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

NACA M24 AIRFOIL (m24-il) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: NACA M24 AIRFOIL (m24-il)
Reynolds number: 500,000
Max Cl/Cd: 100.64 at α=8.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-m24-il-500000-n5.txt
Download as CSV file: xf-m24-il-500000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA M24 AIRFOIL                                
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.3371   0.10618   0.10269   0.0046   0.5981   0.0183
  -8.750  -0.3326   0.10270   0.09920   0.0028   0.5945   0.0184
  -8.500  -0.3377   0.09714   0.09365  -0.0024   0.5921   0.0197
  -8.250  -0.3280   0.09489   0.09138  -0.0027   0.5883   0.0200
  -8.000  -0.3211   0.09221   0.08869  -0.0042   0.5850   0.0202
  -7.750  -0.3162   0.08923   0.08571  -0.0065   0.5816   0.0206
  -7.500  -0.3132   0.08638   0.08285  -0.0081   0.5780   0.0209
  -7.250  -0.3052   0.08327   0.07970  -0.0104   0.5745   0.0214
  -7.000  -0.2956   0.07891   0.07526  -0.0155   0.5716   0.0233
  -6.500  -0.2706   0.07061   0.06681  -0.0217   0.5657   0.0235
  -6.250  -0.2557   0.06683   0.06294  -0.0235   0.5624   0.0235
  -6.000  -0.2398   0.06321   0.05923  -0.0249   0.5590   0.0235
  -5.750  -0.2230   0.05975   0.05566  -0.0260   0.5556   0.0235
  -5.500  -0.2063   0.05681   0.05264  -0.0265   0.5523   0.0234
  -5.250  -0.1893   0.05324   0.04898  -0.0273   0.5489   0.0239
  -5.000  -0.1703   0.05218   0.04789  -0.0273   0.5449   0.0243
  -4.750  -0.1496   0.05052   0.04617  -0.0277   0.5411   0.0250
  -4.500  -0.1274   0.04842   0.04396  -0.0283   0.5377   0.0265
  -4.000  -0.0834   0.04244   0.03770  -0.0289   0.5308   0.0260
  -3.750  -0.0608   0.03907   0.03414  -0.0288   0.5275   0.0252
  -3.500  -0.0379   0.03541   0.03023  -0.0283   0.5243   0.0247
  -3.250  -0.0139   0.03336   0.02803  -0.0281   0.5209   0.0251
  -3.000   0.0105   0.03117   0.02568  -0.0277   0.5173   0.0255
  -2.750   0.0342   0.02760   0.02181  -0.0266   0.5140   0.0252
  -2.500   0.0580   0.02395   0.01781  -0.0254   0.5107   0.0252
  -2.250   0.0813   0.01933   0.01258  -0.0238   0.5078   0.0254
  -2.000   0.1079   0.01747   0.01037  -0.0233   0.5043   0.0257
  -1.750   0.1357   0.01650   0.00919  -0.0232   0.5005   0.0262
  -1.500   0.1639   0.01561   0.00809  -0.0232   0.4967   0.0265
  -1.250   0.1924   0.01486   0.00715  -0.0232   0.4929   0.0267
  -1.000   0.2210   0.01428   0.00642  -0.0233   0.4893   0.0270
  -0.750   0.2497   0.01380   0.00583  -0.0234   0.4854   0.0273
  -0.500   0.2784   0.01340   0.00536  -0.0236   0.4814   0.0276
  -0.250   0.3069   0.01310   0.00497  -0.0237   0.4777   0.0279
   0.000   0.3352   0.01286   0.00465  -0.0238   0.4742   0.0282
   0.250   0.3636   0.01262   0.00440  -0.0239   0.4705   0.0284
   0.500   0.3911   0.01223   0.00399  -0.0239   0.4664   0.0289
   0.750   0.4186   0.01199   0.00374  -0.0239   0.4624   0.0294
   1.000   0.4461   0.01185   0.00358  -0.0239   0.4589   0.0301
   1.250   0.4738   0.01172   0.00347  -0.0240   0.4556   0.0310
   1.500   0.5014   0.01160   0.00337  -0.0240   0.4518   0.0318
   1.750   0.5289   0.01152   0.00326  -0.0240   0.4477   0.0324
   2.000   0.5564   0.01147   0.00319  -0.0240   0.4438   0.0331
   2.250   0.5841   0.01142   0.00314  -0.0241   0.4402   0.0339
   2.500   0.6117   0.01137   0.00310  -0.0242   0.4365   0.0349
   2.750   0.6393   0.01135   0.00308  -0.0242   0.4328   0.0367
   3.000   0.6668   0.01136   0.00308  -0.0244   0.4292   0.0390
   3.250   0.6943   0.01136   0.00311  -0.0244   0.4257   0.0467
   3.500   0.7214   0.01129   0.00320  -0.0245   0.4218   0.0980
   3.750   0.7487   0.01129   0.00330  -0.0246   0.4177   0.1360
   4.000   0.7756   0.01130   0.00340  -0.0247   0.4140   0.1899
   4.250   0.7873   0.01012   0.00340  -0.0219   0.4112   0.7463
   4.750   0.8808   0.01041   0.00430  -0.0298   0.4031   0.9754
   5.000   0.9216   0.01065   0.00450  -0.0329   0.3989   0.9799
   5.250   0.9754   0.01095   0.00478  -0.0388   0.3946   0.9865
   5.500   1.0214   0.01113   0.00497  -0.0431   0.3903   0.9905
   5.750   1.0501   0.01127   0.00510  -0.0436   0.3866   0.9915
   6.000   1.0779   0.01143   0.00525  -0.0440   0.3832   0.9923
   6.250   1.1052   0.01159   0.00541  -0.0444   0.3801   0.9930
   6.500   1.1325   0.01173   0.00558  -0.0446   0.3770   0.9938
   6.750   1.1618   0.01187   0.00575  -0.0454   0.3732   0.9942
   7.000   1.1905   0.01207   0.00594  -0.0461   0.3677   0.9947
   7.250   1.2193   0.01225   0.00613  -0.0469   0.3609   0.9952
   7.500   1.2476   0.01247   0.00634  -0.0476   0.3535   0.9958
   7.750   1.2757   0.01270   0.00658  -0.0482   0.3466   0.9965
   8.000   1.3035   0.01296   0.00685  -0.0489   0.3381   0.9972
   8.250   1.3315   0.01323   0.00712  -0.0496   0.3316   0.9980
   8.500   1.3593   0.01353   0.00744  -0.0503   0.3230   0.9988
   8.750   1.3862   0.01390   0.00780  -0.0510   0.3111   0.9995
   9.000   1.4094   0.01444   0.00828  -0.0511   0.2936   1.0000
   9.250   1.4260   0.01506   0.00882  -0.0499   0.2760   1.0000
   9.500   1.4388   0.01586   0.00953  -0.0482   0.2519   1.0000
   9.750   1.4402   0.01722   0.01068  -0.0451   0.2156   1.0000
  10.000   1.4213   0.01925   0.01248  -0.0392   0.1754   1.0000
  10.250   1.3958   0.02080   0.01402  -0.0321   0.1639   1.0000
  10.500   1.3720   0.02326   0.01640  -0.0268   0.1424   1.0000
  10.750   1.3452   0.02672   0.01972  -0.0228   0.1101   1.0000
  11.000   1.3217   0.03054   0.02342  -0.0200   0.0840   1.0000
  11.250   1.3086   0.03384   0.02669  -0.0183   0.0666   1.0000
  11.500   1.2883   0.03801   0.03077  -0.0168   0.0411   1.0000
  11.750   1.2725   0.04194   0.03467  -0.0156   0.0254   1.0000
  12.000   1.2681   0.04485   0.03761  -0.0149   0.0215   1.0000
  12.250   1.2672   0.04754   0.04036  -0.0144   0.0195   1.0000
  12.500   1.2673   0.05022   0.04310  -0.0141   0.0183   1.0000
  12.750   1.2667   0.05302   0.04597  -0.0139   0.0173   1.0000
  13.000   1.2654   0.05597   0.04899  -0.0137   0.0163   1.0000
  13.250   1.2664   0.05871   0.05180  -0.0136   0.0158   1.0000
  13.500   1.2669   0.06155   0.05472  -0.0135   0.0152   1.0000
  13.750   1.2670   0.06445   0.05770  -0.0135   0.0146   1.0000
  14.000   1.2662   0.06754   0.06085  -0.0136   0.0141   1.0000
  14.250   1.2646   0.07077   0.06415  -0.0138   0.0136   1.0000
  14.500   1.2621   0.07411   0.06756  -0.0139   0.0132   1.0000
  14.750   1.2578   0.07775   0.07128  -0.0142   0.0128   1.0000
  15.000   1.2559   0.08113   0.07474  -0.0145   0.0125   1.0000
  15.250   1.2556   0.08433   0.07802  -0.0149   0.0123   1.0000
  15.500   1.2541   0.08771   0.08147  -0.0153   0.0120   1.0000
  15.750   1.2530   0.09104   0.08488  -0.0157   0.0117   1.0000
  16.000   1.2509   0.09458   0.08850  -0.0163   0.0115   1.0000
  16.250   1.2499   0.09798   0.09196  -0.0168   0.0112   1.0000
  16.500   1.2478   0.10156   0.09561  -0.0174   0.0109   1.0000
  16.750   1.2467   0.10498   0.09911  -0.0181   0.0107   1.0000
  17.000   1.2446   0.10860   0.10279  -0.0188   0.0104   1.0000
  17.250   1.2430   0.11216   0.10641  -0.0196   0.0102   1.0000
  17.500   1.2395   0.11606   0.11037  -0.0206   0.0099   1.0000
<< Back to NACA M24 AIRFOIL (m24-il)

Polar data table (+)

Polar graphs


<< Back to NACA M24 AIRFOIL (m24-il)