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M6 (65%) (m665-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: M6 (65%) (m665-il)
Reynolds number: 200,000
Max Cl/Cd: 60.28 at α=4.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-m665-il-200000.txt
Download as CSV file: xf-m665-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: M6 (65%)                                        
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.5384   0.08674   0.08322  -0.0127   1.0000   0.0445
  -8.250  -0.5460   0.08241   0.07893  -0.0166   1.0000   0.0453
  -8.000  -0.5558   0.07832   0.07483  -0.0193   1.0000   0.0461
  -7.750  -0.5639   0.07390   0.07025  -0.0234   1.0000   0.0472
  -7.500  -0.5651   0.07048   0.06662  -0.0238   1.0000   0.0476
  -7.250  -0.5699   0.06377   0.05986  -0.0238   1.0000   0.0483
  -7.000  -0.5609   0.06015   0.05633  -0.0228   1.0000   0.0494
  -6.750  -0.5518   0.05750   0.05368  -0.0216   1.0000   0.0508
  -6.500  -0.5432   0.05462   0.05073  -0.0207   1.0000   0.0527
  -6.250  -0.5345   0.05144   0.04742  -0.0197   1.0000   0.0555
  -5.750  -0.5210   0.04399   0.03944  -0.0162   1.0000   0.0627
  -5.500  -0.5083   0.04175   0.03717  -0.0146   1.0000   0.0653
  -5.250  -0.5008   0.04026   0.03498  -0.0109   1.0000   0.0743
  -5.000  -0.4874   0.03646   0.03137  -0.0098   1.0000   0.0766
  -4.750  -0.4768   0.03568   0.03017  -0.0066   1.0000   0.0885
  -4.500  -0.4619   0.02751   0.02095  -0.0009   1.0000   0.0480
  -4.250  -0.4478   0.02375   0.01700   0.0016   1.0000   0.0449
  -4.000  -0.4316   0.02146   0.01435   0.0042   1.0000   0.0442
  -3.750  -0.4144   0.02035   0.01295   0.0066   1.0000   0.0465
  -3.500  -0.3976   0.01945   0.01178   0.0089   1.0000   0.0478
  -3.250  -0.3789   0.01733   0.00945   0.0107   1.0000   0.0491
  -3.000  -0.3470   0.01633   0.00845   0.0095   0.9966   0.0539
  -2.750  -0.3048   0.01527   0.00730   0.0067   0.9901   0.0575
  -2.500  -0.2629   0.01415   0.00616   0.0038   0.9839   0.0624
  -2.250  -0.2232   0.01344   0.00548   0.0012   0.9748   0.0716
  -2.000  -0.1834   0.01262   0.00471  -0.0013   0.9665   0.0849
  -1.750  -0.1470   0.01127   0.00405  -0.0034   0.9577   0.2075
  -1.500  -0.0686   0.00945   0.00484  -0.0121   0.9745   0.9592
  -1.250   0.0206   0.00967   0.00483  -0.0240   0.9926   1.0000
  -1.000   0.0711   0.00952   0.00457  -0.0290   0.9774   1.0000
  -0.750   0.1224   0.00937   0.00431  -0.0341   0.9615   1.0000
  -0.500   0.1738   0.00920   0.00405  -0.0392   0.9426   1.0000
  -0.250   0.2206   0.00905   0.00381  -0.0431   0.9176   1.0000
   0.000   0.2569   0.00898   0.00363  -0.0448   0.8900   1.0000
   0.250   0.2833   0.00900   0.00353  -0.0443   0.8630   1.0000
   0.500   0.3064   0.00905   0.00346  -0.0431   0.8396   1.0000
   0.750   0.3282   0.00912   0.00344  -0.0417   0.8173   1.0000
   1.000   0.3502   0.00920   0.00343  -0.0404   0.7979   1.0000
   1.250   0.3724   0.00929   0.00343  -0.0391   0.7796   1.0000
   1.500   0.3945   0.00937   0.00346  -0.0378   0.7607   1.0000
   1.750   0.4166   0.00945   0.00348  -0.0365   0.7417   1.0000
   2.000   0.4388   0.00955   0.00352  -0.0352   0.7233   1.0000
   2.250   0.4612   0.00964   0.00357  -0.0339   0.7049   1.0000
   2.500   0.4840   0.00973   0.00365  -0.0328   0.6876   1.0000
   2.750   0.5067   0.00983   0.00373  -0.0316   0.6701   1.0000
   3.000   0.5292   0.00993   0.00381  -0.0303   0.6507   1.0000
   3.250   0.5515   0.01002   0.00389  -0.0290   0.6283   1.0000
   3.500   0.5735   0.01013   0.00395  -0.0276   0.6041   1.0000
   3.750   0.5945   0.01021   0.00398  -0.0260   0.5691   1.0000
   4.000   0.6154   0.01035   0.00405  -0.0244   0.5314   1.0000
   4.250   0.6360   0.01055   0.00414  -0.0229   0.4849   1.0000
   4.500   0.6553   0.01091   0.00427  -0.0211   0.4203   1.0000
   4.750   0.6722   0.01158   0.00452  -0.0192   0.3202   1.0000
   5.000   0.6788   0.01379   0.00543  -0.0163   0.0953   1.0000
   5.250   0.6943   0.01511   0.00660  -0.0141   0.0608   1.0000
   5.500   0.7118   0.01606   0.00756  -0.0122   0.0517   1.0000
   5.750   0.7301   0.01689   0.00839  -0.0105   0.0465   1.0000
   6.000   0.7451   0.01828   0.00977  -0.0082   0.0437   1.0000
   6.250   0.7644   0.01923   0.01080  -0.0066   0.0421   1.0000
   6.500   0.7840   0.02035   0.01201  -0.0049   0.0408   1.0000
   6.750   0.8043   0.02164   0.01338  -0.0034   0.0400   1.0000
   7.000   0.8254   0.02312   0.01497  -0.0020   0.0394   1.0000
   7.250   0.8466   0.02478   0.01680  -0.0006   0.0389   1.0000
   7.500   0.8667   0.02625   0.01841   0.0007   0.0376   1.0000
   7.750   0.8861   0.02797   0.02034   0.0021   0.0367   1.0000
   8.000   0.9039   0.03050   0.02323   0.0040   0.0373   1.0000
   8.250   0.9175   0.03382   0.02704   0.0065   0.0388   1.0000
   8.500   0.9269   0.03786   0.03153   0.0092   0.0411   1.0000
  12.500   0.6200   0.11996   0.11638   0.0089   0.0682   1.0000
  12.750   0.5994   0.12628   0.12266   0.0037   0.0676   1.0000
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