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NACA 0010-34 a=0.8 c(li)=0.2 (naca001034a08cli02-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: NACA 0010-34 a=0.8 c(li)=0.2 (naca001034a08cli02-il)
Reynolds number: 200,000
Max Cl/Cd: 62.82 at α=3.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-naca001034a08cli02-il-200000-n5.txt
Download as CSV file: xf-naca001034a08cli02-il-200000-n5.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 =     0.200 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.4064   0.07997   0.07672  -0.0414   1.0000   0.0156
  -9.250  -0.4153   0.07530   0.07210  -0.0423   1.0000   0.0150
  -9.000  -0.4267   0.07038   0.06722  -0.0434   1.0000   0.0146
  -8.750  -0.4412   0.06524   0.06214  -0.0449   1.0000   0.0143
  -8.500  -0.4619   0.06005   0.05697  -0.0473   1.0000   0.0145
  -8.000  -0.5150   0.05478   0.05169  -0.0425   1.0000   0.0150
  -7.500  -0.6002   0.06080   0.05724  -0.0362   0.9993   0.0157
  -7.250  -0.5916   0.05478   0.05096  -0.0385   0.9954   0.0137
  -7.000  -0.5809   0.04580   0.04132  -0.0385   0.9904   0.0103
  -6.750  -0.5568   0.04289   0.03806  -0.0389   0.9876   0.0098
  -6.250  -0.5198   0.03561   0.03007  -0.0388   0.9803   0.0092
  -6.000  -0.4971   0.03188   0.02587  -0.0390   0.9778   0.0091
  -5.750  -0.4764   0.02896   0.02252  -0.0381   0.9739   0.0091
  -5.500  -0.4513   0.02644   0.01957  -0.0378   0.9708   0.0093
  -5.250  -0.4225   0.02445   0.01712  -0.0381   0.9686   0.0095
  -5.000  -0.3955   0.02137   0.01373  -0.0386   0.9671   0.0114
  -4.750  -0.3700   0.02016   0.01235  -0.0383   0.9635   0.0127
  -4.500  -0.3427   0.01872   0.01076  -0.0381   0.9604   0.0133
  -4.250  -0.3137   0.01746   0.00937  -0.0384   0.9580   0.0144
  -4.000  -0.2836   0.01640   0.00820  -0.0390   0.9560   0.0159
  -3.750  -0.2515   0.01566   0.00735  -0.0402   0.9544   0.0192
  -3.500  -0.2303   0.01485   0.00637  -0.0390   0.9488   0.0227
  -3.250  -0.2003   0.01436   0.00573  -0.0396   0.9457   0.0298
  -3.000  -0.1712   0.01332   0.00507  -0.0404   0.9432   0.1196
  -2.750  -0.1504   0.01134   0.00467  -0.0404   0.9409   0.4904
  -2.500  -0.1365   0.01059   0.00477  -0.0372   0.9346   0.6911
  -2.250  -0.1113   0.01035   0.00480  -0.0359   0.9313   0.7921
  -2.000  -0.0794   0.01027   0.00467  -0.0366   0.9289   0.8192
  -1.750  -0.0512   0.01024   0.00460  -0.0365   0.9247   0.8421
  -1.500  -0.0220   0.01022   0.00455  -0.0366   0.9202   0.8629
  -1.250   0.0117   0.01020   0.00445  -0.0377   0.9171   0.8805
  -1.000   0.0474   0.01016   0.00435  -0.0393   0.9147   0.8954
  -0.750   0.0764   0.01020   0.00435  -0.0395   0.9085   0.9103
  -0.500   0.1115   0.01017   0.00428  -0.0410   0.9045   0.9222
  -0.250   0.1489   0.01012   0.00418  -0.0430   0.9013   0.9323
   0.000   0.1807   0.01014   0.00418  -0.0439   0.8945   0.9436
   0.250   0.2156   0.01008   0.00411  -0.0454   0.8895   0.9538
   0.500   0.2528   0.01006   0.00409  -0.0475   0.8838   0.9614
   0.750   0.2889   0.01002   0.00406  -0.0494   0.8775   0.9691
   1.000   0.3257   0.00998   0.00403  -0.0514   0.8715   0.9760
   1.250   0.3635   0.00993   0.00403  -0.0537   0.8641   0.9814
   1.500   0.3996   0.00988   0.00403  -0.0557   0.8561   0.9874
   1.750   0.4376   0.00978   0.00398  -0.0580   0.8481   0.9920
   2.000   0.4738   0.00973   0.00403  -0.0600   0.8371   0.9971
   2.250   0.5056   0.00967   0.00403  -0.0610   0.8244   1.0000
   2.500   0.5294   0.00961   0.00401  -0.0601   0.8094   1.0000
   2.750   0.5532   0.00957   0.00405  -0.0592   0.7925   1.0000
   3.000   0.5767   0.00954   0.00407  -0.0582   0.7702   1.0000
   3.250   0.5993   0.00954   0.00389  -0.0566   0.7121   1.0000
   3.500   0.6161   0.00994   0.00382  -0.0539   0.6087   1.0000
   3.750   0.6248   0.01084   0.00402  -0.0500   0.4718   1.0000
   4.000   0.6342   0.01181   0.00441  -0.0467   0.3502   1.0000
   4.250   0.6396   0.01327   0.00505  -0.0430   0.1764   1.0000
   4.500   0.6475   0.01476   0.00582  -0.0399   0.0471   1.0000
   4.750   0.6645   0.01546   0.00649  -0.0378   0.0293   1.0000
   5.000   0.6824   0.01610   0.00723  -0.0360   0.0225   1.0000
   5.250   0.6995   0.01685   0.00810  -0.0339   0.0190   1.0000
   5.500   0.7115   0.01813   0.00947  -0.0311   0.0156   1.0000
   5.750   0.7303   0.01881   0.01025  -0.0294   0.0139   1.0000
   6.000   0.7469   0.01994   0.01150  -0.0273   0.0127   1.0000
   6.250   0.7651   0.02128   0.01295  -0.0255   0.0117   1.0000
   6.500   0.7857   0.02285   0.01471  -0.0241   0.0110   1.0000
   6.750   0.8067   0.02413   0.01613  -0.0230   0.0098   1.0000
   7.000   0.8258   0.02575   0.01784  -0.0219   0.0084   1.0000
   7.250   0.8457   0.02906   0.02150  -0.0206   0.0077   1.0000
   7.500   0.8641   0.03125   0.02403  -0.0189   0.0075   1.0000
   7.750   0.8795   0.03382   0.02698  -0.0168   0.0074   1.0000
   8.000   0.8913   0.03670   0.03026  -0.0142   0.0073   1.0000
   8.250   0.8993   0.03981   0.03376  -0.0114   0.0073   1.0000
   8.500   0.9037   0.04312   0.03744  -0.0083   0.0073   1.0000
   8.750   0.9042   0.04655   0.04122  -0.0049   0.0073   1.0000
   9.000   0.9009   0.05004   0.04502  -0.0015   0.0073   1.0000
   9.250   0.8937   0.05353   0.04879   0.0019   0.0074   1.0000
   9.500   0.8813   0.05674   0.05222   0.0058   0.0074   1.0000
   9.750   0.8663   0.06006   0.05573   0.0091   0.0075   1.0000
  10.000   0.8498   0.06368   0.05953   0.0114   0.0075   1.0000
  10.250   0.8322   0.06773   0.06374   0.0125   0.0076   1.0000
  10.500   0.8139   0.07235   0.06851   0.0123   0.0077   1.0000
  10.750   0.7953   0.07775   0.07403   0.0105   0.0077   1.0000
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