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

EPPLER 908 AIRFOIL (e908-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: EPPLER 908 AIRFOIL (e908-il)
Reynolds number: 50,000
Max Cl/Cd: 22.19 at α=6°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e908-il-50000.txt
Download as CSV file: xf-e908-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 908 AIRFOIL                              
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.250  -0.5086   0.10109   0.09500  -0.0173   1.0000   0.2472
  -7.000  -0.5021   0.09737   0.09131  -0.0154   1.0000   0.2557
  -6.750  -0.5178   0.09537   0.08943  -0.0130   1.0000   0.2651
  -6.500  -0.5440   0.09407   0.08828  -0.0130   1.0000   0.2753
  -6.250  -0.5448   0.09085   0.08514  -0.0111   1.0000   0.2890
  -6.000  -0.5277   0.08709   0.08138  -0.0063   1.0000   0.3095
  -5.000  -0.5040   0.05771   0.05086  -0.0425   1.0000   0.1268
  -4.750  -0.4782   0.05078   0.04326  -0.0456   1.0000   0.0996
  -4.500  -0.4508   0.04549   0.03710  -0.0475   1.0000   0.0881
  -4.250  -0.4266   0.04173   0.03291  -0.0480   1.0000   0.0869
  -4.000  -0.3993   0.03833   0.02885  -0.0485   1.0000   0.0874
  -3.750  -0.3714   0.03532   0.02526  -0.0485   1.0000   0.0877
  -3.500  -0.3434   0.03261   0.02205  -0.0481   1.0000   0.0874
  -3.250  -0.3157   0.03040   0.01936  -0.0472   1.0000   0.0887
  -3.000  -0.2907   0.02837   0.01731  -0.0459   1.0000   0.0950
  -2.750  -0.2662   0.02684   0.01567  -0.0442   1.0000   0.1075
  -2.500  -0.2396   0.02514   0.01423  -0.0436   1.0000   0.1661
  -2.250  -0.2359   0.02096   0.01405  -0.0308   1.0000   1.0000
  -2.000  -0.2151   0.02075   0.01309  -0.0308   1.0000   1.0000
  -1.750  -0.1919   0.02072   0.01244  -0.0311   1.0000   1.0000
  -1.500  -0.1678   0.02078   0.01183  -0.0315   1.0000   1.0000
  -1.250  -0.1436   0.02091   0.01150  -0.0318   1.0000   1.0000
  -1.000  -0.1197   0.02109   0.01129  -0.0320   1.0000   1.0000
  -0.750  -0.0959   0.02131   0.01117  -0.0322   1.0000   1.0000
  -0.500  -0.0725   0.02158   0.01115  -0.0323   1.0000   1.0000
  -0.250  -0.0494   0.02187   0.01117  -0.0324   1.0000   1.0000
   0.000  -0.0264   0.02219   0.01114  -0.0324   1.0000   1.0000
   0.250  -0.0038   0.02255   0.01130  -0.0323   1.0000   1.0000
   0.500   0.0184   0.02293   0.01152  -0.0323   1.0000   1.0000
   0.750   0.0404   0.02335   0.01180  -0.0322   1.0000   1.0000
   1.000   0.0621   0.02380   0.01213  -0.0321   1.0000   1.0000
   1.250   0.0835   0.02427   0.01252  -0.0319   1.0000   1.0000
   1.500   0.1046   0.02478   0.01297  -0.0318   1.0000   1.0000
   1.750   0.1254   0.02532   0.01346  -0.0316   1.0000   1.0000
   2.000   0.1460   0.02589   0.01400  -0.0315   1.0000   1.0000
   2.250   0.1662   0.02650   0.01460  -0.0313   1.0000   1.0000
   2.500   0.1862   0.02714   0.01525  -0.0312   1.0000   1.0000
   2.750   0.2059   0.02782   0.01596  -0.0310   1.0000   1.0000
   3.000   0.2253   0.02853   0.01673  -0.0308   1.0000   1.0000
   3.250   0.2445   0.02929   0.01757  -0.0307   1.0000   1.0000
   3.500   0.2632   0.03009   0.01854  -0.0305   1.0000   1.0000
   3.750   0.2817   0.03093   0.01949  -0.0304   1.0000   1.0000
   4.000   0.2999   0.03183   0.02052  -0.0303   1.0000   1.0000
   4.250   0.3178   0.03277   0.02161  -0.0302   1.0000   1.0000
   4.500   0.3353   0.03377   0.02278  -0.0301   1.0000   1.0000
   4.750   0.3524   0.03483   0.02403  -0.0301   1.0000   1.0000
   5.000   0.3691   0.03596   0.02537  -0.0301   1.0000   1.0000
   5.250   0.3854   0.03717   0.02697  -0.0301   1.0000   1.0000
   5.500   0.4012   0.03846   0.02853  -0.0302   1.0000   1.0000
   5.750   0.4684   0.04162   0.03241  -0.0400   0.9682   1.0000
   6.000   0.8872   0.03999   0.02887  -0.0655   0.0730   1.0000
   6.250   0.9263   0.04517   0.03427  -0.0666   0.0761   1.0000
   6.500   0.9452   0.04718   0.03728  -0.0626   0.0839   1.0000
   6.750   0.9666   0.05094   0.04167  -0.0600   0.0934   1.0000
   7.000   0.9830   0.05501   0.04644  -0.0568   0.1078   1.0000
   7.250   0.9993   0.05958   0.05172  -0.0539   0.1314   1.0000
   7.500   1.0082   0.06459   0.05760  -0.0509   0.1708   1.0000
   7.750   1.0060   0.07097   0.06487  -0.0500   0.2346   1.0000
   8.000   0.9766   0.07540   0.06967  -0.0479   0.2550   1.0000
   8.250   0.9457   0.08032   0.07478  -0.0472   0.2762   1.0000
   8.500   0.9112   0.08472   0.07918  -0.0458   0.2787   1.0000
   8.750   0.8755   0.09092   0.08534  -0.0482   0.2905   1.0000
   9.000   0.8311   0.10119   0.09549  -0.0586   0.3403   1.0000
   9.250   0.7578   0.09433   0.08904  -0.0370   0.2354   1.0000
<< Back to EPPLER 908 AIRFOIL (e908-il)

Polar data table (+)

Polar graphs


<< Back to EPPLER 908 AIRFOIL (e908-il)