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

EPPLER 325 AIRFOIL (e325-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: EPPLER 325 AIRFOIL (e325-il)
Reynolds number: 500,000
Max Cl/Cd: 69.65 at α=9.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e325-il-500000.txt
Download as CSV file: xf-e325-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 325 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.000  -0.5910   0.08024   0.07807  -0.0155   1.0000   0.0141
  -9.750  -0.6107   0.07473   0.07250  -0.0175   1.0000   0.0141
  -9.500  -0.6296   0.07040   0.06809  -0.0172   1.0000   0.0141
  -9.250  -0.6483   0.06655   0.06415  -0.0155   1.0000   0.0141
  -9.000  -0.6634   0.06362   0.06114  -0.0125   1.0000   0.0141
  -8.750  -0.6769   0.06108   0.05853  -0.0087   1.0000   0.0144
  -8.500  -0.6843   0.05813   0.05546  -0.0057   1.0000   0.0145
  -8.250  -0.6887   0.05505   0.05225  -0.0026   1.0000   0.0148
  -8.000  -0.6911   0.05175   0.04879   0.0006   1.0000   0.0154
  -7.750  -0.7134   0.04483   0.04109   0.0090   1.0000   0.0168
  -7.500  -0.7036   0.04211   0.03841   0.0106   1.0000   0.0170
  -7.250  -0.6941   0.04035   0.03662   0.0125   1.0000   0.0174
  -7.000  -0.6845   0.03853   0.03471   0.0149   1.0000   0.0178
  -6.750  -0.6750   0.03650   0.03252   0.0177   1.0000   0.0185
  -6.500  -0.6755   0.03287   0.02829   0.0238   1.0000   0.0210
  -6.250  -0.6400   0.03032   0.02568   0.0209   0.9649   0.0216
  -6.000  -0.5950   0.02132   0.01548   0.0204   0.9177   0.0128
  -5.750  -0.5619   0.01919   0.01294   0.0199   0.8560   0.0121
  -5.500  -0.5369   0.01812   0.01153   0.0209   0.8141   0.0124
  -5.250  -0.5126   0.01751   0.01063   0.0220   0.7811   0.0131
  -5.000  -0.4845   0.01630   0.00924   0.0222   0.7543   0.0129
  -4.750  -0.4585   0.01536   0.00815   0.0228   0.7310   0.0129
  -4.500  -0.4350   0.01462   0.00729   0.0237   0.7095   0.0128
  -4.250  -0.4130   0.01399   0.00653   0.0251   0.6897   0.0132
  -4.000  -0.3933   0.01330   0.00574   0.0267   0.6721   0.0137
  -3.750  -0.3723   0.01284   0.00521   0.0281   0.6555   0.0140
  -3.500  -0.3509   0.01245   0.00474   0.0295   0.6397   0.0155
  -3.000  -0.3094   0.01153   0.00377   0.0326   0.6102   0.0348
  -2.750  -0.2936   0.01079   0.00342   0.0347   0.5972   0.1244
  -2.500  -0.2841   0.00977   0.00305   0.0379   0.5859   0.2936
  -2.250  -0.2966   0.00803   0.00250   0.0455   0.5767   0.5875
  -2.000  -0.2940   0.00748   0.00285   0.0516   0.5663   0.8069
  -1.750  -0.2718   0.00802   0.00336   0.0538   0.5547   0.8458
  -1.500  -0.2447   0.00855   0.00381   0.0548   0.5433   0.8622
  -1.250  -0.2102   0.00923   0.00442   0.0545   0.5311   0.8732
  -1.000  -0.1788   0.00978   0.00489   0.0547   0.5198   0.8829
  -0.750  -0.1350   0.01048   0.00549   0.0523   0.5082   0.8875
  -0.500  -0.1077   0.01092   0.00583   0.0530   0.4980   0.8967
  -0.250  -0.0604   0.01139   0.00619   0.0495   0.4864   0.8984
   0.000  -0.0287   0.01144   0.00615   0.0487   0.4771   0.8994
   0.250   0.0028   0.01147   0.00610   0.0480   0.4676   0.9004
   0.500   0.0329   0.01149   0.00605   0.0474   0.4587   0.9016
   0.750   0.0613   0.01152   0.00599   0.0472   0.4502   0.9031
   1.000   0.0862   0.01147   0.00591   0.0477   0.4427   0.9052
   1.250   0.0833   0.01113   0.00553   0.0537   0.4372   0.9120
   1.500   0.1127   0.01112   0.00546   0.0533   0.4296   0.9128
   1.750   0.1420   0.01111   0.00541   0.0529   0.4218   0.9136
   2.000   0.1713   0.01115   0.00539   0.0525   0.4151   0.9145
   2.250   0.2000   0.01115   0.00537   0.0522   0.4079   0.9154
   2.500   0.2273   0.01120   0.00534   0.0521   0.4012   0.9164
   2.750   0.2557   0.01119   0.00535   0.0519   0.3948   0.9176
   3.000   0.2818   0.01121   0.00533   0.0521   0.3885   0.9190
   3.250   0.3051   0.01119   0.00529   0.0528   0.3826   0.9207
   3.500   0.3256   0.01113   0.00523   0.0541   0.3769   0.9231
   3.750   0.3357   0.01100   0.00506   0.0575   0.3722   0.9267
   4.000   0.3610   0.01097   0.00504   0.0578   0.3666   0.9278
   4.250   0.3888   0.01097   0.00505   0.0576   0.3604   0.9286
   4.500   0.4166   0.01108   0.00511   0.0574   0.3545   0.9295
   4.750   0.4452   0.01109   0.00517   0.0571   0.3483   0.9305
   5.000   0.4721   0.01116   0.00522   0.0570   0.3421   0.9315
   5.250   0.4984   0.01122   0.00529   0.0571   0.3363   0.9326
   5.500   0.5243   0.01125   0.00535   0.0573   0.3303   0.9339
   5.750   0.5484   0.01136   0.00542   0.0578   0.3243   0.9354
   6.000   0.5729   0.01134   0.00548   0.0582   0.3186   0.9370
   6.250   0.5952   0.01138   0.00553   0.0591   0.3126   0.9391
   6.500   0.6144   0.01143   0.00559   0.0605   0.3070   0.9418
   6.750   0.6396   0.01145   0.00566   0.0608   0.3004   0.9431
   7.000   0.6663   0.01159   0.00578   0.0607   0.2937   0.9441
   7.250   0.6941   0.01164   0.00592   0.0604   0.2864   0.9451
   7.500   0.7202   0.01181   0.00607   0.0604   0.2787   0.9463
   7.750   0.7473   0.01188   0.00622   0.0603   0.2704   0.9476
   8.000   0.7725   0.01205   0.00640   0.0604   0.2628   0.9492
   8.250   0.7977   0.01216   0.00657   0.0606   0.2543   0.9509
   8.500   0.8205   0.01230   0.00674   0.0612   0.2458   0.9532
   8.750   0.8389   0.01244   0.00691   0.0627   0.2371   0.9565
   9.000   0.8645   0.01262   0.00713   0.0627   0.2268   0.9581
   9.250   0.8918   0.01288   0.00742   0.0622   0.2149   0.9594
   9.500   0.9183   0.01319   0.00774   0.0618   0.2013   0.9608
   9.750   0.9437   0.01355   0.00810   0.0616   0.1867   0.9625
  10.000   0.9672   0.01398   0.00851   0.0617   0.1701   0.9647
  10.250   0.9864   0.01448   0.00897   0.0625   0.1523   0.9679
  10.500   1.0033   0.01502   0.00947   0.0637   0.1358   0.9715
  10.750   1.0293   0.01573   0.01015   0.0628   0.1176   0.9730
  11.000   1.0537   0.01656   0.01094   0.0620   0.1007   0.9747
  11.250   1.0768   0.01748   0.01183   0.0612   0.0854   0.9769
  11.500   1.0972   0.01839   0.01273   0.0610   0.0726   0.9802
  11.750   1.1151   0.01938   0.01373   0.0610   0.0622   0.9841
  12.000   1.1357   0.02055   0.01491   0.0600   0.0528   0.9866
  12.250   1.1515   0.02189   0.01627   0.0593   0.0447   0.9907
  12.500   1.1656   0.02350   0.01793   0.0580   0.0380   0.9951
  12.750   1.1824   0.02544   0.01994   0.0554   0.0325   0.9990
  13.000   1.1575   0.02687   0.02141   0.0605   0.0317   1.0000
  13.250   1.1392   0.02923   0.02383   0.0632   0.0308   1.0000
  13.500   1.1392   0.03114   0.02579   0.0638   0.0276   1.0000
  13.750   1.1291   0.03435   0.02909   0.0638   0.0266   1.0000
  14.000   1.1169   0.03803   0.03284   0.0634   0.0254   1.0000
  14.250   1.1166   0.04055   0.03544   0.0631   0.0233   1.0000
  14.500   1.1052   0.04441   0.03939   0.0623   0.0226   1.0000
  14.750   1.0943   0.04827   0.04328   0.0613   0.0209   1.0000
  15.000   1.0812   0.05250   0.04762   0.0602   0.0209   1.0000
  15.250   1.0823   0.05509   0.05027   0.0595   0.0187   1.0000
  15.500   1.0730   0.05902   0.05427   0.0583   0.0178   1.0000
  15.750   1.0619   0.06337   0.05869   0.0569   0.0174   1.0000
  16.000   1.0475   0.06832   0.06373   0.0551   0.0171   1.0000
  16.250   1.0335   0.07331   0.06879   0.0532   0.0167   1.0000
  16.500   1.0228   0.07795   0.07353   0.0515   0.0168   1.0000
  16.750   1.0216   0.08138   0.07706   0.0502   0.0159   1.0000
  17.000   1.0186   0.08514   0.08088   0.0486   0.0146   1.0000
  17.250   1.0097   0.08978   0.08560   0.0467   0.0144   1.0000
  17.500   1.0014   0.09444   0.09032   0.0447   0.0139   1.0000
<< Back to EPPLER 325 AIRFOIL (e325-il)

Polar data table (+)

Polar graphs


<< Back to EPPLER 325 AIRFOIL (e325-il)