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

USA 41 AIRFOIL (usa41-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: USA 41 AIRFOIL (usa41-il)
Reynolds number: 100,000
Max Cl/Cd: 59.84 at α=7°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-usa41-il-100000.txt
Download as CSV file: xf-usa41-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: USA 41 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.3437   0.09995   0.09508  -0.0259   1.0000   0.0714
  -7.750  -0.3564   0.09980   0.09507  -0.0273   1.0000   0.0721
  -7.500  -0.3597   0.09889   0.09427  -0.0328   1.0000   0.0726
  -7.250  -0.3388   0.09113   0.08644  -0.0239   1.0000   0.0768
  -7.000  -0.3377   0.08889   0.08427  -0.0232   1.0000   0.0804
  -6.750  -0.3398   0.08725   0.08273  -0.0253   1.0000   0.0844
  -6.500  -0.3400   0.08676   0.08230  -0.0345   1.0000   0.0861
  -6.250  -0.3389   0.08197   0.07760  -0.0295   1.0000   0.0877
  -6.000  -0.3351   0.07903   0.07470  -0.0253   1.0000   0.0904
  -5.750  -0.3320   0.07670   0.07241  -0.0248   1.0000   0.0939
  -5.500  -0.3132   0.07494   0.07052  -0.0398   1.0000   0.1004
  -5.250  -0.3171   0.07096   0.06668  -0.0324   1.0000   0.1018
  -5.000  -0.3144   0.06844   0.06422  -0.0286   1.0000   0.1049
  -4.750  -0.2864   0.06542   0.06099  -0.0400   1.0000   0.1147
  -4.500  -0.2888   0.06244   0.05817  -0.0337   1.0000   0.1171
  -4.250  -0.2619   0.05963   0.05516  -0.0408   1.0000   0.1289
  -4.000  -0.2600   0.05683   0.05246  -0.0365   1.0000   0.1317
  -3.750  -0.2317   0.05386   0.04929  -0.0423   1.0000   0.1433
  -3.500  -0.2046   0.05233   0.04749  -0.0460   1.0000   0.1560
  -3.250  -0.1972   0.04836   0.04369  -0.0439   1.0000   0.1583
  -3.000  -0.1725   0.04612   0.04126  -0.0466   1.0000   0.1716
  -2.750  -0.1359   0.04283   0.03788  -0.0505   0.9953   0.1865
  -2.500  -0.0893   0.03944   0.03429  -0.0563   0.9877   0.2013
  -2.250  -0.0156   0.03240   0.02616  -0.0666   0.9837   0.1305
  -2.000   0.0367   0.02666   0.01929  -0.0707   0.9783   0.0984
  -1.750   0.0847   0.02395   0.01607  -0.0746   0.9716   0.0938
  -1.500   0.1290   0.02202   0.01363  -0.0774   0.9621   0.0923
  -1.250   0.1751   0.02083   0.01213  -0.0806   0.9527   0.0967
  -1.000   0.2233   0.01969   0.01082  -0.0841   0.9437   0.1018
  -0.750   0.2659   0.01881   0.01001  -0.0866   0.9319   0.1083
  -0.500   0.3109   0.01803   0.00930  -0.0895   0.9207   0.1233
  -0.250   0.3619   0.01723   0.00878  -0.0933   0.9123   0.2228
   0.000   0.4091   0.01624   0.00820  -0.0969   0.9015   0.3388
   0.250   0.4644   0.01438   0.00778  -0.1017   0.8912   1.0000
   0.500   0.5048   0.01420   0.00736  -0.1033   0.8765   1.0000
   0.750   0.5413   0.01404   0.00703  -0.1041   0.8604   1.0000
   1.000   0.5752   0.01389   0.00672  -0.1043   0.8434   1.0000
   1.250   0.6054   0.01381   0.00651  -0.1039   0.8248   1.0000
   1.500   0.6328   0.01379   0.00637  -0.1030   0.8039   1.0000
   1.750   0.6614   0.01375   0.00621  -0.1023   0.7849   1.0000
   2.000   0.6878   0.01382   0.00617  -0.1013   0.7640   1.0000
   2.250   0.7144   0.01392   0.00615  -0.1004   0.7438   1.0000
   2.500   0.7414   0.01406   0.00618  -0.0996   0.7247   1.0000
   2.750   0.7665   0.01429   0.00633  -0.0986   0.7038   1.0000
   3.000   0.7925   0.01453   0.00647  -0.0978   0.6849   1.0000
   3.250   0.8184   0.01482   0.00665  -0.0970   0.6668   1.0000
   3.500   0.8432   0.01516   0.00699  -0.0961   0.6477   1.0000
   3.750   0.8683   0.01550   0.00729  -0.0952   0.6299   1.0000
   4.000   0.8936   0.01587   0.00761  -0.0945   0.6130   1.0000
   4.250   0.9189   0.01626   0.00796  -0.0937   0.5966   1.0000
   4.500   0.9442   0.01667   0.00838  -0.0929   0.5808   1.0000
   4.750   0.9691   0.01712   0.00884  -0.0921   0.5653   1.0000
   5.000   0.9937   0.01760   0.00936  -0.0913   0.5499   1.0000
   5.250   1.0185   0.01810   0.00991  -0.0906   0.5353   1.0000
   5.500   1.0430   0.01861   0.01052  -0.0898   0.5205   1.0000
   5.750   1.0668   0.01909   0.01111  -0.0888   0.5048   1.0000
   6.000   1.0885   0.01927   0.01132  -0.0872   0.4816   1.0000
   6.250   1.1077   0.01915   0.01114  -0.0850   0.4507   1.0000
   6.500   1.1257   0.01910   0.01111  -0.0828   0.4185   1.0000
   6.750   1.1431   0.01915   0.01129  -0.0806   0.3831   1.0000
   7.000   1.1579   0.01935   0.01149  -0.0779   0.3298   1.0000
   7.250   1.1521   0.02207   0.01274  -0.0731   0.1186   1.0000
   7.500   1.1562   0.02478   0.01504  -0.0696   0.0801   1.0000
   7.750   1.1649   0.02668   0.01697  -0.0667   0.0698   1.0000
   8.000   1.1746   0.02839   0.01869  -0.0641   0.0624   1.0000
   8.250   1.1860   0.03020   0.02054  -0.0616   0.0580   1.0000
   8.500   1.2023   0.03198   0.02241  -0.0597   0.0548   1.0000
   8.750   1.2228   0.03408   0.02448  -0.0584   0.0521   1.0000
   9.000   1.2501   0.03736   0.02768  -0.0584   0.0488   1.0000
   9.250   1.2715   0.03947   0.03010  -0.0571   0.0469   1.0000
   9.500   1.2946   0.04251   0.03344  -0.0561   0.0463   1.0000
   9.750   1.3129   0.04588   0.03719  -0.0546   0.0463   1.0000
  10.000   1.3261   0.04957   0.04132  -0.0528   0.0467   1.0000
  10.250   1.3346   0.05360   0.04576  -0.0507   0.0474   1.0000
  10.500   1.3381   0.05757   0.05012  -0.0484   0.0477   1.0000
  10.750   1.3366   0.06143   0.05435  -0.0458   0.0478   1.0000
  11.000   1.3306   0.06532   0.05858  -0.0431   0.0480   1.0000
  11.250   1.3210   0.06932   0.06286  -0.0405   0.0483   1.0000
  11.500   1.3102   0.07378   0.06753  -0.0381   0.0487   1.0000
  11.750   1.3081   0.07709   0.07107  -0.0361   0.0501   1.0000
  12.000   1.2786   0.07974   0.07404  -0.0335   0.0508   1.0000
  12.250   1.2388   0.08423   0.07888  -0.0335   0.0516   1.0000
  12.500   1.1878   0.09236   0.08738  -0.0377   0.0542   1.0000
  12.750   1.1545   0.10111   0.09632  -0.0429   0.0564   1.0000
  13.000   1.1269   0.11014   0.10542  -0.0486   0.0577   1.0000
  13.250   1.1093   0.11855   0.11388  -0.0532   0.0592   1.0000
  13.500   0.9049   0.12931   0.12512  -0.0582   0.0576   1.0000
  13.750   0.8917   0.13620   0.13200  -0.0614   0.0596   1.0000
<< Back to USA 41 AIRFOIL (usa41-il)

Polar data table (+)

Polar graphs


<< Back to USA 41 AIRFOIL (usa41-il)