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S2055 (s2055-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: S2055 (s2055-il)
Reynolds number: 100,000
Max Cl/Cd: 48.79 at α=4°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-s2055-il-100000-n5.txt
Download as CSV file: xf-s2055-il-100000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: S2055                                           
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.4775   0.08901   0.08403  -0.0274   1.0000   0.0240
  -8.500  -0.4831   0.08413   0.07923  -0.0299   1.0000   0.0229
  -8.250  -0.5003   0.07639   0.07162  -0.0369   1.0000   0.0209
  -8.000  -0.5033   0.07221   0.06740  -0.0387   1.0000   0.0204
  -7.750  -0.5084   0.06844   0.06363  -0.0389   1.0000   0.0201
  -7.500  -0.5127   0.06454   0.05970  -0.0391   1.0000   0.0198
  -7.250  -0.5155   0.06063   0.05573  -0.0390   1.0000   0.0196
  -7.000  -0.5175   0.05649   0.05147  -0.0385   1.0000   0.0194
  -6.750  -0.5173   0.05244   0.04726  -0.0376   1.0000   0.0192
  -6.500  -0.5149   0.04847   0.04308  -0.0363   1.0000   0.0191
  -6.250  -0.5105   0.04450   0.03883  -0.0348   1.0000   0.0190
  -6.000  -0.5037   0.04058   0.03455  -0.0332   1.0000   0.0190
  -5.750  -0.4938   0.03696   0.03049  -0.0314   1.0000   0.0193
  -5.250  -0.4672   0.03085   0.02341  -0.0281   1.0000   0.0217
  -5.000  -0.4505   0.02884   0.02112  -0.0268   1.0000   0.0230
  -4.750  -0.4316   0.02657   0.01843  -0.0253   1.0000   0.0237
  -4.500  -0.4112   0.02456   0.01606  -0.0239   1.0000   0.0247
  -4.250  -0.3897   0.02277   0.01393  -0.0226   1.0000   0.0263
  -4.000  -0.3678   0.02151   0.01238  -0.0213   1.0000   0.0291
  -3.750  -0.3472   0.02030   0.01113  -0.0203   1.0000   0.0341
  -3.500  -0.3171   0.01911   0.00978  -0.0208   0.9972   0.0387
  -3.250  -0.2846   0.01825   0.00882  -0.0220   0.9932   0.0478
  -3.000  -0.2530   0.01735   0.00781  -0.0229   0.9887   0.0569
  -2.750  -0.2200   0.01660   0.00705  -0.0242   0.9844   0.0765
  -2.500  -0.1888   0.01587   0.00657  -0.0253   0.9798   0.1300
  -2.250  -0.1603   0.01474   0.00628  -0.0263   0.9749   0.3027
  -2.000  -0.1365   0.01336   0.00631  -0.0255   0.9709   0.6425
  -1.750  -0.1122   0.01318   0.00658  -0.0232   0.9658   0.8371
  -1.500  -0.0711   0.01329   0.00660  -0.0249   0.9629   0.9275
  -1.250   0.0094   0.01337   0.00641  -0.0352   0.9695   0.9948
  -1.000   0.0434   0.01338   0.00625  -0.0369   0.9622   1.0000
  -0.750   0.0810   0.01343   0.00615  -0.0393   0.9567   1.0000
  -0.500   0.1111   0.01344   0.00604  -0.0401   0.9477   1.0000
  -0.250   0.1510   0.01346   0.00594  -0.0428   0.9424   1.0000
   0.250   0.2171   0.01339   0.00574  -0.0452   0.9232   1.0000
   0.500   0.2576   0.01325   0.00556  -0.0476   0.9144   1.0000
   0.750   0.2931   0.01307   0.00535  -0.0489   0.9017   1.0000
   1.000   0.3261   0.01289   0.00515  -0.0496   0.8873   1.0000
   1.250   0.3554   0.01275   0.00501  -0.0495   0.8718   1.0000
   1.500   0.3843   0.01265   0.00492  -0.0494   0.8564   1.0000
   1.750   0.4124   0.01257   0.00484  -0.0491   0.8406   1.0000
   2.000   0.4402   0.01251   0.00479  -0.0487   0.8240   1.0000
   2.250   0.4682   0.01246   0.00478  -0.0484   0.8067   1.0000
   2.500   0.4937   0.01246   0.00481  -0.0476   0.7868   1.0000
   2.750   0.5199   0.01246   0.00482  -0.0468   0.7650   1.0000
   3.000   0.5444   0.01250   0.00488  -0.0458   0.7396   1.0000
   3.250   0.5691   0.01255   0.00497  -0.0447   0.7099   1.0000
   3.500   0.5929   0.01266   0.00506  -0.0435   0.6736   1.0000
   3.750   0.6164   0.01281   0.00514  -0.0421   0.6272   1.0000
   4.000   0.6387   0.01309   0.00524  -0.0406   0.5692   1.0000
   4.250   0.6590   0.01354   0.00542  -0.0388   0.5023   1.0000
   4.500   0.6780   0.01415   0.00577  -0.0370   0.4369   1.0000
   4.750   0.6967   0.01482   0.00617  -0.0354   0.3789   1.0000
   5.000   0.7162   0.01548   0.00662  -0.0340   0.3297   1.0000
   5.250   0.7365   0.01610   0.00710  -0.0328   0.2897   1.0000
   5.500   0.7572   0.01672   0.00762  -0.0317   0.2543   1.0000
   5.750   0.7782   0.01733   0.00817  -0.0306   0.2193   1.0000
   6.000   0.7988   0.01801   0.00877  -0.0296   0.1831   1.0000
   6.250   0.8188   0.01881   0.00949  -0.0285   0.1361   1.0000
   6.500   0.8375   0.01986   0.01030  -0.0273   0.0960   1.0000
   6.750   0.8559   0.02101   0.01134  -0.0260   0.0735   1.0000
   7.000   0.8742   0.02216   0.01251  -0.0246   0.0600   1.0000
   7.250   0.8921   0.02335   0.01373  -0.0232   0.0501   1.0000
   7.500   0.9100   0.02455   0.01505  -0.0218   0.0442   1.0000
   7.750   0.9259   0.02611   0.01668  -0.0201   0.0407   1.0000
   8.000   0.9445   0.02761   0.01841  -0.0187   0.0377   1.0000
   8.250   0.9632   0.02914   0.02010  -0.0174   0.0345   1.0000
   8.500   0.9790   0.03091   0.02197  -0.0162   0.0309   1.0000
   8.750   0.9979   0.03292   0.02429  -0.0150   0.0289   1.0000
   9.000   1.0155   0.03532   0.02706  -0.0136   0.0272   1.0000
   9.250   1.0298   0.03745   0.02953  -0.0121   0.0248   1.0000
   9.500   1.0415   0.03908   0.03133  -0.0106   0.0225   1.0000
   9.750   1.0476   0.04230   0.03475  -0.0089   0.0209   1.0000
  10.000   1.0514   0.04544   0.03831  -0.0065   0.0203   1.0000
  10.250   1.0492   0.04878   0.04213  -0.0039   0.0198   1.0000
  10.500   1.0409   0.05201   0.04573  -0.0009   0.0195   1.0000
  10.750   1.0276   0.05536   0.04941   0.0019   0.0192   1.0000
  11.000   1.0119   0.05901   0.05334   0.0037   0.0190   1.0000
  11.250   0.9937   0.06316   0.05776   0.0043   0.0188   1.0000
  11.500   0.9746   0.06799   0.06281   0.0033   0.0189   1.0000
  11.750   0.9544   0.07371   0.06871   0.0007   0.0191   1.0000
  12.000   0.9332   0.08065   0.07581  -0.0037   0.0193   1.0000
  12.250   0.9141   0.08849   0.08375  -0.0092   0.0198   1.0000
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