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S2055 (s2055-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: S2055 (s2055-il)
Reynolds number: 200,000
Max Cl/Cd: 70.13 at α=3.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-s2055-il-200000.txt
Download as CSV file: xf-s2055-il-200000.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 =     0.200 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.4823   0.08647   0.08302  -0.0336   1.0000   0.0431
  -8.250  -0.4935   0.08247   0.07908  -0.0373   1.0000   0.0432
  -8.000  -0.5013   0.07866   0.07524  -0.0396   1.0000   0.0434
  -7.750  -0.5079   0.07511   0.07160  -0.0405   1.0000   0.0436
  -7.500  -0.5122   0.07187   0.06825  -0.0403   1.0000   0.0437
  -7.250  -0.5190   0.06504   0.06153  -0.0396   1.0000   0.0448
  -7.000  -0.5148   0.06238   0.05893  -0.0376   1.0000   0.0458
  -6.750  -0.5121   0.05978   0.05633  -0.0360   1.0000   0.0470
  -6.500  -0.5098   0.05692   0.05341  -0.0347   1.0000   0.0485
  -6.250  -0.5068   0.05384   0.05024  -0.0335   1.0000   0.0505
  -6.000  -0.5012   0.05217   0.04795  -0.0323   1.0000   0.0561
  -5.750  -0.5015   0.04644   0.04205  -0.0311   1.0000   0.0574
  -5.500  -0.4921   0.04319   0.03890  -0.0297   1.0000   0.0589
  -5.250  -0.4814   0.04084   0.03650  -0.0281   1.0000   0.0610
  -5.000  -0.4692   0.03840   0.03387  -0.0265   1.0000   0.0647
  -4.750  -0.4570   0.03533   0.03030  -0.0251   1.0000   0.0719
  -4.500  -0.4297   0.02604   0.01978  -0.0217   1.0000   0.0311
  -4.250  -0.4094   0.02323   0.01653  -0.0203   1.0000   0.0307
  -4.000  -0.3876   0.02132   0.01416  -0.0191   1.0000   0.0329
  -3.750  -0.3638   0.01912   0.01165  -0.0182   1.0000   0.0335
  -3.500  -0.3407   0.01760   0.00998  -0.0172   1.0000   0.0347
  -3.250  -0.3139   0.01654   0.00884  -0.0170   0.9990   0.0380
  -3.000  -0.2775   0.01574   0.00793  -0.0187   0.9956   0.0447
  -2.750  -0.2426   0.01483   0.00701  -0.0203   0.9919   0.0547
  -2.500  -0.2080   0.01395   0.00619  -0.0219   0.9877   0.0735
  -2.250  -0.1728   0.01277   0.00562  -0.0241   0.9842   0.1933
  -2.000  -0.1472   0.01116   0.00554  -0.0247   0.9795   0.5533
  -1.750  -0.1297   0.01057   0.00605  -0.0209   0.9738   0.8690
  -1.500  -0.0755   0.01077   0.00621  -0.0249   0.9740   0.9682
  -1.250  -0.0018   0.01091   0.00613  -0.0342   0.9780   1.0000
  -1.000   0.0400   0.01097   0.00604  -0.0375   0.9726   1.0000
  -0.750   0.0792   0.01099   0.00595  -0.0402   0.9662   1.0000
  -0.500   0.1224   0.01097   0.00583  -0.0435   0.9597   1.0000
  -0.250   0.1687   0.01090   0.00567  -0.0473   0.9535   1.0000
   0.000   0.2171   0.01071   0.00543  -0.0514   0.9459   1.0000
   0.250   0.2707   0.01045   0.00512  -0.0563   0.9399   1.0000
   0.500   0.3128   0.01018   0.00484  -0.0590   0.9314   1.0000
   0.750   0.3471   0.00998   0.00463  -0.0601   0.9217   1.0000
   1.000   0.3848   0.00971   0.00438  -0.0618   0.9138   1.0000
   1.250   0.4134   0.00952   0.00419  -0.0617   0.9022   1.0000
   1.500   0.4394   0.00936   0.00404  -0.0610   0.8895   1.0000
   1.750   0.4649   0.00921   0.00390  -0.0601   0.8759   1.0000
   2.000   0.4896   0.00909   0.00378  -0.0591   0.8614   1.0000
   2.250   0.5141   0.00899   0.00370  -0.0580   0.8458   1.0000
   2.500   0.5359   0.00896   0.00368  -0.0564   0.8269   1.0000
   2.750   0.5587   0.00893   0.00365  -0.0549   0.8068   1.0000
   3.000   0.5809   0.00895   0.00365  -0.0533   0.7838   1.0000
   3.250   0.6030   0.00898   0.00368  -0.0517   0.7573   1.0000
   3.500   0.6242   0.00906   0.00371  -0.0499   0.7243   1.0000
   3.750   0.6445   0.00919   0.00372  -0.0479   0.6784   1.0000
   4.000   0.6626   0.00947   0.00374  -0.0455   0.6082   1.0000
   4.250   0.6784   0.01002   0.00387  -0.0429   0.5187   1.0000
   4.500   0.6941   0.01071   0.00418  -0.0405   0.4391   1.0000
   4.750   0.7122   0.01136   0.00454  -0.0388   0.3787   1.0000
   5.000   0.7320   0.01194   0.00494  -0.0374   0.3341   1.0000
   5.250   0.7526   0.01250   0.00535  -0.0362   0.2957   1.0000
   5.500   0.7735   0.01305   0.00580  -0.0352   0.2590   1.0000
   5.750   0.7947   0.01359   0.00622  -0.0342   0.2185   1.0000
   6.000   0.8153   0.01424   0.00669  -0.0331   0.1671   1.0000
   6.250   0.8328   0.01536   0.00742  -0.0316   0.1033   1.0000
   6.500   0.8500   0.01661   0.00849  -0.0299   0.0731   1.0000
   6.750   0.8674   0.01786   0.00966  -0.0283   0.0576   1.0000
   7.000   0.8864   0.01904   0.01095  -0.0268   0.0501   1.0000
   7.250   0.9031   0.02097   0.01283  -0.0251   0.0443   1.0000
   7.500   0.9255   0.02185   0.01388  -0.0241   0.0401   1.0000
   7.750   0.9469   0.02317   0.01529  -0.0230   0.0366   1.0000
   8.000   0.9679   0.02587   0.01804  -0.0222   0.0336   1.0000
   8.250   0.9887   0.02734   0.01980  -0.0210   0.0309   1.0000
   8.500   1.0089   0.02937   0.02214  -0.0197   0.0289   1.0000
   8.750   1.0270   0.03183   0.02495  -0.0182   0.0274   1.0000
   9.000   1.0427   0.03377   0.02709  -0.0168   0.0255   1.0000
   9.250   1.0493   0.03852   0.03214  -0.0151   0.0236   1.0000
   9.500   1.0526   0.04196   0.03605  -0.0124   0.0231   1.0000
   9.750   1.0541   0.04529   0.03980  -0.0096   0.0229   1.0000
  10.000   1.0492   0.04937   0.04426  -0.0068   0.0229   1.0000
  10.250   1.0386   0.05366   0.04888  -0.0038   0.0230   1.0000
  10.500   1.0323   0.05775   0.05317  -0.0015   0.0236   1.0000
  10.750   1.0198   0.06044   0.05608   0.0015   0.0238   1.0000
  11.000   1.0088   0.06266   0.05851   0.0037   0.0242   1.0000
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