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AG455ct -02f rot. (ag455ct02r-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: AG455ct -02f rot. (ag455ct02r-il)
Reynolds number: 100,000
Max Cl/Cd: 44.68 at α=4.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-ag455ct02r-il-100000.txt
Download as CSV file: xf-ag455ct02r-il-100000.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 =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.750  -0.5542   0.09758   0.09294   0.0154   1.0000   0.0762
  -7.500  -0.5536   0.09412   0.08955   0.0105   1.0000   0.0795
  -7.250  -0.5474   0.08948   0.08490  -0.0074   1.0000   0.0814
  -7.000  -0.5397   0.08504   0.08055   0.0024   1.0000   0.0844
  -6.750  -0.5285   0.08160   0.07712   0.0016   1.0000   0.0901
  -6.500  -0.5123   0.07557   0.07086  -0.0168   1.0000   0.0956
  -6.250  -0.5041   0.07243   0.06794  -0.0076   1.0000   0.1001
  -6.000  -0.4842   0.06696   0.06222  -0.0194   1.0000   0.1098
  -5.750  -0.4722   0.06359   0.05900  -0.0154   1.0000   0.1137
  -5.500  -0.4131   0.04585   0.04162  -0.0177   1.0000   0.1277
  -5.250  -0.3990   0.04081   0.03645  -0.0217   1.0000   0.1392
  -5.000  -0.3829   0.03653   0.03205  -0.0239   1.0000   0.1525
  -4.750  -0.3660   0.03273   0.02817  -0.0249   1.0000   0.1665
  -4.500  -0.3486   0.02928   0.02465  -0.0252   1.0000   0.1812
  -4.250  -0.3303   0.02610   0.02144  -0.0252   1.0000   0.1970
  -4.000  -0.3102   0.02314   0.01840  -0.0255   1.0000   0.2162
  -3.750  -0.2572   0.02488   0.01656  -0.0326   1.0000   0.0557
  -3.500  -0.2314   0.02249   0.01408  -0.0325   1.0000   0.0597
  -3.250  -0.2035   0.02058   0.01186  -0.0319   1.0000   0.0628
  -3.000  -0.1753   0.01874   0.00971  -0.0311   1.0000   0.0655
  -2.750  -0.1485   0.01708   0.00803  -0.0305   1.0000   0.0742
  -2.500  -0.1220   0.01567   0.00668  -0.0298   1.0000   0.0900
  -2.250  -0.0961   0.01422   0.00546  -0.0291   1.0000   0.1322
  -2.000  -0.0745   0.01135   0.00455  -0.0284   1.0000   0.4937
  -1.750  -0.0433   0.00989   0.00420  -0.0265   1.0000   1.0000
  -1.500  -0.0176   0.00987   0.00387  -0.0261   1.0000   1.0000
  -1.250   0.0078   0.00987   0.00362  -0.0257   1.0000   1.0000
  -1.000   0.0329   0.00989   0.00347  -0.0254   1.0000   1.0000
  -0.750   0.0577   0.00995   0.00338  -0.0251   1.0000   1.0000
  -0.500   0.0819   0.01005   0.00338  -0.0249   1.0000   1.0000
  -0.250   0.1049   0.01023   0.00347  -0.0248   1.0000   1.0000
   0.000   0.1267   0.01055   0.00373  -0.0250   0.9998   1.0000
   0.250   0.1891   0.01070   0.00379  -0.0324   0.9735   1.0000
   0.500   0.2441   0.01076   0.00377  -0.0378   0.9439   1.0000
   0.750   0.2884   0.01079   0.00374  -0.0406   0.9090   1.0000
   1.000   0.3222   0.01088   0.00372  -0.0409   0.8726   1.0000
   1.250   0.3488   0.01104   0.00376  -0.0397   0.8367   1.0000
   1.500   0.3734   0.01127   0.00385  -0.0381   0.8035   1.0000
   1.750   0.3975   0.01156   0.00397  -0.0366   0.7725   1.0000
   2.000   0.4216   0.01189   0.00417  -0.0352   0.7417   1.0000
   2.250   0.4458   0.01226   0.00440  -0.0339   0.7122   1.0000
   2.500   0.4704   0.01262   0.00470  -0.0327   0.6820   1.0000
   2.750   0.4952   0.01289   0.00489  -0.0317   0.6523   1.0000
   3.000   0.5202   0.01312   0.00502  -0.0306   0.6226   1.0000
   3.250   0.5453   0.01334   0.00517  -0.0296   0.5911   1.0000
   3.500   0.5705   0.01360   0.00534  -0.0287   0.5585   1.0000
   3.750   0.5956   0.01390   0.00557  -0.0278   0.5259   1.0000
   4.000   0.6208   0.01424   0.00582  -0.0269   0.4918   1.0000
   4.250   0.6460   0.01463   0.00611  -0.0262   0.4572   1.0000
   4.500   0.6711   0.01509   0.00644  -0.0255   0.4229   1.0000
   4.750   0.6961   0.01558   0.00689  -0.0248   0.3878   1.0000
   5.000   0.7210   0.01615   0.00734  -0.0242   0.3544   1.0000
   5.250   0.7460   0.01676   0.00789  -0.0236   0.3202   1.0000
   5.500   0.7707   0.01745   0.00848  -0.0230   0.2887   1.0000
   5.750   0.7952   0.01823   0.00919  -0.0225   0.2591   1.0000
   6.250   0.8439   0.01997   0.01084  -0.0215   0.2044   1.0000
   6.500   0.8680   0.02091   0.01181  -0.0209   0.1791   1.0000
   6.750   0.8918   0.02210   0.01303  -0.0204   0.1571   1.0000
   7.000   0.9150   0.02326   0.01414  -0.0198   0.1364   1.0000
   7.250   0.9384   0.02481   0.01582  -0.0191   0.1187   1.0000
   7.500   0.9610   0.02622   0.01737  -0.0185   0.1018   1.0000
   7.750   0.9836   0.02822   0.01952  -0.0177   0.0896   1.0000
   8.000   1.0051   0.03014   0.02159  -0.0170   0.0782   1.0000
   8.250   1.0264   0.03295   0.02439  -0.0164   0.0706   1.0000
   8.500   1.0456   0.03516   0.02721  -0.0154   0.0634   1.0000
   8.750   1.0631   0.03881   0.03095  -0.0149   0.0583   1.0000
   9.000   1.0753   0.04270   0.03565  -0.0136   0.0563   1.0000
   9.250   1.0822   0.04730   0.04096  -0.0127   0.0542   1.0000
   9.500   1.0866   0.05146   0.04561  -0.0121   0.0516   1.0000
   9.750   1.0984   0.05408   0.04828  -0.0116   0.0485   1.0000
  10.000   1.0916   0.05956   0.05424  -0.0114   0.0485   1.0000
  10.250   1.0491   0.06826   0.06360  -0.0138   0.0513   1.0000
  10.500   1.0155   0.07654   0.07209  -0.0194   0.0543   1.0000
  10.750   0.9937   0.08479   0.08040  -0.0258   0.0561   1.0000
  11.000   0.7813   0.10398   0.09958  -0.0316   0.0799   1.0000
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