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NACA M10 AIRFOIL (m10-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: NACA M10 AIRFOIL (m10-il)
Reynolds number: 50,000
Max Cl/Cd: 34.48 at α=4.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-m10-il-50000-n5.txt
Download as CSV file: xf-m10-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA M10 AIRFOIL                                
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.5718   0.10853   0.10186   0.0064   1.0000   0.0961
  -8.250  -0.5784   0.10573   0.09915   0.0003   1.0000   0.0973
  -8.000  -0.5779   0.10209   0.09552  -0.0061   1.0000   0.0977
  -7.750  -0.5649   0.09622   0.08971  -0.0023   1.0000   0.0995
  -7.500  -0.5564   0.09188   0.08539  -0.0029   1.0000   0.1013
  -7.250  -0.5490   0.08759   0.08110  -0.0055   1.0000   0.1022
  -7.000  -0.5327   0.07817   0.07129  -0.0189   1.0000   0.0550
  -6.750  -0.5217   0.07342   0.06651  -0.0198   1.0000   0.0545
  -6.500  -0.5085   0.06855   0.06144  -0.0227   1.0000   0.0548
  -6.000  -0.4778   0.05897   0.05134  -0.0263   1.0000   0.0551
  -5.750  -0.4611   0.05435   0.04652  -0.0270   1.0000   0.0547
  -5.500  -0.4420   0.04980   0.04152  -0.0279   1.0000   0.0548
  -5.250  -0.4259   0.04703   0.03887  -0.0278   1.0000   0.0598
  -5.000  -0.4043   0.04347   0.03493  -0.0283   1.0000   0.0632
  -4.750  -0.3809   0.03960   0.03054  -0.0287   1.0000   0.0640
  -4.500  -0.3560   0.03621   0.02647  -0.0287   1.0000   0.0671
  -4.250  -0.3326   0.03371   0.02365  -0.0286   1.0000   0.0729
  -4.000  -0.3072   0.03114   0.02063  -0.0283   1.0000   0.0745
  -3.750  -0.2811   0.02887   0.01787  -0.0279   1.0000   0.0763
  -3.500  -0.2547   0.02694   0.01550  -0.0273   1.0000   0.0787
  -3.250  -0.2281   0.02537   0.01348  -0.0267   1.0000   0.0823
  -3.000  -0.2036   0.02419   0.01225  -0.0263   1.0000   0.0907
  -2.750  -0.1774   0.02302   0.01079  -0.0257   1.0000   0.0953
  -2.500  -0.1506   0.02195   0.00949  -0.0250   1.0000   0.0979
  -2.250  -0.1250   0.02098   0.00847  -0.0244   1.0000   0.1018
  -2.000  -0.1004   0.02021   0.00760  -0.0237   1.0000   0.1074
  -1.750  -0.0766   0.01952   0.00681  -0.0229   1.0000   0.1153
  -1.500  -0.0531   0.01886   0.00617  -0.0222   1.0000   0.1299
  -1.250  -0.0299   0.01808   0.00562  -0.0217   1.0000   0.1691
  -1.000   0.0122   0.01497   0.00518  -0.0238   1.0000   1.0000
  -0.750   0.0321   0.01507   0.00497  -0.0226   1.0000   1.0000
  -0.500   0.0514   0.01520   0.00489  -0.0214   1.0000   1.0000
  -0.250   0.0704   0.01539   0.00491  -0.0203   1.0000   1.0000
   0.000   0.0893   0.01562   0.00500  -0.0193   1.0000   1.0000
   0.250   0.1142   0.01589   0.00514  -0.0196   0.9962   1.0000
   0.500   0.1590   0.01611   0.00524  -0.0237   0.9797   1.0000
   0.750   0.2042   0.01632   0.00538  -0.0277   0.9639   1.0000
   1.000   0.2478   0.01649   0.00553  -0.0312   0.9469   1.0000
   1.250   0.2896   0.01665   0.00571  -0.0343   0.9293   1.0000
   1.500   0.3302   0.01679   0.00590  -0.0369   0.9116   1.0000
   1.750   0.3657   0.01696   0.00612  -0.0383   0.8913   1.0000
   2.000   0.4001   0.01711   0.00635  -0.0394   0.8715   1.0000
   2.250   0.4311   0.01730   0.00665  -0.0397   0.8507   1.0000
   2.500   0.4604   0.01749   0.00693  -0.0395   0.8299   1.0000
   2.750   0.4875   0.01771   0.00726  -0.0389   0.8086   1.0000
   3.000   0.5140   0.01794   0.00765  -0.0381   0.7875   1.0000
   3.250   0.5393   0.01820   0.00805  -0.0371   0.7655   1.0000
   3.500   0.5645   0.01845   0.00844  -0.0359   0.7438   1.0000
   3.750   0.5889   0.01873   0.00895  -0.0347   0.7200   1.0000
   4.000   0.6131   0.01899   0.00941  -0.0333   0.6954   1.0000
   4.250   0.6345   0.01908   0.00958  -0.0306   0.6563   1.0000
   4.500   0.6496   0.01898   0.00922  -0.0261   0.5693   1.0000
   4.750   0.6665   0.01933   0.00924  -0.0229   0.4641   1.0000
   5.000   0.6841   0.02024   0.00973  -0.0206   0.3308   1.0000
   5.250   0.6978   0.02308   0.01119  -0.0195   0.1346   1.0000
   5.500   0.7173   0.02509   0.01292  -0.0187   0.0960   1.0000
   5.750   0.7371   0.02687   0.01477  -0.0175   0.0821   1.0000
   6.000   0.7569   0.02860   0.01657  -0.0164   0.0681   1.0000
   6.250   0.7786   0.03051   0.01867  -0.0149   0.0612   1.0000
   6.500   0.8005   0.03251   0.02081  -0.0139   0.0536   1.0000
   6.750   0.8250   0.03488   0.02346  -0.0127   0.0487   1.0000
   7.000   0.8493   0.03751   0.02645  -0.0117   0.0455   1.0000
   7.250   0.8701   0.04015   0.02928  -0.0110   0.0416   1.0000
   7.500   0.8889   0.04354   0.03309  -0.0101   0.0391   1.0000
   7.750   0.9059   0.04714   0.03729  -0.0089   0.0381   1.0000
   8.000   0.9190   0.05108   0.04180  -0.0077   0.0376   1.0000
   8.250   0.9279   0.05523   0.04648  -0.0066   0.0370   1.0000
   8.500   0.9324   0.05949   0.05123  -0.0057   0.0364   1.0000
   8.750   0.9325   0.06386   0.05602  -0.0050   0.0357   1.0000
   9.000   0.9281   0.06834   0.06084  -0.0046   0.0353   1.0000
   9.250   0.9192   0.07293   0.06570  -0.0046   0.0352   1.0000
   9.500   0.9053   0.07755   0.07051  -0.0050   0.0354   1.0000
   9.750   0.8882   0.08259   0.07567  -0.0068   0.0359   1.0000
  10.000   0.8721   0.08858   0.08174  -0.0105   0.0366   1.0000
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