Airfoil Tools
Search 1636 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

NACA 0010-34 (naca001034-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: NACA 0010-34 (naca001034-il)
Reynolds number: 50,000
Max Cl/Cd: 20.77 at α=4.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-naca001034-il-50000.txt
Download as CSV file: xf-naca001034-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 0010-34                                    
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.5067   0.08470   0.07836  -0.0132   1.0000   0.2617
  -8.500  -0.5732   0.07264   0.06647  -0.0246   1.0000   0.2009
  -8.250  -0.6192   0.06536   0.05919  -0.0269   1.0000   0.1747
  -8.000  -0.7109   0.07464   0.06814  -0.0180   1.0000   0.2096
  -7.750  -0.7261   0.06616   0.05928  -0.0206   1.0000   0.1626
  -7.500  -0.7341   0.06039   0.05299  -0.0196   1.0000   0.1385
  -7.250  -0.7331   0.05552   0.04761  -0.0174   1.0000   0.1233
  -7.000  -0.7284   0.05137   0.04287  -0.0146   1.0000   0.1124
  -6.750  -0.7210   0.04752   0.03847  -0.0117   1.0000   0.1062
  -6.500  -0.7115   0.04506   0.03516  -0.0081   1.0000   0.1025
  -6.250  -0.6960   0.04196   0.03163  -0.0059   1.0000   0.1024
  -6.000  -0.6766   0.03880   0.02812  -0.0040   1.0000   0.1018
  -5.750  -0.6537   0.03577   0.02475  -0.0024   1.0000   0.1008
  -5.500  -0.6281   0.03316   0.02175  -0.0011   1.0000   0.1010
  -5.250  -0.5991   0.03085   0.01909  -0.0001   1.0000   0.1032
  -5.000  -0.5687   0.02859   0.01680   0.0002   1.0000   0.1147
  -4.750  -0.5367   0.02656   0.01481   0.0006   1.0000   0.1276
  -4.500  -0.5139   0.02474   0.01307   0.0022   1.0000   0.1459
  -4.250  -0.3028   0.02841   0.01900  -0.0059   1.0000   0.9411
  -4.000  -0.2270   0.02704   0.01700  -0.0166   1.0000   0.9717
  -3.750  -0.1509   0.02530   0.01468  -0.0283   1.0000   0.9988
  -3.500  -0.1347   0.02442   0.01364  -0.0285   1.0000   1.0000
  -3.250  -0.1213   0.02367   0.01276  -0.0279   1.0000   1.0000
  -3.000  -0.1080   0.02301   0.01198  -0.0272   1.0000   1.0000
  -2.750  -0.0949   0.02243   0.01130  -0.0262   1.0000   1.0000
  -2.500  -0.0822   0.02193   0.01067  -0.0250   1.0000   1.0000
  -2.250  -0.0702   0.02149   0.01016  -0.0235   1.0000   1.0000
  -2.000  -0.0589   0.02112   0.00973  -0.0218   1.0000   1.0000
  -1.750  -0.0486   0.02080   0.00937  -0.0198   1.0000   1.0000
  -1.500  -0.0393   0.02054   0.00907  -0.0176   1.0000   1.0000
  -1.250  -0.0308   0.02033   0.00883  -0.0151   1.0000   1.0000
  -1.000  -0.0232   0.02015   0.00865  -0.0124   1.0000   1.0000
  -0.750  -0.0164   0.02003   0.00849  -0.0095   1.0000   1.0000
  -0.500  -0.0105   0.01994   0.00839  -0.0065   1.0000   1.0000
  -0.250  -0.0051   0.01988   0.00834  -0.0033   1.0000   1.0000
   0.000   0.0000   0.01987   0.00832   0.0000   1.0000   1.0000
   0.250   0.0051   0.01988   0.00834   0.0033   1.0000   1.0000
   0.500   0.0105   0.01994   0.00839   0.0065   1.0000   1.0000
   0.750   0.0164   0.02002   0.00849   0.0095   1.0000   1.0000
   1.000   0.0232   0.02015   0.00864   0.0124   1.0000   1.0000
   1.250   0.0308   0.02032   0.00883   0.0151   1.0000   1.0000
   1.500   0.0394   0.02054   0.00907   0.0176   1.0000   1.0000
   1.750   0.0486   0.02080   0.00936   0.0198   1.0000   1.0000
   2.000   0.0589   0.02111   0.00972   0.0218   1.0000   1.0000
   2.250   0.0702   0.02149   0.01015   0.0235   1.0000   1.0000
   2.500   0.0823   0.02192   0.01066   0.0250   1.0000   1.0000
   2.750   0.0950   0.02242   0.01129   0.0262   1.0000   1.0000
   3.000   0.1082   0.02300   0.01197   0.0271   1.0000   1.0000
   3.250   0.1215   0.02366   0.01275   0.0279   1.0000   1.0000
   3.500   0.1349   0.02440   0.01362   0.0285   1.0000   1.0000
   3.750   0.1509   0.02528   0.01466   0.0283   0.9989   1.0000
   4.000   0.2270   0.02702   0.01698   0.0166   0.9717   1.0000
   4.250   0.3031   0.02840   0.01899   0.0058   0.9412   1.0000
   4.500   0.5139   0.02474   0.01307  -0.0022   0.1460   1.0000
   4.750   0.5367   0.02656   0.01480  -0.0006   0.1276   1.0000
   5.000   0.5687   0.02859   0.01680  -0.0002   0.1147   1.0000
   5.250   0.5991   0.03085   0.01909   0.0001   0.1032   1.0000
   5.500   0.6281   0.03316   0.02174   0.0011   0.1010   1.0000
   5.750   0.6537   0.03577   0.02475   0.0024   0.1008   1.0000
   6.000   0.6766   0.03880   0.02812   0.0040   0.1018   1.0000
   6.250   0.6960   0.04196   0.03164   0.0059   0.1024   1.0000
   6.500   0.7115   0.04507   0.03516   0.0081   0.1025   1.0000
   6.750   0.7209   0.04752   0.03847   0.0117   0.1063   1.0000
   7.000   0.7284   0.05137   0.04287   0.0146   0.1124   1.0000
   7.250   0.7331   0.05551   0.04761   0.0174   0.1233   1.0000
   7.500   0.7342   0.06039   0.05299   0.0196   0.1385   1.0000
   7.750   0.7261   0.06618   0.05930   0.0206   0.1627   1.0000
   8.000   0.6661   0.06047   0.05411   0.0254   0.1699   1.0000
   8.250   0.6185   0.06526   0.05909   0.0269   0.1748   1.0000
   8.500   0.5738   0.07250   0.06634   0.0247   0.2007   1.0000
   8.750   0.5065   0.08459   0.07825   0.0132   0.2618   1.0000
   9.000   0.5755   0.10446   0.09763  -0.0113   0.3786   1.0000
<< Back to NACA 0010-34 (naca001034-il)

Polar data table (+)

Polar graphs


<< Back to NACA 0010-34 (naca001034-il)