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

USA 45 M AIRFOIL (usa45m-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: USA 45 M AIRFOIL (usa45m-il)
Reynolds number: 50,000
Max Cl/Cd: 29.5 at α=6°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-usa45m-il-50000-n5.txt
Download as CSV file: xf-usa45m-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: USA 45 M AIRFOIL                                
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.3838   0.10206   0.09540  -0.0234   1.0000   0.1691
  -8.500  -0.3595   0.09776   0.09107  -0.0209   1.0000   0.1765
  -8.250  -0.3875   0.09644   0.08997  -0.0238   1.0000   0.1848
  -8.000  -0.3589   0.09223   0.08572  -0.0210   1.0000   0.1932
  -7.500  -0.3648   0.08685   0.08054  -0.0204   1.0000   0.2080
  -7.000  -0.4202   0.07046   0.06390  -0.0358   1.0000   0.0915
  -6.750  -0.4135   0.06734   0.06084  -0.0343   1.0000   0.0887
  -6.500  -0.4141   0.06388   0.05734  -0.0334   1.0000   0.0855
  -6.250  -0.4276   0.05919   0.05193  -0.0330   1.0000   0.0782
  -6.000  -0.4204   0.05631   0.04908  -0.0314   1.0000   0.0774
  -5.750  -0.4161   0.05365   0.04628  -0.0296   1.0000   0.0772
  -5.500  -0.4116   0.05118   0.04355  -0.0277   1.0000   0.0775
  -5.250  -0.4059   0.04889   0.04095  -0.0258   1.0000   0.0778
  -5.000  -0.3736   0.04553   0.03720  -0.0287   0.9887   0.0776
  -4.750  -0.3364   0.04228   0.03360  -0.0321   0.9751   0.0768
  -4.500  -0.2986   0.03941   0.03030  -0.0350   0.9610   0.0763
  -4.250  -0.2603   0.03691   0.02736  -0.0376   0.9462   0.0761
  -4.000  -0.2215   0.03479   0.02479  -0.0398   0.9308   0.0767
  -3.750  -0.1817   0.03300   0.02254  -0.0420   0.9148   0.0787
  -3.500  -0.1428   0.03101   0.02036  -0.0441   0.8984   0.0804
  -3.250  -0.1066   0.02950   0.01861  -0.0452   0.8786   0.0811
  -3.000  -0.0692   0.02811   0.01704  -0.0464   0.8583   0.0820
  -2.750  -0.0305   0.02686   0.01560  -0.0476   0.8378   0.0833
  -2.500   0.0046   0.02584   0.01443  -0.0481   0.8139   0.0851
  -2.250   0.0406   0.02497   0.01338  -0.0487   0.7905   0.0890
  -2.000   0.0727   0.02430   0.01247  -0.0486   0.7644   0.0933
  -1.750   0.1029   0.02355   0.01161  -0.0485   0.7400   0.0973
  -1.500   0.1294   0.02307   0.01092  -0.0476   0.7153   0.1011
  -1.250   0.1554   0.02269   0.01029  -0.0467   0.6925   0.1061
  -1.000   0.1800   0.02229   0.00979  -0.0457   0.6707   0.1140
  -0.750   0.2045   0.02190   0.00935  -0.0448   0.6514   0.1323
  -0.500   0.3147   0.01894   0.00905  -0.0583   0.6247   1.0000
  -0.250   0.3363   0.01915   0.00888  -0.0570   0.6082   1.0000
   0.250   0.3799   0.01962   0.00874  -0.0546   0.5798   1.0000
   0.500   0.4022   0.01988   0.00872  -0.0535   0.5674   1.0000
   0.750   0.4252   0.02016   0.00879  -0.0526   0.5557   1.0000
   1.000   0.4481   0.02047   0.00892  -0.0518   0.5444   1.0000
   1.250   0.4715   0.02078   0.00904  -0.0510   0.5344   1.0000
   1.500   0.4947   0.02111   0.00921  -0.0501   0.5248   1.0000
   1.750   0.5176   0.02148   0.00946  -0.0493   0.5155   1.0000
   2.000   0.5414   0.02183   0.00964  -0.0485   0.5075   1.0000
   2.250   0.5637   0.02225   0.01000  -0.0476   0.4985   1.0000
   2.500   0.5876   0.02262   0.01023  -0.0469   0.4915   1.0000
   2.750   0.6095   0.02310   0.01070  -0.0460   0.4831   1.0000
   3.000   0.6333   0.02351   0.01098  -0.0452   0.4764   1.0000
   3.250   0.6550   0.02403   0.01153  -0.0443   0.4685   1.0000
   3.500   0.6784   0.02448   0.01191  -0.0435   0.4619   1.0000
   3.750   0.7004   0.02504   0.01247  -0.0427   0.4550   1.0000
   4.000   0.7228   0.02557   0.01301  -0.0418   0.4482   1.0000
   4.250   0.7460   0.02609   0.01348  -0.0410   0.4420   1.0000
   4.500   0.7667   0.02675   0.01423  -0.0401   0.4348   1.0000
   4.750   0.7909   0.02724   0.01466  -0.0394   0.4291   1.0000
   5.000   0.8102   0.02800   0.01555  -0.0383   0.4218   1.0000
   5.250   0.8327   0.02858   0.01614  -0.0375   0.4155   1.0000
   5.500   0.8539   0.02927   0.01688  -0.0366   0.4092   1.0000
   5.750   0.8736   0.03001   0.01776  -0.0356   0.4020   1.0000
   6.000   0.8988   0.03047   0.01816  -0.0350   0.3968   1.0000
   6.250   0.9137   0.03152   0.01945  -0.0336   0.3888   1.0000
   6.500   0.9369   0.03205   0.02001  -0.0328   0.3828   1.0000
   6.750   0.9531   0.03306   0.02120  -0.0315   0.3757   1.0000
   7.000   0.9732   0.03376   0.02199  -0.0305   0.3689   1.0000
   7.250   0.9925   0.03455   0.02288  -0.0295   0.3626   1.0000
   7.500   1.0075   0.03562   0.02416  -0.0281   0.3553   1.0000
   7.750   1.0338   0.03591   0.02444  -0.0276   0.3500   1.0000
   8.000   1.0396   0.03757   0.02642  -0.0255   0.3418   1.0000
   8.250   1.0623   0.03808   0.02699  -0.0247   0.3362   1.0000
   8.500   1.0697   0.03968   0.02884  -0.0228   0.3292   1.0000
   8.750   1.0856   0.04063   0.02996  -0.0215   0.3229   1.0000
   9.000   1.1028   0.04156   0.03101  -0.0204   0.3175   1.0000
   9.250   1.1011   0.04375   0.03348  -0.0180   0.3105   1.0000
   9.500   1.1247   0.04419   0.03400  -0.0173   0.3057   1.0000
   9.750   1.1126   0.04708   0.03714  -0.0145   0.2993   1.0000
  10.000   1.1125   0.04912   0.03936  -0.0125   0.2937   1.0000
  10.250   1.1472   0.04856   0.03891  -0.0121   0.2891   1.0000
  10.500   1.0499   0.05823   0.04867  -0.0077   0.2822   1.0000
  10.750   1.0453   0.06133   0.05187  -0.0072   0.2774   1.0000
  11.000   0.9500   0.07728   0.06769  -0.0132   0.2643   1.0000
  11.250   0.9681   0.07796   0.06850  -0.0123   0.2614   1.0000
  11.750   0.9251   0.09203   0.08260  -0.0172   0.2473   1.0000
<< Back to USA 45 M AIRFOIL (usa45m-il)

Polar data table (+)

Polar graphs


<< Back to USA 45 M AIRFOIL (usa45m-il)