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

NACA 64-108 AIRFOIL (n64108-il) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: NACA 64-108 AIRFOIL (n64108-il)
Reynolds number: 500,000
Max Cl/Cd: 50.28 at α=5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-n64108-il-500000-n5.txt
Download as CSV file: xf-n64108-il-500000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 64-108 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
  ------ -------- --------- --------- -------- -------- --------
  -9.750  -0.6546   0.08420   0.08203  -0.0052   1.0000   0.0052
  -9.500  -0.6629   0.07773   0.07559  -0.0098   1.0000   0.0052
  -9.250  -0.6773   0.06924   0.06712  -0.0183   1.0000   0.0051
  -9.000  -0.6969   0.06346   0.06128  -0.0222   1.0000   0.0051
  -8.750  -0.7121   0.05696   0.05464  -0.0244   1.0000   0.0051
  -8.500  -0.7217   0.05013   0.04757  -0.0255   1.0000   0.0051
  -8.250  -0.7288   0.04216   0.03923  -0.0255   1.0000   0.0051
  -8.000  -0.7505   0.02634   0.02218  -0.0231   1.0000   0.0054
  -7.750  -0.7332   0.02365   0.01914  -0.0225   1.0000   0.0057
  -7.500  -0.7113   0.02241   0.01771  -0.0221   1.0000   0.0059
  -7.250  -0.6886   0.02127   0.01640  -0.0218   1.0000   0.0063
  -7.000  -0.6662   0.01969   0.01458  -0.0213   1.0000   0.0068
  -6.750  -0.6434   0.01816   0.01280  -0.0207   1.0000   0.0074
  -6.500  -0.6188   0.01744   0.01194  -0.0203   1.0000   0.0083
  -6.250  -0.5942   0.01673   0.01105  -0.0199   1.0000   0.0088
  -6.000  -0.5729   0.01472   0.00881  -0.0190   1.0000   0.0097
  -5.750  -0.5490   0.01407   0.00810  -0.0186   1.0000   0.0105
  -5.500  -0.5230   0.01348   0.00746  -0.0185   0.9979   0.0113
  -5.250  -0.4906   0.01286   0.00675  -0.0197   0.9881   0.0124
  -5.000  -0.4578   0.01244   0.00626  -0.0211   0.9786   0.0138
  -4.750  -0.4254   0.01189   0.00560  -0.0223   0.9684   0.0144
  -4.500  -0.3946   0.01106   0.00467  -0.0232   0.9571   0.0153
  -4.250  -0.3653   0.01039   0.00392  -0.0237   0.9449   0.0169
  -4.000  -0.3372   0.01004   0.00352  -0.0239   0.9321   0.0187
  -3.750  -0.3104   0.00971   0.00312  -0.0237   0.9191   0.0206
  -3.500  -0.2839   0.00946   0.00276  -0.0234   0.9066   0.0229
  -3.250  -0.2576   0.00921   0.00244  -0.0230   0.8943   0.0265
  -3.000  -0.2313   0.00899   0.00218  -0.0227   0.8825   0.0329
  -2.750  -0.2047   0.00881   0.00197  -0.0225   0.8708   0.0418
  -2.500  -0.1783   0.00851   0.00175  -0.0223   0.8595   0.0768
  -2.250  -0.1535   0.00767   0.00147  -0.0223   0.8483   0.2344
  -2.000  -0.1293   0.00676   0.00122  -0.0222   0.8375   0.4316
  -1.750  -0.1030   0.00646   0.00113  -0.0220   0.8273   0.5121
  -1.500  -0.0768   0.00619   0.00108  -0.0218   0.8165   0.5899
  -1.250  -0.0512   0.00595   0.00111  -0.0213   0.8059   0.6693
  -1.000  -0.0246   0.00586   0.00111  -0.0209   0.7955   0.7068
  -0.500   0.0303   0.00583   0.00106  -0.0206   0.7752   0.7350
  -0.250   0.0581   0.00584   0.00104  -0.0206   0.7648   0.7454
   0.000   0.0857   0.00584   0.00104  -0.0205   0.7545   0.7558
   0.250   0.1132   0.00585   0.00105  -0.0204   0.7446   0.7664
   0.500   0.1409   0.00586   0.00107  -0.0203   0.7346   0.7771
   0.750   0.1685   0.00588   0.00110  -0.0202   0.7243   0.7877
   1.000   0.1956   0.00593   0.00112  -0.0200   0.7055   0.7985
   1.250   0.2222   0.00600   0.00115  -0.0196   0.6793   0.8095
   1.500   0.2488   0.00608   0.00120  -0.0193   0.6510   0.8206
   1.750   0.2743   0.00629   0.00123  -0.0187   0.5955   0.8317
   2.000   0.2993   0.00659   0.00129  -0.0182   0.5238   0.8430
   2.250   0.3246   0.00690   0.00140  -0.0178   0.4552   0.8550
   2.750   0.3683   0.00888   0.00208  -0.0166   0.1017   0.8802
   3.000   0.3929   0.00928   0.00233  -0.0161   0.0520   0.8935
   3.250   0.4180   0.00951   0.00255  -0.0156   0.0377   0.9077
   3.500   0.4432   0.00972   0.00281  -0.0150   0.0310   0.9229
   3.750   0.4686   0.00997   0.00311  -0.0145   0.0257   0.9393
   4.000   0.4966   0.01020   0.00341  -0.0146   0.0222   0.9568
   4.500   0.5580   0.01122   0.00453  -0.0163   0.0165   1.0000
   4.750   0.5841   0.01164   0.00502  -0.0161   0.0159   1.0000
   5.000   0.6099   0.01213   0.00558  -0.0159   0.0152   1.0000
   5.250   0.6352   0.01273   0.00625  -0.0156   0.0144   1.0000
   5.500   0.6609   0.01324   0.00682  -0.0154   0.0132   1.0000
   5.750   0.6865   0.01374   0.00735  -0.0152   0.0122   1.0000
   6.000   0.7116   0.01436   0.00803  -0.0149   0.0114   1.0000
   6.250   0.7355   0.01530   0.00907  -0.0145   0.0107   1.0000
   6.500   0.7575   0.01699   0.01095  -0.0137   0.0098   1.0000
   6.750   0.7832   0.01744   0.01149  -0.0135   0.0092   1.0000
   7.000   0.8080   0.01821   0.01238  -0.0132   0.0084   1.0000
   7.250   0.8317   0.01936   0.01371  -0.0127   0.0077   1.0000
   7.500   0.8554   0.02037   0.01487  -0.0123   0.0071   1.0000
   7.750   0.8789   0.02126   0.01588  -0.0120   0.0067   1.0000
   8.000   0.9021   0.02209   0.01686  -0.0117   0.0063   1.0000
   8.250   0.9181   0.02516   0.02032  -0.0106   0.0058   1.0000
   8.500   0.9353   0.02781   0.02337  -0.0095   0.0056   1.0000
   8.750   0.9433   0.03294   0.02912  -0.0077   0.0051   1.0000
   9.000   0.9174   0.04657   0.04367  -0.0040   0.0046   1.0000
   9.250   0.9014   0.05457   0.05206  -0.0025   0.0044   1.0000
   9.500   0.8840   0.06089   0.05860  -0.0019   0.0044   1.0000
   9.750   0.8581   0.06601   0.06383  -0.0013   0.0045   1.0000
  10.000   0.8347   0.07196   0.06987  -0.0043   0.0045   1.0000
  10.250   0.8111   0.08123   0.07922  -0.0123   0.0049   1.0000
<< Back to NACA 64-108 AIRFOIL (n64108-il)

Polar data table (+)

Polar graphs


<< Back to NACA 64-108 AIRFOIL (n64108-il)