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

EPPLER 678 AIRFOIL (e678-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: EPPLER 678 AIRFOIL (e678-il)
Reynolds number: 200,000
Max Cl/Cd: 79.34 at α=8°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e678-il-200000.txt
Download as CSV file: xf-e678-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 678 AIRFOIL                              
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.2006   0.10232   0.09890  -0.0928   0.9490   0.0539
  -8.750  -0.2182   0.09673   0.09337  -0.0993   0.9384   0.0547
  -8.500  -0.2335   0.09023   0.08689  -0.1067   0.9286   0.0549
  -8.250  -0.1998   0.08848   0.08514  -0.1009   0.9286   0.0566
  -8.000  -0.1813   0.08562   0.08227  -0.1025   0.9251   0.0583
  -7.750  -0.1671   0.08118   0.07781  -0.1077   0.9223   0.0600
  -7.500  -0.1773   0.07709   0.07376  -0.1093   0.9107   0.0617
  -7.250  -0.1912   0.07085   0.06751  -0.1154   0.8980   0.0626
  -7.000  -0.2146   0.06326   0.05917  -0.1266   0.8813   0.0666
  -6.750  -0.1958   0.05641   0.05252  -0.1288   0.8798   0.0687
  -6.500  -0.1872   0.05429   0.05039  -0.1273   0.8700   0.0703
  -6.250  -0.1645   0.05081   0.04676  -0.1301   0.8664   0.0739
  -6.000  -0.1537   0.03797   0.03246  -0.1339   0.8590   0.0466
  -5.750  -0.1298   0.03763   0.03216  -0.1348   0.8532   0.0573
  -5.500  -0.0905   0.03011   0.02326  -0.1367   0.8517   0.0362
  -5.250  -0.0541   0.02759   0.02041  -0.1387   0.8504   0.0359
  -5.000  -0.0162   0.02574   0.01825  -0.1407   0.8493   0.0362
  -4.750  -0.0006   0.02485   0.01721  -0.1386   0.8399   0.0368
  -4.500   0.0328   0.02320   0.01557  -0.1399   0.8378   0.0386
  -4.250   0.0688   0.02221   0.01456  -0.1416   0.8362   0.0423
  -4.000   0.1061   0.02097   0.01329  -0.1436   0.8352   0.0473
  -3.750   0.1247   0.02068   0.01300  -0.1424   0.8273   0.0524
  -3.500   0.1600   0.01958   0.01198  -0.1444   0.8245   0.0730
  -3.250   0.2009   0.01819   0.01108  -0.1479   0.8232   0.1749
  -3.000   0.2447   0.01667   0.01061  -0.1526   0.8224   0.4292
  -2.750   0.2804   0.01599   0.01053  -0.1540   0.8211   0.5959
  -2.500   0.3150   0.01578   0.01037  -0.1546   0.8199   0.6608
  -2.250   0.3240   0.01637   0.01094  -0.1511   0.8105   0.6905
  -2.000   0.3540   0.01636   0.01087  -0.1509   0.8084   0.7187
  -1.750   0.3853   0.01632   0.01076  -0.1508   0.8068   0.7420
  -1.500   0.4170   0.01622   0.01060  -0.1508   0.8055   0.7602
  -1.250   0.4218   0.01688   0.01124  -0.1465   0.7968   0.7740
  -1.000   0.4507   0.01676   0.01105  -0.1462   0.7946   0.7890
  -0.750   0.4824   0.01658   0.01080  -0.1464   0.7930   0.8032
  -0.500   0.5128   0.01638   0.01055  -0.1463   0.7917   0.8163
  -0.250   0.5129   0.01703   0.01121  -0.1410   0.7840   0.8276
   0.000   0.5399   0.01685   0.01099  -0.1405   0.7816   0.8393
   0.250   0.5740   0.01658   0.01065  -0.1416   0.7799   0.8500
   0.500   0.6065   0.01627   0.01028  -0.1422   0.7785   0.8588
   0.750   0.6211   0.01655   0.01055  -0.1399   0.7736   0.8680
   1.000   0.6434   0.01665   0.01063  -0.1391   0.7695   0.8776
   1.250   0.6722   0.01640   0.01035  -0.1391   0.7672   0.8853
   1.500   0.7114   0.01614   0.01003  -0.1415   0.7655   0.8927
   1.750   0.7503   0.01577   0.00962  -0.1436   0.7639   0.8986
   2.000   0.7589   0.01625   0.01012  -0.1405   0.7578   0.9073
   2.250   0.7848   0.01614   0.01000  -0.1402   0.7542   0.9138
   2.500   0.8257   0.01588   0.00970  -0.1429   0.7518   0.9196
   2.750   0.8732   0.01560   0.00939  -0.1469   0.7497   0.9241
   3.000   0.8791   0.01594   0.00977  -0.1431   0.7438   0.9329
   3.250   0.9056   0.01595   0.00979  -0.1431   0.7394   0.9397
   3.500   0.9493   0.01579   0.00961  -0.1464   0.7364   0.9451
   3.750   0.9983   0.01566   0.00943  -0.1508   0.7336   0.9495
   4.000   0.9965   0.01613   0.01001  -0.1456   0.7264   0.9620
   4.250   1.0363   0.01610   0.00999  -0.1482   0.7221   0.9690
   4.500   1.0907   0.01599   0.00984  -0.1537   0.7184   0.9730
   4.750   1.1067   0.01646   0.01044  -0.1522   0.7111   1.0000
   5.000   1.1463   0.01657   0.01057  -0.1550   0.7057   1.0000
   5.250   1.1891   0.01672   0.01071  -0.1584   0.7006   1.0000
   5.500   1.2059   0.01716   0.01125  -0.1570   0.6929   1.0000
   5.750   1.2527   0.01717   0.01125  -0.1610   0.6871   1.0000
   6.000   1.2688   0.01761   0.01180  -0.1594   0.6790   1.0000
   6.250   1.3075   0.01764   0.01184  -0.1618   0.6719   1.0000
   6.500   1.3281   0.01796   0.01227  -0.1610   0.6638   1.0000
   6.750   1.3615   0.01799   0.01235  -0.1623   0.6560   1.0000
   7.000   1.3803   0.01823   0.01269  -0.1611   0.6472   1.0000
   7.250   1.4147   0.01814   0.01262  -0.1625   0.6388   1.0000
   7.500   1.4221   0.01838   0.01300  -0.1589   0.6287   1.0000
   7.750   1.4469   0.01833   0.01299  -0.1585   0.6193   1.0000
   8.000   1.4598   0.01840   0.01313  -0.1558   0.6083   1.0000
   8.250   1.4648   0.01867   0.01352  -0.1518   0.5963   1.0000
   8.500   1.4729   0.01890   0.01382  -0.1484   0.5832   1.0000
   8.750   1.4788   0.01921   0.01419  -0.1446   0.5676   1.0000
   9.000   1.4837   0.01957   0.01457  -0.1407   0.5481   1.0000
   9.250   1.4854   0.02017   0.01518  -0.1365   0.5244   1.0000
   9.500   1.4874   0.02091   0.01582  -0.1324   0.4935   1.0000
   9.750   1.4879   0.02192   0.01663  -0.1283   0.4578   1.0000
  10.000   1.4861   0.02327   0.01776  -0.1240   0.4232   1.0000
  10.250   1.4832   0.02483   0.01916  -0.1199   0.3906   1.0000
  10.500   1.4794   0.02658   0.02075  -0.1160   0.3598   1.0000
  10.750   1.4750   0.02848   0.02252  -0.1122   0.3310   1.0000
  11.000   1.4703   0.03054   0.02446  -0.1087   0.3025   1.0000
  11.250   1.4664   0.03270   0.02651  -0.1055   0.2754   1.0000
  11.500   1.4625   0.03501   0.02872  -0.1026   0.2497   1.0000
  11.750   1.4593   0.03740   0.03103  -0.1000   0.2267   1.0000
  12.000   1.4580   0.03978   0.03335  -0.0978   0.2036   1.0000
  12.250   1.4558   0.04237   0.03586  -0.0958   0.1827   1.0000
  12.500   1.4554   0.04492   0.03835  -0.0942   0.1613   1.0000
  12.750   1.4530   0.04775   0.04110  -0.0925   0.1402   1.0000
  13.000   1.4513   0.05063   0.04390  -0.0911   0.1170   1.0000
  13.250   1.4458   0.05399   0.04711  -0.0897   0.0913   1.0000
  13.500   1.4347   0.05806   0.05096  -0.0881   0.0649   1.0000
  13.750   1.4237   0.06229   0.05507  -0.0867   0.0473   1.0000
  14.000   1.4177   0.06610   0.05889  -0.0856   0.0397   1.0000
  14.250   1.4105   0.07011   0.06293  -0.0848   0.0358   1.0000
  14.500   1.4104   0.07336   0.06631  -0.0842   0.0330   1.0000
  14.750   1.4088   0.07685   0.06987  -0.0838   0.0310   1.0000
  15.000   1.4034   0.08085   0.07392  -0.0835   0.0296   1.0000
  15.250   1.4028   0.08428   0.07745  -0.0832   0.0285   1.0000
  15.500   1.4054   0.08731   0.08061  -0.0829   0.0274   1.0000
  15.750   1.4082   0.09033   0.08372  -0.0828   0.0265   1.0000
  16.000   1.4118   0.09324   0.08671  -0.0827   0.0255   1.0000
<< Back to EPPLER 678 AIRFOIL (e678-il)

Polar data table (+)

Polar graphs


<< Back to EPPLER 678 AIRFOIL (e678-il)