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AG16 (ag16-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: AG16 (ag16-il)
Reynolds number: 200,000
Max Cl/Cd: 63.61 at α=4.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-ag16-il-200000.txt
Download as CSV file: xf-ag16-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: AG16                                            
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.5446   0.09185   0.08855   0.0050   1.0000   0.0412
  -7.750  -0.5414   0.08787   0.08460   0.0023   1.0000   0.0427
  -7.500  -0.5384   0.08344   0.08023  -0.0019   1.0000   0.0445
  -7.250  -0.5224   0.07441   0.07118  -0.0216   1.0000   0.0472
  -7.000  -0.5075   0.06673   0.06325  -0.0329   1.0000   0.0476
  -6.750  -0.5037   0.06006   0.05670  -0.0336   1.0000   0.0491
  -6.500  -0.4900   0.05741   0.05404  -0.0333   1.0000   0.0506
  -6.250  -0.4729   0.05330   0.04985  -0.0361   1.0000   0.0530
  -6.000  -0.4499   0.04460   0.04035  -0.0463   1.0000   0.0614
  -5.750  -0.4268   0.03279   0.02798  -0.0482   1.0000   0.0316
  -5.500  -0.4022   0.02834   0.02298  -0.0493   1.0000   0.0317
  -5.250  -0.3763   0.02442   0.01848  -0.0497   1.0000   0.0311
  -5.000  -0.3492   0.02071   0.01413  -0.0495   1.0000   0.0292
  -4.750  -0.3221   0.01830   0.01132  -0.0490   1.0000   0.0291
  -4.500  -0.2954   0.01656   0.00934  -0.0485   1.0000   0.0304
  -4.250  -0.2689   0.01517   0.00780  -0.0479   1.0000   0.0331
  -4.000  -0.2428   0.01386   0.00644  -0.0475   1.0000   0.0401
  -3.750  -0.2162   0.01269   0.00524  -0.0471   1.0000   0.0505
  -3.500  -0.1892   0.01158   0.00424  -0.0470   1.0000   0.0887
  -3.250  -0.1626   0.01062   0.00369  -0.0472   1.0000   0.1724
  -3.000  -0.1364   0.01010   0.00346  -0.0471   1.0000   0.2534
  -2.750  -0.1102   0.00976   0.00330  -0.0469   1.0000   0.3197
  -2.500  -0.0841   0.00946   0.00318  -0.0467   1.0000   0.3776
  -2.250  -0.0580   0.00921   0.00305  -0.0465   1.0000   0.4315
  -2.000  -0.0319   0.00897   0.00298  -0.0462   1.0000   0.4870
  -1.750  -0.0060   0.00872   0.00295  -0.0459   1.0000   0.5495
  -1.500   0.0196   0.00845   0.00293  -0.0455   1.0000   0.6192
  -1.250   0.0441   0.00812   0.00292  -0.0447   1.0000   0.7068
  -1.000   0.0626   0.00763   0.00293  -0.0421   1.0000   0.8658
  -0.750   0.0970   0.00759   0.00288  -0.0437   0.9957   1.0000
  -0.500   0.1457   0.00763   0.00281  -0.0482   0.9837   1.0000
  -0.250   0.1927   0.00763   0.00271  -0.0521   0.9706   1.0000
   0.000   0.2360   0.00760   0.00262  -0.0551   0.9548   1.0000
   0.250   0.2724   0.00758   0.00254  -0.0564   0.9327   1.0000
   0.500   0.3046   0.00757   0.00247  -0.0567   0.9080   1.0000
   0.750   0.3322   0.00759   0.00241  -0.0560   0.8800   1.0000
   1.000   0.3576   0.00766   0.00238  -0.0548   0.8502   1.0000
   1.250   0.3827   0.00776   0.00236  -0.0536   0.8192   1.0000
   1.500   0.4078   0.00790   0.00238  -0.0525   0.7873   1.0000
   1.750   0.4333   0.00807   0.00242  -0.0516   0.7544   1.0000
   2.000   0.4589   0.00827   0.00247  -0.0507   0.7216   1.0000
   2.250   0.4848   0.00848   0.00255  -0.0499   0.6879   1.0000
   2.500   0.5108   0.00872   0.00265  -0.0493   0.6542   1.0000
   2.750   0.5369   0.00897   0.00280  -0.0486   0.6202   1.0000
   3.000   0.5630   0.00923   0.00294  -0.0481   0.5853   1.0000
   3.250   0.5892   0.00952   0.00310  -0.0476   0.5500   1.0000
   3.500   0.6153   0.00983   0.00329  -0.0471   0.5144   1.0000
   3.750   0.6415   0.01015   0.00354  -0.0466   0.4774   1.0000
   4.000   0.6675   0.01051   0.00378  -0.0462   0.4398   1.0000
   4.250   0.6934   0.01090   0.00405  -0.0458   0.4018   1.0000
   4.500   0.7192   0.01131   0.00435  -0.0455   0.3633   1.0000
   4.750   0.7450   0.01175   0.00469  -0.0451   0.3249   1.0000
   5.000   0.7707   0.01224   0.00511  -0.0448   0.2872   1.0000
   5.250   0.7961   0.01277   0.00553  -0.0445   0.2505   1.0000
   5.500   0.8214   0.01335   0.00600  -0.0442   0.2138   1.0000
   5.750   0.8466   0.01397   0.00654  -0.0438   0.1785   1.0000
   6.000   0.8712   0.01473   0.00716  -0.0435   0.1421   1.0000
   6.250   0.8953   0.01562   0.00796  -0.0431   0.1073   1.0000
   6.500   0.9182   0.01680   0.00898  -0.0425   0.0786   1.0000
   6.750   0.9417   0.01791   0.01012  -0.0419   0.0585   1.0000
   7.000   0.9637   0.01931   0.01152  -0.0411   0.0446   1.0000
   7.250   0.9859   0.02065   0.01294  -0.0402   0.0346   1.0000
   7.500   1.0051   0.02302   0.01541  -0.0389   0.0293   1.0000
   7.750   1.0278   0.02423   0.01676  -0.0381   0.0244   1.0000
   8.000   1.0439   0.02787   0.02056  -0.0368   0.0211   1.0000
   8.250   1.0651   0.02987   0.02289  -0.0357   0.0201   1.0000
   8.500   1.0834   0.03273   0.02621  -0.0343   0.0194   1.0000
   8.750   1.0981   0.03625   0.03019  -0.0329   0.0190   1.0000
   9.000   1.1080   0.04038   0.03483  -0.0314   0.0190   1.0000
   9.250   1.1120   0.04505   0.04001  -0.0298   0.0192   1.0000
   9.500   1.1098   0.05004   0.04546  -0.0284   0.0195   1.0000
   9.750   1.1015   0.05513   0.05092  -0.0273   0.0199   1.0000
  10.000   1.0869   0.06001   0.05610  -0.0264   0.0202   1.0000
  10.250   1.0676   0.06461   0.06090  -0.0262   0.0204   1.0000
  10.500   1.0475   0.07006   0.06653  -0.0287   0.0206   1.0000
  10.750   1.0273   0.07690   0.07352  -0.0337   0.0207   1.0000
  11.500   0.9572   0.11383   0.11082  -0.0616   0.0236   1.0000
  11.750   0.9248   0.13001   0.12689  -0.0698   0.0273   1.0000
  12.000   0.9186   0.13652   0.13337  -0.0722   0.0291   1.0000
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