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E62 (5.62%) (e62-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: E62 (5.62%) (e62-il)
Reynolds number: 200,000
Max Cl/Cd: 107.6 at α=4.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e62-il-200000.txt
Download as CSV file: xf-e62-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: E62  (5.62%)                                    
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -7.750  -0.3695   0.11232   0.10906  -0.0225   1.0000   0.0218
  -7.500  -0.3773   0.11106   0.10786  -0.0206   1.0000   0.0218
  -7.250  -0.3841   0.10964   0.10649  -0.0187   1.0000   0.0218
  -7.000  -0.3873   0.10785   0.10475  -0.0182   1.0000   0.0219
  -6.750  -0.3886   0.10578   0.10272  -0.0182   1.0000   0.0219
  -6.500  -0.3939   0.10067   0.09768  -0.0160   1.0000   0.0224
  -6.250  -0.3906   0.09709   0.09411  -0.0132   0.9996   0.0231
  -6.000  -0.3686   0.09299   0.09000  -0.0171   0.9960   0.0240
  -5.750  -0.3452   0.08902   0.08601  -0.0225   0.9912   0.0253
  -5.500  -0.3174   0.08485   0.08183  -0.0295   0.9863   0.0267
  -5.250  -0.2892   0.08058   0.07754  -0.0371   0.9801   0.0291
  -5.000  -0.2269   0.07587   0.07273  -0.0571   0.9740   0.0319
  -4.750  -0.1899   0.06944   0.06623  -0.0681   0.9687   0.0329
  -4.500  -0.1843   0.06613   0.06295  -0.0649   0.9643   0.0346
  -4.250  -0.1509   0.06232   0.05909  -0.0706   0.9613   0.0371
  -4.000  -0.1139   0.05821   0.05491  -0.0782   0.9562   0.0401
  -3.750  -0.0249   0.05274   0.04900  -0.0994   0.9526   0.0449
  -3.500  -0.0005   0.04691   0.04324  -0.1032   0.9502   0.0470
  -3.250   0.0362   0.04392   0.04018  -0.1077   0.9483   0.0506
  -3.000   0.1089   0.03814   0.03388  -0.1215   0.9478   0.0599
  -2.750   0.1322   0.03592   0.03171  -0.1225   0.9416   0.0629
  -2.500   0.1861   0.03229   0.02765  -0.1300   0.9398   0.0745
  -1.750   0.3422   0.02096   0.01437  -0.1445   0.9386   0.0346
  -1.500   0.3833   0.01976   0.01298  -0.1473   0.9367   0.0394
  -1.250   0.4260   0.01824   0.01114  -0.1498   0.9355   0.0369
  -1.000   0.4574   0.01740   0.01023  -0.1504   0.9299   0.0371
  -0.750   0.4952   0.01666   0.00943  -0.1524   0.9263   0.0396
  -0.500   0.5365   0.01600   0.00870  -0.1549   0.9238   0.0498
  -0.250   0.5785   0.01411   0.00870  -0.1586   0.9229   0.6512
   0.000   0.5987   0.01352   0.00847  -0.1563   0.9150   1.0000
   0.250   0.6376   0.01334   0.00811  -0.1584   0.9112   1.0000
   0.500   0.6797   0.01305   0.00770  -0.1610   0.9089   1.0000
   0.750   0.7078   0.01302   0.00758  -0.1610   0.9005   1.0000
   1.000   0.7487   0.01267   0.00718  -0.1633   0.8975   1.0000
   1.250   0.7934   0.01224   0.00670  -0.1663   0.8953   1.0000
   1.500   0.8225   0.01213   0.00658  -0.1664   0.8863   1.0000
   1.750   0.8653   0.01171   0.00615  -0.1690   0.8826   1.0000
   2.000   0.8965   0.01156   0.00603  -0.1693   0.8734   1.0000
   2.250   0.9391   0.01114   0.00562  -0.1719   0.8679   1.0000
   2.500   0.9693   0.01102   0.00553  -0.1720   0.8567   1.0000
   2.750   1.0010   0.01090   0.00544  -0.1724   0.8452   1.0000
   3.000   1.0328   0.01079   0.00540  -0.1728   0.8324   1.0000
   3.250   1.0635   0.01072   0.00536  -0.1729   0.8178   1.0000
   3.500   1.0931   0.01070   0.00536  -0.1728   0.8018   1.0000
   3.750   1.1219   0.01072   0.00540  -0.1725   0.7840   1.0000
   4.000   1.1467   0.01082   0.00553  -0.1715   0.7622   1.0000
   4.250   1.1714   0.01090   0.00564  -0.1702   0.7326   1.0000
   4.500   1.1922   0.01108   0.00567  -0.1680   0.6852   1.0000
   4.750   1.2103   0.01144   0.00582  -0.1655   0.6233   1.0000
   5.000   1.2216   0.01228   0.00609  -0.1617   0.5188   1.0000
   5.250   1.2291   0.01356   0.00671  -0.1577   0.4053   1.0000
   5.500   1.2369   0.01506   0.00759  -0.1541   0.2920   1.0000
   5.750   1.2446   0.01689   0.00866  -0.1509   0.1677   1.0000
   6.000   1.2476   0.01963   0.01042  -0.1469   0.0383   1.0000
   6.250   1.2639   0.02095   0.01182  -0.1445   0.0297   1.0000
   6.500   1.2808   0.02206   0.01304  -0.1425   0.0255   1.0000
   6.750   1.2914   0.02388   0.01495  -0.1395   0.0226   1.0000
   7.000   1.3068   0.02525   0.01645  -0.1372   0.0219   1.0000
   7.250   1.3226   0.02687   0.01822  -0.1350   0.0212   1.0000
   7.500   1.3413   0.02879   0.02026  -0.1333   0.0207   1.0000
   7.750   1.3650   0.03110   0.02273  -0.1324   0.0205   1.0000
   8.000   1.3929   0.03397   0.02581  -0.1321   0.0207   1.0000
   8.250   1.4203   0.03750   0.02967  -0.1317   0.0213   1.0000
   8.500   1.4432   0.04237   0.03494  -0.1308   0.0222   1.0000
   8.750   1.4657   0.04532   0.03813  -0.1296   0.0235   1.0000
  16.500   1.2338   0.23949   0.23679  -0.1727   0.0324   1.0000
  16.750   1.2391   0.24445   0.24178  -0.1762   0.0323   1.0000
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