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AG12 (ag12-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: AG12 (ag12-il)
Reynolds number: 50,000
Max Cl/Cd: 33.19 at α=4.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-ag12-il-50000.txt
Download as CSV file: xf-ag12-il-50000.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 =     0.050 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.500  -0.5444   0.10114   0.09472   0.0178   1.0000   0.2148
  -7.250  -0.5505   0.09947   0.09317   0.0137   1.0000   0.2251
  -7.000  -0.5426   0.09576   0.08951   0.0130   1.0000   0.2387
  -6.750  -0.5365   0.09238   0.08618   0.0108   1.0000   0.2528
  -6.500  -0.5245   0.08801   0.08186   0.0121   1.0000   0.2675
  -6.000  -0.5015   0.08019   0.07412   0.0128   1.0000   0.2977
  -5.750  -0.4901   0.07658   0.07055   0.0127   1.0000   0.3155
  -5.500  -0.4816   0.07319   0.06716   0.0108   1.0000   0.3374
  -5.250  -0.4678   0.06947   0.06347   0.0119   1.0000   0.3567
  -5.000  -0.4563   0.06596   0.06001   0.0117   1.0000   0.3808
  -4.750  -0.3578   0.04183   0.03383  -0.0426   1.0000   0.1174
  -4.500  -0.3295   0.03730   0.02890  -0.0440   1.0000   0.1118
  -4.250  -0.2992   0.03327   0.02412  -0.0455   1.0000   0.1136
  -4.000  -0.2704   0.02998   0.02038  -0.0461   1.0000   0.1197
  -3.750  -0.2405   0.02700   0.01690  -0.0461   1.0000   0.1250
  -3.500  -0.2114   0.02466   0.01415  -0.0459   1.0000   0.1419
  -3.250  -0.1836   0.02247   0.01188  -0.0453   1.0000   0.1679
  -3.000  -0.1570   0.02060   0.01018  -0.0445   1.0000   0.2227
  -2.750  -0.1310   0.01872   0.00881  -0.0434   1.0000   0.3298
  -2.500  -0.1093   0.01701   0.00792  -0.0412   1.0000   0.5032
  -2.250  -0.0960   0.01544   0.00726  -0.0364   1.0000   0.6829
  -2.000  -0.0640   0.01387   0.00601  -0.0351   1.0000   1.0000
  -1.750  -0.0335   0.01382   0.00534  -0.0363   1.0000   1.0000
  -1.500  -0.0058   0.01380   0.00490  -0.0366   1.0000   1.0000
  -1.250   0.0209   0.01381   0.00454  -0.0366   1.0000   1.0000
  -1.000   0.0470   0.01384   0.00430  -0.0365   1.0000   1.0000
  -0.750   0.0727   0.01389   0.00414  -0.0363   1.0000   1.0000
  -0.500   0.0982   0.01396   0.00405  -0.0362   1.0000   1.0000
  -0.250   0.1233   0.01406   0.00401  -0.0360   1.0000   1.0000
   0.000   0.1482   0.01418   0.00401  -0.0358   1.0000   1.0000
   0.250   0.1728   0.01434   0.00410  -0.0356   1.0000   1.0000
   0.500   0.1969   0.01455   0.00426  -0.0355   1.0000   1.0000
   0.750   0.2206   0.01480   0.00450  -0.0355   1.0000   1.0000
   1.000   0.2439   0.01513   0.00484  -0.0356   1.0000   1.0000
   1.250   0.2667   0.01556   0.00528  -0.0359   1.0000   1.0000
   1.500   0.2886   0.01611   0.00586  -0.0365   1.0000   1.0000
   1.750   0.3094   0.01683   0.00664  -0.0373   1.0000   1.0000
   2.000   0.3288   0.01776   0.00763  -0.0386   1.0000   1.0000
   2.250   0.4211   0.01852   0.00871  -0.0523   0.9556   1.0000
   2.500   0.5003   0.01864   0.00915  -0.0612   0.9066   1.0000
   2.750   0.5527   0.01855   0.00926  -0.0634   0.8589   1.0000
   3.000   0.5824   0.01869   0.00954  -0.0614   0.8097   1.0000
   3.250   0.6062   0.01890   0.00974  -0.0580   0.7634   1.0000
   3.500   0.6272   0.01928   0.01008  -0.0547   0.7154   1.0000
   3.750   0.6484   0.01970   0.01042  -0.0515   0.6686   1.0000
   4.000   0.6700   0.02024   0.01094  -0.0486   0.6202   1.0000
   4.250   0.6920   0.02085   0.01143  -0.0460   0.5722   1.0000
   4.500   0.7143   0.02158   0.01206  -0.0438   0.5229   1.0000
   4.750   0.7367   0.02236   0.01268  -0.0416   0.4748   1.0000
   5.000   0.7593   0.02330   0.01354  -0.0398   0.4251   1.0000
   5.250   0.7820   0.02433   0.01436  -0.0380   0.3786   1.0000
   5.500   0.8047   0.02557   0.01563  -0.0366   0.3318   1.0000
   5.750   0.8277   0.02699   0.01699  -0.0352   0.2894   1.0000
   6.000   0.8503   0.02851   0.01843  -0.0339   0.2497   1.0000
   6.250   0.8735   0.03040   0.02030  -0.0327   0.2167   1.0000
   6.500   0.8964   0.03249   0.02232  -0.0316   0.1876   1.0000
   6.750   0.9185   0.03502   0.02516  -0.0306   0.1631   1.0000
   7.000   0.9398   0.03808   0.02850  -0.0296   0.1444   1.0000
   7.250   0.9597   0.04177   0.03270  -0.0286   0.1312   1.0000
   7.500   0.9768   0.04521   0.03653  -0.0278   0.1176   1.0000
   7.750   0.9888   0.05069   0.04279  -0.0273   0.1137   1.0000
   8.000   0.9972   0.05643   0.04910  -0.0271   0.1116   1.0000
   8.250   1.0077   0.06103   0.05385  -0.0267   0.1060   1.0000
   8.500   1.0042   0.06748   0.06082  -0.0278   0.1055   1.0000
   8.750   0.9980   0.07397   0.06763  -0.0293   0.1058   1.0000
   9.000   0.9915   0.08037   0.07420  -0.0310   0.1065   1.0000
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