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USA 50 AIRFOIL (usa50-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5


Details Polar file
Airfoil: USA 50 AIRFOIL (usa50-il)
Reynolds number: 1,000,000
Max Cl/Cd: 83.65 at α=2°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-usa50-il-1000000-n5.txt
Download as CSV file: xf-usa50-il-1000000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: USA 50 AIRFOIL                                  
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     1.000 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.4965   0.08623   0.08469  -0.0158   1.0000   0.0026
  -8.000  -0.5014   0.08268   0.08116  -0.0166   1.0000   0.0025
  -7.750  -0.5103   0.07968   0.07819  -0.0167   1.0000   0.0025
  -7.250  -0.5200   0.07108   0.06962  -0.0212   1.0000   0.0023
  -6.500  -0.4587   0.05055   0.04879  -0.0427   0.9939   0.0031
  -6.250  -0.4379   0.04416   0.04221  -0.0470   0.9907   0.0028
  -6.000  -0.4133   0.03757   0.03535  -0.0503   0.9882   0.0027
  -5.750  -0.3937   0.03122   0.02866  -0.0509   0.9841   0.0025
  -5.500  -0.3824   0.01714   0.01339  -0.0481   0.9792   0.0021
  -5.250  -0.3620   0.01410   0.00983  -0.0465   0.9752   0.0021
  -5.000  -0.3376   0.01230   0.00769  -0.0458   0.9715   0.0021
  -4.750  -0.3089   0.01118   0.00636  -0.0460   0.9689   0.0022
  -4.500  -0.2855   0.01048   0.00554  -0.0449   0.9628   0.0026
  -4.250  -0.2589   0.00974   0.00467  -0.0446   0.9574   0.0029
  -4.000  -0.2333   0.00912   0.00393  -0.0441   0.9511   0.0031
  -3.750  -0.2043   0.00859   0.00330  -0.0443   0.9456   0.0035
  -3.500  -0.1708   0.00823   0.00293  -0.0457   0.9411   0.0044
  -3.250  -0.1360   0.00783   0.00245  -0.0473   0.9345   0.0053
  -3.000  -0.0937   0.00741   0.00192  -0.0506   0.9282   0.0063
  -2.750  -0.0525   0.00711   0.00155  -0.0537   0.9166   0.0092
  -2.250   0.0149   0.00669   0.00110  -0.0565   0.8818   0.0487
  -2.000   0.0425   0.00665   0.00100  -0.0564   0.8549   0.0609
  -1.750   0.0662   0.00670   0.00090  -0.0554   0.8075   0.0725
  -1.500   0.0864   0.00685   0.00082  -0.0536   0.7453   0.0853
  -1.250   0.1084   0.00693   0.00076  -0.0524   0.7004   0.1038
  -1.000   0.1304   0.00687   0.00070  -0.0512   0.6585   0.1629
  -0.750   0.1512   0.00666   0.00065  -0.0499   0.6139   0.2780
  -0.500   0.1731   0.00655   0.00064  -0.0488   0.5756   0.3650
  -0.250   0.1955   0.00631   0.00063  -0.0479   0.5552   0.4723
   0.000   0.2176   0.00603   0.00065  -0.0468   0.5421   0.5881
   0.250   0.2393   0.00578   0.00067  -0.0455   0.5325   0.6885
   0.500   0.2563   0.00535   0.00070  -0.0430   0.5235   0.8363
   0.750   0.3171   0.00524   0.00084  -0.0506   0.5128   0.9462
   1.000   0.3627   0.00536   0.00091  -0.0547   0.5012   0.9665
   1.250   0.3956   0.00549   0.00099  -0.0559   0.4893   0.9797
   1.500   0.4266   0.00561   0.00106  -0.0567   0.4743   0.9873
   1.750   0.4640   0.00573   0.00113  -0.0590   0.4537   0.9921
   2.000   0.4944   0.00591   0.00120  -0.0597   0.4197   0.9949
   2.250   0.5236   0.00654   0.00134  -0.0605   0.3060   0.9977
   2.500   0.5562   0.00701   0.00151  -0.0619   0.2331   1.0000
   2.750   0.5714   0.00793   0.00185  -0.0597   0.0919   1.0000
   3.000   0.5928   0.00831   0.00208  -0.0584   0.0520   1.0000
   3.250   0.6138   0.00874   0.00235  -0.0571   0.0113   1.0000
   3.500   0.6372   0.00893   0.00258  -0.0561   0.0084   1.0000
   3.750   0.6599   0.00920   0.00290  -0.0551   0.0063   1.0000
   4.000   0.6828   0.00947   0.00321  -0.0540   0.0050   1.0000
   4.250   0.7057   0.00972   0.00347  -0.0530   0.0041   1.0000
   4.500   0.7268   0.01018   0.00399  -0.0516   0.0034   1.0000
   4.750   0.7488   0.01055   0.00442  -0.0504   0.0031   1.0000
   5.000   0.7695   0.01107   0.00504  -0.0488   0.0028   1.0000
   5.250   0.7890   0.01170   0.00576  -0.0471   0.0026   1.0000
   5.500   0.8083   0.01235   0.00650  -0.0453   0.0025   1.0000
   5.750   0.8291   0.01283   0.00702  -0.0440   0.0023   1.0000
   6.000   0.8505   0.01324   0.00746  -0.0429   0.0021   1.0000
   6.250   0.8652   0.01447   0.00883  -0.0403   0.0019   1.0000
   6.500   0.8848   0.01521   0.00967  -0.0386   0.0018   1.0000
   6.750   0.9024   0.01637   0.01101  -0.0366   0.0016   1.0000
   7.000   0.9187   0.01807   0.01291  -0.0342   0.0015   1.0000
   7.250   0.9348   0.02040   0.01551  -0.0316   0.0013   1.0000
   7.500   0.9497   0.02293   0.01833  -0.0289   0.0012   1.0000
   7.750   0.9640   0.02506   0.02071  -0.0265   0.0011   1.0000
   8.000   0.9722   0.02856   0.02459  -0.0229   0.0011   1.0000
   8.250   0.9739   0.03308   0.02955  -0.0184   0.0012   1.0000
   8.500   0.9738   0.03731   0.03414  -0.0142   0.0012   1.0000
   8.750   0.9690   0.04197   0.03912  -0.0099   0.0012   1.0000
   9.000   0.9634   0.04606   0.04348  -0.0062   0.0013   1.0000
   9.250   0.9550   0.04985   0.04748  -0.0027   0.0014   1.0000
   9.500   0.9368   0.05377   0.05158   0.0017   0.0014   1.0000
   9.750   0.9207   0.05665   0.05458   0.0050   0.0014   1.0000
  10.000   0.8989   0.06075   0.05882   0.0065   0.0014   1.0000
  10.250   0.8826   0.06481   0.06301   0.0056   0.0014   1.0000
  10.500   0.8624   0.07092   0.06922   0.0016   0.0014   1.0000
  10.750   0.8470   0.07808   0.07648  -0.0046   0.0014   1.0000
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