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AG17 (ag17-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: AG17 (ag17-il)
Reynolds number: 100,000
Max Cl/Cd: 48.43 at α=4.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-ag17-il-100000-n5.txt
Download as CSV file: xf-ag17-il-100000-n5.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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.750  -0.5295   0.09129   0.08654   0.0033   1.0000   0.0323
  -7.500  -0.5301   0.08368   0.07903  -0.0050   1.0000   0.0267
  -7.250  -0.5189   0.07903   0.07436  -0.0089   1.0000   0.0244
  -7.000  -0.5092   0.07409   0.06944  -0.0133   1.0000   0.0237
  -6.750  -0.4965   0.06840   0.06374  -0.0192   1.0000   0.0231
  -6.500  -0.4804   0.06189   0.05717  -0.0262   1.0000   0.0225
  -6.250  -0.4615   0.05461   0.04974  -0.0334   1.0000   0.0218
  -6.000  -0.4405   0.04724   0.04208  -0.0396   1.0000   0.0212
  -5.750  -0.4176   0.04049   0.03487  -0.0442   1.0000   0.0207
  -5.500  -0.3928   0.03474   0.02853  -0.0472   1.0000   0.0204
  -5.250  -0.3668   0.03018   0.02332  -0.0488   1.0000   0.0205
  -5.000  -0.3400   0.02664   0.01916  -0.0495   1.0000   0.0209
  -4.750  -0.3130   0.02386   0.01585  -0.0496   1.0000   0.0216
  -4.500  -0.2860   0.02193   0.01349  -0.0494   1.0000   0.0240
  -4.250  -0.2596   0.02006   0.01135  -0.0492   1.0000   0.0271
  -4.000  -0.2335   0.01856   0.00972  -0.0489   1.0000   0.0298
  -3.750  -0.2071   0.01727   0.00823  -0.0484   1.0000   0.0337
  -3.500  -0.1810   0.01625   0.00714  -0.0482   1.0000   0.0426
  -3.250  -0.1543   0.01522   0.00605  -0.0479   1.0000   0.0556
  -3.000  -0.1278   0.01432   0.00526  -0.0477   1.0000   0.0900
  -2.750  -0.1017   0.01348   0.00473  -0.0477   1.0000   0.1565
  -2.500  -0.0759   0.01283   0.00442  -0.0476   1.0000   0.2438
  -2.250  -0.0503   0.01236   0.00416  -0.0474   1.0000   0.3271
  -2.000  -0.0250   0.01199   0.00400  -0.0470   1.0000   0.4055
  -1.750   0.0004   0.01164   0.00387  -0.0465   1.0000   0.4841
  -1.500   0.0255   0.01129   0.00376  -0.0459   1.0000   0.5673
  -1.250   0.0489   0.01088   0.00365  -0.0448   1.0000   0.6686
  -1.000   0.0682   0.01023   0.00355  -0.0421   1.0000   0.8706
  -0.750   0.1016   0.01021   0.00342  -0.0435   0.9940   1.0000
  -0.500   0.1447   0.01029   0.00333  -0.0468   0.9789   1.0000
  -0.250   0.1856   0.01035   0.00324  -0.0495   0.9612   1.0000
   0.000   0.2254   0.01040   0.00318  -0.0518   0.9421   1.0000
   0.250   0.2620   0.01044   0.00314  -0.0533   0.9192   1.0000
   0.500   0.2966   0.01049   0.00310  -0.0543   0.8943   1.0000
   0.750   0.3283   0.01055   0.00307  -0.0546   0.8666   1.0000
   1.000   0.3576   0.01063   0.00306  -0.0543   0.8368   1.0000
   1.250   0.3851   0.01074   0.00307  -0.0536   0.8054   1.0000
   1.500   0.4116   0.01088   0.00310  -0.0528   0.7726   1.0000
   1.750   0.4377   0.01105   0.00317  -0.0519   0.7387   1.0000
   2.000   0.4639   0.01124   0.00325  -0.0510   0.7039   1.0000
   2.250   0.4898   0.01147   0.00335  -0.0502   0.6679   1.0000
   2.500   0.5157   0.01172   0.00348  -0.0494   0.6306   1.0000
   2.750   0.5414   0.01201   0.00368  -0.0487   0.5926   1.0000
   3.000   0.5671   0.01232   0.00386  -0.0480   0.5539   1.0000
   3.250   0.5927   0.01267   0.00408  -0.0473   0.5149   1.0000
   3.500   0.6183   0.01304   0.00434  -0.0467   0.4755   1.0000
   3.750   0.6438   0.01345   0.00463  -0.0461   0.4364   1.0000
   4.000   0.6694   0.01389   0.00501  -0.0456   0.3980   1.0000
   4.250   0.6948   0.01436   0.00539  -0.0451   0.3608   1.0000
   4.500   0.7201   0.01487   0.00581  -0.0447   0.3245   1.0000
   4.750   0.7454   0.01539   0.00627  -0.0443   0.2896   1.0000
   5.000   0.7705   0.01597   0.00678  -0.0439   0.2559   1.0000
   5.250   0.7956   0.01658   0.00740  -0.0435   0.2237   1.0000
   5.500   0.8205   0.01723   0.00803  -0.0431   0.1930   1.0000
   5.750   0.8451   0.01795   0.00871  -0.0427   0.1639   1.0000
   6.000   0.8694   0.01874   0.00948  -0.0423   0.1369   1.0000
   6.250   0.8934   0.01960   0.01035  -0.0419   0.1128   1.0000
   6.500   0.9168   0.02060   0.01132  -0.0414   0.0915   1.0000
   6.750   0.9400   0.02167   0.01252  -0.0408   0.0734   1.0000
   7.000   0.9623   0.02289   0.01376  -0.0403   0.0582   1.0000
   7.250   0.9840   0.02427   0.01526  -0.0395   0.0473   1.0000
   7.500   1.0050   0.02578   0.01692  -0.0387   0.0382   1.0000
   7.750   1.0241   0.02763   0.01886  -0.0378   0.0322   1.0000
   8.000   1.0444   0.02942   0.02094  -0.0367   0.0273   1.0000
   8.250   1.0611   0.03154   0.02315  -0.0359   0.0230   1.0000
   8.500   1.0794   0.03394   0.02592  -0.0347   0.0211   1.0000
   8.750   1.0954   0.03688   0.02927  -0.0334   0.0195   1.0000
   9.000   1.1093   0.03983   0.03270  -0.0322   0.0180   1.0000
   9.250   1.1216   0.04218   0.03533  -0.0313   0.0163   1.0000
   9.500   1.1289   0.04511   0.03847  -0.0305   0.0150   1.0000
   9.750   1.1287   0.04937   0.04314  -0.0293   0.0143   1.0000
  10.000   1.1260   0.05339   0.04761  -0.0283   0.0140   1.0000
  10.250   1.1159   0.05761   0.05219  -0.0275   0.0139   1.0000
  10.500   1.1006   0.06206   0.05692  -0.0276   0.0138   1.0000
  10.750   1.0838   0.06743   0.06253  -0.0301   0.0139   1.0000
  11.000   1.0664   0.07409   0.06938  -0.0349   0.0140   1.0000
  11.250   1.0486   0.08210   0.07755  -0.0414   0.0141   1.0000
  11.500   1.0303   0.09115   0.08671  -0.0482   0.0143   1.0000
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