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DSMA-523A AIRFOIL (dsma523a-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: DSMA-523A AIRFOIL (dsma523a-il)
Reynolds number: 200,000
Max Cl/Cd: 27.09 at α=6.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-dsma523a-il-200000.txt
Download as CSV file: xf-dsma523a-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: DSMA-523A AIRFOIL                               
  
 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
  ------ -------- --------- --------- -------- -------- --------
 -10.750  -0.6074   0.10701   0.10281  -0.0256   1.0000   0.0838
 -10.500  -0.5976   0.10538   0.10118  -0.0244   1.0000   0.0857
 -10.250  -0.7302   0.08422   0.07997  -0.0441   1.0000   0.0910
 -10.000  -0.6314   0.09214   0.08805  -0.0318   1.0000   0.0932
  -9.750  -0.6169   0.09177   0.08768  -0.0293   1.0000   0.0955
  -9.500  -0.9121   0.04735   0.04100  -0.0469   1.0000   0.0547
  -9.250  -0.9006   0.04405   0.03744  -0.0468   1.0000   0.0546
  -8.750  -0.8685   0.03826   0.03137  -0.0464   1.0000   0.0556
  -8.500  -0.8494   0.03599   0.02896  -0.0463   1.0000   0.0560
  -8.250  -0.8288   0.03369   0.02640  -0.0464   1.0000   0.0564
  -8.000  -0.8066   0.03172   0.02423  -0.0465   1.0000   0.0570
  -7.750  -0.7832   0.02995   0.02222  -0.0466   1.0000   0.0577
  -7.500  -0.7585   0.02840   0.02043  -0.0467   1.0000   0.0589
  -7.250  -0.7311   0.02733   0.01894  -0.0475   1.0000   0.0607
  -7.000  -0.7083   0.02551   0.01717  -0.0468   1.0000   0.0623
  -6.750  -0.6841   0.02440   0.01607  -0.0463   1.0000   0.0640
  -6.500  -0.6586   0.02339   0.01496  -0.0460   1.0000   0.0658
  -6.250  -0.6318   0.02258   0.01396  -0.0460   1.0000   0.0676
  -6.000  -0.6099   0.02136   0.01293  -0.0448   1.0000   0.0700
  -5.750  -0.5848   0.02062   0.01221  -0.0444   1.0000   0.0730
  -5.500  -0.5600   0.01976   0.01140  -0.0438   1.0000   0.0757
  -5.250  -0.5339   0.01906   0.01082  -0.0438   1.0000   0.0788
  -5.000  -0.5060   0.01846   0.01024  -0.0440   1.0000   0.0828
  -4.750  -0.4767   0.01788   0.00979  -0.0448   1.0000   0.0876
  -4.500  -0.4462   0.01735   0.00931  -0.0458   1.0000   0.0928
  -4.250  -0.4151   0.01693   0.00895  -0.0469   1.0000   0.0997
  -4.000  -0.3819   0.01643   0.00854  -0.0486   1.0000   0.1080
  -3.750  -0.3478   0.01598   0.00818  -0.0504   1.0000   0.1194
  -3.500  -0.3108   0.01544   0.00782  -0.0530   1.0000   0.1438
  -3.250  -0.2510   0.01394   0.00725  -0.0617   1.0000   0.3019
  -3.000  -0.2257   0.01393   0.00877  -0.0605   1.0000   0.5532
  -2.750  -0.2138   0.01486   0.00980  -0.0560   1.0000   0.6122
  -2.500  -0.1847   0.01530   0.01014  -0.0562   1.0000   0.6430
  -2.250  -0.1596   0.01576   0.01055  -0.0554   1.0000   0.6576
  -2.000  -0.1427   0.01636   0.01115  -0.0525   1.0000   0.6655
  -1.750  -0.1186   0.01676   0.01152  -0.0515   1.0000   0.6754
  -1.500  -0.1032   0.01731   0.01210  -0.0482   1.0000   0.6817
  -1.250  -0.0799   0.01774   0.01252  -0.0471   1.0000   0.6919
  -1.000  -0.0675   0.01825   0.01308  -0.0431   1.0000   0.6967
  -0.750  -0.0413   0.01857   0.01339  -0.0428   1.0000   0.7058
  -0.500  -0.0250   0.01892   0.01379  -0.0400   1.0000   0.7110
  -0.250  -0.0150   0.01947   0.01441  -0.0355   1.0000   0.7191
   0.000  -0.0045   0.02003   0.01503  -0.0312   1.0000   0.7309
   0.250   0.0287   0.02089   0.01597  -0.0306   0.9921   0.7472
   0.500   0.0752   0.02127   0.01640  -0.0326   0.9783   0.7565
   0.750   0.1329   0.02111   0.01628  -0.0373   0.9620   0.7653
   1.000   0.1670   0.02054   0.01577  -0.0367   0.9450   0.7716
   1.250   0.2265   0.01984   0.01512  -0.0422   0.9291   0.7797
   1.500   0.2626   0.01882   0.01419  -0.0417   0.9123   0.7842
   1.750   0.2841   0.01810   0.01354  -0.0390   0.8837   0.7883
   2.000   0.3783   0.01849   0.01194  -0.0482   0.4216   0.8004
   2.250   0.3781   0.01951   0.01204  -0.0423   0.2304   0.8060
   2.500   0.3984   0.02043   0.01240  -0.0412   0.1424   0.8122
   2.750   0.4300   0.02102   0.01284  -0.0427   0.1217   0.8164
   3.000   0.4474   0.02121   0.01292  -0.0404   0.1123   0.8184
   3.250   0.4741   0.02142   0.01312  -0.0404   0.1046   0.8193
   3.500   0.5019   0.02210   0.01367  -0.0409   0.0977   0.8194
   3.750   0.5301   0.02240   0.01401  -0.0411   0.0936   0.8203
   4.000   0.5578   0.02282   0.01439  -0.0412   0.0892   0.8215
   4.250   0.5888   0.02388   0.01531  -0.0423   0.0845   0.8218
   4.500   0.6207   0.02435   0.01585  -0.0433   0.0817   0.8218
   4.750   0.6511   0.02495   0.01649  -0.0440   0.0788   0.8227
   5.000   0.6814   0.02556   0.01709  -0.0447   0.0757   0.8235
   5.250   0.7146   0.02704   0.01849  -0.0463   0.0725   0.8235
   5.500   0.7457   0.02785   0.01947  -0.0471   0.0708   0.8234
   5.750   0.7766   0.02882   0.02061  -0.0479   0.0690   0.8233
   6.000   0.8068   0.02992   0.02188  -0.0486   0.0671   0.8233
   6.250   0.8365   0.03088   0.02289  -0.0493   0.0651   0.8232
   6.500   0.8652   0.03207   0.02407  -0.0501   0.0632   0.8235
   6.750   0.8908   0.03477   0.02696  -0.0503   0.0619   0.8243
   7.000   0.9146   0.03633   0.02889  -0.0497   0.0612   0.8248
   7.250   0.9369   0.03846   0.03140  -0.0491   0.0606   0.8247
   7.500   0.9563   0.04093   0.03430  -0.0481   0.0599   0.8247
   7.750   0.9728   0.04382   0.03759  -0.0469   0.0593   0.8246
   8.000   0.9854   0.04740   0.04160  -0.0454   0.0595   0.8246
   8.250   0.9956   0.05149   0.04601  -0.0440   0.0605   0.8247
   8.500   1.0150   0.05503   0.04965  -0.0438   0.0630   0.8248
   8.750   0.9364   0.07429   0.07052  -0.0361   0.0865   0.8243
   9.000   0.9514   0.07759   0.07378  -0.0363   0.0856   0.8244
   9.250   0.9592   0.08670   0.08278  -0.0380   0.0842   0.8245
   9.500   0.9368   0.09033   0.08671  -0.0363   0.0839   0.8245
   9.750   0.8914   0.09467   0.09132  -0.0351   0.0836   0.8248
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