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AG455ct -02f rot. (ag455ct02r-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: AG455ct -02f rot. (ag455ct02r-il)
Reynolds number: 50,000
Max Cl/Cd: 33.95 at α=5.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-ag455ct02r-il-50000-n5.txt
Download as CSV file: xf-ag455ct02r-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: AG455ct -02f rot.                               
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.4520   0.09707   0.09077   0.0068   1.0000   0.1037
  -8.000  -0.4531   0.09048   0.08415   0.0031   1.0000   0.0639
  -7.750  -0.4576   0.08516   0.07887  -0.0004   1.0000   0.0531
  -7.500  -0.4541   0.08089   0.07463  -0.0008   1.0000   0.0507
  -7.250  -0.5395   0.08761   0.08111   0.0014   1.0000   0.0522
  -7.000  -0.5305   0.08295   0.07646  -0.0015   1.0000   0.0491
  -6.500  -0.5034   0.06997   0.06327  -0.0159   1.0000   0.0414
  -6.250  -0.4898   0.06542   0.05866  -0.0181   1.0000   0.0406
  -6.000  -0.4741   0.06059   0.05366  -0.0210   1.0000   0.0398
  -5.750  -0.4559   0.05560   0.04845  -0.0241   1.0000   0.0395
  -5.500  -0.4354   0.05065   0.04318  -0.0269   1.0000   0.0398
  -5.250  -0.4128   0.04589   0.03797  -0.0293   1.0000   0.0405
  -5.000  -0.3886   0.04163   0.03317  -0.0309   1.0000   0.0413
  -4.750  -0.3634   0.03784   0.02881  -0.0318   1.0000   0.0418
  -4.500  -0.3374   0.03441   0.02483  -0.0321   1.0000   0.0421
  -4.250  -0.3114   0.03121   0.02121  -0.0323   1.0000   0.0428
  -4.000  -0.2855   0.02860   0.01830  -0.0322   1.0000   0.0446
  -3.750  -0.2591   0.02674   0.01614  -0.0319   1.0000   0.0496
  -3.500  -0.2312   0.02488   0.01377  -0.0313   1.0000   0.0548
  -3.250  -0.2053   0.02301   0.01188  -0.0306   1.0000   0.0597
  -3.000  -0.1787   0.02163   0.01035  -0.0300   1.0000   0.0714
  -2.750  -0.1522   0.02030   0.00894  -0.0293   1.0000   0.0865
  -2.500  -0.1254   0.01904   0.00772  -0.0290   1.0000   0.1217
  -2.250  -0.0995   0.01745   0.00673  -0.0290   1.0000   0.2239
  -2.000  -0.0802   0.01541   0.00624  -0.0277   1.0000   0.5190
  -1.750  -0.0464   0.01398   0.00570  -0.0262   1.0000   1.0000
  -1.500  -0.0211   0.01395   0.00528  -0.0258   1.0000   1.0000
  -1.250   0.0040   0.01394   0.00493  -0.0255   1.0000   1.0000
  -1.000   0.0286   0.01396   0.00469  -0.0251   1.0000   1.0000
  -0.750   0.0530   0.01400   0.00454  -0.0247   1.0000   1.0000
  -0.500   0.0768   0.01407   0.00446  -0.0243   1.0000   1.0000
  -0.250   0.1000   0.01419   0.00446  -0.0240   1.0000   1.0000
   0.000   0.1237   0.01437   0.00453  -0.0240   0.9984   1.0000
   0.250   0.1734   0.01458   0.00460  -0.0288   0.9654   1.0000
   0.500   0.2195   0.01476   0.00467  -0.0326   0.9301   1.0000
   0.750   0.2621   0.01492   0.00472  -0.0353   0.8934   1.0000
   1.000   0.3006   0.01508   0.00478  -0.0369   0.8573   1.0000
   1.250   0.3329   0.01529   0.00487  -0.0371   0.8215   1.0000
   1.500   0.3615   0.01554   0.00499  -0.0365   0.7887   1.0000
   1.750   0.3878   0.01583   0.00513  -0.0354   0.7581   1.0000
   2.000   0.4130   0.01617   0.00534  -0.0342   0.7289   1.0000
   2.250   0.4379   0.01655   0.00560  -0.0330   0.7009   1.0000
   2.500   0.4628   0.01691   0.00596  -0.0321   0.6706   1.0000
   2.750   0.4876   0.01723   0.00625  -0.0311   0.6410   1.0000
   3.000   0.5124   0.01755   0.00653  -0.0300   0.6114   1.0000
   3.250   0.5371   0.01786   0.00680  -0.0290   0.5816   1.0000
   3.500   0.5619   0.01819   0.00708  -0.0280   0.5515   1.0000
   3.750   0.5867   0.01855   0.00748  -0.0271   0.5197   1.0000
   4.000   0.6114   0.01894   0.00784  -0.0262   0.4879   1.0000
   4.250   0.6361   0.01937   0.00824  -0.0253   0.4563   1.0000
   4.500   0.6607   0.01986   0.00868  -0.0245   0.4246   1.0000
   4.750   0.6858   0.02040   0.00929  -0.0239   0.3921   1.0000
   5.000   0.7106   0.02099   0.00988  -0.0233   0.3605   1.0000
   5.250   0.7351   0.02165   0.01053  -0.0226   0.3300   1.0000
   5.500   0.7594   0.02239   0.01126  -0.0221   0.3008   1.0000
   5.750   0.7834   0.02319   0.01206  -0.0215   0.2724   1.0000
   6.000   0.8071   0.02407   0.01304  -0.0210   0.2451   1.0000
   6.250   0.8307   0.02503   0.01408  -0.0204   0.2193   1.0000
   6.500   0.8538   0.02609   0.01523  -0.0199   0.1952   1.0000
   6.750   0.8763   0.02724   0.01641  -0.0193   0.1736   1.0000
   7.000   0.8986   0.02853   0.01784  -0.0187   0.1533   1.0000
   7.250   0.9198   0.02988   0.01924  -0.0182   0.1351   1.0000
   7.500   0.9414   0.03150   0.02110  -0.0175   0.1188   1.0000
   7.750   0.9618   0.03318   0.02298  -0.0169   0.1036   1.0000
   8.000   0.9815   0.03513   0.02511  -0.0161   0.0916   1.0000
   8.250   1.0003   0.03716   0.02725  -0.0154   0.0818   1.0000
   8.500   1.0178   0.03924   0.02961  -0.0148   0.0719   1.0000
   8.750   1.0342   0.04248   0.03331  -0.0140   0.0661   1.0000
   9.000   1.0488   0.04490   0.03602  -0.0134   0.0598   1.0000
   9.250   1.0587   0.04844   0.04002  -0.0128   0.0549   1.0000
   9.500   1.0646   0.05267   0.04477  -0.0124   0.0520   1.0000
   9.750   1.0669   0.05690   0.04939  -0.0122   0.0500   1.0000
  10.000   1.0689   0.06047   0.05317  -0.0121   0.0480   1.0000
  10.250   1.0665   0.06444   0.05725  -0.0122   0.0464   1.0000
  10.500   1.0499   0.06962   0.06276  -0.0133   0.0461   1.0000
  10.750   1.0315   0.07534   0.06867  -0.0162   0.0462   1.0000
  11.000   1.0135   0.08203   0.07550  -0.0209   0.0465   1.0000
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