Airfoil Tools
Search 1636 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

AG37 (ag37-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: AG37 (ag37-il)
Reynolds number: 50,000
Max Cl/Cd: 36.42 at α=4.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-ag37-il-50000-n5.txt
Download as CSV file: xf-ag37-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: AG37                                            
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.4005   0.09283   0.08633  -0.0117   1.0000   0.1264
  -9.250  -0.4003   0.08899   0.08254  -0.0126   1.0000   0.1270
  -9.000  -0.5053   0.09046   0.08363  -0.0188   1.0000   0.0657
  -8.750  -0.4959   0.08665   0.07982  -0.0175   1.0000   0.0637
  -8.250  -0.4842   0.07394   0.06694  -0.0307   1.0000   0.0529
  -7.750  -0.4639   0.06560   0.05851  -0.0326   1.0000   0.0513
  -7.500  -0.4521   0.06117   0.05397  -0.0343   1.0000   0.0502
  -7.250  -0.4386   0.05661   0.04922  -0.0361   1.0000   0.0491
  -7.000  -0.4232   0.05205   0.04440  -0.0376   1.0000   0.0482
  -6.750  -0.4058   0.04765   0.03965  -0.0388   1.0000   0.0475
  -6.500  -0.3866   0.04352   0.03510  -0.0395   1.0000   0.0472
  -6.250  -0.3657   0.03999   0.03111  -0.0398   1.0000   0.0482
  -6.000  -0.3433   0.03682   0.02741  -0.0398   1.0000   0.0504
  -5.750  -0.3195   0.03394   0.02393  -0.0395   1.0000   0.0521
  -5.500  -0.2950   0.03138   0.02079  -0.0389   1.0000   0.0529
  -5.250  -0.2711   0.02899   0.01813  -0.0382   1.0000   0.0544
  -5.000  -0.2474   0.02729   0.01629  -0.0376   1.0000   0.0582
  -4.750  -0.2225   0.02581   0.01453  -0.0367   1.0000   0.0637
  -4.500  -0.1974   0.02426   0.01274  -0.0357   1.0000   0.0675
  -4.250  -0.1730   0.02302   0.01143  -0.0349   1.0000   0.0742
  -4.000  -0.1482   0.02196   0.01027  -0.0341   1.0000   0.0847
  -3.750  -0.1230   0.02096   0.00919  -0.0334   1.0000   0.0982
  -3.500  -0.0978   0.01997   0.00828  -0.0329   1.0000   0.1263
  -3.250  -0.0723   0.01852   0.00748  -0.0329   1.0000   0.2273
  -3.000  -0.0566   0.01627   0.00722  -0.0304   1.0000   0.6315
  -2.750  -0.0217   0.01554   0.00681  -0.0302   1.0000   1.0000
  -2.500   0.0007   0.01563   0.00655  -0.0295   1.0000   1.0000
  -2.250   0.0231   0.01575   0.00638  -0.0288   1.0000   1.0000
  -2.000   0.0453   0.01591   0.00631  -0.0282   1.0000   1.0000
  -1.750   0.0672   0.01612   0.00632  -0.0276   1.0000   1.0000
  -1.500   0.0920   0.01637   0.00641  -0.0278   0.9979   1.0000
  -1.250   0.1410   0.01660   0.00642  -0.0325   0.9799   1.0000
  -1.000   0.1886   0.01677   0.00645  -0.0368   0.9601   1.0000
  -0.750   0.2363   0.01689   0.00645  -0.0409   0.9402   1.0000
  -0.500   0.2782   0.01694   0.00644  -0.0436   0.9158   1.0000
  -0.250   0.3163   0.01697   0.00642  -0.0454   0.8885   1.0000
   0.000   0.3554   0.01696   0.00637  -0.0472   0.8625   1.0000
   0.250   0.3927   0.01698   0.00632  -0.0484   0.8359   1.0000
   0.500   0.4265   0.01705   0.00632  -0.0489   0.8079   1.0000
   0.750   0.4570   0.01719   0.00640  -0.0488   0.7792   1.0000
   1.000   0.4852   0.01738   0.00651  -0.0482   0.7506   1.0000
   1.250   0.5119   0.01762   0.00668  -0.0475   0.7225   1.0000
   1.500   0.5379   0.01789   0.00688  -0.0466   0.6949   1.0000
   1.750   0.5633   0.01817   0.00715  -0.0458   0.6674   1.0000
   2.000   0.5885   0.01847   0.00739  -0.0449   0.6412   1.0000
   2.250   0.6138   0.01881   0.00766  -0.0440   0.6165   1.0000
   2.500   0.6391   0.01920   0.00798  -0.0431   0.5929   1.0000
   2.750   0.6642   0.01964   0.00844  -0.0423   0.5689   1.0000
   3.000   0.6896   0.02012   0.00893  -0.0417   0.5445   1.0000
   3.250   0.7149   0.02061   0.00942  -0.0410   0.5209   1.0000
   3.500   0.7398   0.02110   0.00997  -0.0404   0.4960   1.0000
   3.750   0.7642   0.02159   0.01058  -0.0397   0.4704   1.0000
   4.000   0.7882   0.02205   0.01113  -0.0389   0.4439   1.0000
   4.250   0.8117   0.02251   0.01165  -0.0380   0.4166   1.0000
   4.500   0.8346   0.02299   0.01218  -0.0371   0.3875   1.0000
   4.750   0.8569   0.02353   0.01278  -0.0361   0.3564   1.0000
   5.000   0.8784   0.02414   0.01338  -0.0351   0.3236   1.0000
   5.250   0.8993   0.02487   0.01411  -0.0342   0.2885   1.0000
   5.500   0.9193   0.02572   0.01493  -0.0332   0.2520   1.0000
   5.750   0.9387   0.02675   0.01593  -0.0322   0.2146   1.0000
   6.000   0.9568   0.02800   0.01711  -0.0313   0.1811   1.0000
   6.500   0.9920   0.03105   0.02009  -0.0292   0.1295   1.0000
   6.750   1.0086   0.03276   0.02176  -0.0281   0.1136   1.0000
   7.000   1.0255   0.03452   0.02360  -0.0270   0.1009   1.0000
   7.250   1.0423   0.03633   0.02561  -0.0259   0.0900   1.0000
   7.500   1.0597   0.03840   0.02782  -0.0247   0.0832   1.0000
   7.750   1.0761   0.04041   0.03000  -0.0236   0.0763   1.0000
   8.000   1.0915   0.04270   0.03252  -0.0225   0.0703   1.0000
   8.250   1.1067   0.04525   0.03540  -0.0214   0.0661   1.0000
   8.500   1.1210   0.04772   0.03796  -0.0204   0.0626   1.0000
   8.750   1.1293   0.05090   0.04160  -0.0191   0.0594   1.0000
   9.000   1.1330   0.05416   0.04533  -0.0176   0.0563   1.0000
   9.250   1.1345   0.05753   0.04905  -0.0162   0.0543   1.0000
   9.500   1.1325   0.06113   0.05295  -0.0149   0.0530   1.0000
   9.750   1.1253   0.06473   0.05681  -0.0134   0.0522   1.0000
  10.000   1.1147   0.06854   0.06082  -0.0124   0.0515   1.0000
  10.250   1.1013   0.07284   0.06533  -0.0123   0.0510   1.0000
  10.500   1.0786   0.07861   0.07135  -0.0142   0.0511   1.0000
  10.750   1.0358   0.08883   0.08190  -0.0211   0.0527   1.0000
  11.000   0.9844   0.10403   0.09723  -0.0325   0.0550   1.0000
<< Back to AG37 (ag37-il)

Polar data table (+)

Polar graphs


<< Back to AG37 (ag37-il)