Airfoil Tools
Search 1636 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

NACA 0008-34 (naca000834-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9


Details Polar file
Airfoil: NACA 0008-34 (naca000834-il)
Reynolds number: 1,000,000
Max Cl/Cd: 42.99 at α=3.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-naca000834-il-1000000.txt
Download as CSV file: xf-naca000834-il-1000000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 0008-34                                    
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     1.000 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.6711   0.08448   0.08297  -0.0083   1.0000   0.0053
  -9.250  -0.6737   0.07614   0.07462  -0.0175   1.0000   0.0052
  -9.000  -0.6815   0.07058   0.06899  -0.0217   1.0000   0.0052
  -8.750  -0.6921   0.06625   0.06458  -0.0223   1.0000   0.0052
  -8.500  -0.7032   0.06112   0.05933  -0.0224   1.0000   0.0052
  -8.250  -0.7104   0.05604   0.05410  -0.0219   1.0000   0.0053
  -8.000  -0.7131   0.05148   0.04937  -0.0208   1.0000   0.0054
  -7.750  -0.7114   0.04748   0.04520  -0.0194   1.0000   0.0055
  -7.500  -0.7064   0.04413   0.04168  -0.0177   1.0000   0.0057
  -7.250  -0.6994   0.04133   0.03870  -0.0158   1.0000   0.0058
  -7.000  -0.6919   0.03855   0.03575  -0.0134   1.0000   0.0060
  -6.750  -0.6836   0.03589   0.03290  -0.0108   1.0000   0.0063
  -6.500  -0.6747   0.03333   0.03015  -0.0079   1.0000   0.0066
  -6.250  -0.6649   0.03087   0.02749  -0.0049   1.0000   0.0071
  -6.000  -0.6530   0.02861   0.02502  -0.0018   1.0000   0.0079
  -5.500  -0.6243   0.02205   0.01778   0.0041   0.9994   0.0082
  -5.250  -0.5954   0.01938   0.01480   0.0040   0.9981   0.0084
  -5.000  -0.5662   0.01673   0.01184   0.0039   0.9969   0.0082
  -4.750  -0.5341   0.01581   0.01079   0.0029   0.9957   0.0089
  -4.500  -0.5042   0.01352   0.00829   0.0026   0.9948   0.0086
  -4.250  -0.4747   0.01179   0.00640   0.0024   0.9937   0.0085
  -4.000  -0.4462   0.01067   0.00518   0.0022   0.9915   0.0089
  -3.750  -0.4152   0.01005   0.00450   0.0014   0.9892   0.0099
  -3.500  -0.3852   0.00896   0.00326   0.0007   0.9870   0.0122
  -3.250  -0.3517   0.00860   0.00286  -0.0008   0.9854   0.0145
  -3.000  -0.3171   0.00832   0.00254  -0.0024   0.9840   0.0174
  -2.750  -0.2823   0.00786   0.00205  -0.0041   0.9827   0.0345
  -2.500  -0.2573   0.00654   0.00164  -0.0043   0.9787   0.2772
  -2.250  -0.2321   0.00550   0.00139  -0.0045   0.9739   0.4913
  -2.000  -0.2047   0.00485   0.00124  -0.0048   0.9699   0.6295
  -1.750  -0.1817   0.00439   0.00116  -0.0037   0.9631   0.7401
  -1.500  -0.1568   0.00414   0.00112  -0.0029   0.9569   0.8052
  -1.250  -0.1310   0.00405   0.00109  -0.0024   0.9499   0.8330
  -1.000  -0.1051   0.00396   0.00106  -0.0018   0.9430   0.8570
  -0.750  -0.0790   0.00392   0.00104  -0.0013   0.9349   0.8717
  -0.500  -0.0526   0.00389   0.00101  -0.0009   0.9274   0.8858
  -0.250  -0.0264   0.00387   0.00101  -0.0004   0.9180   0.8980
   0.000   0.0000   0.00387   0.00101   0.0000   0.9084   0.9084
   0.250   0.0264   0.00387   0.00101   0.0004   0.8981   0.9180
   0.500   0.0525   0.00389   0.00101   0.0009   0.8859   0.9274
   0.750   0.0790   0.00392   0.00104   0.0013   0.8718   0.9350
   1.000   0.1051   0.00396   0.00106   0.0018   0.8570   0.9430
   1.250   0.1310   0.00405   0.00109   0.0024   0.8335   0.9499
   1.500   0.1567   0.00414   0.00112   0.0029   0.8049   0.9569
   1.750   0.1817   0.00439   0.00116   0.0037   0.7406   0.9631
   2.000   0.2046   0.00486   0.00124   0.0048   0.6285   0.9699
   2.250   0.2321   0.00550   0.00139   0.0045   0.4907   0.9739
   2.500   0.2573   0.00653   0.00164   0.0044   0.2794   0.9787
   2.750   0.2823   0.00785   0.00204   0.0041   0.0356   0.9827
   3.000   0.3170   0.00832   0.00253   0.0025   0.0174   0.9840
   3.250   0.3516   0.00860   0.00286   0.0008   0.0145   0.9854
   3.500   0.3852   0.00896   0.00326  -0.0007   0.0122   0.9871
   3.750   0.4153   0.01004   0.00448  -0.0014   0.0099   0.9893
   4.000   0.4462   0.01067   0.00518  -0.0022   0.0089   0.9915
   4.250   0.4746   0.01179   0.00639  -0.0024   0.0085   0.9938
   4.500   0.5041   0.01352   0.00828  -0.0026   0.0086   0.9948
   4.750   0.5340   0.01584   0.01083  -0.0029   0.0090   0.9957
   5.000   0.5662   0.01669   0.01180  -0.0039   0.0082   0.9969
   5.250   0.5954   0.01936   0.01477  -0.0040   0.0084   0.9981
   5.500   0.6241   0.02214   0.01789  -0.0041   0.0082   0.9995
   6.000   0.6529   0.02860   0.02501   0.0018   0.0079   1.0000
   6.250   0.6648   0.03087   0.02749   0.0049   0.0071   1.0000
   6.500   0.6747   0.03334   0.03016   0.0079   0.0066   1.0000
   6.750   0.6836   0.03589   0.03291   0.0108   0.0063   1.0000
   7.000   0.6919   0.03855   0.03575   0.0134   0.0060   1.0000
   7.250   0.6994   0.04132   0.03868   0.0158   0.0058   1.0000
   7.500   0.7065   0.04414   0.04169   0.0177   0.0057   1.0000
   7.750   0.7115   0.04749   0.04521   0.0194   0.0055   1.0000
   8.000   0.7132   0.05150   0.04939   0.0208   0.0054   1.0000
   8.250   0.7106   0.05606   0.05412   0.0218   0.0053   1.0000
   8.500   0.7034   0.06115   0.05936   0.0223   0.0052   1.0000
   8.750   0.6922   0.06632   0.06465   0.0222   0.0052   1.0000
   9.000   0.6816   0.07068   0.06909   0.0215   0.0052   1.0000
   9.250   0.6743   0.07623   0.07471   0.0173   0.0052   1.0000
<< Back to NACA 0008-34 (naca000834-il)

Polar data table (+)

Polar graphs


<< Back to NACA 0008-34 (naca000834-il)