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=500,000 Ncrit=5


Details Polar file
Airfoil: EPPLER 908 AIRFOIL (e908-il)
Reynolds number: 500,000
Max Cl/Cd: 91.8 at α=2°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-e908-il-500000-n5.txt
Download as CSV file: xf-e908-il-500000-n5.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 =     0.500 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.3114   0.06896   0.06657  -0.0904   0.9690   0.0022
  -8.000  -0.3160   0.06494   0.06258  -0.0929   0.9631   0.0022
  -7.750  -0.3117   0.05714   0.05478  -0.1042   0.9592   0.0021
  -7.500  -0.3095   0.05111   0.04864  -0.1088   0.9549   0.0021
  -7.250  -0.3070   0.04588   0.04325  -0.1108   0.9504   0.0020
  -7.000  -0.2957   0.04014   0.03727  -0.1134   0.9481   0.0020
  -6.750  -0.2804   0.03427   0.03104  -0.1156   0.9465   0.0019
  -6.500  -0.2612   0.02924   0.02553  -0.1170   0.9454   0.0018
  -6.250  -0.2599   0.02662   0.02262  -0.1128   0.9394   0.0018
  -6.000  -0.2409   0.02302   0.01856  -0.1121   0.9373   0.0017
  -5.750  -0.2174   0.01981   0.01487  -0.1118   0.9360   0.0017
  -5.500  -0.1910   0.01734   0.01198  -0.1118   0.9351   0.0017
  -5.250  -0.1633   0.01533   0.00964  -0.1119   0.9343   0.0019
  -5.000  -0.1346   0.01391   0.00800  -0.1122   0.9337   0.0021
  -4.750  -0.1049   0.01311   0.00706  -0.1129   0.9331   0.0025
  -4.250  -0.0663   0.01184   0.00564  -0.1099   0.9272   0.0042
  -4.000  -0.0385   0.01167   0.00547  -0.1103   0.9256   0.0066
  -3.750  -0.0097   0.01099   0.00465  -0.1106   0.9242   0.0071
  -3.500   0.0204   0.01026   0.00368  -0.1111   0.9230   0.0149
  -3.250   0.0509   0.00964   0.00324  -0.1121   0.9219   0.0768
  -3.000   0.0827   0.00911   0.00298  -0.1135   0.9209   0.1607
  -2.750   0.1152   0.00872   0.00276  -0.1150   0.9200   0.2219
  -2.500   0.1340   0.00831   0.00269  -0.1136   0.9162   0.3261
  -2.250   0.1575   0.00779   0.00259  -0.1132   0.9133   0.4557
  -2.000   0.1853   0.00743   0.00246  -0.1135   0.9113   0.5534
  -1.750   0.2158   0.00727   0.00242  -0.1143   0.9098   0.6110
  -1.500   0.2475   0.00719   0.00235  -0.1153   0.9086   0.6377
  -1.250   0.2797   0.00711   0.00229  -0.1164   0.9074   0.6580
  -1.000   0.3050   0.00711   0.00230  -0.1160   0.9047   0.6715
  -0.750   0.3277   0.00712   0.00234  -0.1150   0.9011   0.6822
  -0.500   0.3561   0.00709   0.00233  -0.1152   0.8987   0.6941
  -0.250   0.3866   0.00703   0.00231  -0.1159   0.8968   0.7098
   0.000   0.4181   0.00696   0.00230  -0.1169   0.8951   0.7295
   0.250   0.4470   0.00691   0.00233  -0.1172   0.8929   0.7462
   0.500   0.4668   0.00693   0.00244  -0.1155   0.8881   0.7594
   0.750   0.4952   0.00690   0.00246  -0.1157   0.8850   0.7703
   1.000   0.5262   0.00685   0.00251  -0.1165   0.8826   0.7782
   1.250   0.5554   0.00682   0.00254  -0.1170   0.8798   0.7854
   1.500   0.5765   0.00683   0.00264  -0.1155   0.8741   0.7929
   1.750   0.6069   0.00669   0.00256  -0.1160   0.8662   0.7995
   2.000   0.6215   0.00677   0.00185  -0.1119   0.6896   0.8075
   2.250   0.6025   0.00835   0.00232  -0.1020   0.4562   0.8196
   2.500   0.5864   0.01129   0.00333  -0.0941   0.0064   0.8334
   2.750   0.6092   0.01148   0.00359  -0.0931   0.0027   0.8465
   3.250   0.6507   0.01194   0.00446  -0.0899   0.0024   0.8862
   3.500   0.6681   0.01228   0.00502  -0.0875   0.0022   0.9288
   3.750   0.6953   0.01294   0.00583  -0.0875   0.0022   1.0000
   4.000   0.7133   0.01380   0.00679  -0.0854   0.0022   1.0000
   4.250   0.7319   0.01477   0.00784  -0.0835   0.0022   1.0000
   4.500   0.7538   0.01607   0.00924  -0.0822   0.0022   1.0000
   4.750   0.7825   0.01773   0.01101  -0.0822   0.0023   1.0000
   5.000   0.8166   0.01985   0.01331  -0.0831   0.0025   1.0000
   5.250   0.8504   0.02258   0.01629  -0.0837   0.0027   1.0000
   7.250   0.9552   0.03775   0.03416  -0.0595   0.0047   1.0000
   7.500   0.9576   0.04155   0.03818  -0.0560   0.0046   1.0000
   7.750   0.9568   0.04552   0.04237  -0.0523   0.0045   1.0000
   8.000   0.9529   0.04948   0.04651  -0.0485   0.0044   1.0000
   8.250   0.9455   0.05330   0.05050  -0.0446   0.0043   1.0000
   8.500   0.9318   0.05668   0.05403  -0.0399   0.0043   1.0000
   8.750   0.9148   0.06002   0.05750  -0.0355   0.0043   1.0000
   9.000   0.8950   0.06379   0.06140  -0.0318   0.0043   1.0000
   9.250   0.8740   0.06789   0.06563  -0.0291   0.0043   1.0000
   9.500   0.8503   0.07280   0.07067  -0.0274   0.0044   1.0000
   9.750   0.8258   0.07829   0.07627  -0.0273   0.0044   1.0000
  10.000   0.7998   0.08490   0.08298  -0.0290   0.0045   1.0000
<< Back to EPPLER 908 AIRFOIL (e908-il)

Polar data table (+)

Polar graphs


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