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AG455ct -02f rot. (ag455ct02r-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: AG455ct -02f rot. (ag455ct02r-il)
Reynolds number: 500,000
Max Cl/Cd: 78.74 at α=5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-ag455ct02r-il-500000.txt
Download as CSV file: xf-ag455ct02r-il-500000.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 =     0.500 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.5692   0.09723   0.09508   0.0172   1.0000   0.0139
  -8.000  -0.5650   0.09311   0.09098   0.0145   1.0000   0.0139
  -7.750  -0.5607   0.08886   0.08676   0.0113   1.0000   0.0139
  -7.500  -0.5532   0.08406   0.08197   0.0064   1.0000   0.0140
  -7.250  -0.5519   0.07567   0.07361  -0.0017   1.0000   0.0143
  -7.000  -0.5421   0.07040   0.06832  -0.0066   1.0000   0.0147
  -6.750  -0.5284   0.06594   0.06382  -0.0107   1.0000   0.0150
  -6.500  -0.5123   0.06152   0.05935  -0.0147   1.0000   0.0153
  -6.250  -0.4941   0.05687   0.05463  -0.0188   1.0000   0.0158
  -4.250  -0.3004   0.01806   0.01304  -0.0313   1.0000   0.0142
  -4.000  -0.2738   0.01609   0.01078  -0.0309   1.0000   0.0141
  -3.750  -0.2475   0.01379   0.00817  -0.0304   1.0000   0.0143
  -3.500  -0.2217   0.01215   0.00641  -0.0301   1.0000   0.0159
  -3.250  -0.1951   0.01108   0.00523  -0.0296   1.0000   0.0161
  -3.000  -0.1686   0.01019   0.00426  -0.0290   1.0000   0.0165
  -2.750  -0.1420   0.00951   0.00354  -0.0286   1.0000   0.0177
  -2.500  -0.1153   0.00900   0.00299  -0.0281   1.0000   0.0203
  -2.250  -0.0883   0.00837   0.00231  -0.0278   1.0000   0.0287
  -2.000  -0.0495   0.00767   0.00187  -0.0302   0.9884   0.0899
  -1.750  -0.0087   0.00706   0.00168  -0.0333   0.9606   0.2065
  -1.500   0.0264   0.00647   0.00152  -0.0348   0.9001   0.3575
  -1.250   0.0486   0.00597   0.00139  -0.0335   0.8240   0.5515
  -1.000   0.0681   0.00535   0.00139  -0.0313   0.7783   0.7887
  -0.750   0.0915   0.00509   0.00136  -0.0293   0.7502   0.9628
  -0.500   0.1262   0.00521   0.00128  -0.0306   0.7285   1.0000
  -0.250   0.1534   0.00535   0.00124  -0.0303   0.7112   1.0000
   0.000   0.1809   0.00546   0.00122  -0.0301   0.6951   1.0000
   0.250   0.2085   0.00554   0.00121  -0.0299   0.6780   1.0000
   0.500   0.2363   0.00560   0.00118  -0.0297   0.6591   1.0000
   0.750   0.2639   0.00568   0.00115  -0.0295   0.6385   1.0000
   1.000   0.2915   0.00578   0.00114  -0.0293   0.6163   1.0000
   1.250   0.3190   0.00590   0.00114  -0.0291   0.5928   1.0000
   1.500   0.3464   0.00603   0.00116  -0.0289   0.5674   1.0000
   1.750   0.3739   0.00617   0.00119  -0.0287   0.5415   1.0000
   2.000   0.4014   0.00633   0.00123  -0.0286   0.5153   1.0000
   2.250   0.4289   0.00650   0.00129  -0.0284   0.4886   1.0000
   2.500   0.4563   0.00668   0.00138  -0.0283   0.4614   1.0000
   2.750   0.4837   0.00688   0.00146  -0.0282   0.4343   1.0000
   3.000   0.5111   0.00709   0.00157  -0.0280   0.4067   1.0000
   3.250   0.5385   0.00731   0.00169  -0.0279   0.3796   1.0000
   3.500   0.5658   0.00754   0.00184  -0.0278   0.3522   1.0000
   3.750   0.5930   0.00779   0.00199  -0.0277   0.3256   1.0000
   4.000   0.6202   0.00806   0.00216  -0.0276   0.2988   1.0000
   4.250   0.6473   0.00833   0.00235  -0.0275   0.2734   1.0000
   4.500   0.6744   0.00862   0.00257  -0.0274   0.2472   1.0000
   4.750   0.7014   0.00892   0.00280  -0.0273   0.2234   1.0000
   5.000   0.7283   0.00925   0.00305  -0.0272   0.1991   1.0000
   5.250   0.7550   0.00961   0.00332  -0.0271   0.1769   1.0000
   5.500   0.7818   0.00995   0.00361  -0.0270   0.1555   1.0000
   5.750   0.8083   0.01035   0.00396  -0.0268   0.1362   1.0000
   6.000   0.8348   0.01072   0.00430  -0.0267   0.1183   1.0000
   6.250   0.8612   0.01113   0.00467  -0.0265   0.1018   1.0000
   6.500   0.8873   0.01159   0.00508  -0.0264   0.0860   1.0000
   6.750   0.9133   0.01209   0.00555  -0.0262   0.0715   1.0000
   7.000   0.9389   0.01263   0.00607  -0.0260   0.0581   1.0000
   7.250   0.9643   0.01324   0.00666  -0.0257   0.0465   1.0000
   7.500   0.9899   0.01378   0.00722  -0.0255   0.0381   1.0000
   7.750   1.0151   0.01440   0.00790  -0.0252   0.0311   1.0000
   8.000   1.0387   0.01540   0.00896  -0.0247   0.0251   1.0000
   8.250   1.0636   0.01603   0.00970  -0.0243   0.0216   1.0000
   8.500   1.0844   0.01760   0.01138  -0.0236   0.0177   1.0000
   8.750   1.1092   0.01817   0.01205  -0.0233   0.0161   1.0000
   9.000   1.1323   0.01907   0.01307  -0.0228   0.0144   1.0000
   9.250   1.1538   0.02026   0.01435  -0.0222   0.0130   1.0000
   9.500   1.1666   0.02326   0.01764  -0.0208   0.0117   1.0000
   9.750   1.1893   0.02403   0.01854  -0.0203   0.0110   1.0000
  10.000   1.2102   0.02506   0.01975  -0.0197   0.0101   1.0000
  10.250   1.2292   0.02638   0.02122  -0.0190   0.0094   1.0000
  10.500   1.2464   0.02788   0.02287  -0.0183   0.0088   1.0000
  10.750   1.2607   0.02973   0.02488  -0.0174   0.0083   1.0000
  11.000   1.2670   0.03274   0.02815  -0.0161   0.0079   1.0000
  11.250   1.2607   0.03736   0.03316  -0.0144   0.0076   1.0000
  11.750   1.2380   0.04524   0.04161  -0.0124   0.0075   1.0000
  12.000   1.2304   0.04928   0.04588  -0.0145   0.0074   1.0000
  12.250   1.2182   0.05496   0.05177  -0.0185   0.0074   1.0000
  12.500   1.2045   0.06154   0.05855  -0.0234   0.0074   1.0000
  12.750   1.1865   0.06937   0.06655  -0.0291   0.0074   1.0000
  13.000   1.1620   0.07866   0.07600  -0.0354   0.0075   1.0000
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