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

KC-135 Winglet (supercritical Whitcomb) (kc135winglet-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: KC-135 Winglet (supercritical Whitcomb) (kc135winglet-il)
Reynolds number: 50,000
Max Cl/Cd: 40.81 at α=4.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-kc135winglet-il-50000.txt
Download as CSV file: xf-kc135winglet-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: KC-135 Winglet (supercritical Whitcomb)         
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.4256   0.09875   0.09196  -0.0185   1.0000   0.2494
  -7.750  -0.4249   0.09622   0.08952  -0.0173   1.0000   0.2648
  -7.500  -0.4273   0.09402   0.08742  -0.0156   1.0000   0.2812
  -7.250  -0.4348   0.09225   0.08578  -0.0136   1.0000   0.2976
  -7.000  -0.4465   0.09078   0.08447  -0.0123   1.0000   0.3139
  -6.750  -0.4358   0.08774   0.08148  -0.0086   1.0000   0.3426
  -6.500  -0.4303   0.08522   0.07903  -0.0052   1.0000   0.3718
  -6.250  -0.4080   0.08156   0.07534  -0.0014   1.0000   0.4104
  -6.000  -0.4109   0.07996   0.07386   0.0028   1.0000   0.4444
  -5.750  -0.3968   0.07724   0.07118   0.0068   1.0000   0.4884
  -5.500  -0.3662   0.07353   0.06744   0.0094   1.0000   0.5447
  -5.250  -0.3633   0.07196   0.06597   0.0143   1.0000   0.5916
  -5.000  -0.3170   0.06670   0.06064   0.0135   1.0000   0.6459
  -4.250  -0.2466   0.04004   0.03179  -0.0699   1.0000   0.1654
  -4.000  -0.2026   0.03618   0.02718  -0.0734   1.0000   0.1343
  -3.750  -0.1642   0.03335   0.02372  -0.0754   1.0000   0.1184
  -3.500  -0.1278   0.03137   0.02103  -0.0763   1.0000   0.1081
  -3.250  -0.0968   0.02928   0.01868  -0.0767   1.0000   0.1038
  -3.000  -0.0665   0.02771   0.01678  -0.0765   1.0000   0.1006
  -2.750  -0.0397   0.02646   0.01536  -0.0756   1.0000   0.1006
  -2.500  -0.0143   0.02544   0.01425  -0.0744   1.0000   0.1050
  -2.250   0.0122   0.02464   0.01330  -0.0734   1.0000   0.1103
  -2.000   0.0438   0.02359   0.01224  -0.0740   1.0000   0.1193
  -1.750   0.0826   0.02255   0.01123  -0.0764   1.0000   0.1423
  -1.500   0.0992   0.02028   0.01190  -0.0714   1.0000   0.8004
  -1.250   0.0774   0.01975   0.01156  -0.0584   1.0000   0.9114
  -1.000   0.0756   0.01922   0.01079  -0.0524   1.0000   1.0000
  -0.750   0.1085   0.01953   0.01053  -0.0543   1.0000   1.0000
  -0.500   0.1408   0.01987   0.01053  -0.0562   1.0000   1.0000
  -0.250   0.1723   0.02025   0.01065  -0.0581   1.0000   1.0000
   0.000   0.2032   0.02066   0.01086  -0.0598   1.0000   1.0000
   0.250   0.2333   0.02112   0.01114  -0.0615   1.0000   1.0000
   0.500   0.2628   0.02160   0.01150  -0.0630   1.0000   1.0000
   0.750   0.2916   0.02213   0.01194  -0.0644   1.0000   1.0000
   1.000   0.3198   0.02270   0.01244  -0.0658   1.0000   1.0000
   1.250   0.3473   0.02331   0.01301  -0.0670   1.0000   1.0000
   1.500   0.3742   0.02397   0.01366  -0.0682   1.0000   1.0000
   1.750   0.4003   0.02468   0.01438  -0.0692   1.0000   1.0000
   2.000   0.4257   0.02544   0.01519  -0.0703   1.0000   1.0000
   2.250   0.4505   0.02627   0.01608  -0.0712   1.0000   1.0000
   2.500   0.4744   0.02717   0.01706  -0.0722   1.0000   1.0000
   2.750   0.4976   0.02815   0.01815  -0.0731   1.0000   1.0000
   3.000   0.5199   0.02923   0.01940  -0.0740   1.0000   1.0000
   3.250   0.5412   0.03043   0.02074  -0.0750   1.0000   1.0000
   3.500   0.6090   0.03156   0.02219  -0.0840   0.9737   1.0000
   3.750   0.7200   0.03050   0.02173  -0.0966   0.9150   1.0000
   4.000   0.7965   0.02799   0.01978  -0.1003   0.8613   1.0000
   4.250   0.8627   0.02372   0.01618  -0.0983   0.7889   1.0000
   4.500   0.8918   0.02185   0.01245  -0.0875   0.4345   1.0000
   4.750   0.9086   0.02483   0.01414  -0.0853   0.3212   1.0000
   5.000   0.9414   0.02697   0.01593  -0.0857   0.2738   1.0000
   5.250   0.9738   0.02887   0.01770  -0.0862   0.2430   1.0000
   5.500   1.0075   0.03103   0.01970  -0.0869   0.2231   1.0000
   5.750   1.0392   0.03334   0.02213  -0.0872   0.2089   1.0000
   6.000   1.0661   0.03558   0.02459  -0.0870   0.1936   1.0000
   6.250   1.0928   0.03812   0.02747  -0.0865   0.1820   1.0000
   6.500   1.1157   0.04072   0.03042  -0.0857   0.1690   1.0000
   6.750   1.1360   0.04391   0.03411  -0.0845   0.1592   1.0000
   7.000   1.1561   0.04709   0.03751  -0.0836   0.1494   1.0000
   7.250   1.1704   0.05111   0.04228  -0.0818   0.1453   1.0000
   7.500   1.1823   0.05563   0.04738  -0.0800   0.1438   1.0000
   7.750   1.1892   0.06055   0.05289  -0.0781   0.1435   1.0000
   8.000   1.1916   0.06562   0.05846  -0.0762   0.1432   1.0000
   8.250   1.1878   0.07107   0.06439  -0.0744   0.1441   1.0000
   8.500   1.1786   0.07690   0.07060  -0.0728   0.1461   1.0000
   8.750   1.1696   0.08282   0.07677  -0.0717   0.1489   1.0000
   9.000   1.1292   0.09099   0.08529  -0.0720   0.1612   1.0000
   9.250   1.0945   0.09995   0.09433  -0.0742   0.1780   1.0000
<< Back to KC-135 Winglet (supercritical Whitcomb) (kc135winglet-il)

Polar data table (+)

Polar graphs


<< Back to KC-135 Winglet (supercritical Whitcomb) (kc135winglet-il)