.00942   0.00271  -0.0015   0.9220   0.0530
  -2.000  -0.2165   0.00920   0.00257  -0.0011   0.9128   0.0746
  -1.750  -0.1900   0.00892   0.00243  -0.0007   0.9029   0.1133
  -1.500  -0.1632   0.00854   0.00229  -0.0005   0.8909   0.1793
  -1.250  -0.1372   0.00807   0.00214  -0.0001   0.8744   0.2775
  -1.000  -0.1113   0.00749   0.00200   0.0001   0.8553   0.4158
  -0.750  -0.0847   0.00702   0.00192   0.0004   0.8369   0.5354
  -0.500  -0.0577   0.00674   0.00189   0.0008   0.8190   0.6188
  -0.250  -0.0311   0.00654   0.00186   0.0013   0.7964   0.6852
   0.000  -0.0043   0.00644   0.00183   0.0019   0.7696   0.7365
   0.250   0.0221   0.00634   0.00181   0.0025   0.7399   0.7842
   0.500   0.0478   0.00625   0.00178   0.0034   0.7027   0.8335
   0.750   0.0729   0.00626   0.00174   0.0044   0.6524   0.8736
   1.000   0.0981   0.00638   0.00170   0.0053   0.5858   0.9123
   1.250   0.1276   0.00663   0.00168   0.0052   0.5013   0.9549
   1.500   0.1633   0.00700   0.00169   0.0033   0.4075   0.9920
   1.750   0.1932   0.00738   0.00176   0.0026   0.3330   1.0000
   2.000   0.2219   0.00773   0.00184   0.0023   0.2746   1.0000
   2.250   0.2505   0.00802   0.00194   0.0020   0.2311   1.0000
   2.500   0.2792   0.00829   0.00205   0.0018   0.1960   1.0000
   3.000   0.3365   0.00879   0.00229   0.0013   0.1427   1.0000
   3.250   0.3652   0.00904   0.00243   0.0011   0.1230   1.0000
   3.500   0.3938   0.00928   0.00259   0.0009   0.1063   1.0000
   4.000   0.4509   0.00977   0.00294   0.0006   0.0810   1.0000
   4.250   0.4793   0.01002   0.00314   0.0004   0.0710   1.0000
   4.500   0.5077   0.01029   0.00335   0.0003   0.0627   1.0000
   4.750   0.5361   0.01056   0.00358   0.0001   0.0558   1.0000
   5.000   0.5644   0.01083   0.00382   0.0000   0.0498   1.0000
   5.250   0.5926   0.01111   0.00408  -0.0001   0.0448   1.0000
   5.500   0.6207   0.01141   0.00436  -0.0002   0.0406   1.0000
   5.750   0.6488   0.01171   0.00466  -0.0003   0.0371   1.0000
   6.000   0.6767   0.01203   0.00498  -0.0004   0.0341   1.0000
   6.250   0.7046   0.01236   0.00532  -0.0005   0.0314   1.0000
   6.500   0.7323   0.01273   0.00568  -0.0006   0.0290   1.0000
   6.750   0.7599   0.01308   0.00607  -0.0006   0.0271   1.0000
   7.000   0.7873   0.01350   0.00649  -0.0007   0.0255   1.0000
   7.250   0.8147   0.01389   0.00693  -0.0007   0.0240   1.0000
   7.500   0.8417   0.01436   0.00740  -0.0008   0.0227   1.0000
   7.750   0.8688   0.01479   0.00789  -0.0008   0.0215   1.0000
   8.000   0.8955   0.01525   0.00839  -0.0008   0.0204   1.0000
   8.250   0.9220   0.01580   0.00898  -0.0007   0.0196   1.0000
   8.500   0.9483   0.01634   0.00959  -0.0007   0.0188   1.0000
   8.750   0.9744   0.01689   0.01020  -0.0006   0.0181   1.0000
   9.000   1.0002   0.01749   0.01085  -0.0005   0.0175   1.0000
   9.250   1.0253   0.01821   0.01162  -0.0004   0.0170   1.0000
   9.500   1.0505   0.01887   0.01240  -0.0002   0.0165   1.0000
   9.750   1.0753   0.01956   0.01318   0.0000   0.0159   1.0000
  10.000   1.0999   0.02023   0.01393   0.0001   0.0154   1.0000
  10.250   1.1240   0.02096   0.01473   0.0003   0.0150   1.0000
  10.500   1.1467   0.02190   0.01573   0.0006   0.0146   1.0000
  10.750   1.1695   0.02281   0.01679   0.0010   0.0143   1.0000
  11.000   1.1915   0.02380   0.01793   0.0014   0.0140   1.0000
  11.250   1.2127   0.02485   0.01912   0.0018   0.0137   1.0000
  11.500   1.2331   0.02595   0.02036   0.0023   0.0134   1.0000
  11.750   1.2527   0.02708   0.02162   0.0028   0.0131   1.0000
  12.000   1.2712   0.02827   0.02293   0.0034   0.0129   1.0000
  12.250   1.2885   0.02952   0.02431   0.0041   0.0127   1.0000
  12.500   1.3039   0.03090   0.02582   0.0048   0.0125   1.0000
  12.750   1.3163   0.03252   0.02756   0.0058   0.0123   1.0000
  13.000   1.3248   0.03444   0.02963   0.0069   0.0122   1.0000
  13.250   1.3305   0.03646   0.03188   0.0081   0.0121   1.0000
  13.500   1.3286   0.03868   0.03431   0.0100   0.0120   1.0000
  13.750   1.3218   0.04151   0.03734   0.0110   0.0119   1.0000
  14.000   1.3127   0.04521   0.04126   0.0104   0.0117   1.0000
  14.250   1.2998   0.05027   0.04654   0.0078   0.0117   1.0000
  14.500   1.2805   0.05780   0.05430   0.0020   0.0117   1.0000
  14.750   1.2490   0.06991   0.06668  -0.0080   0.0118   1.0000
  15.000   1.2056   0.08451   0.08148  -0.0179   0.0119   1.0000
  15.250   1.1585   0.09871   0.09581  -0.0262   0.0121   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)