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

LDS-2 AIRFOIL (lds2-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: LDS-2 AIRFOIL (lds2-il)
Reynolds number: 50,000
Max Cl/Cd: 33.27 at α=6.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-lds2-il-50000-n5.txt
Download as CSV file: xf-lds2-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: LDS-2 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
  ------ -------- --------- --------- -------- -------- --------
 -10.250  -0.4425   0.09988   0.09234  -0.0396   1.0000   0.0616
 -10.000  -0.4465   0.09510   0.08763  -0.0418   1.0000   0.0604
  -9.500  -0.5101   0.07748   0.07008  -0.0560   1.0000   0.0550
  -9.250  -0.5215   0.07409   0.06672  -0.0558   1.0000   0.0547
  -9.000  -0.5384   0.07072   0.06335  -0.0548   1.0000   0.0545
  -8.750  -0.5529   0.06739   0.05996  -0.0535   1.0000   0.0544
  -8.500  -0.5652   0.06419   0.05668  -0.0517   1.0000   0.0542
  -8.250  -0.5753   0.06113   0.05349  -0.0495   1.0000   0.0540
  -8.000  -0.5834   0.05824   0.05045  -0.0470   1.0000   0.0538
  -7.750  -0.5900   0.05544   0.04745  -0.0442   1.0000   0.0537
  -7.500  -0.5948   0.05272   0.04448  -0.0413   1.0000   0.0537
  -7.250  -0.5964   0.05017   0.04165  -0.0384   1.0000   0.0537
  -7.000  -0.5958   0.04765   0.03881  -0.0356   1.0000   0.0539
  -6.750  -0.5927   0.04527   0.03605  -0.0329   1.0000   0.0544
  -6.500  -0.5873   0.04315   0.03346  -0.0302   1.0000   0.0552
  -6.250  -0.5778   0.04108   0.03134  -0.0284   1.0000   0.0566
  -6.000  -0.5665   0.03947   0.02963  -0.0265   1.0000   0.0583
  -5.750  -0.5447   0.03751   0.02736  -0.0264   0.9967   0.0600
  -5.500  -0.5132   0.03538   0.02483  -0.0276   0.9903   0.0616
  -5.250  -0.4795   0.03360   0.02264  -0.0290   0.9843   0.0645
  -5.000  -0.4483   0.03205   0.02089  -0.0300   0.9776   0.0680
  -4.750  -0.4153   0.03071   0.01947  -0.0312   0.9719   0.0719
  -4.500  -0.3837   0.02960   0.01817  -0.0319   0.9651   0.0766
  -4.250  -0.3518   0.02853   0.01708  -0.0330   0.9591   0.0831
  -4.000  -0.3229   0.02766   0.01606  -0.0333   0.9517   0.0905
  -3.750  -0.2901   0.02671   0.01510  -0.0347   0.9458   0.1025
  -3.500  -0.2639   0.02584   0.01420  -0.0349   0.9377   0.1186
  -3.250  -0.2317   0.02449   0.01320  -0.0366   0.9321   0.1675
  -3.000  -0.2168   0.02272   0.01299  -0.0351   0.9241   0.4352
  -2.750  -0.1925   0.02243   0.01312  -0.0337   0.9173   0.5811
  -2.500  -0.1746   0.02232   0.01345  -0.0304   0.9099   0.6876
  -2.250  -0.1514   0.02247   0.01396  -0.0268   0.9038   0.7922
  -2.000  -0.1063   0.02262   0.01397  -0.0290   0.9000   0.8380
  -1.750  -0.0642   0.02274   0.01388  -0.0316   0.8941   0.8682
  -1.500  -0.0136   0.02287   0.01380  -0.0361   0.8893   0.8981
  -1.250   0.0461   0.02297   0.01366  -0.0424   0.8860   0.9249
  -1.000   0.1084   0.02300   0.01347  -0.0494   0.8832   0.9489
  -0.750   0.1603   0.02307   0.01338  -0.0549   0.8772   0.9748
  -0.500   0.2189   0.02304   0.01318  -0.0618   0.8727   0.9993
  -0.250   0.2446   0.02304   0.01304  -0.0623   0.8652   1.0000
   0.000   0.2587   0.02320   0.01309  -0.0607   0.8546   1.0000
   0.250   0.2788   0.02338   0.01315  -0.0599   0.8456   1.0000
   0.500   0.3032   0.02352   0.01319  -0.0597   0.8371   1.0000
   0.750   0.3211   0.02380   0.01338  -0.0583   0.8275   1.0000
   1.000   0.3490   0.02395   0.01346  -0.0585   0.8201   1.0000
   1.250   0.3649   0.02433   0.01378  -0.0567   0.8103   1.0000
   1.500   0.3940   0.02449   0.01391  -0.0570   0.8035   1.0000
   1.750   0.4089   0.02494   0.01432  -0.0550   0.7935   1.0000
   2.000   0.4391   0.02510   0.01447  -0.0554   0.7870   1.0000
   2.250   0.4538   0.02557   0.01494  -0.0534   0.7763   1.0000
   2.500   0.4872   0.02556   0.01493  -0.0539   0.7686   1.0000
   2.750   0.5078   0.02575   0.01514  -0.0524   0.7561   1.0000
   3.000   0.5289   0.02593   0.01535  -0.0509   0.7435   1.0000
   3.250   0.5547   0.02597   0.01542  -0.0500   0.7320   1.0000
   3.500   0.5886   0.02570   0.01520  -0.0500   0.7217   1.0000
   3.750   0.6110   0.02568   0.01526  -0.0484   0.7069   1.0000
   4.000   0.6356   0.02552   0.01514  -0.0469   0.6913   1.0000
   4.250   0.6608   0.02530   0.01498  -0.0455   0.6751   1.0000
   4.500   0.6864   0.02504   0.01478  -0.0440   0.6582   1.0000
   4.750   0.7060   0.02498   0.01481  -0.0418   0.6387   1.0000
   5.000   0.7259   0.02487   0.01478  -0.0396   0.6172   1.0000
   5.250   0.7506   0.02451   0.01445  -0.0378   0.5941   1.0000
   5.500   0.7669   0.02452   0.01455  -0.0351   0.5660   1.0000
   5.750   0.7850   0.02448   0.01453  -0.0327   0.5341   1.0000
   6.000   0.8013   0.02454   0.01457  -0.0300   0.4952   1.0000
   6.250   0.8178   0.02467   0.01456  -0.0274   0.4474   1.0000
   6.500   0.8331   0.02504   0.01462  -0.0246   0.3943   1.0000
   6.750   0.8433   0.02586   0.01509  -0.0217   0.3430   1.0000
   7.000   0.8516   0.02694   0.01582  -0.0188   0.3031   1.0000
   7.250   0.8601   0.02812   0.01676  -0.0162   0.2713   1.0000
   7.500   0.8689   0.02930   0.01780  -0.0137   0.2451   1.0000
   7.750   0.8786   0.03048   0.01885  -0.0114   0.2246   1.0000
   8.000   0.8895   0.03168   0.01992  -0.0094   0.2073   1.0000
   8.250   0.9025   0.03287   0.02104  -0.0076   0.1920   1.0000
   8.500   0.9183   0.03404   0.02221  -0.0062   0.1787   1.0000
   8.750   0.9353   0.03523   0.02343  -0.0050   0.1667   1.0000
   9.000   0.9548   0.03644   0.02463  -0.0041   0.1566   1.0000
   9.250   0.9763   0.03765   0.02586  -0.0035   0.1475   1.0000
   9.500   0.9972   0.03896   0.02734  -0.0028   0.1388   1.0000
   9.750   1.0198   0.04026   0.02858  -0.0024   0.1319   1.0000
  10.000   1.0408   0.04185   0.03044  -0.0018   0.1253   1.0000
  10.250   1.0596   0.04329   0.03193  -0.0010   0.1196   1.0000
  10.500   1.0782   0.04504   0.03386  -0.0003   0.1146   1.0000
  10.750   1.0926   0.04698   0.03607   0.0008   0.1102   1.0000
  11.000   1.1076   0.04869   0.03788   0.0018   0.1061   1.0000
  11.250   1.1213   0.05068   0.03996   0.0027   0.1025   1.0000
  11.500   1.1215   0.05315   0.04281   0.0049   0.0996   1.0000
  11.750   1.1228   0.05571   0.04567   0.0066   0.0971   1.0000
  12.000   1.1244   0.05824   0.04843   0.0081   0.0948   1.0000
  12.250   1.1298   0.06050   0.05079   0.0092   0.0925   1.0000
  12.500   1.1335   0.06309   0.05346   0.0102   0.0903   1.0000
  12.750   1.1136   0.06700   0.05773   0.0117   0.0893   1.0000
  13.000   1.0904   0.07155   0.06260   0.0123   0.0885   1.0000
  13.250   1.0634   0.07695   0.06828   0.0119   0.0880   1.0000
  13.500   1.0310   0.08361   0.07519   0.0100   0.0879   1.0000
  13.750   0.9893   0.09250   0.08428   0.0060   0.0882   1.0000
  14.000   0.9393   0.10465   0.09655  -0.0008   0.0889   1.0000
<< Back to LDS-2 AIRFOIL (lds2-il)

Polar data table (+)

Polar graphs


<< Back to LDS-2 AIRFOIL (lds2-il)