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 BL52.44 AIRFOIL (kc135a-il) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: KC-135 BL52.44 AIRFOIL (kc135a-il)
Reynolds number: 500,000
Max Cl/Cd: 73.18 at α=7°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-kc135a-il-500000-n5.txt
Download as CSV file: xf-kc135a-il-500000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: KC-135 BL52.44 AIRFOIL                          
  
 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
  ------ -------- --------- --------- -------- -------- --------
 -16.500  -0.8556   0.08467   0.08146  -0.0444   1.0000   0.0129
 -16.250  -0.8844   0.07590   0.07248  -0.0495   1.0000   0.0130
 -16.000  -0.9280   0.06581   0.06204  -0.0546   1.0000   0.0128
 -15.750  -0.9356   0.06170   0.05783  -0.0564   1.0000   0.0130
 -15.500  -0.9451   0.05757   0.05357  -0.0579   1.0000   0.0130
 -15.250  -0.9437   0.05502   0.05096  -0.0588   1.0000   0.0132
 -15.000  -0.9622   0.05004   0.04572  -0.0599   1.0000   0.0131
 -14.750  -0.9635   0.04735   0.04294  -0.0603   1.0000   0.0132
 -14.500  -0.9650   0.04467   0.04013  -0.0605   1.0000   0.0133
 -14.250  -0.9651   0.04223   0.03757  -0.0605   1.0000   0.0134
 -14.000  -0.9625   0.04017   0.03541  -0.0603   1.0000   0.0136
 -13.750  -0.9623   0.03784   0.03293  -0.0597   1.0000   0.0137
 -13.500  -0.9581   0.03602   0.03099  -0.0591   1.0000   0.0138
 -13.250  -0.9531   0.03433   0.02919  -0.0582   1.0000   0.0140
 -13.000  -0.9483   0.03259   0.02732  -0.0571   1.0000   0.0140
 -12.750  -0.9419   0.03108   0.02567  -0.0559   1.0000   0.0142
 -12.500  -0.9350   0.02961   0.02409  -0.0545   1.0000   0.0143
 -12.250  -0.9267   0.02834   0.02271  -0.0531   1.0000   0.0144
 -12.000  -0.9179   0.02715   0.02140  -0.0514   1.0000   0.0146
 -11.750  -0.9085   0.02606   0.02022  -0.0497   1.0000   0.0148
 -11.500  -0.8989   0.02507   0.01913  -0.0478   1.0000   0.0149
 -11.250  -0.8893   0.02415   0.01812  -0.0457   1.0000   0.0150
 -11.000  -0.8801   0.02328   0.01717  -0.0434   1.0000   0.0151
 -10.750  -0.8727   0.02244   0.01629  -0.0408   1.0000   0.0153
 -10.500  -0.8663   0.02171   0.01554  -0.0377   1.0000   0.0155
 -10.250  -0.8404   0.02087   0.01467  -0.0384   0.9843   0.0157
 -10.000  -0.8038   0.02000   0.01375  -0.0412   0.9656   0.0160
  -9.750  -0.7578   0.01908   0.01275  -0.0460   0.9434   0.0165
  -9.500  -0.7117   0.01819   0.01176  -0.0507   0.9114   0.0170
  -9.250  -0.6863   0.01760   0.01097  -0.0509   0.8694   0.0174
  -9.000  -0.6723   0.01720   0.01038  -0.0486   0.8349   0.0178
  -8.750  -0.6588   0.01689   0.00991  -0.0461   0.8092   0.0182
  -8.500  -0.6449   0.01646   0.00939  -0.0438   0.7902   0.0186
  -8.250  -0.6275   0.01612   0.00898  -0.0420   0.7751   0.0192
  -8.000  -0.6100   0.01575   0.00852  -0.0403   0.7623   0.0196
  -7.750  -0.5914   0.01541   0.00809  -0.0388   0.7514   0.0202
  -7.500  -0.5723   0.01505   0.00766  -0.0373   0.7411   0.0210
  -7.250  -0.5522   0.01473   0.00724  -0.0360   0.7328   0.0217
  -7.000  -0.5323   0.01436   0.00685  -0.0346   0.7247   0.0226
  -6.750  -0.5109   0.01408   0.00652  -0.0335   0.7176   0.0236
  -6.500  -0.4889   0.01380   0.00618  -0.0325   0.7104   0.0247
  -6.250  -0.4669   0.01353   0.00585  -0.0315   0.7035   0.0259
  -6.000  -0.4445   0.01324   0.00555  -0.0306   0.6970   0.0273
  -5.750  -0.4213   0.01301   0.00527  -0.0298   0.6902   0.0289
  -5.250  -0.3743   0.01254   0.00474  -0.0282   0.6789   0.0325
  -5.000  -0.3500   0.01235   0.00451  -0.0276   0.6733   0.0347
  -4.750  -0.3261   0.01214   0.00427  -0.0269   0.6681   0.0371
  -4.500  -0.3014   0.01194   0.00406  -0.0263   0.6631   0.0401
  -4.250  -0.2766   0.01174   0.00385  -0.0258   0.6578   0.0434
  -4.000  -0.2519   0.01156   0.00365  -0.0252   0.6524   0.0476
  -3.750  -0.2271   0.01138   0.00345  -0.0246   0.6473   0.0529
  -3.500  -0.2019   0.01118   0.00327  -0.0242   0.6420   0.0605
  -3.250  -0.1774   0.01095   0.00308  -0.0236   0.6369   0.0766
  -3.000  -0.1555   0.01052   0.00284  -0.0226   0.6324   0.1326
  -2.750  -0.1499   0.00880   0.00225  -0.0192   0.6280   0.4217
  -2.500  -0.1262   0.00856   0.00219  -0.0184   0.6230   0.4760
  -2.250  -0.1005   0.00848   0.00215  -0.0180   0.6181   0.5050
  -2.000  -0.0746   0.00843   0.00213  -0.0175   0.6134   0.5280
  -1.750  -0.0478   0.00839   0.00212  -0.0172   0.6082   0.5450
  -1.500  -0.0205   0.00839   0.00210  -0.0171   0.6033   0.5573
  -1.250   0.0062   0.00839   0.00208  -0.0168   0.5989   0.5685
  -1.000   0.0334   0.00837   0.00208  -0.0166   0.5942   0.5793
  -0.750   0.0609   0.00837   0.00207  -0.0164   0.5890   0.5872
  -0.500   0.0885   0.00840   0.00204  -0.0164   0.5838   0.5918
  -0.250   0.1160   0.00839   0.00202  -0.0162   0.5787   0.5958
   0.000   0.1436   0.00838   0.00201  -0.0162   0.5732   0.5999
   0.250   0.1711   0.00840   0.00201  -0.0161   0.5680   0.6041
   0.500   0.1988   0.00844   0.00200  -0.0160   0.5634   0.6081
   0.750   0.2264   0.00843   0.00202  -0.0160   0.5583   0.6122
   1.000   0.2534   0.00845   0.00202  -0.0158   0.5512   0.6168
   1.250   0.2805   0.00848   0.00203  -0.0156   0.5421   0.6216
   1.500   0.3073   0.00853   0.00204  -0.0154   0.5330   0.6261
   1.750   0.3345   0.00854   0.00207  -0.0153   0.5265   0.6306
   2.000   0.3614   0.00856   0.00211  -0.0151   0.5197   0.6359
   2.250   0.3883   0.00862   0.00216  -0.0149   0.5138   0.6415
   2.500   0.4155   0.00864   0.00221  -0.0148   0.5073   0.6470
   2.750   0.4420   0.00868   0.00227  -0.0145   0.5004   0.6528
   3.000   0.4689   0.00872   0.00233  -0.0144   0.4935   0.6593
   3.250   0.4952   0.00877   0.00240  -0.0141   0.4857   0.6661
   3.500   0.5214   0.00881   0.00248  -0.0138   0.4774   0.6739
   3.750   0.5474   0.00888   0.00256  -0.0134   0.4678   0.6819
   4.000   0.5731   0.00893   0.00265  -0.0131   0.4568   0.6910
   4.250   0.5982   0.00901   0.00275  -0.0126   0.4441   0.7011
   4.500   0.6226   0.00910   0.00286  -0.0120   0.4280   0.7133
   4.750   0.6460   0.00922   0.00299  -0.0111   0.4067   0.7266
   5.000   0.6676   0.00942   0.00315  -0.0101   0.3799   0.7416
   5.250   0.6878   0.00969   0.00336  -0.0087   0.3512   0.7594
   5.500   0.7075   0.00996   0.00361  -0.0073   0.3248   0.7819
   5.750   0.7270   0.01021   0.00388  -0.0058   0.3008   0.8090
   6.000   0.7461   0.01047   0.00417  -0.0043   0.2775   0.8416
   6.250   0.7679   0.01075   0.00448  -0.0033   0.2581   0.8783
   6.500   0.7987   0.01108   0.00485  -0.0042   0.2406   0.9139
   6.750   0.8386   0.01152   0.00527  -0.0073   0.2224   0.9399
   7.000   0.8774   0.01199   0.00569  -0.0102   0.2043   0.9578
   7.250   0.9115   0.01255   0.00614  -0.0123   0.1816   0.9704
   7.500   0.9433   0.01305   0.00655  -0.0138   0.1627   0.9803
   7.750   0.9734   0.01351   0.00696  -0.0149   0.1492   0.9894
   8.000   1.0012   0.01399   0.00739  -0.0156   0.1376   0.9985
   8.250   1.0217   0.01435   0.00774  -0.0147   0.1286   1.0000
   8.500   1.0341   0.01473   0.00810  -0.0121   0.1208   1.0000
   8.750   1.0454   0.01515   0.00849  -0.0093   0.1121   1.0000
   9.000   1.0546   0.01553   0.00886  -0.0062   0.1047   1.0000
   9.250   1.0609   0.01598   0.00927  -0.0025   0.0973   1.0000
   9.500   1.0697   0.01643   0.00972   0.0005   0.0904   1.0000
   9.750   1.0785   0.01697   0.01025   0.0033   0.0834   1.0000
  10.250   1.0969   0.01826   0.01151   0.0082   0.0691   1.0000
  10.500   1.1055   0.01904   0.01228   0.0104   0.0614   1.0000
  10.750   1.1138   0.01993   0.01315   0.0125   0.0544   1.0000
  11.000   1.1217   0.02091   0.01412   0.0143   0.0477   1.0000
  11.250   1.1293   0.02201   0.01520   0.0160   0.0419   1.0000
  11.500   1.1382   0.02309   0.01630   0.0174   0.0382   1.0000
  11.750   1.1471   0.02425   0.01749   0.0186   0.0353   1.0000
  12.000   1.1554   0.02552   0.01878   0.0198   0.0326   1.0000
  12.250   1.1637   0.02684   0.02013   0.0208   0.0305   1.0000
  12.500   1.1728   0.02815   0.02149   0.0216   0.0289   1.0000
  12.750   1.1801   0.02964   0.02301   0.0224   0.0274   1.0000
  13.000   1.1874   0.03117   0.02460   0.0231   0.0265   1.0000
  13.250   1.1956   0.03267   0.02617   0.0237   0.0255   1.0000
  13.500   1.2028   0.03426   0.02782   0.0242   0.0246   1.0000
  13.750   1.2089   0.03600   0.02962   0.0247   0.0239   1.0000
  14.000   1.2133   0.03793   0.03160   0.0251   0.0230   1.0000
  14.250   1.2165   0.03999   0.03372   0.0255   0.0225   1.0000
  14.500   1.2208   0.04202   0.03583   0.0257   0.0221   1.0000
  14.750   1.2264   0.04397   0.03786   0.0258   0.0213   1.0000
  15.000   1.2296   0.04620   0.04017   0.0259   0.0208   1.0000
  15.250   1.2322   0.04857   0.04261   0.0258   0.0202   1.0000
  15.500   1.2339   0.05110   0.04520   0.0257   0.0197   1.0000
  15.750   1.2343   0.05382   0.04800   0.0254   0.0193   1.0000
  16.000   1.2318   0.05696   0.05121   0.0249   0.0190   1.0000
  16.250   1.2315   0.05993   0.05426   0.0244   0.0186   1.0000
  16.500   1.2310   0.06298   0.05741   0.0238   0.0183   1.0000
  16.750   1.2308   0.06605   0.06058   0.0231   0.0179   1.0000
  17.000   1.2301   0.06923   0.06385   0.0223   0.0174   1.0000
  17.250   1.2268   0.07282   0.06753   0.0213   0.0171   1.0000
  17.500   1.2228   0.07658   0.07138   0.0202   0.0168   1.0000
  17.750   1.2175   0.08059   0.07548   0.0189   0.0165   1.0000
  18.000   1.2109   0.08484   0.07982   0.0175   0.0163   1.0000
  18.250   1.2034   0.08934   0.08441   0.0158   0.0161   1.0000
  18.500   1.1938   0.09422   0.08940   0.0140   0.0161   1.0000
<< Back to KC-135 BL52.44 AIRFOIL (kc135a-il)

Polar data table (+)

Polar graphs


<< Back to KC-135 BL52.44 AIRFOIL (kc135a-il)