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=50,000 Ncrit=5


Details Polar file
Airfoil: NACA M24 AIRFOIL (m24-il)
Reynolds number: 50,000
Max Cl/Cd: 19.91 at α=3.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-m24-il-50000-n5.txt
Download as CSV file: xf-m24-il-50000-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 =     0.050 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.2515   0.11582   0.10963  -0.0264   0.7946   0.0960
  -8.750  -0.2560   0.11474   0.10854  -0.0306   0.7875   0.0969
  -8.500  -0.2610   0.11366   0.10746  -0.0346   0.7799   0.0972
  -8.250  -0.2477   0.10837   0.10218  -0.0343   0.7733   0.0981
  -8.000  -0.2284   0.10308   0.09683  -0.0323   0.7671   0.1002
  -7.750  -0.2187   0.09982   0.09356  -0.0331   0.7598   0.1024
  -7.500  -0.2149   0.09727   0.09097  -0.0336   0.7541   0.1046
  -7.250  -0.2088   0.09479   0.08846  -0.0357   0.7473   0.1077
  -7.000  -0.2063   0.09348   0.08707  -0.0396   0.7410   0.1116
  -6.750  -0.2044   0.09382   0.08717  -0.0444   0.7359   0.1131
  -6.500  -0.1883   0.08734   0.08085  -0.0424   0.7290   0.1153
  -6.250  -0.1765   0.08404   0.07753  -0.0419   0.7233   0.1185
  -6.000  -0.1656   0.08149   0.07489  -0.0426   0.7183   0.1222
  -5.750  -0.1523   0.08047   0.07369  -0.0464   0.7114   0.1280
  -5.500  -0.1394   0.07857   0.07160  -0.0482   0.7062   0.1297
  -5.250  -0.1268   0.07432   0.06737  -0.0471   0.7012   0.1317
  -5.000  -0.1117   0.07151   0.06454  -0.0474   0.6946   0.1350
  -4.750  -0.0966   0.06936   0.06225  -0.0477   0.6896   0.1402
  -4.500  -0.0789   0.06805   0.06064  -0.0492   0.6848   0.1465
  -4.250  -0.0639   0.06487   0.05752  -0.0488   0.6782   0.1510
  -3.750  -0.0304   0.06069   0.05302  -0.0484   0.6691   0.1673
  -3.250   0.0071   0.05706   0.04912  -0.0487   0.6571   0.1829
  -3.000   0.0493   0.05365   0.04484  -0.0502   0.6536   0.0948
  -2.500   0.0963   0.04984   0.04038  -0.0500   0.6416   0.0811
  -2.250   0.1183   0.04803   0.03833  -0.0491   0.6378   0.0804
  -2.000   0.1408   0.04684   0.03690  -0.0490   0.6316   0.0807
  -1.750   0.1639   0.04558   0.03538  -0.0485   0.6259   0.0808
  -1.500   0.1882   0.04409   0.03359  -0.0477   0.6219   0.0804
  -1.250   0.2116   0.04301   0.03228  -0.0473   0.6166   0.0796
  -1.000   0.2350   0.04214   0.03119  -0.0470   0.6103   0.0790
  -0.750   0.2613   0.04095   0.02968  -0.0464   0.6062   0.0786
  -0.500   0.2889   0.03984   0.02824  -0.0459   0.6025   0.0786
  -0.250   0.3120   0.03967   0.02790  -0.0462   0.5948   0.0792
   0.000   0.3405   0.03892   0.02681  -0.0459   0.5905   0.0815
   0.250   0.3727   0.03795   0.02562  -0.0461   0.5874   0.0846
   0.500   0.4022   0.03832   0.02590  -0.0480   0.5792   0.0870
   0.750   0.4409   0.03780   0.02514  -0.0498   0.5747   0.0899
   1.000   0.4788   0.03706   0.02415  -0.0508   0.5716   0.0939
   1.250   0.5002   0.03780   0.02485  -0.0512   0.5639   0.0985
   1.500   0.5242   0.03781   0.02486  -0.0508   0.5590   0.1074
   1.750   0.5495   0.03742   0.02438  -0.0499   0.5559   0.1183
   2.000   0.5620   0.03847   0.02543  -0.0489   0.5487   0.1299
   2.250   0.5785   0.03879   0.02583  -0.0476   0.5434   0.1559
   2.500   0.5981   0.03834   0.02563  -0.0460   0.5402   0.2301
   2.750   0.7287   0.03760   0.02607  -0.0661   0.5348   1.0000
   3.000   0.7398   0.03906   0.02744  -0.0649   0.5281   1.0000
   3.250   0.7606   0.03940   0.02762  -0.0637   0.5246   1.0000
   3.500   0.7848   0.03942   0.02746  -0.0625   0.5222   1.0000
   3.750   0.7749   0.04277   0.03091  -0.0606   0.5120   1.0000
   4.000   0.7946   0.04320   0.03123  -0.0593   0.5087   1.0000
   4.250   0.8049   0.04450   0.03247  -0.0577   0.5039   1.0000
   4.500   0.7940   0.04757   0.03557  -0.0552   0.4951   1.0000
   4.750   0.8149   0.04787   0.03579  -0.0539   0.4923   1.0000
   5.000   0.8359   0.04821   0.03605  -0.0527   0.4898   1.0000
   5.500   0.8184   0.05403   0.04189  -0.0482   0.4754   1.0000
   6.000   0.7880   0.06047   0.04831  -0.0427   0.4609   1.0000
   6.250   0.7988   0.06182   0.04962  -0.0414   0.4575   1.0000
   6.500   0.7726   0.06633   0.05416  -0.0392   0.4482   1.0000
   6.750   0.7903   0.06716   0.05497  -0.0381   0.4452   1.0000
   7.000   0.8127   0.06761   0.05540  -0.0372   0.4433   1.0000
   7.250   0.7778   0.07340   0.06122  -0.0357   0.4336   1.0000
   7.500   0.7933   0.07455   0.06237  -0.0349   0.4303   1.0000
   7.750   0.8151   0.07513   0.06294  -0.0341   0.4281   1.0000
   8.000   0.7892   0.08034   0.06820  -0.0332   0.4198   1.0000
   8.250   0.7992   0.08215   0.07003  -0.0326   0.4161   1.0000
   8.500   0.8183   0.08308   0.07099  -0.0319   0.4134   1.0000
   8.750   0.8065   0.08703   0.07498  -0.0314   0.4068   1.0000
   9.000   0.8097   0.08960   0.07760  -0.0309   0.4021   1.0000
   9.250   0.8253   0.09096   0.07899  -0.0303   0.3991   1.0000
   9.500   0.8472   0.09168   0.07977  -0.0297   0.3968   1.0000
   9.750   0.8242   0.09686   0.08501  -0.0297   0.3884   1.0000
  10.000   0.8362   0.09861   0.08681  -0.0292   0.3847   1.0000
  10.250   0.8565   0.09951   0.08778  -0.0287   0.3821   1.0000
  10.500   0.8419   0.10394   0.09227  -0.0287   0.3747   1.0000
  10.750   0.8504   0.10605   0.09445  -0.0285   0.3702   1.0000
  11.000   0.8690   0.10714   0.09563  -0.0280   0.3672   1.0000
  11.250   0.8615   0.11094   0.09950  -0.0281   0.3607   1.0000
  11.500   0.8659   0.11351   0.10216  -0.0280   0.3556   1.0000
  11.750   0.8827   0.11482   0.10356  -0.0277   0.3523   1.0000
<< Back to NACA M24 AIRFOIL (m24-il)

Polar data table (+)

Polar graphs


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