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AG16 (ag16-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: AG16 (ag16-il)
Reynolds number: 100,000
Max Cl/Cd: 48.38 at α=4.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-ag16-il-100000-n5.txt
Download as CSV file: xf-ag16-il-100000-n5.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 =     0.100 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.5533   0.09086   0.08611   0.0013   1.0000   0.0231
  -8.000  -0.5515   0.08637   0.08168  -0.0014   1.0000   0.0229
  -7.750  -0.5495   0.08172   0.07709  -0.0047   1.0000   0.0225
  -7.500  -0.5440   0.07625   0.07166  -0.0100   1.0000   0.0221
  -7.250  -0.5356   0.06977   0.06519  -0.0172   1.0000   0.0217
  -7.000  -0.5236   0.06179   0.05715  -0.0261   1.0000   0.0212
  -6.750  -0.5099   0.05302   0.04818  -0.0346   1.0000   0.0207
  -6.500  -0.4941   0.04515   0.03990  -0.0409   1.0000   0.0203
  -6.250  -0.4748   0.03856   0.03273  -0.0450   1.0000   0.0201
  -6.000  -0.4524   0.03336   0.02687  -0.0471   1.0000   0.0202
  -5.750  -0.4280   0.02933   0.02220  -0.0481   1.0000   0.0205
  -5.500  -0.4025   0.02618   0.01848  -0.0485   1.0000   0.0212
  -5.250  -0.3764   0.02369   0.01547  -0.0484   1.0000   0.0223
  -5.000  -0.3502   0.02189   0.01331  -0.0481   1.0000   0.0250
  -4.750  -0.3250   0.02028   0.01160  -0.0480   1.0000   0.0286
  -4.500  -0.2989   0.01882   0.00997  -0.0476   1.0000   0.0317
  -4.250  -0.2729   0.01753   0.00852  -0.0472   1.0000   0.0368
  -4.000  -0.2467   0.01660   0.00750  -0.0469   1.0000   0.0479
  -3.750  -0.2205   0.01549   0.00644  -0.0468   1.0000   0.0682
  -3.500  -0.1942   0.01455   0.00563  -0.0467   1.0000   0.1105
  -3.250  -0.1683   0.01380   0.00516  -0.0467   1.0000   0.1738
  -3.000  -0.1425   0.01327   0.00483  -0.0465   1.0000   0.2397
  -2.750  -0.1168   0.01287   0.00456  -0.0463   1.0000   0.2959
  -2.500  -0.0912   0.01254   0.00435  -0.0459   1.0000   0.3502
  -2.250  -0.0659   0.01224   0.00415  -0.0455   1.0000   0.4043
  -2.000  -0.0406   0.01198   0.00402  -0.0450   1.0000   0.4578
  -1.750  -0.0154   0.01173   0.00392  -0.0446   1.0000   0.5128
  -1.500   0.0098   0.01149   0.00385  -0.0440   1.0000   0.5712
  -1.250   0.0343   0.01124   0.00377  -0.0433   1.0000   0.6374
  -1.000   0.0621   0.01091   0.00373  -0.0431   0.9951   0.7261
  -0.750   0.1023   0.01034   0.00356  -0.0447   0.9782   1.0000
  -0.500   0.1435   0.01040   0.00346  -0.0475   0.9616   1.0000
  -0.250   0.1833   0.01045   0.00336  -0.0499   0.9432   1.0000
   0.000   0.2204   0.01049   0.00330  -0.0516   0.9219   1.0000
   0.250   0.2559   0.01053   0.00325  -0.0528   0.8988   1.0000
   0.500   0.2880   0.01058   0.00321  -0.0533   0.8726   1.0000
   1.000   0.3457   0.01075   0.00317  -0.0526   0.8143   1.0000
   1.250   0.3726   0.01087   0.00318  -0.0518   0.7833   1.0000
   1.500   0.3990   0.01102   0.00323  -0.0510   0.7519   1.0000
   1.750   0.4254   0.01120   0.00329  -0.0502   0.7204   1.0000
   2.000   0.4518   0.01140   0.00337  -0.0495   0.6878   1.0000
   2.250   0.4780   0.01162   0.00348  -0.0488   0.6553   1.0000
   2.500   0.5041   0.01187   0.00362  -0.0481   0.6229   1.0000
   2.750   0.5302   0.01214   0.00383  -0.0475   0.5896   1.0000
   3.000   0.5561   0.01243   0.00402  -0.0469   0.5564   1.0000
   3.250   0.5822   0.01274   0.00424  -0.0463   0.5228   1.0000
   3.500   0.6081   0.01308   0.00449  -0.0458   0.4890   1.0000
   3.750   0.6339   0.01345   0.00482  -0.0452   0.4549   1.0000
   4.000   0.6597   0.01384   0.00514  -0.0447   0.4203   1.0000
   4.250   0.6853   0.01426   0.00549  -0.0443   0.3857   1.0000
   4.500   0.7108   0.01473   0.00588  -0.0438   0.3509   1.0000
   4.750   0.7363   0.01522   0.00632  -0.0434   0.3159   1.0000
   5.000   0.7614   0.01577   0.00685  -0.0430   0.2815   1.0000
   5.250   0.7865   0.01635   0.00739  -0.0426   0.2476   1.0000
   5.500   0.8113   0.01700   0.00798  -0.0422   0.2143   1.0000
   5.750   0.8359   0.01770   0.00863  -0.0418   0.1824   1.0000
   6.000   0.8603   0.01846   0.00937  -0.0414   0.1510   1.0000
   6.250   0.8841   0.01934   0.01019  -0.0410   0.1211   1.0000
   6.500   0.9075   0.02033   0.01122  -0.0405   0.0956   1.0000
   6.750   0.9302   0.02147   0.01234  -0.0400   0.0739   1.0000
   7.000   0.9524   0.02272   0.01360  -0.0394   0.0570   1.0000
   7.250   0.9737   0.02415   0.01513  -0.0386   0.0457   1.0000
   7.500   0.9942   0.02571   0.01682  -0.0377   0.0369   1.0000
   7.750   1.0125   0.02760   0.01875  -0.0368   0.0309   1.0000
   8.000   1.0328   0.02934   0.02080  -0.0357   0.0266   1.0000
   8.250   1.0515   0.03097   0.02255  -0.0350   0.0224   1.0000
   8.500   1.0671   0.03370   0.02558  -0.0337   0.0205   1.0000
   8.750   1.0838   0.03643   0.02870  -0.0324   0.0192   1.0000
   9.000   1.0980   0.03947   0.03215  -0.0311   0.0179   1.0000
   9.250   1.1099   0.04231   0.03538  -0.0300   0.0165   1.0000
   9.500   1.1198   0.04473   0.03808  -0.0291   0.0151   1.0000
   9.750   1.1257   0.04745   0.04101  -0.0282   0.0141   1.0000
  10.000   1.1229   0.05149   0.04536  -0.0272   0.0135   1.0000
  10.250   1.1137   0.05583   0.05001  -0.0261   0.0133   1.0000
  10.500   1.1011   0.05980   0.05428  -0.0255   0.0132   1.0000
  10.750   1.0859   0.06451   0.05925  -0.0267   0.0132   1.0000
  11.000   1.0696   0.07025   0.06522  -0.0300   0.0133   1.0000
  11.250   1.0528   0.07711   0.07227  -0.0350   0.0133   1.0000
  11.500   1.0375   0.08479   0.08018  -0.0411   0.0135   1.0000
  11.750   0.9673   0.11148   0.10724  -0.0598   0.0175   1.0000
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