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AG12 (ag12-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: AG12 (ag12-il)
Reynolds number: 200,000
Max Cl/Cd: 61.37 at α=4.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-ag12-il-200000.txt
Download as CSV file: xf-ag12-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: AG12                                            
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.5626   0.10455   0.10111   0.0174   1.0000   0.0345
  -8.250  -0.5594   0.10074   0.09734   0.0139   1.0000   0.0358
  -8.000  -0.5565   0.09699   0.09364   0.0084   1.0000   0.0367
  -7.750  -0.5516   0.09291   0.08961   0.0027   1.0000   0.0371
  -7.500  -0.5401   0.08797   0.08467  -0.0042   1.0000   0.0373
  -7.250  -0.5259   0.08223   0.07892  -0.0120   1.0000   0.0374
  -7.000  -0.5278   0.07581   0.07257  -0.0116   1.0000   0.0386
  -6.750  -0.5179   0.07285   0.06961  -0.0109   1.0000   0.0399
  -6.500  -0.5038   0.06899   0.06575  -0.0138   1.0000   0.0416
  -6.250  -0.4858   0.06403   0.06074  -0.0193   1.0000   0.0439
  -6.000  -0.4481   0.05540   0.05159  -0.0352   1.0000   0.0497
  -5.750  -0.4347   0.04799   0.04411  -0.0385   1.0000   0.0513
  -5.500  -0.4170   0.04528   0.04146  -0.0385   1.0000   0.0535
  -5.250  -0.3895   0.04015   0.03578  -0.0432   1.0000   0.0642
  -5.000  -0.3701   0.03822   0.03403  -0.0426   1.0000   0.0694
  -4.750  -0.3321   0.02652   0.02081  -0.0461   1.0000   0.0393
  -4.500  -0.3043   0.02257   0.01647  -0.0461   1.0000   0.0324
  -4.250  -0.2751   0.01894   0.01215  -0.0458   1.0000   0.0295
  -4.000  -0.2472   0.01681   0.00967  -0.0453   1.0000   0.0295
  -3.750  -0.2199   0.01524   0.00789  -0.0449   1.0000   0.0312
  -3.500  -0.1927   0.01417   0.00669  -0.0444   1.0000   0.0357
  -3.250  -0.1659   0.01283   0.00533  -0.0440   1.0000   0.0433
  -3.000  -0.1386   0.01176   0.00426  -0.0437   1.0000   0.0589
  -2.750  -0.1111   0.01063   0.00340  -0.0437   1.0000   0.1200
  -2.500  -0.0844   0.00976   0.00303  -0.0439   1.0000   0.2362
  -2.250  -0.0581   0.00921   0.00287  -0.0438   1.0000   0.3479
  -2.000  -0.0322   0.00877   0.00276  -0.0435   1.0000   0.4468
  -1.750  -0.0066   0.00834   0.00267  -0.0430   1.0000   0.5456
  -1.500   0.0175   0.00783   0.00260  -0.0421   1.0000   0.6647
  -1.250   0.0385   0.00698   0.00245  -0.0395   1.0000   1.0000
  -1.000   0.0655   0.00702   0.00235  -0.0394   1.0000   1.0000
  -0.750   0.0922   0.00708   0.00230  -0.0393   1.0000   1.0000
  -0.500   0.1188   0.00716   0.00230  -0.0393   1.0000   1.0000
  -0.250   0.1454   0.00727   0.00233  -0.0393   1.0000   1.0000
   0.000   0.1878   0.00734   0.00234  -0.0426   0.9891   1.0000
   0.250   0.2358   0.00735   0.00231  -0.0467   0.9713   1.0000
   0.500   0.2785   0.00734   0.00226  -0.0495   0.9471   1.0000
   0.750   0.3159   0.00732   0.00220  -0.0509   0.9165   1.0000
   1.000   0.3457   0.00736   0.00215  -0.0506   0.8788   1.0000
   1.250   0.3712   0.00747   0.00213  -0.0493   0.8379   1.0000
   1.500   0.3956   0.00764   0.00214  -0.0479   0.7961   1.0000
   1.750   0.4203   0.00786   0.00220  -0.0467   0.7542   1.0000
   2.000   0.4454   0.00810   0.00225  -0.0457   0.7122   1.0000
   2.250   0.4708   0.00836   0.00234  -0.0448   0.6706   1.0000
   2.500   0.4965   0.00865   0.00244  -0.0440   0.6295   1.0000
   2.750   0.5223   0.00894   0.00258  -0.0434   0.5891   1.0000
   3.000   0.5483   0.00925   0.00277  -0.0428   0.5492   1.0000
   3.250   0.5744   0.00958   0.00295  -0.0423   0.5098   1.0000
   3.500   0.6005   0.00992   0.00315  -0.0419   0.4710   1.0000
   3.750   0.6266   0.01028   0.00338  -0.0415   0.4323   1.0000
   4.000   0.6527   0.01066   0.00367  -0.0411   0.3933   1.0000
   4.250   0.6788   0.01106   0.00396  -0.0408   0.3546   1.0000
   4.500   0.7049   0.01149   0.00429  -0.0405   0.3162   1.0000
   4.750   0.7307   0.01196   0.00465  -0.0402   0.2775   1.0000
   5.000   0.7564   0.01250   0.00506  -0.0399   0.2404   1.0000
   5.250   0.7820   0.01307   0.00556  -0.0396   0.2030   1.0000
   5.500   0.8074   0.01372   0.00612  -0.0393   0.1689   1.0000
   5.750   0.8324   0.01445   0.00674  -0.0390   0.1369   1.0000
   6.000   0.8571   0.01530   0.00749  -0.0386   0.1095   1.0000
   6.250   0.8818   0.01617   0.00834  -0.0382   0.0858   1.0000
   6.500   0.9051   0.01740   0.00951  -0.0376   0.0674   1.0000
   6.750   0.9291   0.01847   0.01070  -0.0370   0.0520   1.0000
   7.000   0.9516   0.02002   0.01231  -0.0362   0.0414   1.0000
   7.250   0.9725   0.02198   0.01430  -0.0353   0.0328   1.0000
   7.500   0.9960   0.02357   0.01614  -0.0343   0.0286   1.0000
   7.750   1.0174   0.02522   0.01788  -0.0336   0.0243   1.0000
   8.000   1.0365   0.02828   0.02130  -0.0324   0.0218   1.0000
   8.250   1.0561   0.03105   0.02448  -0.0313   0.0206   1.0000
   8.500   1.0725   0.03454   0.02847  -0.0300   0.0198   1.0000
   8.750   1.0842   0.03881   0.03329  -0.0286   0.0196   1.0000
   9.000   1.0896   0.04385   0.03890  -0.0273   0.0198   1.0000
   9.250   1.0884   0.04937   0.04494  -0.0263   0.0202   1.0000
   9.500   1.0807   0.05499   0.05098  -0.0258   0.0207   1.0000
   9.750   1.0670   0.06044   0.05676  -0.0260   0.0210   1.0000
  10.000   1.0480   0.06531   0.06183  -0.0267   0.0212   1.0000
  10.250   1.0282   0.07144   0.06815  -0.0313   0.0214   1.0000
  10.500   1.0086   0.07990   0.07675  -0.0394   0.0215   1.0000
  10.750   0.9881   0.09062   0.08756  -0.0488   0.0218   1.0000
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