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=100,000 Ncrit=5


Details Polar file
Airfoil: NACA M24 AIRFOIL (m24-il)
Reynolds number: 100,000
Max Cl/Cd: 45.23 at α=6.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-m24-il-100000-n5.txt
Download as CSV file: xf-m24-il-100000-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 =     0.100 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.2700   0.09977   0.09461  -0.0133   0.6902   0.0554
  -7.750  -0.2678   0.09692   0.09172  -0.0166   0.6862   0.0568
  -7.500  -0.2690   0.09496   0.08974  -0.0213   0.6810   0.0578
  -7.250  -0.2644   0.09324   0.08789  -0.0273   0.6763   0.0583
  -7.000  -0.2547   0.08845   0.08310  -0.0264   0.6720   0.0590
  -6.750  -0.2428   0.08460   0.07924  -0.0248   0.6675   0.0601
  -6.500  -0.2308   0.08157   0.07619  -0.0254   0.6621   0.0616
  -6.250  -0.2188   0.07870   0.07325  -0.0267   0.6576   0.0631
  -6.000  -0.2059   0.07592   0.07035  -0.0283   0.6539   0.0650
  -5.750  -0.1861   0.07405   0.06828  -0.0325   0.6490   0.0683
  -5.500  -0.1632   0.07288   0.06676  -0.0361   0.6442   0.0692
  -5.250  -0.1507   0.06834   0.06220  -0.0361   0.6401   0.0698
  -5.000  -0.1387   0.06458   0.05844  -0.0351   0.6366   0.0708
  -4.750  -0.1226   0.06182   0.05568  -0.0350   0.6313   0.0728
  -4.500  -0.1038   0.05947   0.05324  -0.0354   0.6265   0.0757
  -4.250  -0.0803   0.05761   0.05113  -0.0365   0.6225   0.0801
  -4.000  -0.0484   0.05767   0.05059  -0.0378   0.6188   0.0824
  -3.750  -0.0317   0.05349   0.04644  -0.0380   0.6138   0.0831
  -3.500  -0.0138   0.05040   0.04331  -0.0378   0.6094   0.0840
  -3.250   0.0061   0.04797   0.04076  -0.0374   0.6056   0.0849
  -3.000   0.0280   0.04584   0.03847  -0.0372   0.6018   0.0858
  -2.500   0.0834   0.03943   0.03121  -0.0364   0.5929   0.0578
  -2.250   0.1057   0.03756   0.02919  -0.0360   0.5892   0.0565
  -2.000   0.1302   0.03568   0.02707  -0.0355   0.5856   0.0548
  -1.750   0.1563   0.03376   0.02491  -0.0352   0.5804   0.0532
  -1.500   0.1833   0.03163   0.02240  -0.0345   0.5762   0.0516
  -1.250   0.2108   0.02961   0.01992  -0.0338   0.5727   0.0506
  -1.000   0.2382   0.02821   0.01820  -0.0334   0.5690   0.0506
  -0.750   0.2653   0.02751   0.01743  -0.0336   0.5637   0.0521
  -0.500   0.2930   0.02662   0.01632  -0.0334   0.5593   0.0537
  -0.250   0.3217   0.02557   0.01497  -0.0332   0.5559   0.0542
   0.000   0.3507   0.02470   0.01390  -0.0333   0.5517   0.0545
   0.250   0.3798   0.02400   0.01306  -0.0335   0.5466   0.0552
   0.500   0.4090   0.02334   0.01225  -0.0336   0.5424   0.0561
   0.750   0.4383   0.02275   0.01149  -0.0336   0.5391   0.0573
   1.000   0.4665   0.02243   0.01117  -0.0339   0.5345   0.0595
   1.250   0.4954   0.02226   0.01107  -0.0343   0.5296   0.0635
   1.500   0.5267   0.02198   0.01070  -0.0349   0.5256   0.0677
   1.750   0.5557   0.02167   0.01029  -0.0350   0.5225   0.0717
   2.000   0.5819   0.02171   0.01036  -0.0350   0.5173   0.0789
   2.250   0.6075   0.02166   0.01033  -0.0347   0.5129   0.0942
   2.500   0.6325   0.02150   0.01027  -0.0342   0.5093   0.1404
   3.000   0.8053   0.02063   0.01103  -0.0586   0.4978   1.0000
   3.250   0.8298   0.02082   0.01111  -0.0582   0.4941   1.0000
   3.500   0.8545   0.02094   0.01110  -0.0576   0.4912   1.0000
   3.750   0.8781   0.02131   0.01147  -0.0573   0.4866   1.0000
   4.000   0.9016   0.02166   0.01182  -0.0570   0.4819   1.0000
   4.250   0.9254   0.02186   0.01196  -0.0564   0.4782   1.0000
   4.500   0.9497   0.02199   0.01200  -0.0559   0.4753   1.0000
   4.750   0.9720   0.02247   0.01254  -0.0555   0.4710   1.0000
   5.000   0.9941   0.02293   0.01304  -0.0550   0.4665   1.0000
   5.250   1.0170   0.02320   0.01330  -0.0544   0.4629   1.0000
   5.500   1.0408   0.02336   0.01341  -0.0537   0.4600   1.0000
   5.750   1.0620   0.02384   0.01395  -0.0531   0.4559   1.0000
   6.000   1.0819   0.02443   0.01463  -0.0524   0.4513   1.0000
   6.250   1.1034   0.02482   0.01504  -0.0517   0.4478   1.0000
   6.500   1.1263   0.02506   0.01529  -0.0510   0.4450   1.0000
   6.750   1.1484   0.02539   0.01562  -0.0503   0.4421   1.0000
   7.000   1.1637   0.02629   0.01670  -0.0492   0.4371   1.0000
   7.250   1.1826   0.02681   0.01730  -0.0482   0.4332   1.0000
   7.500   1.2041   0.02711   0.01763  -0.0474   0.4302   1.0000
   7.750   1.2276   0.02732   0.01783  -0.0468   0.4279   1.0000
   8.000   1.2382   0.02846   0.01917  -0.0451   0.4235   1.0000
   8.250   1.2509   0.02936   0.02022  -0.0436   0.4193   1.0000
   8.500   1.2688   0.02985   0.02077  -0.0424   0.4160   1.0000
   8.750   1.2915   0.03005   0.02100  -0.0417   0.4135   1.0000
   9.000   1.3055   0.03084   0.02191  -0.0402   0.4103   1.0000
   9.250   1.3019   0.03262   0.02391  -0.0371   0.4054   1.0000
   9.500   1.3106   0.03358   0.02498  -0.0351   0.4018   1.0000
   9.750   1.3316   0.03380   0.02526  -0.0342   0.3989   1.0000
  10.000   1.3483   0.03414   0.02567  -0.0328   0.3949   1.0000
  10.250   1.2843   0.03902   0.03072  -0.0247   0.3889   1.0000
  10.500   1.3046   0.03895   0.03071  -0.0236   0.3846   1.0000
  10.750   1.3273   0.03846   0.03027  -0.0225   0.3786   1.0000
  11.000   1.2067   0.05189   0.04382  -0.0181   0.3689   1.0000
  11.250   1.1998   0.05503   0.04702  -0.0172   0.3629   1.0000
  11.500   1.1592   0.06254   0.05459  -0.0169   0.3525   1.0000
  11.750   1.1747   0.06300   0.05513  -0.0161   0.3483   1.0000
  12.250   1.3020   0.05084   0.04309  -0.0136   0.3352   1.0000
  12.750   1.0752   0.08835   0.08075  -0.0175   0.3107   1.0000
  13.750   1.1433   0.08774   0.08053  -0.0148   0.2833   1.0000
  14.000   1.1898   0.08305   0.07597  -0.0134   0.2761   1.0000
  14.250   1.1513   0.09205   0.08504  -0.0149   0.2663   1.0000
  14.500   1.1605   0.09340   0.08650  -0.0148   0.2565   1.0000
  14.750   1.1653   0.09551   0.08870  -0.0150   0.2444   1.0000
  15.000   1.1619   0.09907   0.09233  -0.0156   0.2313   1.0000
  15.250   1.1548   0.10331   0.09666  -0.0164   0.2170   1.0000
  15.500   1.1507   0.10707   0.10046  -0.0172   0.1980   1.0000
  15.750   1.1680   0.10624   0.09880  -0.0163   0.1217   1.0000
  16.000   1.1457   0.11272   0.10491  -0.0178   0.0901   1.0000
  16.250   1.1297   0.11855   0.11058  -0.0192   0.0733   1.0000
<< Back to NACA M24 AIRFOIL (m24-il)

Polar data table (+)

Polar graphs


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