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S6062 8% (s6062-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: S6062 8% (s6062-il)
Reynolds number: 200,000
Max Cl/Cd: 65.25 at α=4.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-s6062-il-200000.txt
Download as CSV file: xf-s6062-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: S6062 8%                                        
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.5256   0.09278   0.08940  -0.0169   1.0000   0.0441
  -8.500  -0.5288   0.08819   0.08486  -0.0221   1.0000   0.0443
  -8.250  -0.5342   0.08303   0.07969  -0.0279   1.0000   0.0444
  -8.000  -0.5361   0.07836   0.07493  -0.0319   1.0000   0.0445
  -7.750  -0.5328   0.07277   0.06951  -0.0292   1.0000   0.0460
  -7.500  -0.5249   0.06999   0.06675  -0.0287   1.0000   0.0472
  -7.250  -0.5188   0.06629   0.06303  -0.0304   1.0000   0.0486
  -7.000  -0.5126   0.06208   0.05875  -0.0328   1.0000   0.0504
  -6.750  -0.5053   0.05744   0.05399  -0.0356   1.0000   0.0529
  -6.500  -0.5031   0.05234   0.04815  -0.0394   1.0000   0.0574
  -6.250  -0.4914   0.04730   0.04330  -0.0392   1.0000   0.0591
  -6.000  -0.4776   0.04468   0.04069  -0.0385   1.0000   0.0611
  -5.750  -0.4649   0.04196   0.03783  -0.0378   1.0000   0.0646
  -5.500  -0.4563   0.03847   0.03385  -0.0368   1.0000   0.0720
  -5.000  -0.4234   0.02796   0.02189  -0.0310   1.0000   0.0331
  -4.750  -0.4089   0.02367   0.01733  -0.0294   1.0000   0.0294
  -4.500  -0.3902   0.02094   0.01415  -0.0275   1.0000   0.0273
  -4.250  -0.3697   0.01890   0.01174  -0.0258   1.0000   0.0264
  -4.000  -0.3490   0.01750   0.01014  -0.0244   1.0000   0.0270
  -3.750  -0.3285   0.01668   0.00922  -0.0231   1.0000   0.0294
  -3.500  -0.3074   0.01578   0.00821  -0.0219   1.0000   0.0312
  -3.250  -0.2729   0.01465   0.00697  -0.0232   0.9966   0.0324
  -3.000  -0.2352   0.01331   0.00560  -0.0254   0.9917   0.0372
  -2.750  -0.1972   0.01179   0.00451  -0.0277   0.9864   0.1274
  -2.500  -0.1624   0.01041   0.00431  -0.0303   0.9809   0.3936
  -2.250  -0.1332   0.00931   0.00437  -0.0306   0.9739   0.6771
  -2.000  -0.1029   0.00902   0.00452  -0.0299   0.9675   0.8301
  -1.750  -0.0657   0.00899   0.00454  -0.0307   0.9617   0.9118
  -1.500  -0.0082   0.00902   0.00445  -0.0360   0.9609   0.9603
  -1.250   0.0568   0.00897   0.00423  -0.0433   0.9613   0.9859
  -1.000   0.1180   0.00884   0.00397  -0.0502   0.9599   1.0000
  -0.750   0.1574   0.00873   0.00376  -0.0527   0.9499   1.0000
  -0.500   0.1932   0.00862   0.00357  -0.0544   0.9386   1.0000
  -0.250   0.2222   0.00854   0.00343  -0.0545   0.9251   1.0000
   0.000   0.2456   0.00850   0.00332  -0.0535   0.9101   1.0000
   0.250   0.2666   0.00849   0.00324  -0.0518   0.8948   1.0000
   0.500   0.2872   0.00848   0.00318  -0.0501   0.8797   1.0000
   0.750   0.3079   0.00849   0.00313  -0.0483   0.8645   1.0000
   1.000   0.3291   0.00851   0.00309  -0.0466   0.8493   1.0000
   1.250   0.3509   0.00853   0.00307  -0.0450   0.8337   1.0000
   1.500   0.3734   0.00855   0.00304  -0.0436   0.8179   1.0000
   1.750   0.3965   0.00859   0.00303  -0.0423   0.8015   1.0000
   2.000   0.4202   0.00865   0.00305  -0.0412   0.7833   1.0000
   2.250   0.4443   0.00870   0.00309  -0.0401   0.7650   1.0000
   2.500   0.4687   0.00877   0.00311  -0.0391   0.7467   1.0000
   2.750   0.4935   0.00886   0.00318  -0.0382   0.7254   1.0000
   3.000   0.5182   0.00896   0.00322  -0.0373   0.7044   1.0000
   3.250   0.5432   0.00907   0.00331  -0.0364   0.6797   1.0000
   3.500   0.5681   0.00921   0.00345  -0.0356   0.6535   1.0000
   3.750   0.5931   0.00937   0.00357  -0.0348   0.6257   1.0000
   4.000   0.6179   0.00957   0.00372  -0.0340   0.5952   1.0000
   4.250   0.6414   0.00983   0.00383  -0.0330   0.5471   1.0000
   4.500   0.6632   0.01025   0.00398  -0.0317   0.4731   1.0000
   4.750   0.6861   0.01074   0.00424  -0.0308   0.4123   1.0000
   5.000   0.7084   0.01135   0.00459  -0.0300   0.3430   1.0000
   5.250   0.7294   0.01219   0.00505  -0.0292   0.2573   1.0000
   5.500   0.7484   0.01347   0.00573  -0.0283   0.1430   1.0000
   5.750   0.7639   0.01559   0.00715  -0.0268   0.0430   1.0000
   6.000   0.7861   0.01655   0.00820  -0.0259   0.0365   1.0000
   6.250   0.8056   0.01793   0.00965  -0.0246   0.0333   1.0000
   6.500   0.8271   0.01905   0.01087  -0.0235   0.0320   1.0000
   6.750   0.8484   0.02037   0.01229  -0.0224   0.0309   1.0000
   7.000   0.8705   0.02192   0.01396  -0.0214   0.0303   1.0000
   7.250   0.8935   0.02352   0.01568  -0.0205   0.0293   1.0000
   7.500   0.9163   0.02494   0.01723  -0.0197   0.0276   1.0000
   7.750   0.9389   0.02694   0.01942  -0.0188   0.0270   1.0000
   8.000   0.9609   0.02948   0.02228  -0.0178   0.0272   1.0000
   8.250   0.9798   0.03288   0.02616  -0.0163   0.0281   1.0000
   8.500   0.9932   0.03728   0.03114  -0.0143   0.0299   1.0000
   8.750   1.0012   0.04235   0.03667  -0.0126   0.0320   1.0000
  11.000   0.8252   0.10932   0.10603  -0.0306   0.0642   1.0000
  11.250   0.8190   0.11552   0.11220  -0.0342   0.0623   1.0000
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