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AG10 (ag10-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: AG10 (ag10-il)
Reynolds number: 50,000
Max Cl/Cd: 30.96 at α=4.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-ag10-il-50000-n5.txt
Download as CSV file: xf-ag10-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: AG10                                            
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.6173   0.12378   0.11673   0.0283   1.0000   0.0891
  -8.750  -0.6205   0.12164   0.11467   0.0241   1.0000   0.0900
  -8.500  -0.6226   0.11912   0.11223   0.0197   1.0000   0.0904
  -8.250  -0.6084   0.11283   0.10596   0.0218   1.0000   0.0916
  -8.000  -0.5974   0.10824   0.10137   0.0226   1.0000   0.0935
  -7.750  -0.5907   0.10448   0.09762   0.0216   1.0000   0.0953
  -7.500  -0.5851   0.10092   0.09409   0.0198   1.0000   0.0976
  -7.250  -0.5786   0.09731   0.09053   0.0158   1.0000   0.1013
  -7.000  -0.5704   0.09410   0.08729   0.0034   1.0000   0.1047
  -6.750  -0.5601   0.08898   0.08224   0.0054   1.0000   0.1066
  -6.500  -0.5482   0.08506   0.07834   0.0065   1.0000   0.1107
  -6.000  -0.5053   0.07142   0.06435  -0.0089   1.0000   0.0611
  -5.750  -0.4885   0.06685   0.05972  -0.0110   1.0000   0.0580
  -5.250  -0.4360   0.05443   0.04663  -0.0220   1.0000   0.0501
  -5.000  -0.4127   0.04997   0.04194  -0.0241   1.0000   0.0496
  -4.750  -0.3851   0.04511   0.03660  -0.0270   1.0000   0.0502
  -4.500  -0.3605   0.04150   0.03274  -0.0284   1.0000   0.0524
  -4.250  -0.3341   0.03820   0.02911  -0.0295   1.0000   0.0540
  -4.000  -0.3052   0.03460   0.02502  -0.0306   1.0000   0.0547
  -3.750  -0.2764   0.03168   0.02162  -0.0313   1.0000   0.0577
  -3.500  -0.2462   0.02887   0.01810  -0.0317   1.0000   0.0620
  -3.250  -0.2179   0.02646   0.01531  -0.0317   1.0000   0.0650
  -3.000  -0.1903   0.02488   0.01351  -0.0316   1.0000   0.0722
  -2.750  -0.1620   0.02309   0.01138  -0.0312   1.0000   0.0784
  -2.500  -0.1345   0.02186   0.00993  -0.0307   1.0000   0.0895
  -2.250  -0.1077   0.02065   0.00863  -0.0301   1.0000   0.1017
  -2.000  -0.0799   0.01963   0.00746  -0.0297   1.0000   0.1193
  -1.750  -0.0530   0.01867   0.00659  -0.0293   1.0000   0.1443
  -1.500  -0.0262   0.01775   0.00582  -0.0290   1.0000   0.1830
  -1.250   0.0001   0.01671   0.00517  -0.0288   1.0000   0.2578
  -1.000   0.0235   0.01377   0.00459  -0.0270   1.0000   1.0000
  -0.750   0.0501   0.01377   0.00418  -0.0264   1.0000   1.0000
  -0.500   0.0762   0.01379   0.00392  -0.0260   1.0000   1.0000
  -0.250   0.1021   0.01381   0.00374  -0.0255   1.0000   1.0000
   0.000   0.1278   0.01385   0.00364  -0.0250   1.0000   1.0000
   0.250   0.1532   0.01391   0.00360  -0.0246   1.0000   1.0000
   0.500   0.1785   0.01398   0.00361  -0.0241   1.0000   1.0000
   0.750   0.2035   0.01407   0.00369  -0.0236   1.0000   1.0000
   1.000   0.2282   0.01419   0.00381  -0.0232   1.0000   1.0000
   1.250   0.2527   0.01433   0.00400  -0.0228   1.0000   1.0000
   1.500   0.2770   0.01450   0.00426  -0.0224   1.0000   1.0000
   1.750   0.3155   0.01471   0.00460  -0.0250   0.9783   1.0000
   2.000   0.3693   0.01487   0.00487  -0.0298   0.9147   1.0000
   2.250   0.4135   0.01504   0.00506  -0.0319   0.8284   1.0000
   2.500   0.4438   0.01538   0.00522  -0.0308   0.7342   1.0000
   2.750   0.4670   0.01590   0.00540  -0.0285   0.6473   1.0000
   3.000   0.4896   0.01653   0.00568  -0.0266   0.5733   1.0000
   3.250   0.5131   0.01719   0.00606  -0.0252   0.5117   1.0000
   3.500   0.5375   0.01787   0.00656  -0.0242   0.4596   1.0000
   3.750   0.5624   0.01856   0.00709  -0.0234   0.4139   1.0000
   4.000   0.5875   0.01926   0.00767  -0.0227   0.3732   1.0000
   4.250   0.6130   0.01996   0.00833  -0.0222   0.3345   1.0000
   4.500   0.6384   0.02070   0.00905  -0.0217   0.2985   1.0000
   4.750   0.6643   0.02146   0.00984  -0.0213   0.2628   1.0000
   5.000   0.6898   0.02230   0.01066  -0.0210   0.2295   1.0000
   5.250   0.7149   0.02323   0.01154  -0.0206   0.1978   1.0000
   5.500   0.7400   0.02426   0.01266  -0.0202   0.1669   1.0000
   5.750   0.7645   0.02545   0.01386  -0.0197   0.1390   1.0000
   6.000   0.7884   0.02681   0.01523  -0.0193   0.1152   1.0000
   6.250   0.8125   0.02843   0.01704  -0.0187   0.0956   1.0000
   6.500   0.8360   0.03022   0.01900  -0.0181   0.0811   1.0000
   6.750   0.8590   0.03227   0.02124  -0.0175   0.0705   1.0000
   7.000   0.8808   0.03431   0.02335  -0.0170   0.0626   1.0000
   7.250   0.9030   0.03718   0.02679  -0.0163   0.0566   1.0000
   7.500   0.9228   0.03955   0.02924  -0.0159   0.0521   1.0000
   7.750   0.9407   0.04350   0.03401  -0.0154   0.0482   1.0000
   8.000   0.9553   0.04774   0.03883  -0.0152   0.0462   1.0000
   8.250   0.9669   0.05205   0.04361  -0.0151   0.0446   1.0000
   8.500   0.9773   0.05573   0.04757  -0.0152   0.0429   1.0000
   8.750   0.9818   0.06038   0.05251  -0.0156   0.0416   1.0000
   9.000   0.9759   0.06667   0.05929  -0.0173   0.0413   1.0000
   9.250   0.9653   0.07305   0.06599  -0.0198   0.0413   1.0000
   9.500   0.9516   0.07939   0.07250  -0.0231   0.0414   1.0000
   9.750   0.9376   0.08596   0.07912  -0.0277   0.0418   1.0000
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