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

E63 (4.25%) (e63-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: E63 (4.25%) (e63-il)
Reynolds number: 100,000
Max Cl/Cd: 77.65 at α=4.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-e63-il-100000-n5.txt
Download as CSV file: xf-e63-il-100000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: E63  (4.25%)                                    
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.500  -0.3661   0.11169   0.10686  -0.0194   1.0000   0.0254
  -7.250  -0.3715   0.11068   0.10593  -0.0184   1.0000   0.0255
  -7.000  -0.3763   0.10938   0.10470  -0.0175   1.0000   0.0256
  -6.750  -0.3758   0.10765   0.10303  -0.0182   1.0000   0.0257
  -6.500  -0.3724   0.10551   0.10095  -0.0195   1.0000   0.0258
  -6.250  -0.3668   0.10309   0.09858  -0.0212   1.0000   0.0258
  -6.000  -0.3657   0.09850   0.09406  -0.0189   1.0000   0.0262
  -5.750  -0.3635   0.09462   0.09021  -0.0156   1.0000   0.0269
  -5.500  -0.3598   0.09178   0.08741  -0.0149   1.0000   0.0276
  -5.250  -0.3546   0.08912   0.08479  -0.0153   1.0000   0.0283
  -5.000  -0.3479   0.08642   0.08212  -0.0163   1.0000   0.0291
  -4.750  -0.3349   0.08334   0.07906  -0.0190   0.9993   0.0302
  -4.500  -0.3038   0.07925   0.07495  -0.0265   0.9954   0.0320
  -4.250  -0.2656   0.07502   0.07067  -0.0364   0.9914   0.0344
  -4.000  -0.1902   0.06993   0.06539  -0.0583   0.9875   0.0360
  -3.750  -0.1644   0.06482   0.06030  -0.0624   0.9850   0.0368
  -3.500  -0.1448   0.06114   0.05662  -0.0637   0.9816   0.0381
  -3.250  -0.1073   0.05713   0.05247  -0.0703   0.9783   0.0402
  -3.000  -0.0535   0.05255   0.04775  -0.0812   0.9762   0.0439
  -2.500   0.0794   0.04109   0.03566  -0.1058   0.9755   0.0247
  -2.250   0.1469   0.03533   0.02944  -0.1169   0.9767   0.0194
  -2.000   0.1987   0.03220   0.02598  -0.1239   0.9764   0.0211
  -1.750   0.2547   0.02899   0.02221  -0.1308   0.9765   0.0230
  -1.500   0.3078   0.02607   0.01866  -0.1364   0.9767   0.0219
  -1.250   0.3556   0.02387   0.01584  -0.1404   0.9763   0.0214
  -1.000   0.3987   0.02229   0.01378  -0.1434   0.9751   0.0219
  -0.750   0.4393   0.02119   0.01242  -0.1458   0.9735   0.0231
  -0.500   0.4759   0.02041   0.01147  -0.1476   0.9694   0.0250
  -0.250   0.5145   0.01985   0.01084  -0.1500   0.9656   0.0376
   0.000   0.5601   0.01810   0.01060  -0.1544   0.9657   0.5315
   0.500   0.6175   0.01733   0.01004  -0.1541   0.9516   1.0000
   0.750   0.6505   0.01737   0.00989  -0.1552   0.9437   1.0000
   1.000   0.6884   0.01735   0.00972  -0.1572   0.9376   1.0000
   1.250   0.7198   0.01734   0.00964  -0.1579   0.9281   1.0000
   1.500   0.7543   0.01728   0.00956  -0.1591   0.9197   1.0000
   1.750   0.7911   0.01713   0.00937  -0.1607   0.9118   1.0000
   2.250   0.8558   0.01685   0.00914  -0.1619   0.8898   1.0000
   2.500   0.8892   0.01661   0.00896  -0.1626   0.8780   1.0000
   2.750   0.9226   0.01634   0.00883  -0.1632   0.8644   1.0000
   3.000   0.9567   0.01606   0.00864  -0.1639   0.8487   1.0000
   3.250   0.9950   0.01568   0.00836  -0.1652   0.8307   1.0000
   3.500   1.0322   0.01537   0.00817  -0.1664   0.8057   1.0000
   3.750   1.0764   0.01502   0.00803  -0.1688   0.7717   1.0000
   4.000   1.1243   0.01478   0.00779  -0.1719   0.7176   1.0000
   4.250   1.1647   0.01500   0.00782  -0.1736   0.6410   1.0000
   4.500   1.1881   0.01607   0.00814  -0.1720   0.5054   1.0000
   4.750   1.2000   0.01774   0.00887  -0.1688   0.3571   1.0000
   5.000   1.2159   0.01915   0.00984  -0.1668   0.2572   1.0000
   5.250   1.2318   0.02078   0.01086  -0.1651   0.1566   1.0000
   5.500   1.2481   0.02252   0.01209  -0.1633   0.0743   1.0000
   5.750   1.2656   0.02423   0.01362  -0.1614   0.0402   1.0000
   6.000   1.2836   0.02582   0.01530  -0.1594   0.0326   1.0000
   6.250   1.3021   0.02722   0.01697  -0.1575   0.0292   1.0000
   6.500   1.3194   0.02874   0.01861  -0.1557   0.0246   1.0000
   6.750   1.3358   0.03048   0.02054  -0.1536   0.0220   1.0000
   7.000   1.3535   0.03239   0.02269  -0.1517   0.0208   1.0000
   7.250   1.3729   0.03458   0.02511  -0.1500   0.0198   1.0000
   7.500   1.3944   0.03715   0.02810  -0.1486   0.0190   1.0000
   7.750   1.4168   0.04011   0.03142  -0.1474   0.0185   1.0000
   8.000   1.4376   0.04351   0.03525  -0.1459   0.0181   1.0000
   8.250   1.4543   0.04728   0.03953  -0.1438   0.0179   1.0000
   8.500   1.4655   0.05149   0.04432  -0.1411   0.0178   1.0000
   8.750   1.4714   0.05557   0.04892  -0.1379   0.0173   1.0000
   9.000   1.4742   0.05858   0.05223  -0.1350   0.0160   1.0000
   9.500   1.4556   0.06775   0.06215  -0.1272   0.0146   1.0000
   9.750   1.4410   0.07134   0.06612  -0.1225   0.0144   1.0000
  10.000   1.4229   0.07537   0.07050  -0.1182   0.0143   1.0000
  10.250   1.4032   0.07990   0.07535  -0.1149   0.0143   1.0000
  10.500   1.3831   0.08483   0.08055  -0.1129   0.0143   1.0000
  10.750   1.3641   0.09005   0.08600  -0.1122   0.0144   1.0000
  11.000   1.3436   0.09603   0.09220  -0.1130   0.0144   1.0000
  11.250   1.3234   0.10272   0.09908  -0.1154   0.0144   1.0000
  11.500   1.3036   0.11035   0.10689  -0.1196   0.0145   1.0000
  11.750   1.2885   0.11804   0.11467  -0.1242   0.0148   1.0000
<< Back to E63 (4.25%) (e63-il)

Polar data table (+)

Polar graphs


<< Back to E63 (4.25%) (e63-il)