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AG17 (ag17-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: AG17 (ag17-il)
Reynolds number: 100,000
Max Cl/Cd: 49.55 at α=4°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-ag17-il-100000.txt
Download as CSV file: xf-ag17-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: AG17                                            
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.4342   0.10077   0.09621   0.0020   1.0000   0.0855
  -8.500  -0.4442   0.09791   0.09342  -0.0018   1.0000   0.0879
  -8.250  -0.4570   0.09519   0.09079  -0.0066   1.0000   0.0886
  -8.000  -0.4337   0.08887   0.08444  -0.0019   1.0000   0.0919
  -7.750  -0.4279   0.08514   0.08074  -0.0020   1.0000   0.0958
  -7.500  -0.4313   0.08157   0.07722  -0.0044   1.0000   0.0998
  -7.250  -0.4476   0.07814   0.07390  -0.0136   1.0000   0.1019
  -7.000  -0.5110   0.08527   0.08076  -0.0015   1.0000   0.0963
  -6.750  -0.4984   0.07886   0.07431  -0.0218   1.0000   0.1027
  -6.500  -0.4896   0.07586   0.07140  -0.0132   1.0000   0.1054
  -6.250  -0.4733   0.07163   0.06714  -0.0189   1.0000   0.1137
  -6.000  -0.4589   0.06667   0.06217  -0.0227   1.0000   0.1188
  -5.750  -0.4377   0.06133   0.05667  -0.0313   1.0000   0.1306
  -5.500  -0.4232   0.05837   0.05380  -0.0287   1.0000   0.1353
  -5.250  -0.4021   0.05386   0.04918  -0.0331   1.0000   0.1471
  -5.000  -0.3527   0.03812   0.03207  -0.0485   1.0000   0.0772
  -4.750  -0.3227   0.03171   0.02500  -0.0499   1.0000   0.0582
  -4.500  -0.2936   0.02706   0.01960  -0.0507   1.0000   0.0534
  -4.250  -0.2663   0.02466   0.01680  -0.0507   1.0000   0.0571
  -4.000  -0.2377   0.02229   0.01383  -0.0504   1.0000   0.0600
  -3.750  -0.2092   0.02025   0.01130  -0.0497   1.0000   0.0622
  -3.500  -0.1828   0.01809   0.00916  -0.0493   1.0000   0.0704
  -3.250  -0.1560   0.01668   0.00767  -0.0488   1.0000   0.0876
  -3.000  -0.1296   0.01511   0.00635  -0.0484   1.0000   0.1253
  -2.750  -0.1032   0.01348   0.00534  -0.0483   1.0000   0.2332
  -2.500  -0.0781   0.01247   0.00499  -0.0479   1.0000   0.3851
  -2.250  -0.0543   0.01184   0.00474  -0.0470   1.0000   0.5049
  -2.000  -0.0314   0.01123   0.00453  -0.0455   1.0000   0.6231
  -1.750  -0.0147   0.01039   0.00430  -0.0420   1.0000   0.7915
  -1.500   0.0142   0.00999   0.00389  -0.0418   1.0000   1.0000
  -1.250   0.0412   0.01004   0.00365  -0.0420   1.0000   1.0000
  -1.000   0.0676   0.01012   0.00351  -0.0420   1.0000   1.0000
  -0.750   0.0935   0.01022   0.00345  -0.0419   1.0000   1.0000
  -0.500   0.1192   0.01035   0.00345  -0.0419   1.0000   1.0000
  -0.250   0.1447   0.01051   0.00351  -0.0419   1.0000   1.0000
   0.000   0.1699   0.01072   0.00361  -0.0420   1.0000   1.0000
   0.250   0.1947   0.01098   0.00381  -0.0422   1.0000   1.0000
   0.500   0.2190   0.01130   0.00409  -0.0425   1.0000   1.0000
   0.750   0.2569   0.01165   0.00442  -0.0455   0.9931   1.0000
   1.000   0.3158   0.01181   0.00460  -0.0520   0.9722   1.0000
   1.250   0.3725   0.01184   0.00467  -0.0575   0.9511   1.0000
   1.500   0.4208   0.01177   0.00466  -0.0609   0.9269   1.0000
   1.750   0.4577   0.01172   0.00467  -0.0617   0.8969   1.0000
   2.000   0.4874   0.01171   0.00466  -0.0609   0.8637   1.0000
   2.250   0.5132   0.01175   0.00466  -0.0591   0.8292   1.0000
   2.500   0.5369   0.01185   0.00470  -0.0571   0.7918   1.0000
   2.750   0.5602   0.01202   0.00483  -0.0551   0.7522   1.0000
   3.000   0.5837   0.01224   0.00494  -0.0533   0.7108   1.0000
   3.250   0.6073   0.01253   0.00509  -0.0516   0.6679   1.0000
   3.500   0.6312   0.01286   0.00528  -0.0501   0.6230   1.0000
   3.750   0.6553   0.01325   0.00552  -0.0488   0.5758   1.0000
   4.000   0.6793   0.01371   0.00585  -0.0476   0.5294   1.0000
   4.250   0.7036   0.01422   0.00621  -0.0465   0.4810   1.0000
   4.500   0.7277   0.01480   0.00661  -0.0456   0.4345   1.0000
   4.750   0.7518   0.01544   0.00709  -0.0447   0.3887   1.0000
   5.000   0.7759   0.01614   0.00765  -0.0440   0.3439   1.0000
   5.250   0.7998   0.01691   0.00834  -0.0432   0.3012   1.0000
   5.500   0.8236   0.01776   0.00907  -0.0426   0.2602   1.0000
   5.750   0.8471   0.01873   0.00992  -0.0419   0.2216   1.0000
   6.000   0.8702   0.01981   0.01083  -0.0412   0.1857   1.0000
   6.250   0.8935   0.02107   0.01206  -0.0404   0.1535   1.0000
   6.500   0.9162   0.02235   0.01329  -0.0397   0.1250   1.0000
   6.750   0.9389   0.02412   0.01510  -0.0388   0.1034   1.0000
   7.000   0.9613   0.02574   0.01675  -0.0380   0.0845   1.0000
   7.250   0.9839   0.02789   0.01898  -0.0370   0.0710   1.0000
   7.500   1.0060   0.03030   0.02167  -0.0360   0.0605   1.0000
   7.750   1.0266   0.03361   0.02526  -0.0349   0.0539   1.0000
   8.000   1.0463   0.03574   0.02769  -0.0339   0.0470   1.0000
   8.500   1.0727   0.04538   0.03834  -0.0315   0.0437   1.0000
   8.750   1.0793   0.05023   0.04385  -0.0301   0.0433   1.0000
   9.000   1.0805   0.05527   0.04947  -0.0290   0.0428   1.0000
   9.250   1.0758   0.06053   0.05521  -0.0284   0.0425   1.0000
   9.500   1.0660   0.06590   0.06095  -0.0282   0.0426   1.0000
   9.750   1.0519   0.07121   0.06651  -0.0286   0.0429   1.0000
  10.000   1.0349   0.07628   0.07172  -0.0292   0.0434   1.0000
  10.250   1.0194   0.08206   0.07757  -0.0314   0.0440   1.0000
  10.500   0.9583   0.10290   0.09868  -0.0520   0.0540   1.0000
  10.750   0.9293   0.11835   0.11402  -0.0630   0.0633   1.0000
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