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NACA 65(1)-212 a=0.6 (naca651212a06-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: NACA 65(1)-212 a=0.6 (naca651212a06-il)
Reynolds number: 200,000
Max Cl/Cd: 66.71 at α=4.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-naca651212a06-il-200000.txt
Download as CSV file: xf-naca651212a06-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 65(1)-212 a=0.6                            
  
 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
  ------ -------- --------- --------- -------- -------- --------
 -10.000  -0.5139   0.08465   0.08128  -0.0444   1.0000   0.0579
  -9.750  -0.5304   0.07783   0.07446  -0.0506   1.0000   0.0579
  -9.500  -0.5444   0.07287   0.06949  -0.0537   1.0000   0.0576
  -9.250  -0.5761   0.06959   0.06612  -0.0546   1.0000   0.0584
  -9.000  -0.6010   0.06766   0.06408  -0.0522   1.0000   0.0587
  -8.750  -0.6233   0.06678   0.06300  -0.0493   0.9946   0.0592
  -8.500  -0.6171   0.05839   0.05451  -0.0531   0.9874   0.0612
  -8.250  -0.5879   0.05519   0.05143  -0.0558   0.9843   0.0645
  -8.000  -0.5695   0.05164   0.04755  -0.0587   0.9767   0.0700
  -7.750  -0.5556   0.04664   0.04211  -0.0617   0.9701   0.0756
  -7.500  -0.5322   0.04383   0.03928  -0.0630   0.9630   0.0794
  -7.250  -0.5159   0.04082   0.03578  -0.0639   0.9555   0.0896
  -7.000  -0.4942   0.03840   0.03335  -0.0642   0.9477   0.0942
  -6.750  -0.4725   0.03291   0.02651  -0.0605   0.9394   0.0588
  -6.500  -0.4526   0.02839   0.02126  -0.0584   0.9320   0.0477
  -6.250  -0.4280   0.02584   0.01850  -0.0579   0.9262   0.0463
  -6.000  -0.4050   0.02392   0.01633  -0.0569   0.9191   0.0461
  -5.750  -0.3801   0.02271   0.01489  -0.0561   0.9140   0.0468
  -5.500  -0.3562   0.02084   0.01289  -0.0555   0.9074   0.0487
  -5.250  -0.3315   0.01954   0.01152  -0.0548   0.9020   0.0498
  -5.000  -0.3080   0.01858   0.01054  -0.0540   0.8968   0.0515
  -4.750  -0.2852   0.01782   0.00977  -0.0531   0.8904   0.0536
  -4.500  -0.2625   0.01722   0.00912  -0.0520   0.8856   0.0570
  -4.250  -0.2405   0.01663   0.00848  -0.0511   0.8801   0.0604
  -4.000  -0.2207   0.01587   0.00774  -0.0498   0.8742   0.0654
  -3.750  -0.1984   0.01541   0.00720  -0.0487   0.8699   0.0722
  -3.500  -0.1772   0.01482   0.00668  -0.0477   0.8641   0.0905
  -3.250  -0.1761   0.01230   0.00607  -0.0441   0.8580   0.4927
  -3.000  -0.1587   0.01209   0.00631  -0.0413   0.8544   0.6461
  -2.750  -0.1361   0.01225   0.00652  -0.0400   0.8486   0.6901
  -2.500  -0.1125   0.01240   0.00666  -0.0388   0.8438   0.7190
  -2.250  -0.0907   0.01268   0.00697  -0.0366   0.8404   0.7541
  -2.000  -0.0706   0.01279   0.00711  -0.0343   0.8360   0.7878
  -1.750  -0.0485   0.01265   0.00703  -0.0328   0.8308   0.8106
  -1.500  -0.0229   0.01255   0.00694  -0.0320   0.8270   0.8234
  -1.250   0.0043   0.01252   0.00685  -0.0317   0.8241   0.8317
  -1.000   0.0307   0.01270   0.00705  -0.0314   0.8191   0.8379
  -0.750   0.0570   0.01282   0.00716  -0.0311   0.8145   0.8454
  -0.500   0.0843   0.01296   0.00728  -0.0306   0.8111   0.8510
  -0.250   0.1114   0.01309   0.00737  -0.0302   0.8085   0.8586
   0.000   0.1363   0.01340   0.00770  -0.0298   0.8027   0.8664
   0.250   0.1619   0.01355   0.00783  -0.0295   0.7984   0.8776
   0.500   0.1860   0.01363   0.00786  -0.0289   0.7951   0.8923
   0.750   0.2094   0.01373   0.00796  -0.0281   0.7911   0.9052
   1.000   0.2329   0.01383   0.00809  -0.0275   0.7857   0.9170
   1.250   0.2598   0.01380   0.00808  -0.0273   0.7820   0.9282
   1.500   0.2942   0.01374   0.00804  -0.0285   0.7793   0.9358
   1.750   0.3283   0.01386   0.00820  -0.0300   0.7752   0.9449
   2.000   0.3656   0.01399   0.00840  -0.0323   0.7704   0.9532
   2.250   0.4074   0.01401   0.00846  -0.0354   0.7670   0.9590
   2.500   0.4512   0.01397   0.00846  -0.0386   0.7641   0.9644
   2.750   0.4922   0.01407   0.00866  -0.0417   0.7581   0.9713
   3.000   0.5364   0.01398   0.00865  -0.0452   0.7525   0.9766
   3.250   0.5767   0.01363   0.00833  -0.0473   0.7442   0.9838
   3.500   0.6185   0.01256   0.00724  -0.0487   0.7231   0.9901
   3.750   0.6590   0.01152   0.00612  -0.0500   0.6932   0.9980
   4.000   0.6834   0.01116   0.00575  -0.0492   0.6739   1.0000
   4.250   0.7012   0.01098   0.00562  -0.0472   0.6511   1.0000
   4.500   0.7192   0.01080   0.00540  -0.0450   0.6147   1.0000
   4.750   0.7311   0.01096   0.00517  -0.0415   0.5070   1.0000
   5.000   0.7135   0.01337   0.00602  -0.0343   0.2331   1.0000
   5.250   0.7029   0.01551   0.00718  -0.0286   0.0856   1.0000
   5.500   0.7114   0.01638   0.00798  -0.0253   0.0677   1.0000
   5.750   0.7194   0.01726   0.00884  -0.0221   0.0594   1.0000
   6.000   0.7310   0.01799   0.00963  -0.0195   0.0552   1.0000
   6.250   0.7432   0.01879   0.01045  -0.0170   0.0518   1.0000
   6.500   0.7544   0.01984   0.01145  -0.0145   0.0485   1.0000
   6.750   0.7692   0.02106   0.01267  -0.0126   0.0460   1.0000
   7.000   0.7891   0.02195   0.01363  -0.0114   0.0439   1.0000
   7.250   0.8124   0.02312   0.01483  -0.0107   0.0422   1.0000
   7.500   0.8388   0.02445   0.01621  -0.0104   0.0410   1.0000
   7.750   0.8673   0.02602   0.01785  -0.0105   0.0400   1.0000
   8.000   0.8943   0.02766   0.01957  -0.0106   0.0386   1.0000
   8.250   0.9251   0.03074   0.02273  -0.0114   0.0371   1.0000
   8.500   0.9496   0.03309   0.02530  -0.0110   0.0371   1.0000
   8.750   0.9709   0.03523   0.02771  -0.0100   0.0373   1.0000
  12.500   0.6399   0.11431   0.11112  -0.0024   0.0744   1.0000
  12.750   0.7459   0.12304   0.11955  -0.0097   0.0641   1.0000
  13.000   0.7511   0.12628   0.12280  -0.0104   0.0619   1.0000
  13.250   0.7705   0.12586   0.12242  -0.0070   0.0600   1.0000
  13.500   0.5931   0.13556   0.13234  -0.0150   0.0666   1.0000
  13.750   0.5905   0.13889   0.13568  -0.0169   0.0639   1.0000
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