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=500,000 Ncrit=5


Details Polar file
Airfoil: E63 (4.25%) (e63-il)
Reynolds number: 500,000
Max Cl/Cd: 159.98 at α=1°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-e63-il-500000-n5.txt
Download as CSV file: xf-e63-il-500000-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 =     0.500 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.000  -0.2521   0.09042   0.08811  -0.0460   0.9843   0.0066
  -6.750  -0.2488   0.08807   0.08578  -0.0460   0.9793   0.0066
  -6.500  -0.2360   0.08508   0.08281  -0.0487   0.9761   0.0068
  -6.250  -0.2148   0.08135   0.07908  -0.0536   0.9742   0.0071
  -6.000  -0.2003   0.07811   0.07584  -0.0567   0.9701   0.0074
  -5.750  -0.1822   0.07475   0.07249  -0.0607   0.9661   0.0075
  -5.500  -0.1560   0.07095   0.06868  -0.0668   0.9637   0.0081
  -5.250  -0.1259   0.06697   0.06469  -0.0739   0.9621   0.0083
  -5.000  -0.1072   0.06383   0.06155  -0.0776   0.9561   0.0087
  -4.750  -0.0728   0.05986   0.05755  -0.0855   0.9533   0.0090
  -4.500  -0.0309   0.05559   0.05324  -0.0951   0.9516   0.0093
  -4.250   0.0117   0.05122   0.04880  -0.1040   0.9505   0.0094
  -3.750   0.0902   0.04300   0.04044  -0.1188   0.9438   0.0096
  -3.500   0.1391   0.03839   0.03573  -0.1278   0.9422   0.0096
  -3.250   0.1911   0.03373   0.03092  -0.1367   0.9412   0.0097
  -3.000   0.2438   0.02919   0.02618  -0.1449   0.9407   0.0097
  -2.500   0.3619   0.01835   0.01450  -0.1635   0.9415   0.0062
  -2.250   0.4181   0.01401   0.00945  -0.1695   0.9425   0.0048
  -2.000   0.4592   0.01224   0.00727  -0.1723   0.9417   0.0046
  -1.750   0.4973   0.01096   0.00570  -0.1744   0.9407   0.0049
  -1.500   0.5339   0.01011   0.00471  -0.1763   0.9392   0.0056
  -1.250   0.5698   0.00957   0.00411  -0.1780   0.9373   0.0077
  -1.000   0.5976   0.00921   0.00371  -0.1779   0.9309   0.0091
  -0.750   0.6318   0.00873   0.00315  -0.1791   0.9270   0.0094
  -0.500   0.6680   0.00833   0.00265  -0.1808   0.9235   0.0102
  -0.250   0.7010   0.00778   0.00246  -0.1821   0.9159   0.1582
   0.000   0.7417   0.00712   0.00238  -0.1853   0.9114   0.4129
   0.250   0.7795   0.00666   0.00233  -0.1876   0.9018   0.5986
   0.500   0.8184   0.00597   0.00220  -0.1899   0.8899   0.8571
   0.750   0.8624   0.00566   0.00190  -0.1933   0.8725   1.0000
   1.000   0.9071   0.00567   0.00174  -0.1970   0.8411   1.0000
   1.250   0.9402   0.00589   0.00173  -0.1979   0.7989   1.0000
   1.500   0.9658   0.00621   0.00181  -0.1973   0.7540   1.0000
   1.750   0.9882   0.00662   0.00199  -0.1960   0.7045   1.0000
   2.000   1.0091   0.00709   0.00218  -0.1944   0.6504   1.0000
   2.250   1.0297   0.00761   0.00241  -0.1928   0.5929   1.0000
   2.500   1.0500   0.00820   0.00267  -0.1912   0.5281   1.0000
   2.750   1.0663   0.00929   0.00310  -0.1890   0.4027   1.0000
   3.000   1.0850   0.01039   0.00356  -0.1875   0.2842   1.0000
   3.250   1.1062   0.01124   0.00403  -0.1865   0.2014   1.0000
   3.500   1.1286   0.01196   0.00444  -0.1856   0.1400   1.0000
   3.750   1.1513   0.01264   0.00486  -0.1847   0.0904   1.0000
   4.000   1.1741   0.01331   0.00533  -0.1839   0.0498   1.0000
   4.250   1.1963   0.01412   0.00592  -0.1829   0.0145   1.0000
   4.500   1.2206   0.01454   0.00636  -0.1822   0.0121   1.0000
   4.750   1.2451   0.01495   0.00688  -0.1815   0.0110   1.0000
   5.000   1.2692   0.01541   0.00742  -0.1807   0.0100   1.0000
   5.250   1.2928   0.01598   0.00809  -0.1798   0.0091   1.0000
   5.500   1.3147   0.01685   0.00913  -0.1786   0.0079   1.0000
   5.750   1.3359   0.01779   0.01020  -0.1772   0.0075   1.0000
   6.000   1.3575   0.01862   0.01115  -0.1759   0.0074   1.0000
   6.250   1.3781   0.01961   0.01226  -0.1745   0.0073   1.0000
   6.500   1.3980   0.02077   0.01356  -0.1729   0.0072   1.0000
   6.750   1.4174   0.02210   0.01505  -0.1712   0.0071   1.0000
   7.000   1.4365   0.02369   0.01689  -0.1695   0.0071   1.0000
   7.250   1.4564   0.02528   0.01866  -0.1680   0.0070   1.0000
   7.500   1.4771   0.02650   0.02004  -0.1667   0.0066   1.0000
   7.750   1.4974   0.02769   0.02140  -0.1654   0.0061   1.0000
   8.000   1.5174   0.02867   0.02252  -0.1641   0.0054   1.0000
   8.250   1.5374   0.02906   0.02300  -0.1631   0.0045   1.0000
   8.500   1.5564   0.02969   0.02372  -0.1618   0.0040   1.0000
   8.750   1.5734   0.03063   0.02476  -0.1604   0.0035   1.0000
   9.000   1.5839   0.03353   0.02798  -0.1576   0.0030   1.0000
   9.250   1.6028   0.03418   0.02881  -0.1562   0.0025   1.0000
   9.500   1.6132   0.03688   0.03190  -0.1533   0.0020   1.0000
   9.750   1.6225   0.03907   0.03438  -0.1505   0.0016   1.0000
  10.000   1.6295   0.04097   0.03650  -0.1475   0.0014   1.0000
  10.250   1.6342   0.04235   0.03805  -0.1441   0.0013   1.0000
  10.500   1.6329   0.04471   0.04065  -0.1399   0.0013   1.0000
  10.750   1.6228   0.04842   0.04471  -0.1349   0.0013   1.0000
  11.000   1.6101   0.05230   0.04893  -0.1302   0.0012   1.0000
  11.250   1.6035   0.05515   0.05198  -0.1269   0.0012   1.0000
  11.500   1.5814   0.06041   0.05756  -0.1229   0.0011   1.0000
  11.750   1.5592   0.06615   0.06362  -0.1202   0.0012   1.0000
  12.000   1.5363   0.07225   0.06998  -0.1191   0.0011   1.0000
  12.250   1.5093   0.07989   0.07789  -0.1196   0.0012   1.0000
  12.500   1.4893   0.08700   0.08519  -0.1220   0.0011   1.0000
  12.750   1.4645   0.09645   0.09485  -0.1266   0.0012   1.0000
  13.000   1.4342   0.10936   0.10798  -0.1348   0.0012   1.0000
<< Back to E63 (4.25%) (e63-il)

Polar data table (+)

Polar graphs


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