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NACA M2 AIRFOIL (m2-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: NACA M2 AIRFOIL (m2-il)
Reynolds number: 100,000
Max Cl/Cd: 27.62 at α=5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-m2-il-100000-n5.txt
Download as CSV file: xf-m2-il-100000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA M2 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.000  -0.7461   0.08470   0.07979  -0.0045   1.0000   0.0284
  -9.750  -0.7609   0.07698   0.07203  -0.0107   1.0000   0.0282
  -9.500  -0.7788   0.07091   0.06586  -0.0137   1.0000   0.0281
  -9.250  -0.7950   0.06523   0.06005  -0.0147   1.0000   0.0280
  -9.000  -0.8074   0.05917   0.05370  -0.0151   1.0000   0.0281
  -8.750  -0.8151   0.05327   0.04737  -0.0147   1.0000   0.0283
  -8.500  -0.8169   0.04783   0.04134  -0.0137   1.0000   0.0288
  -8.250  -0.8101   0.04417   0.03745  -0.0130   1.0000   0.0302
  -8.000  -0.7950   0.04238   0.03552  -0.0124   1.0000   0.0321
  -7.750  -0.7814   0.03931   0.03207  -0.0115   1.0000   0.0339
  -7.500  -0.7671   0.03548   0.02767  -0.0103   1.0000   0.0355
  -7.250  -0.7492   0.03239   0.02394  -0.0092   1.0000   0.0384
  -7.000  -0.7295   0.02973   0.02088  -0.0084   1.0000   0.0408
  -6.750  -0.7075   0.02801   0.01898  -0.0078   1.0000   0.0430
  -6.500  -0.6844   0.02672   0.01746  -0.0072   1.0000   0.0466
  -6.250  -0.6603   0.02510   0.01546  -0.0065   1.0000   0.0498
  -6.000  -0.6362   0.02346   0.01363  -0.0058   1.0000   0.0519
  -5.750  -0.6124   0.02262   0.01279  -0.0054   1.0000   0.0556
  -5.500  -0.5879   0.02163   0.01168  -0.0048   1.0000   0.0593
  -5.250  -0.5635   0.02052   0.01042  -0.0041   1.0000   0.0620
  -5.000  -0.5400   0.01955   0.00942  -0.0033   1.0000   0.0650
  -4.750  -0.5163   0.01882   0.00868  -0.0027   1.0000   0.0693
  -4.500  -0.4927   0.01811   0.00787  -0.0018   1.0000   0.0754
  -4.250  -0.4690   0.01756   0.00736  -0.0012   1.0000   0.0856
  -4.000  -0.4455   0.01697   0.00675  -0.0005   1.0000   0.0999
  -3.750  -0.4221   0.01638   0.00615   0.0003   1.0000   0.1155
  -3.500  -0.3986   0.01588   0.00564   0.0010   1.0000   0.1305
  -3.250  -0.3750   0.01542   0.00518   0.0016   1.0000   0.1463
  -3.000  -0.3515   0.01493   0.00478   0.0023   1.0000   0.1670
  -2.750  -0.3286   0.01429   0.00437   0.0029   1.0000   0.2085
  -2.500  -0.3113   0.01266   0.00412   0.0042   1.0000   0.4852
  -2.250  -0.2915   0.01200   0.00408   0.0062   1.0000   0.6300
  -2.000  -0.2708   0.01169   0.00413   0.0084   1.0000   0.7247
  -1.750  -0.2484   0.01158   0.00418   0.0102   1.0000   0.7906
  -1.500  -0.2220   0.01160   0.00430   0.0114   1.0000   0.8466
  -1.250  -0.1916   0.01169   0.00440   0.0116   1.0000   0.8944
  -1.000  -0.1604   0.01177   0.00444   0.0111   1.0000   0.9313
  -0.750  -0.1218   0.01183   0.00445   0.0090   1.0000   0.9591
  -0.500  -0.0732   0.01188   0.00445   0.0046   1.0000   0.9809
  -0.250  -0.0165   0.01188   0.00444  -0.0016   1.0000   0.9981
   0.000   0.0000   0.01187   0.00441   0.0000   1.0000   1.0000
   0.250   0.0165   0.01188   0.00444   0.0016   0.9981   1.0000
   0.500   0.0732   0.01188   0.00445  -0.0046   0.9809   1.0000
   0.750   0.1218   0.01183   0.00445  -0.0090   0.9591   1.0000
   1.000   0.1603   0.01177   0.00444  -0.0111   0.9314   1.0000
   1.250   0.1916   0.01169   0.00440  -0.0116   0.8944   1.0000
   1.500   0.2220   0.01160   0.00430  -0.0114   0.8466   1.0000
   1.750   0.2483   0.01158   0.00418  -0.0102   0.7906   1.0000
   2.000   0.2708   0.01169   0.00413  -0.0084   0.7245   1.0000
   2.250   0.2915   0.01200   0.00408  -0.0062   0.6300   1.0000
   2.500   0.3112   0.01266   0.00411  -0.0042   0.4850   1.0000
   2.750   0.3286   0.01429   0.00437  -0.0029   0.2086   1.0000
   3.000   0.3515   0.01493   0.00478  -0.0023   0.1671   1.0000
   3.250   0.3749   0.01542   0.00518  -0.0016   0.1464   1.0000
   3.500   0.3986   0.01588   0.00564  -0.0010   0.1305   1.0000
   3.750   0.4221   0.01638   0.00615  -0.0003   0.1155   1.0000
   4.000   0.4455   0.01697   0.00675   0.0005   0.0999   1.0000
   4.250   0.4689   0.01756   0.00736   0.0012   0.0856   1.0000
   4.500   0.4927   0.01811   0.00787   0.0018   0.0754   1.0000
   4.750   0.5163   0.01882   0.00868   0.0027   0.0693   1.0000
   5.000   0.5400   0.01955   0.00941   0.0033   0.0650   1.0000
   5.250   0.5635   0.02052   0.01039   0.0041   0.0620   1.0000
   5.500   0.5879   0.02163   0.01168   0.0048   0.0593   1.0000
   5.750   0.6124   0.02262   0.01279   0.0054   0.0556   1.0000
   6.000   0.6362   0.02347   0.01363   0.0058   0.0519   1.0000
   6.250   0.6603   0.02510   0.01546   0.0065   0.0498   1.0000
   6.500   0.6844   0.02672   0.01747   0.0072   0.0466   1.0000
   6.750   0.7075   0.02801   0.01898   0.0077   0.0430   1.0000
   7.000   0.7296   0.02973   0.02088   0.0083   0.0408   1.0000
   7.250   0.7493   0.03240   0.02395   0.0092   0.0384   1.0000
   7.500   0.7672   0.03549   0.02768   0.0103   0.0355   1.0000
   7.750   0.7815   0.03932   0.03208   0.0115   0.0339   1.0000
   8.000   0.7951   0.04241   0.03556   0.0124   0.0321   1.0000
   8.250   0.8102   0.04419   0.03747   0.0130   0.0302   1.0000
   8.500   0.8170   0.04785   0.04136   0.0137   0.0288   1.0000
   8.750   0.8152   0.05330   0.04741   0.0147   0.0283   1.0000
   9.000   0.8075   0.05922   0.05375   0.0151   0.0281   1.0000
   9.250   0.7953   0.06527   0.06010   0.0146   0.0280   1.0000
   9.500   0.7790   0.07098   0.06594   0.0136   0.0280   1.0000
   9.750   0.7613   0.07707   0.07212   0.0106   0.0282   1.0000
  10.000   0.7466   0.08481   0.07990   0.0043   0.0284   1.0000
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