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

E182 (8.47%) (e182-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: E182 (8.47%) (e182-il)
Reynolds number: 100,000
Max Cl/Cd: 48.5 at α=5.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e182-il-100000.txt
Download as CSV file: xf-e182-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: E182 (8.47%)                                    
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.5630   0.09066   0.08641   0.0064   1.0000   0.1046
  -7.750  -0.5861   0.08626   0.08207  -0.0055   1.0000   0.1079
  -7.500  -0.5760   0.08112   0.07699  -0.0048   1.0000   0.1103
  -7.250  -0.5609   0.07794   0.07384  -0.0031   1.0000   0.1142
  -7.000  -0.5735   0.07275   0.06839  -0.0145   1.0000   0.1227
  -6.750  -0.5518   0.06872   0.06457  -0.0109   1.0000   0.1259
  -6.500  -0.5488   0.06439   0.05998  -0.0166   1.0000   0.1372
  -6.250  -0.5294   0.06072   0.05647  -0.0148   1.0000   0.1412
  -6.000  -0.5174   0.05676   0.05240  -0.0171   1.0000   0.1536
  -5.750  -0.5020   0.05347   0.04904  -0.0180   1.0000   0.1682
  -5.500  -0.4793   0.04027   0.03412  -0.0233   1.0000   0.0819
  -5.250  -0.4566   0.03523   0.02855  -0.0227   1.0000   0.0667
  -5.000  -0.4402   0.03175   0.02488  -0.0220   1.0000   0.0641
  -4.750  -0.4335   0.02929   0.02207  -0.0191   1.0000   0.0610
  -4.500  -0.4264   0.02697   0.01918  -0.0159   1.0000   0.0578
  -4.250  -0.3954   0.02461   0.01623  -0.0164   0.9940   0.0561
  -4.000  -0.3521   0.02208   0.01338  -0.0193   0.9852   0.0568
  -3.750  -0.3092   0.02014   0.01133  -0.0221   0.9765   0.0601
  -3.500  -0.2649   0.01854   0.00965  -0.0253   0.9687   0.0710
  -3.250  -0.2253   0.01698   0.00818  -0.0274   0.9595   0.0951
  -3.000  -0.1997   0.01431   0.00706  -0.0278   0.9485   0.3403
  -2.750  -0.1546   0.01238   0.00720  -0.0286   0.9455   0.8421
  -2.500  -0.0358   0.01246   0.00674  -0.0436   0.9551   0.9897
  -2.250   0.0162   0.01230   0.00621  -0.0490   0.9443   1.0000
  -2.000   0.0445   0.01237   0.00604  -0.0497   0.9280   1.0000
  -1.750   0.0680   0.01251   0.00597  -0.0493   0.9130   1.0000
  -1.500   0.0898   0.01267   0.00595  -0.0485   0.8995   1.0000
  -1.250   0.1109   0.01286   0.00598  -0.0473   0.8875   1.0000
  -1.000   0.1320   0.01306   0.00603  -0.0462   0.8759   1.0000
  -0.750   0.1538   0.01328   0.00613  -0.0453   0.8648   1.0000
  -0.500   0.1755   0.01352   0.00626  -0.0443   0.8547   1.0000
  -0.250   0.1961   0.01375   0.00638  -0.0428   0.8458   1.0000
   0.000   0.2195   0.01401   0.00657  -0.0423   0.8356   1.0000
   0.250   0.2422   0.01429   0.00677  -0.0414   0.8264   1.0000
   0.500   0.2637   0.01454   0.00695  -0.0400   0.8182   1.0000
   0.750   0.2877   0.01484   0.00723  -0.0396   0.8084   1.0000
   1.000   0.3103   0.01514   0.00748  -0.0386   0.8001   1.0000
   1.250   0.3331   0.01541   0.00773  -0.0376   0.7913   1.0000
   1.500   0.3567   0.01575   0.00807  -0.0370   0.7818   1.0000
   1.750   0.3785   0.01599   0.00829  -0.0355   0.7740   1.0000
   2.000   0.4021   0.01630   0.00865  -0.0348   0.7636   1.0000
   2.250   0.4252   0.01662   0.00899  -0.0339   0.7535   1.0000
   2.500   0.4471   0.01682   0.00920  -0.0323   0.7446   1.0000
   2.750   0.4698   0.01705   0.00947  -0.0311   0.7339   1.0000
   3.000   0.4927   0.01730   0.00981  -0.0299   0.7222   1.0000
   3.250   0.5152   0.01748   0.01005  -0.0285   0.7104   1.0000
   3.500   0.5374   0.01758   0.01020  -0.0268   0.6985   1.0000
   3.750   0.5596   0.01760   0.01026  -0.0249   0.6863   1.0000
   4.000   0.5819   0.01755   0.01027  -0.0230   0.6733   1.0000
   4.250   0.6046   0.01750   0.01033  -0.0212   0.6584   1.0000
   4.500   0.6275   0.01740   0.01034  -0.0195   0.6416   1.0000
   4.750   0.6505   0.01720   0.01025  -0.0176   0.6237   1.0000
   5.000   0.6736   0.01688   0.01001  -0.0155   0.6052   1.0000
   5.250   0.6969   0.01657   0.00986  -0.0136   0.5821   1.0000
   5.500   0.7192   0.01584   0.00915  -0.0110   0.5469   1.0000
   5.750   0.7406   0.01527   0.00855  -0.0087   0.4799   1.0000
   6.000   0.7590   0.01565   0.00840  -0.0067   0.3509   1.0000
   6.250   0.7661   0.01866   0.00993  -0.0051   0.1340   1.0000
   6.500   0.7761   0.02160   0.01221  -0.0032   0.0678   1.0000
   6.750   0.7923   0.02337   0.01393  -0.0015   0.0584   1.0000
   7.000   0.8113   0.02495   0.01551   0.0000   0.0531   1.0000
   7.250   0.8313   0.02722   0.01765   0.0013   0.0500   1.0000
   7.500   0.8550   0.02922   0.01982   0.0025   0.0481   1.0000
   7.750   0.8780   0.03113   0.02201   0.0035   0.0455   1.0000
   8.000   0.9000   0.03340   0.02453   0.0045   0.0435   1.0000
   8.250   0.9204   0.03634   0.02784   0.0057   0.0435   1.0000
   8.500   0.9373   0.03974   0.03172   0.0070   0.0442   1.0000
   8.750   0.9500   0.04365   0.03612   0.0084   0.0454   1.0000
   9.000   0.9591   0.04779   0.04068   0.0096   0.0469   1.0000
   9.250   0.9654   0.05244   0.04563   0.0106   0.0484   1.0000
   9.500   0.9700   0.05625   0.04994   0.0121   0.0517   1.0000
   9.750   0.9455   0.06249   0.05685   0.0134   0.0566   1.0000
  10.000   0.9456   0.06795   0.06250   0.0139   0.0634   1.0000
  10.250   0.9100   0.07217   0.06700   0.0140   0.0648   1.0000
  10.500   0.8770   0.07788   0.07286   0.0111   0.0656   1.0000
  10.750   0.8466   0.08500   0.08008   0.0060   0.0661   1.0000
  11.000   0.8126   0.09552   0.09062  -0.0024   0.0675   1.0000
<< Back to E182 (8.47%) (e182-il)

Polar data table (+)

Polar graphs


<< Back to E182 (8.47%) (e182-il)