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AG17 (ag17-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: AG17 (ag17-il)
Reynolds number: 200,000
Max Cl/Cd: 64.09 at α=4°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-ag17-il-200000.txt
Download as CSV file: xf-ag17-il-200000.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 =     0.200 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.750  -0.4488   0.12584   0.12243   0.0106   1.0000   0.0308
 -10.500  -0.4467   0.12242   0.11903   0.0097   1.0000   0.0318
  -8.250  -0.4507   0.08696   0.08380  -0.0073   1.0000   0.0372
  -8.000  -0.4438   0.08203   0.07888  -0.0046   1.0000   0.0387
  -7.750  -0.4379   0.07856   0.07542  -0.0043   1.0000   0.0400
  -7.500  -0.4352   0.07482   0.07170  -0.0051   1.0000   0.0411
  -7.250  -0.4343   0.07089   0.06781  -0.0065   1.0000   0.0425
  -7.000  -0.4347   0.06674   0.06370  -0.0086   1.0000   0.0437
  -6.750  -0.4324   0.06176   0.05875  -0.0130   1.0000   0.0454
  -6.500  -0.4254   0.05499   0.05196  -0.0230   1.0000   0.0479
  -6.250  -0.4107   0.04647   0.04307  -0.0372   1.0000   0.0493
  -4.750  -0.3131   0.02593   0.02028  -0.0503   1.0000   0.0351
  -4.500  -0.2858   0.02149   0.01525  -0.0506   1.0000   0.0311
  -4.250  -0.2574   0.01839   0.01155  -0.0502   1.0000   0.0292
  -4.000  -0.2302   0.01645   0.00932  -0.0497   1.0000   0.0296
  -3.750  -0.2034   0.01499   0.00765  -0.0491   1.0000   0.0317
  -3.500  -0.1767   0.01421   0.00673  -0.0485   1.0000   0.0366
  -3.250  -0.1503   0.01269   0.00524  -0.0482   1.0000   0.0451
  -3.000  -0.1231   0.01161   0.00418  -0.0480   1.0000   0.0658
  -2.750  -0.0957   0.01045   0.00342  -0.0481   1.0000   0.1475
  -2.500  -0.0691   0.00973   0.00318  -0.0483   1.0000   0.2652
  -2.250  -0.0428   0.00928   0.00303  -0.0482   1.0000   0.3637
  -2.000  -0.0167   0.00892   0.00294  -0.0480   1.0000   0.4505
  -1.750   0.0092   0.00857   0.00288  -0.0477   1.0000   0.5361
  -1.500   0.0343   0.00817   0.00284  -0.0471   1.0000   0.6371
  -1.250   0.0541   0.00753   0.00279  -0.0448   1.0000   0.8064
  -1.000   0.0791   0.00732   0.00269  -0.0440   1.0000   1.0000
  -0.750   0.1057   0.00743   0.00269  -0.0441   1.0000   1.0000
  -0.500   0.1400   0.00755   0.00270  -0.0458   0.9959   1.0000
  -0.250   0.1886   0.00759   0.00263  -0.0502   0.9835   1.0000
   0.000   0.2351   0.00758   0.00256  -0.0539   0.9699   1.0000
   0.250   0.2766   0.00756   0.00249  -0.0565   0.9521   1.0000
   0.500   0.3132   0.00753   0.00242  -0.0579   0.9301   1.0000
   0.750   0.3443   0.00753   0.00236  -0.0579   0.9042   1.0000
   1.000   0.3713   0.00757   0.00232  -0.0570   0.8755   1.0000
   1.250   0.3964   0.00765   0.00230  -0.0557   0.8450   1.0000
   1.500   0.4212   0.00776   0.00230  -0.0545   0.8120   1.0000
   1.750   0.4462   0.00792   0.00234  -0.0533   0.7784   1.0000
   2.000   0.4714   0.00810   0.00238  -0.0523   0.7432   1.0000
   2.250   0.4967   0.00832   0.00245  -0.0513   0.7072   1.0000
   2.500   0.5224   0.00855   0.00254  -0.0505   0.6690   1.0000
   2.750   0.5481   0.00881   0.00268  -0.0498   0.6295   1.0000
   3.000   0.5738   0.00910   0.00282  -0.0491   0.5884   1.0000
   3.250   0.5996   0.00942   0.00298  -0.0485   0.5463   1.0000
   3.500   0.6253   0.00977   0.00317  -0.0480   0.5035   1.0000
   3.750   0.6510   0.01016   0.00339  -0.0475   0.4604   1.0000
   4.000   0.6768   0.01056   0.00369  -0.0470   0.4176   1.0000
   4.250   0.7025   0.01100   0.00398  -0.0466   0.3757   1.0000
   4.500   0.7281   0.01147   0.00432  -0.0463   0.3357   1.0000
   4.750   0.7538   0.01196   0.00469  -0.0459   0.2971   1.0000
   5.000   0.7794   0.01247   0.00510  -0.0456   0.2597   1.0000
   5.250   0.8048   0.01303   0.00559  -0.0453   0.2244   1.0000
   5.500   0.8300   0.01364   0.00610  -0.0450   0.1898   1.0000
   5.750   0.8550   0.01433   0.00669  -0.0446   0.1579   1.0000
   6.000   0.8798   0.01507   0.00736  -0.0443   0.1268   1.0000
   6.250   0.9040   0.01595   0.00816  -0.0438   0.1004   1.0000
   6.500   0.9276   0.01698   0.00913  -0.0433   0.0781   1.0000
   6.750   0.9512   0.01806   0.01026  -0.0426   0.0605   1.0000
   7.000   0.9737   0.01936   0.01158  -0.0419   0.0471   1.0000
   7.250   0.9962   0.02067   0.01297  -0.0411   0.0367   1.0000
   7.500   1.0162   0.02293   0.01536  -0.0399   0.0308   1.0000
   7.750   1.0386   0.02420   0.01674  -0.0391   0.0255   1.0000
   8.000   1.0564   0.02732   0.02009  -0.0378   0.0222   1.0000
   8.250   1.0769   0.02971   0.02285  -0.0366   0.0207   1.0000
   8.500   1.0946   0.03280   0.02637  -0.0352   0.0197   1.0000
   8.750   1.1086   0.03654   0.03060  -0.0337   0.0193   1.0000
   9.000   1.1172   0.04104   0.03566  -0.0322   0.0193   1.0000
   9.250   1.1190   0.04620   0.04143  -0.0306   0.0196   1.0000
   9.500   1.1141   0.05165   0.04736  -0.0293   0.0201   1.0000
   9.750   1.1031   0.05701   0.05310  -0.0284   0.0205   1.0000
  10.000   1.0862   0.06182   0.05818  -0.0278   0.0208   1.0000
  10.250   1.0662   0.06673   0.06329  -0.0290   0.0210   1.0000
  10.500   1.0461   0.07299   0.06967  -0.0334   0.0212   1.0000
  10.750   1.0262   0.08117   0.07800  -0.0408   0.0213   1.0000
  11.000   1.0056   0.09158   0.08852  -0.0498   0.0215   1.0000
  11.250   0.9837   0.10319   0.10017  -0.0579   0.0219   1.0000
  11.500   0.9645   0.11325   0.11019  -0.0631   0.0226   1.0000
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