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NACA 0010-34 a=0.8 c(li)=0.2 (naca001034a08cli02-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: NACA 0010-34 a=0.8 c(li)=0.2 (naca001034a08cli02-il)
Reynolds number: 50,000
Max Cl/Cd: 26.49 at α=5.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-naca001034a08cli02-il-50000.txt
Download as CSV file: xf-naca001034a08cli02-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 0010-34 a=0.8 c(li)=0.2                    
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.500  -0.4998   0.12648   0.11950  -0.0164   1.0000   0.2251
 -10.250  -0.5164   0.12532   0.11846  -0.0181   1.0000   0.2346
 -10.000  -0.5073   0.12126   0.11442  -0.0173   1.0000   0.2480
  -9.750  -0.5008   0.11748   0.11058  -0.0167   1.0000   0.2614
  -9.500  -0.4976   0.11409   0.10724  -0.0163   1.0000   0.2752
  -9.250  -0.5000   0.11126   0.10448  -0.0159   1.0000   0.2897
  -9.000  -0.4805   0.10660   0.09981  -0.0140   1.0000   0.3098
  -8.750  -0.4830   0.10411   0.09739  -0.0130   1.0000   0.3299
  -8.500  -0.3775   0.09138   0.08526  -0.0177   1.0000   0.3600
  -8.250  -0.3772   0.08796   0.08190  -0.0163   1.0000   0.3739
  -8.000  -0.3616   0.08334   0.07727  -0.0155   1.0000   0.3828
  -7.750  -0.3640   0.07985   0.07384  -0.0142   1.0000   0.3926
  -6.000  -0.5979   0.05674   0.04990  -0.0260   1.0000   0.1754
  -5.750  -0.5871   0.05103   0.04332  -0.0259   1.0000   0.1356
  -5.500  -0.5727   0.04691   0.03865  -0.0243   1.0000   0.1210
  -5.250  -0.5566   0.04371   0.03458  -0.0221   1.0000   0.1100
  -5.000  -0.5379   0.04017   0.03079  -0.0207   1.0000   0.1067
  -4.750  -0.5166   0.03705   0.02709  -0.0190   1.0000   0.1009
  -4.500  -0.4925   0.03490   0.02417  -0.0171   1.0000   0.0964
  -4.250  -0.4671   0.03229   0.02123  -0.0160   1.0000   0.0954
  -4.000  -0.4402   0.03017   0.01874  -0.0147   1.0000   0.0961
  -3.750  -0.1416   0.02356   0.01482  -0.0434   1.0000   1.0000
  -3.500  -0.1426   0.02285   0.01391  -0.0407   1.0000   1.0000
  -3.250  -0.1442   0.02224   0.01312  -0.0377   1.0000   1.0000
  -3.000  -0.1459   0.02172   0.01242  -0.0343   1.0000   1.0000
  -2.750  -0.1475   0.02127   0.01180  -0.0307   1.0000   1.0000
  -2.500  -0.1485   0.02088   0.01125  -0.0270   1.0000   1.0000
  -2.250  -0.1484   0.02056   0.01076  -0.0233   1.0000   1.0000
  -2.000  -0.1463   0.02031   0.01027  -0.0198   1.0000   1.0000
  -1.750  -0.1417   0.02013   0.00989  -0.0166   1.0000   1.0000
  -1.500  -0.1343   0.02003   0.00958  -0.0138   1.0000   1.0000
  -1.250  -0.1248   0.01999   0.00933  -0.0114   1.0000   1.0000
  -1.000  -0.1131   0.02001   0.00916  -0.0093   1.0000   1.0000
  -0.750  -0.0995   0.02008   0.00904  -0.0075   1.0000   1.0000
  -0.500  -0.0844   0.02021   0.00896  -0.0060   1.0000   1.0000
  -0.250  -0.0682   0.02038   0.00898  -0.0047   1.0000   1.0000
   0.000  -0.0512   0.02059   0.00905  -0.0036   1.0000   1.0000
   0.250  -0.0336   0.02084   0.00919  -0.0026   1.0000   1.0000
   0.500  -0.0155   0.02113   0.00938  -0.0018   1.0000   1.0000
   0.750   0.0028   0.02146   0.00960  -0.0010   1.0000   1.0000
   1.000   0.0213   0.02182   0.00991  -0.0004   1.0000   1.0000
   1.250   0.0398   0.02223   0.01027   0.0002   1.0000   1.0000
   1.500   0.0584   0.02268   0.01070   0.0007   1.0000   1.0000
   1.750   0.0770   0.02318   0.01119   0.0012   1.0000   1.0000
   2.000   0.0954   0.02372   0.01174   0.0015   1.0000   1.0000
   2.250   0.1137   0.02431   0.01235   0.0018   1.0000   1.0000
   2.500   0.1317   0.02496   0.01304   0.0021   1.0000   1.0000
   2.750   0.1495   0.02567   0.01380   0.0023   1.0000   1.0000
   3.000   0.1669   0.02644   0.01467   0.0024   1.0000   1.0000
   3.250   0.1840   0.02728   0.01559   0.0024   1.0000   1.0000
   3.500   0.2006   0.02819   0.01661   0.0024   1.0000   1.0000
   3.750   0.2372   0.02957   0.01816  -0.0016   0.9911   1.0000
   4.000   0.2940   0.03126   0.02019  -0.0094   0.9702   1.0000
   4.250   0.3444   0.03268   0.02192  -0.0155   0.9478   1.0000
   4.500   0.3958   0.03395   0.02357  -0.0213   0.9246   1.0000
   4.750   0.4483   0.03499   0.02511  -0.0268   0.9002   1.0000
   5.000   0.4983   0.03570   0.02632  -0.0312   0.8740   1.0000
   5.250   0.5416   0.03612   0.02729  -0.0339   0.8458   1.0000
   5.500   0.6659   0.02514   0.01410  -0.0192   0.1829   1.0000
   5.750   0.6775   0.02770   0.01612  -0.0163   0.1333   1.0000
   6.000   0.7096   0.03013   0.01849  -0.0159   0.1098   1.0000
   6.250   0.7590   0.03324   0.02165  -0.0178   0.0967   1.0000
   6.500   0.7937   0.03638   0.02495  -0.0181   0.0875   1.0000
   6.750   0.8241   0.03966   0.02870  -0.0173   0.0868   1.0000
   7.000   0.8489   0.04328   0.03290  -0.0159   0.0879   1.0000
   7.250   0.8707   0.04747   0.03741  -0.0145   0.0899   1.0000
   7.500   0.8836   0.05045   0.04112  -0.0114   0.0928   1.0000
   7.750   0.8891   0.05424   0.04565  -0.0079   0.0976   1.0000
   8.000   0.8958   0.05850   0.05030  -0.0054   0.1009   1.0000
   8.250   0.8999   0.06262   0.05491  -0.0027   0.1072   1.0000
   8.500   0.8982   0.06779   0.06045  -0.0004   0.1156   1.0000
   8.750   0.8983   0.07350   0.06658   0.0012   0.1322   1.0000
   9.000   0.7763   0.06464   0.05839   0.0132   0.1262   1.0000
   9.250   0.7577   0.06945   0.06332   0.0151   0.1338   1.0000
   9.500   0.7160   0.07424   0.06820   0.0160   0.1359   1.0000
   9.750   0.6814   0.08019   0.07405   0.0145   0.1379   1.0000
  10.000   0.6600   0.08726   0.08114   0.0116   0.1472   1.0000
  10.250   0.6229   0.09556   0.08935   0.0059   0.1508   1.0000
  10.500   0.5948   0.10456   0.09828   0.0000   0.1776   1.0000
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