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KC-135 BL351.6 AIRFOIL (kc135d-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: KC-135 BL351.6 AIRFOIL (kc135d-il)
Reynolds number: 200,000
Max Cl/Cd: 74.69 at α=4.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-kc135d-il-200000.txt
Download as CSV file: xf-kc135d-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: KC-135 BL351.6 AIRFOIL                          
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.4386   0.08674   0.08326  -0.0295   1.0000   0.0268
  -8.250  -0.4401   0.08316   0.07973  -0.0315   1.0000   0.0270
  -8.000  -0.4453   0.07950   0.07613  -0.0337   1.0000   0.0276
  -7.750  -0.4531   0.07593   0.07260  -0.0357   1.0000   0.0276
  -7.500  -0.4563   0.07241   0.06908  -0.0369   1.0000   0.0279
  -7.250  -0.4585   0.06912   0.06576  -0.0373   1.0000   0.0284
  -7.000  -0.4602   0.06600   0.06262  -0.0370   1.0000   0.0290
  -6.750  -0.4624   0.06319   0.05975  -0.0358   1.0000   0.0300
  -6.500  -0.4644   0.06160   0.05802  -0.0334   1.0000   0.0316
  -6.250  -0.4679   0.06124   0.05745  -0.0298   1.0000   0.0324
  -6.000  -0.4707   0.06051   0.05653  -0.0262   1.0000   0.0327
  -5.750  -0.4579   0.05335   0.04917  -0.0286   0.9962   0.0336
  -5.500  -0.4350   0.04815   0.04400  -0.0317   0.9917   0.0350
  -5.250  -0.4069   0.04442   0.04011  -0.0346   0.9865   0.0370
  -5.000  -0.3752   0.04111   0.03655  -0.0373   0.9817   0.0399
  -4.750  -0.3342   0.04172   0.03646  -0.0381   0.9746   0.0455
  -4.500  -0.3091   0.03482   0.02944  -0.0410   0.9709   0.0480
  -4.250  -0.2823   0.03212   0.02667  -0.0424   0.9646   0.0517
  -4.000  -0.2477   0.03028   0.02425  -0.0436   0.9594   0.0614
  -3.750  -0.1746   0.01182   0.00581  -0.0508   0.9447   0.0889
  -3.500  -0.1464   0.00893   0.00302  -0.0529   0.9421   0.1073
  -2.750  -0.0939   0.01994   0.01333  -0.0512   0.9350   0.1798
  -2.500  -0.0371   0.01837   0.01076  -0.0499   0.9325   0.0576
  -2.250   0.0002   0.01652   0.00880  -0.0511   0.9293   0.0525
  -2.000   0.0303   0.01549   0.00770  -0.0511   0.9231   0.0517
  -1.750   0.0607   0.01473   0.00692  -0.0515   0.9171   0.0553
  -1.500   0.0913   0.01401   0.00618  -0.0518   0.9120   0.0575
  -1.250   0.1157   0.01354   0.00568  -0.0510   0.9038   0.0592
  -1.000   0.1437   0.01288   0.00498  -0.0509   0.8986   0.0648
  -0.750   0.1674   0.01265   0.00469  -0.0501   0.8899   0.0757
  -0.500   0.1826   0.01093   0.00435  -0.0481   0.8829   0.5065
  -0.250   0.2356   0.00996   0.00481  -0.0504   0.8810   0.9536
   0.000   0.3105   0.01004   0.00475  -0.0595   0.8787   0.9825
   0.250   0.3808   0.00983   0.00443  -0.0680   0.8728   0.9990
   0.500   0.3997   0.00970   0.00421  -0.0664   0.8588   1.0000
   0.750   0.4165   0.00961   0.00404  -0.0642   0.8452   1.0000
   1.000   0.4351   0.00959   0.00397  -0.0625   0.8334   1.0000
   1.250   0.4554   0.00960   0.00393  -0.0610   0.8231   1.0000
   1.500   0.4767   0.00960   0.00386  -0.0595   0.8128   1.0000
   1.750   0.4978   0.00962   0.00385  -0.0581   0.8017   1.0000
   2.000   0.5194   0.00968   0.00391  -0.0569   0.7911   1.0000
   2.250   0.5419   0.00972   0.00394  -0.0557   0.7809   1.0000
   2.500   0.5650   0.00975   0.00395  -0.0546   0.7705   1.0000
   2.750   0.5877   0.00980   0.00400  -0.0534   0.7590   1.0000
   3.000   0.6103   0.00985   0.00406  -0.0522   0.7465   1.0000
   3.250   0.6331   0.00989   0.00415  -0.0511   0.7333   1.0000
   3.500   0.6561   0.00994   0.00421  -0.0499   0.7191   1.0000
   3.750   0.6790   0.00997   0.00427  -0.0486   0.7032   1.0000
   4.000   0.7017   0.01002   0.00433  -0.0474   0.6852   1.0000
   4.250   0.7238   0.01006   0.00444  -0.0460   0.6626   1.0000
   4.500   0.7455   0.01012   0.00451  -0.0445   0.6345   1.0000
   4.750   0.7648   0.01024   0.00452  -0.0424   0.5852   1.0000
   5.000   0.7790   0.01069   0.00454  -0.0394   0.4861   1.0000
   5.250   0.7896   0.01163   0.00489  -0.0363   0.3810   1.0000
   5.500   0.8046   0.01244   0.00540  -0.0342   0.3189   1.0000
   5.750   0.8218   0.01316   0.00593  -0.0325   0.2746   1.0000
   6.000   0.8401   0.01381   0.00643  -0.0310   0.2334   1.0000
   6.250   0.8554   0.01475   0.00701  -0.0292   0.1547   1.0000
   6.500   0.8620   0.01676   0.00829  -0.0260   0.0592   1.0000
   6.750   0.8773   0.01791   0.00944  -0.0239   0.0477   1.0000
   7.000   0.8906   0.01923   0.01080  -0.0216   0.0410   1.0000
   7.250   0.9073   0.02030   0.01197  -0.0198   0.0377   1.0000
   7.500   0.9237   0.02157   0.01328  -0.0180   0.0350   1.0000
   7.750   0.9409   0.02322   0.01495  -0.0164   0.0330   1.0000
   8.000   0.9618   0.02571   0.01751  -0.0155   0.0308   1.0000
   8.250   0.9832   0.02701   0.01905  -0.0144   0.0292   1.0000
   8.500   1.0055   0.02912   0.02139  -0.0135   0.0286   1.0000
   8.750   1.0259   0.03162   0.02419  -0.0123   0.0282   1.0000
   9.000   1.0429   0.03456   0.02749  -0.0107   0.0283   1.0000
   9.250   1.0552   0.03799   0.03133  -0.0087   0.0290   1.0000
   9.500   1.0623   0.04204   0.03578  -0.0064   0.0299   1.0000
   9.750   1.0643   0.04673   0.04082  -0.0040   0.0307   1.0000
  10.000   1.0741   0.05108   0.04538  -0.0025   0.0318   1.0000
  16.500   0.6758   0.18304   0.17990  -0.0437   0.0418   1.0000
  16.750   0.6730   0.18634   0.18320  -0.0467   0.0416   1.0000
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