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

EPPLER 664 AIRFOIL (e664-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: EPPLER 664 AIRFOIL (e664-il)
Reynolds number: 200,000
Max Cl/Cd: 58.27 at α=7.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-e664-il-200000-n5.txt
Download as CSV file: xf-e664-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 664 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -13.000  -0.3473   0.09692   0.09319  -0.0867   0.9398   0.0156
 -12.750  -0.3664   0.08384   0.08005  -0.0955   0.9320   0.0153
 -12.500  -0.4089   0.06537   0.06129  -0.1118   0.9209   0.0149
 -12.250  -0.4310   0.05679   0.05236  -0.1197   0.9071   0.0148
 -12.000  -0.4487   0.05105   0.04627  -0.1232   0.8935   0.0147
 -11.750  -0.4638   0.04676   0.04162  -0.1240   0.8815   0.0147
 -11.500  -0.4782   0.04330   0.03782  -0.1229   0.8708   0.0148
 -11.250  -0.4845   0.04071   0.03494  -0.1213   0.8627   0.0147
 -11.000  -0.4935   0.03810   0.03198  -0.1188   0.8548   0.0148
 -10.750  -0.4955   0.03595   0.02949  -0.1165   0.8487   0.0149
 -10.500  -0.4960   0.03388   0.02709  -0.1138   0.8426   0.0150
 -10.250  -0.4903   0.03219   0.02512  -0.1117   0.8373   0.0151
 -10.000  -0.4830   0.03060   0.02316  -0.1095   0.8329   0.0154
  -9.750  -0.4704   0.02923   0.02162  -0.1080   0.8292   0.0158
  -9.500  -0.4550   0.02806   0.02035  -0.1069   0.8254   0.0161
  -9.250  -0.4390   0.02709   0.01925  -0.1057   0.8219   0.0166
  -9.000  -0.4204   0.02611   0.01812  -0.1048   0.8189   0.0171
  -8.750  -0.3984   0.02508   0.01693  -0.1044   0.8164   0.0174
  -8.500  -0.3747   0.02406   0.01575  -0.1041   0.8141   0.0178
  -8.250  -0.3524   0.02316   0.01476  -0.1036   0.8111   0.0182
  -8.000  -0.3309   0.02235   0.01385  -0.1029   0.8081   0.0186
  -7.750  -0.3102   0.02164   0.01304  -0.1020   0.8052   0.0190
  -7.500  -0.2914   0.02091   0.01225  -0.1010   0.8027   0.0195
  -7.250  -0.2735   0.02029   0.01160  -0.0998   0.8005   0.0202
  -6.750  -0.2388   0.01933   0.01053  -0.0970   0.7959   0.0218
  -6.500  -0.2219   0.01895   0.01011  -0.0955   0.7933   0.0237
  -6.250  -0.2054   0.01853   0.00968  -0.0940   0.7907   0.0254
  -6.000  -0.1871   0.01815   0.00924  -0.0926   0.7881   0.0274
  -5.750  -0.1682   0.01774   0.00880  -0.0914   0.7858   0.0301
  -5.500  -0.1489   0.01735   0.00839  -0.0902   0.7840   0.0357
  -5.250  -0.1297   0.01695   0.00802  -0.0890   0.7823   0.0489
  -5.000  -0.1132   0.01657   0.00778  -0.0874   0.7800   0.0735
  -4.750  -0.0969   0.01621   0.00758  -0.0859   0.7774   0.1035
  -4.500  -0.0805   0.01583   0.00737  -0.0843   0.7748   0.1424
  -4.250  -0.0669   0.01528   0.00715  -0.0824   0.7724   0.2103
  -4.000  -0.0586   0.01446   0.00685  -0.0798   0.7702   0.3263
  -3.750  -0.0533   0.01340   0.00654  -0.0765   0.7680   0.4872
  -3.500  -0.0263   0.01461   0.00868  -0.0728   0.7667   0.6791
  -3.250  -0.0043   0.01480   0.00877  -0.0719   0.7647   0.7121
  -3.000   0.0166   0.01510   0.00901  -0.0707   0.7620   0.7323
  -2.750   0.0397   0.01545   0.00927  -0.0696   0.7594   0.7461
  -2.500   0.0634   0.01572   0.00945  -0.0689   0.7570   0.7575
  -2.250   0.0891   0.01615   0.00980  -0.0681   0.7549   0.7680
  -2.000   0.1180   0.01691   0.01051  -0.0670   0.7531   0.7778
  -1.750   0.1450   0.01742   0.01094  -0.0662   0.7513   0.7893
  -1.500   0.1762   0.01786   0.01132  -0.0660   0.7498   0.7956
  -1.250   0.1923   0.01790   0.01133  -0.0643   0.7467   0.8032
  -1.000   0.2152   0.01798   0.01139  -0.0634   0.7437   0.8050
  -0.750   0.2392   0.01801   0.01138  -0.0628   0.7410   0.8071
  -0.500   0.2639   0.01798   0.01131  -0.0625   0.7386   0.8094
  -0.250   0.2894   0.01789   0.01116  -0.0624   0.7365   0.8120
   0.000   0.3159   0.01776   0.01095  -0.0627   0.7346   0.8150
   0.250   0.3401   0.01764   0.01078  -0.0628   0.7322   0.8179
   0.500   0.3581   0.01771   0.01089  -0.0611   0.7282   0.8193
   0.750   0.3810   0.01770   0.01088  -0.0605   0.7249   0.8206
   1.000   0.4072   0.01762   0.01076  -0.0604   0.7218   0.8217
   1.250   0.4363   0.01749   0.01060  -0.0609   0.7192   0.8231
   1.500   0.4644   0.01740   0.01048  -0.0612   0.7164   0.8245
   1.750   0.4806   0.01744   0.01056  -0.0595   0.7112   0.8263
   2.000   0.5052   0.01736   0.01049  -0.0593   0.7073   0.8280
   2.250   0.5341   0.01722   0.01033  -0.0599   0.7042   0.8297
   2.500   0.5668   0.01705   0.01012  -0.0613   0.7016   0.8310
   2.750   0.5842   0.01709   0.01022  -0.0600   0.6957   0.8329
   3.000   0.6079   0.01701   0.01017  -0.0595   0.6910   0.8337
   3.250   0.6372   0.01685   0.01002  -0.0600   0.6875   0.8344
   3.500   0.6604   0.01681   0.01001  -0.0593   0.6828   0.8355
   3.750   0.6785   0.01682   0.01010  -0.0577   0.6765   0.8369
   4.000   0.7073   0.01665   0.00994  -0.0581   0.6719   0.8378
   4.250   0.7274   0.01662   0.00997  -0.0570   0.6654   0.8390
   4.500   0.7494   0.01654   0.00995  -0.0562   0.6586   0.8401
   4.750   0.7757   0.01641   0.00984  -0.0562   0.6527   0.8411
   5.000   0.7927   0.01641   0.00992  -0.0545   0.6438   0.8426
   5.250   0.8153   0.01633   0.00989  -0.0539   0.6356   0.8440
   5.500   0.8375   0.01625   0.00985  -0.0531   0.6261   0.8456
   5.750   0.8537   0.01625   0.00990  -0.0514   0.6148   0.8471
   6.000   0.8721   0.01620   0.00988  -0.0499   0.6027   0.8482
   6.250   0.8905   0.01615   0.00987  -0.0483   0.5892   0.8491
   6.500   0.9103   0.01614   0.00986  -0.0470   0.5738   0.8500
   6.750   0.9303   0.01619   0.00987  -0.0458   0.5556   0.8510
   7.000   0.9496   0.01631   0.00993  -0.0445   0.5350   0.8520
   7.250   0.9656   0.01657   0.01014  -0.0427   0.5132   0.8532
   7.500   0.9792   0.01695   0.01044  -0.0407   0.4896   0.8548
   7.750   0.9907   0.01745   0.01086  -0.0384   0.4651   0.8564
   8.000   1.0001   0.01807   0.01139  -0.0360   0.4402   0.8579
   8.250   1.0091   0.01878   0.01202  -0.0337   0.4161   0.8593
   8.500   1.0170   0.01960   0.01275  -0.0315   0.3928   0.8607
   8.750   1.0258   0.02045   0.01354  -0.0295   0.3700   0.8620
   9.000   1.0324   0.02140   0.01442  -0.0273   0.3469   0.8632
   9.250   1.0383   0.02237   0.01534  -0.0250   0.3249   0.8642
   9.500   1.0446   0.02340   0.01632  -0.0228   0.3030   0.8655
   9.750   1.0503   0.02451   0.01737  -0.0207   0.2824   0.8669
  10.000   1.0574   0.02561   0.01843  -0.0189   0.2617   0.8684
  10.250   1.0643   0.02679   0.01956  -0.0172   0.2413   0.8698
  10.500   1.0710   0.02804   0.02075  -0.0156   0.2222   0.8710
  10.750   1.0784   0.02931   0.02198  -0.0142   0.2031   0.8722
  11.000   1.0850   0.03069   0.02329  -0.0129   0.1821   0.8734
  11.250   1.0905   0.03220   0.02472  -0.0115   0.1622   0.8746
  11.500   1.0974   0.03367   0.02612  -0.0104   0.1430   0.8757
  11.750   1.1042   0.03522   0.02760  -0.0094   0.1249   0.8768
  12.000   1.1099   0.03687   0.02918  -0.0084   0.1077   0.8778
  12.250   1.1163   0.03842   0.03070  -0.0073   0.0941   0.8788
  12.500   1.1234   0.03997   0.03225  -0.0064   0.0827   0.8799
  12.750   1.1313   0.04150   0.03381  -0.0055   0.0733   0.8812
  13.000   1.1385   0.04316   0.03550  -0.0048   0.0649   0.8826
  13.250   1.1453   0.04491   0.03727  -0.0041   0.0569   0.8840
  13.500   1.1512   0.04682   0.03920  -0.0035   0.0490   0.8853
  14.000   1.1612   0.05101   0.04343  -0.0026   0.0345   0.8877
  14.250   1.1657   0.05325   0.04573  -0.0023   0.0292   0.8889
  14.500   1.1692   0.05568   0.04821  -0.0021   0.0247   0.8900
  14.750   1.1718   0.05829   0.05087  -0.0020   0.0221   0.8911
  15.250   1.1756   0.06379   0.05658  -0.0021   0.0188   0.8933
  15.500   1.1769   0.06669   0.05959  -0.0022   0.0177   0.8945
  15.750   1.1777   0.06973   0.06277  -0.0025   0.0170   0.8957
  16.000   1.1779   0.07295   0.06612  -0.0030   0.0163   0.8969
  16.250   1.1767   0.07643   0.06973  -0.0037   0.0157   0.8981
  16.500   1.1745   0.08016   0.07357  -0.0045   0.0154   0.8994
  16.750   1.1707   0.08419   0.07772  -0.0056   0.0149   0.9008
  17.000   1.1690   0.08805   0.08172  -0.0067   0.0146   0.9022
  17.250   1.1664   0.09210   0.08592  -0.0080   0.0142   0.9037
  17.500   1.1642   0.09619   0.09015  -0.0095   0.0138   0.9053
  18.000   1.1573   0.10483   0.09907  -0.0129   0.0132   0.9082
  18.250   1.1536   0.10925   0.10362  -0.0147   0.0129   0.9097
  18.500   1.1496   0.11379   0.10828  -0.0167   0.0126   0.9113
  18.750   1.1450   0.11848   0.11308  -0.0190   0.0124   0.9130
  19.000   1.1412   0.12307   0.11778  -0.0212   0.0122   0.9148
<< Back to EPPLER 664 AIRFOIL (e664-il)

Polar data table (+)

Polar graphs


<< Back to EPPLER 664 AIRFOIL (e664-il)