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=200,000 Ncrit=5


Details Polar file
Airfoil: KC-135 Winglet (supercritical Whitcomb) (kc135winglet-il)
Reynolds number: 200,000
Max Cl/Cd: 73.15 at α=2°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-kc135winglet-il-200000-n5.txt
Download as CSV file: xf-kc135winglet-il-200000-n5.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 =     0.200 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.4359   0.09085   0.08732  -0.0375   1.0000   0.0225
  -8.250  -0.4421   0.08666   0.08321  -0.0361   1.0000   0.0228
  -8.000  -0.4440   0.08363   0.08024  -0.0349   1.0000   0.0231
  -7.750  -0.4465   0.08074   0.07740  -0.0346   1.0000   0.0234
  -7.500  -0.4493   0.07767   0.07438  -0.0354   1.0000   0.0236
  -7.250  -0.4495   0.07395   0.07069  -0.0381   0.9997   0.0238
  -7.000  -0.4261   0.06842   0.06509  -0.0462   0.9960   0.0244
  -6.750  -0.4008   0.06288   0.05944  -0.0542   0.9918   0.0251
  -6.500  -0.3721   0.05735   0.05377  -0.0621   0.9880   0.0259
  -6.250  -0.3410   0.05209   0.04830  -0.0692   0.9845   0.0270
  -6.000  -0.3073   0.04726   0.04318  -0.0752   0.9804   0.0286
  -5.750  -0.2621   0.04450   0.03976  -0.0811   0.9775   0.0306
  -5.500  -0.2262   0.04101   0.03588  -0.0853   0.9757   0.0307
  -5.250  -0.1967   0.03431   0.02905  -0.0899   0.9734   0.0320
  -5.000  -0.1610   0.03038   0.02479  -0.0919   0.9704   0.0215
  -4.750  -0.1215   0.02601   0.01991  -0.0950   0.9685   0.0160
  -4.500  -0.0841   0.02339   0.01696  -0.0977   0.9666   0.0150
  -4.250  -0.0456   0.02099   0.01416  -0.1000   0.9650   0.0142
  -4.000  -0.0078   0.01902   0.01182  -0.1019   0.9636   0.0138
  -3.750   0.0284   0.01786   0.01048  -0.1038   0.9623   0.0149
  -3.500   0.0626   0.01681   0.00924  -0.1051   0.9600   0.0160
  -3.250   0.0947   0.01574   0.00808  -0.1059   0.9564   0.0158
  -3.000   0.1291   0.01479   0.00708  -0.1074   0.9538   0.0156
  -2.750   0.1649   0.01398   0.00625  -0.1092   0.9515   0.0156
  -2.500   0.2020   0.01330   0.00555  -0.1114   0.9494   0.0156
  -2.250   0.2396   0.01279   0.00496  -0.1137   0.9476   0.0158
  -2.000   0.2718   0.01247   0.00457  -0.1147   0.9428   0.0162
  -1.750   0.3053   0.01222   0.00424  -0.1160   0.9387   0.0168
  -1.500   0.3402   0.01198   0.00397  -0.1175   0.9355   0.0190
  -1.250   0.3864   0.01011   0.00407  -0.1232   0.9349   0.6157
  -1.000   0.4161   0.01009   0.00419  -0.1232   0.9295   0.6870
  -0.750   0.4446   0.01013   0.00430  -0.1229   0.9231   0.7287
  -0.500   0.4737   0.01017   0.00441  -0.1226   0.9188   0.7618
  -0.250   0.4975   0.01031   0.00461  -0.1212   0.9115   0.7880
   0.000   0.5287   0.01026   0.00454  -0.1217   0.9056   0.7938
   0.250   0.5576   0.01017   0.00442  -0.1216   0.8927   0.7986
   0.500   0.5852   0.01001   0.00422  -0.1208   0.8729   0.8028
   0.750   0.6124   0.00990   0.00403  -0.1200   0.8470   0.8074
   1.000   0.6392   0.00988   0.00393  -0.1193   0.8192   0.8124
   1.250   0.6656   0.00989   0.00390  -0.1186   0.7953   0.8167
   1.500   0.6929   0.00994   0.00391  -0.1182   0.7732   0.8216
   1.750   0.7201   0.01003   0.00395  -0.1178   0.7444   0.8266
   2.000   0.7447   0.01018   0.00395  -0.1167   0.6909   0.8311
   2.250   0.7653   0.01078   0.00396  -0.1148   0.5678   0.8368
   2.500   0.7829   0.01185   0.00433  -0.1129   0.4278   0.8423
   2.750   0.8029   0.01277   0.00475  -0.1117   0.3241   0.8480
   3.250   0.8479   0.01415   0.00553  -0.1102   0.1946   0.8598
   3.500   0.8716   0.01473   0.00595  -0.1097   0.1554   0.8667
   3.750   0.8950   0.01519   0.00633  -0.1089   0.1315   0.8734
   4.000   0.9193   0.01563   0.00673  -0.1084   0.1158   0.8815
   4.250   0.9425   0.01604   0.00713  -0.1075   0.1045   0.8902
   4.500   0.9653   0.01649   0.00758  -0.1067   0.0945   0.9007
   4.750   0.9879   0.01683   0.00799  -0.1057   0.0863   0.9135
   5.000   1.0083   0.01722   0.00840  -0.1043   0.0799   0.9321
   5.250   1.0300   0.01749   0.00880  -0.1031   0.0744   1.0000
   5.500   1.0558   0.01810   0.00938  -0.1031   0.0682   1.0000
   5.750   1.0821   0.01858   0.00996  -0.1031   0.0635   1.0000
   6.000   1.1076   0.01913   0.01058  -0.1029   0.0587   1.0000
   6.250   1.1322   0.01980   0.01127  -0.1026   0.0544   1.0000
   6.500   1.1571   0.02037   0.01197  -0.1023   0.0497   1.0000
   6.750   1.1804   0.02114   0.01273  -0.1018   0.0445   1.0000
   7.000   1.2043   0.02183   0.01357  -0.1012   0.0401   1.0000
   7.250   1.2274   0.02257   0.01438  -0.1007   0.0360   1.0000
   7.500   1.2483   0.02364   0.01549  -0.0997   0.0334   1.0000
   7.750   1.2700   0.02462   0.01664  -0.0988   0.0316   1.0000
   8.000   1.2910   0.02568   0.01784  -0.0978   0.0301   1.0000
   8.250   1.3116   0.02676   0.01909  -0.0968   0.0288   1.0000
   8.500   1.3317   0.02788   0.02034  -0.0958   0.0277   1.0000
   8.750   1.3507   0.02909   0.02168  -0.0946   0.0268   1.0000
   9.000   1.3680   0.03060   0.02333  -0.0932   0.0260   1.0000
   9.250   1.3846   0.03236   0.02527  -0.0918   0.0253   1.0000
   9.500   1.4020   0.03375   0.02697  -0.0904   0.0243   1.0000
   9.750   1.4184   0.03492   0.02837  -0.0890   0.0227   1.0000
  10.000   1.4333   0.03605   0.02968  -0.0874   0.0214   1.0000
  10.250   1.4463   0.03721   0.03100  -0.0856   0.0204   1.0000
  10.500   1.4555   0.03869   0.03261  -0.0834   0.0194   1.0000
  10.750   1.4584   0.04100   0.03513  -0.0805   0.0185   1.0000
  11.000   1.4611   0.04284   0.03730  -0.0773   0.0179   1.0000
  11.250   1.4619   0.04489   0.03967  -0.0743   0.0170   1.0000
  11.500   1.4601   0.04719   0.04226  -0.0714   0.0162   1.0000
  11.750   1.4577   0.04949   0.04479  -0.0688   0.0154   1.0000
  12.000   1.4537   0.05202   0.04756  -0.0666   0.0147   1.0000
  12.250   1.4479   0.05483   0.05057  -0.0647   0.0142   1.0000
  12.500   1.4393   0.05818   0.05411  -0.0634   0.0138   1.0000
  12.750   1.4268   0.06229   0.05844  -0.0626   0.0135   1.0000
  13.000   1.4111   0.06720   0.06355  -0.0628   0.0132   1.0000
  13.250   1.3918   0.07317   0.06974  -0.0641   0.0131   1.0000
  13.500   1.3697   0.08040   0.07721  -0.0670   0.0130   1.0000
  13.750   1.3442   0.08952   0.08661  -0.0721   0.0132   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)