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

NACA M25 AIRFOIL (m25-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5


Details Polar file
Airfoil: NACA M25 AIRFOIL (m25-il)
Reynolds number: 1,000,000
Max Cl/Cd: 107.76 at α=5.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-m25-il-1000000-n5.txt
Download as CSV file: xf-m25-il-1000000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA M25 AIRFOIL                                
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     1.000 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.500  -0.2127   0.08630   0.08353   0.0050   0.5221   0.0052
  -7.250  -0.2082   0.08304   0.08026   0.0039   0.5195   0.0052
  -7.000  -0.2039   0.07981   0.07702   0.0026   0.5171   0.0052
  -6.750  -0.2003   0.07662   0.07383   0.0013   0.5147   0.0052
  -6.500  -0.1918   0.07347   0.07068  -0.0005   0.5124   0.0046
  -6.000  -0.2333   0.08098   0.07797  -0.0015   0.5147   0.0035
  -5.750  -0.2131   0.07731   0.07427  -0.0050   0.5122   0.0034
  -5.500  -0.1917   0.07398   0.07091  -0.0084   0.5095   0.0035
  -5.250  -0.1695   0.07119   0.06808  -0.0113   0.5063   0.0037
  -5.000  -0.1460   0.06821   0.06506  -0.0144   0.5034   0.0038
  -4.750  -0.1211   0.06508   0.06187  -0.0175   0.5005   0.0041
  -4.500  -0.0949   0.06173   0.05846  -0.0207   0.4979   0.0043
  -4.250  -0.0670   0.05808   0.05474  -0.0240   0.4953   0.0048
  -4.000  -0.0370   0.05391   0.05049  -0.0275   0.4926   0.0052
  -3.750  -0.0070   0.05027   0.04676  -0.0305   0.4897   0.0054
  -3.500   0.0207   0.04805   0.04446  -0.0325   0.4866   0.0057
  -3.250   0.0493   0.04568   0.04200  -0.0345   0.4837   0.0062
  -3.000   0.0800   0.04260   0.03883  -0.0366   0.4812   0.0068
  -2.750   0.1117   0.03924   0.03534  -0.0386   0.4785   0.0080
  -2.500   0.1399   0.03763   0.03366  -0.0398   0.4754   0.0085
  -2.250   0.1695   0.03554   0.03146  -0.0411   0.4723   0.0093
  -2.000   0.2029   0.03192   0.02764  -0.0425   0.4695   0.0111
  -1.750   0.2302   0.03107   0.02674  -0.0432   0.4668   0.0116
  -1.500   0.2590   0.02968   0.02527  -0.0439   0.4641   0.0124
  -1.250   0.2947   0.02589   0.02121  -0.0445   0.4615   0.0153
  -1.000   0.3217   0.02513   0.02038  -0.0451   0.4582   0.0157
  -0.750   0.3494   0.02430   0.01947  -0.0456   0.4549   0.0161
  -0.500   0.3779   0.02329   0.01838  -0.0460   0.4522   0.0166
  -0.250   0.4069   0.02215   0.01715  -0.0464   0.4493   0.0176
   0.000   0.4369   0.02076   0.01561  -0.0466   0.4464   0.0189
   0.250   0.4674   0.01927   0.01395  -0.0466   0.4434   0.0198
   0.500   0.4971   0.01797   0.01249  -0.0466   0.4403   0.0202
   1.000   0.5561   0.01475   0.00890  -0.0463   0.4346   0.0212
   1.250   0.5843   0.01355   0.00754  -0.0465   0.4314   0.0219
   1.500   0.6124   0.01319   0.00711  -0.0468   0.4282   0.0225
   1.750   0.6406   0.01283   0.00669  -0.0470   0.4254   0.0233
   2.000   0.6689   0.01229   0.00606  -0.0472   0.4226   0.0243
   2.250   0.6973   0.01167   0.00534  -0.0473   0.4191   0.0250
   2.500   0.7256   0.01117   0.00473  -0.0475   0.4155   0.0253
   2.750   0.7538   0.01080   0.00427  -0.0477   0.4119   0.0256
   3.000   0.7820   0.01052   0.00396  -0.0479   0.4092   0.0261
   3.250   0.8100   0.01019   0.00357  -0.0480   0.4061   0.0257
   3.500   0.8377   0.00993   0.00327  -0.0482   0.4027   0.0253
   3.750   0.8653   0.00974   0.00305  -0.0483   0.3994   0.0250
   4.000   0.8928   0.00961   0.00289  -0.0485   0.3962   0.0248
   4.250   0.9204   0.00950   0.00278  -0.0486   0.3930   0.0246
   4.500   0.9480   0.00944   0.00272  -0.0489   0.3884   0.0245
   4.750   0.9756   0.00946   0.00272  -0.0491   0.3827   0.0245
   5.000   1.0033   0.00950   0.00273  -0.0494   0.3724   0.0246
   5.250   1.0309   0.00961   0.00280  -0.0498   0.3596   0.0251
   5.500   1.0582   0.00982   0.00294  -0.0501   0.3424   0.0264
   5.750   1.0853   0.01015   0.00317  -0.0506   0.3200   0.0281
   6.000   1.1118   0.01063   0.00352  -0.0511   0.2889   0.0290
   6.250   1.1264   0.01405   0.00582  -0.0521   0.0796   0.0304
   6.500   1.1491   0.01504   0.00663  -0.0524   0.0185   0.0320
   6.750   1.1746   0.01537   0.00699  -0.0526   0.0146   0.0350
   7.000   1.1996   0.01579   0.00744  -0.0527   0.0108   0.0375
   7.250   1.2244   0.01617   0.00792  -0.0529   0.0095   0.0705
   7.500   1.2486   0.01655   0.00847  -0.0530   0.0084   0.1893
   8.000   1.3255   0.01682   0.01013  -0.0604   0.0059   1.0000
   8.250   1.3450   0.01743   0.01079  -0.0599   0.0053   1.0000
   8.500   1.3635   0.01812   0.01153  -0.0592   0.0049   1.0000
   8.750   1.3803   0.01889   0.01234  -0.0585   0.0045   1.0000
   9.000   1.3919   0.02010   0.01362  -0.0575   0.0040   1.0000
   9.250   1.4032   0.02146   0.01506  -0.0572   0.0038   1.0000
   9.500   1.4068   0.02317   0.01684  -0.0561   0.0036   1.0000
   9.750   1.4108   0.02513   0.01888  -0.0554   0.0034   1.0000
  10.000   1.4145   0.02735   0.02119  -0.0550   0.0033   1.0000
  10.250   1.4170   0.02978   0.02372  -0.0547   0.0031   1.0000
  10.500   1.4192   0.03225   0.02627  -0.0544   0.0030   1.0000
  10.750   1.4202   0.03487   0.02896  -0.0540   0.0029   1.0000
  11.000   1.4195   0.03767   0.03184  -0.0537   0.0028   1.0000
  11.250   1.4186   0.04047   0.03471  -0.0533   0.0027   1.0000
  11.500   1.4160   0.04350   0.03782  -0.0530   0.0026   1.0000
  11.750   1.4116   0.04673   0.04113  -0.0526   0.0026   1.0000
  12.000   1.4050   0.05026   0.04474  -0.0523   0.0025   1.0000
  12.250   1.3953   0.05428   0.04886  -0.0520   0.0024   1.0000
  12.500   1.3847   0.05854   0.05322  -0.0519   0.0023   1.0000
  12.750   1.3822   0.06183   0.05659  -0.0518   0.0023   1.0000
  13.000   1.3794   0.06519   0.06003  -0.0517   0.0023   1.0000
  13.250   1.3763   0.06864   0.06357  -0.0517   0.0022   1.0000
  13.500   1.3736   0.07198   0.06699  -0.0516   0.0022   1.0000
  13.750   1.3712   0.07531   0.07040  -0.0515   0.0021   1.0000
  14.000   1.3698   0.07844   0.07363  -0.0513   0.0021   1.0000
  14.250   1.3690   0.08144   0.07672  -0.0510   0.0020   1.0000
  14.500   1.3700   0.08415   0.07949  -0.0506   0.0019   1.0000
  14.750   1.3727   0.08662   0.08204  -0.0502   0.0018   1.0000
  15.000   1.3764   0.08898   0.08448  -0.0499   0.0017   1.0000
  15.250   1.3805   0.09136   0.08694  -0.0497   0.0016   1.0000
  15.500   1.3844   0.09377   0.08942  -0.0495   0.0016   1.0000
  15.750   1.3878   0.09630   0.09202  -0.0494   0.0015   1.0000
  16.000   1.3911   0.09878   0.09459  -0.0493   0.0015   1.0000
  16.250   1.3944   0.10123   0.09714  -0.0490   0.0014   1.0000
  16.500   1.3969   0.10393   0.09992  -0.0490   0.0014   1.0000
  16.750   1.3988   0.10669   0.10279  -0.0490   0.0014   1.0000
  17.000   1.3992   0.10969   0.10590  -0.0491   0.0013   1.0000
  17.250   1.3979   0.11286   0.10923  -0.0490   0.0013   1.0000
  17.500   1.3813   0.11725   0.11403  -0.0448   0.0012   1.0000
  17.750   1.3670   0.12312   0.12014  -0.0449   0.0012   1.0000
  18.000   1.3504   0.12993   0.12713  -0.0466   0.0011   1.0000
  18.250   1.3318   0.13751   0.13491  -0.0493   0.0011   1.0000
  18.500   1.3136   0.14535   0.14293  -0.0528   0.0011   1.0000
<< Back to NACA M25 AIRFOIL (m25-il)

Polar data table (+)

Polar graphs


<< Back to NACA M25 AIRFOIL (m25-il)