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

DSMA-523B AIRFOIL (dsma523b-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: DSMA-523B AIRFOIL (dsma523b-il)
Reynolds number: 100,000
Max Cl/Cd: 27.86 at α=5.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-dsma523b-il-100000-n5.txt
Download as CSV file: xf-dsma523b-il-100000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: DSMA-523B 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
  ------ -------- --------- --------- -------- -------- --------
 -12.000  -0.7187   0.09047   0.08410  -0.0407   1.0000   0.0453
 -11.750  -0.7593   0.07974   0.07322  -0.0474   1.0000   0.0450
 -11.500  -0.7942   0.07248   0.06580  -0.0506   1.0000   0.0448
 -11.250  -0.8254   0.06715   0.06033  -0.0513   1.0000   0.0447
 -11.000  -0.8545   0.06320   0.05623  -0.0500   1.0000   0.0446
 -10.750  -0.8759   0.05879   0.05160  -0.0501   1.0000   0.0446
 -10.500  -0.8869   0.05410   0.04653  -0.0519   1.0000   0.0449
 -10.250  -0.8892   0.04961   0.04150  -0.0537   1.0000   0.0454
 -10.000  -0.8736   0.04792   0.03981  -0.0535   1.0000   0.0461
  -9.750  -0.8585   0.04589   0.03765  -0.0538   1.0000   0.0470
  -9.500  -0.8424   0.04350   0.03500  -0.0546   1.0000   0.0481
  -9.250  -0.8240   0.04092   0.03206  -0.0557   1.0000   0.0492
  -9.000  -0.8029   0.03839   0.02909  -0.0568   1.0000   0.0503
  -8.750  -0.7797   0.03623   0.02652  -0.0577   1.0000   0.0512
  -8.500  -0.7575   0.03473   0.02499  -0.0576   1.0000   0.0522
  -8.250  -0.7342   0.03333   0.02350  -0.0576   1.0000   0.0534
  -8.000  -0.7095   0.03200   0.02196  -0.0579   1.0000   0.0555
  -7.750  -0.6839   0.03070   0.02039  -0.0583   1.0000   0.0575
  -7.500  -0.6609   0.02954   0.01928  -0.0578   1.0000   0.0593
  -7.250  -0.6368   0.02848   0.01817  -0.0575   1.0000   0.0611
  -7.000  -0.6116   0.02753   0.01707  -0.0573   1.0000   0.0631
  -6.750  -0.5881   0.02652   0.01617  -0.0568   1.0000   0.0655
  -6.500  -0.5630   0.02568   0.01536  -0.0567   1.0000   0.0684
  -6.250  -0.5371   0.02488   0.01456  -0.0566   1.0000   0.0711
  -6.000  -0.5101   0.02409   0.01388  -0.0570   1.0000   0.0738
  -5.750  -0.4818   0.02345   0.01326  -0.0576   1.0000   0.0777
  -5.500  -0.4524   0.02284   0.01276  -0.0586   1.0000   0.0820
  -5.000  -0.3928   0.02193   0.01193  -0.0604   1.0000   0.0931
  -4.750  -0.3626   0.02153   0.01160  -0.0613   1.0000   0.1010
  -4.500  -0.3321   0.02117   0.01131  -0.0622   1.0000   0.1123
  -4.250  -0.3009   0.02077   0.01109  -0.0633   1.0000   0.1317
  -4.000  -0.2685   0.02035   0.01092  -0.0648   1.0000   0.1686
  -3.750  -0.2320   0.01980   0.01076  -0.0674   1.0000   0.2358
  -3.500  -0.1867   0.01899   0.01062  -0.0725   1.0000   0.3582
  -3.250  -0.1649   0.01907   0.01146  -0.0706   1.0000   0.4763
  -3.000  -0.1692   0.02019   0.01299  -0.0613   1.0000   0.5287
  -2.500  -0.1101   0.02084   0.01355  -0.0621   1.0000   0.6128
  -2.250  -0.0798   0.02117   0.01381  -0.0628   1.0000   0.6342
  -2.000  -0.0508   0.02153   0.01412  -0.0631   1.0000   0.6509
  -1.750  -0.0260   0.02197   0.01454  -0.0622   1.0000   0.6647
  -1.500  -0.0094   0.02249   0.01511  -0.0592   1.0000   0.6755
  -1.250   0.0113   0.02300   0.01564  -0.0572   1.0000   0.6898
  -1.000   0.0341   0.02347   0.01614  -0.0559   1.0000   0.7042
  -0.750   0.0465   0.02380   0.01656  -0.0519   1.0000   0.7112
  -0.500   0.0704   0.02404   0.01684  -0.0511   1.0000   0.7191
  -0.250   0.0971   0.02425   0.01708  -0.0511   1.0000   0.7261
   0.000   0.1155   0.02444   0.01736  -0.0490   1.0000   0.7312
   0.250   0.1667   0.02454   0.01750  -0.0535   0.9897   0.7398
   0.500   0.2367   0.02400   0.01703  -0.0608   0.9669   0.7468
   0.750   0.2734   0.02289   0.01601  -0.0612   0.9337   0.7513
   1.000   0.3219   0.02184   0.01503  -0.0641   0.8998   0.7547
   1.250   0.3452   0.02150   0.01478  -0.0628   0.8511   0.7571
   1.500   0.4748   0.02198   0.01248  -0.0809   0.2836   0.7576
   1.750   0.5056   0.02291   0.01275  -0.0821   0.1711   0.7600
   2.000   0.5377   0.02356   0.01309  -0.0835   0.1316   0.7618
   2.250   0.5636   0.02396   0.01338  -0.0832   0.1154   0.7629
   2.500   0.5889   0.02436   0.01375  -0.0826   0.1057   0.7644
   2.750   0.6144   0.02484   0.01418  -0.0821   0.0985   0.7660
   3.000   0.6411   0.02536   0.01468  -0.0820   0.0925   0.7673
   3.250   0.6688   0.02586   0.01518  -0.0821   0.0870   0.7684
   3.500   0.6956   0.02658   0.01584  -0.0820   0.0832   0.7699
   3.750   0.7239   0.02728   0.01657  -0.0822   0.0798   0.7715
   4.000   0.7532   0.02793   0.01726  -0.0826   0.0759   0.7729
   4.250   0.7825   0.02865   0.01795  -0.0832   0.0725   0.7738
   4.500   0.8124   0.02975   0.01895  -0.0839   0.0702   0.7748
   4.750   0.8432   0.03062   0.01998  -0.0845   0.0679   0.7757
   5.000   0.8733   0.03152   0.02100  -0.0851   0.0650   0.7766
   5.250   0.9028   0.03240   0.02193  -0.0857   0.0625   0.7776
   5.500   0.9325   0.03354   0.02304  -0.0865   0.0608   0.7788
   5.750   0.9627   0.03501   0.02467  -0.0871   0.0593   0.7802
   6.000   0.9889   0.03638   0.02637  -0.0868   0.0574   0.7812
   6.250   1.0138   0.03763   0.02784  -0.0863   0.0554   0.7820
   6.500   1.0383   0.03879   0.02914  -0.0860   0.0537   0.7827
   6.750   1.0624   0.04008   0.03053  -0.0856   0.0526   0.7835
   7.000   1.0858   0.04173   0.03229  -0.0853   0.0517   0.7842
   7.250   1.1046   0.04415   0.03520  -0.0839   0.0508   0.7850
   7.500   1.1201   0.04695   0.03851  -0.0821   0.0500   0.7857
   7.750   1.1323   0.05003   0.04210  -0.0800   0.0492   0.7865
   8.000   1.1417   0.05317   0.04570  -0.0778   0.0483   0.7873
   8.250   1.1490   0.05625   0.04918  -0.0754   0.0474   0.7881
   8.500   1.1562   0.05893   0.05215  -0.0733   0.0464   0.7892
   8.750   1.1633   0.06138   0.05482  -0.0713   0.0457   0.7903
   9.000   1.1681   0.06399   0.05763  -0.0692   0.0452   0.7915
   9.250   1.1690   0.06694   0.06079  -0.0669   0.0449   0.7925
   9.500   1.1613   0.07052   0.06464  -0.0640   0.0447   0.7932
   9.750   1.1423   0.07441   0.06879  -0.0602   0.0446   0.7938
  10.000   1.1132   0.07932   0.07401  -0.0567   0.0446   0.7942
  10.250   0.9511   0.11459   0.11013  -0.0761   0.0468   0.7933
<< Back to DSMA-523B AIRFOIL (dsma523b-il)

Polar data table (+)

Polar graphs


<< Back to DSMA-523B AIRFOIL (dsma523b-il)