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NASA/LANGLEY RC-08(B)3 AIRFOIL (rc08b3-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: NASA/LANGLEY RC-08(B)3 AIRFOIL (rc08b3-il)
Reynolds number: 200,000
Max Cl/Cd: 57.7 at α=3.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-rc08b3-il-200000.txt
Download as CSV file: xf-rc08b3-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NASA/LANGLEY RC-08(B)3 AIRFOIL                  
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.250  -0.5988   0.09813   0.09463  -0.0133   1.0000   0.0453
  -9.000  -0.6053   0.09292   0.08941  -0.0192   1.0000   0.0454
  -8.750  -0.6119   0.08865   0.08507  -0.0217   1.0000   0.0455
  -8.500  -0.6157   0.08440   0.08070  -0.0236   1.0000   0.0456
  -8.250  -0.6132   0.07752   0.07400  -0.0220   1.0000   0.0470
  -8.000  -0.6064   0.07454   0.07101  -0.0211   1.0000   0.0481
  -7.750  -0.6013   0.07114   0.06759  -0.0214   1.0000   0.0493
  -7.500  -0.5965   0.06740   0.06378  -0.0222   1.0000   0.0509
  -7.250  -0.5912   0.06343   0.05970  -0.0230   1.0000   0.0531
  -7.000  -0.5876   0.06102   0.05664  -0.0236   1.0000   0.0579
  -6.750  -0.5863   0.05403   0.04964  -0.0232   1.0000   0.0594
  -6.500  -0.5735   0.05116   0.04687  -0.0223   1.0000   0.0611
  -6.250  -0.5623   0.04863   0.04429  -0.0209   1.0000   0.0634
  -6.000  -0.5524   0.04602   0.04151  -0.0191   1.0000   0.0670
  -5.750  -0.5500   0.04304   0.03796  -0.0156   1.0000   0.0728
  -5.500  -0.5386   0.04018   0.03516  -0.0140   1.0000   0.0745
  -5.250  -0.5274   0.03816   0.03308  -0.0117   1.0000   0.0773
  -5.000  -0.5204   0.03667   0.03105  -0.0081   1.0000   0.0867
  -4.750  -0.5024   0.02105   0.01597  -0.0051   1.0000   0.0882
  -4.500  -0.4870   0.02970   0.02312  -0.0015   1.0000   0.0576
  -4.250  -0.4708   0.02434   0.01706   0.0015   1.0000   0.0467
  -4.000  -0.4506   0.02371   0.01623   0.0034   1.0000   0.0457
  -3.750  -0.4291   0.02090   0.01319   0.0046   1.0000   0.0442
  -3.500  -0.4064   0.01927   0.01132   0.0058   1.0000   0.0439
  -3.250  -0.3832   0.01805   0.00991   0.0068   1.0000   0.0442
  -3.000  -0.3598   0.01712   0.00886   0.0078   1.0000   0.0449
  -2.750  -0.3358   0.01597   0.00765   0.0085   1.0000   0.0467
  -2.500  -0.3133   0.01519   0.00693   0.0093   1.0000   0.0510
  -2.250  -0.2875   0.01460   0.00632   0.0095   0.9993   0.0551
  -2.000  -0.2505   0.01377   0.00552   0.0075   0.9957   0.0640
  -1.750  -0.2151   0.01270   0.00484   0.0056   0.9920   0.1341
  -1.500  -0.0507   0.01027   0.00528  -0.0191   1.0000   1.0000
  -1.250  -0.0471   0.01039   0.00533  -0.0150   1.0000   1.0000
  -1.000  -0.0429   0.01054   0.00540  -0.0112   1.0000   1.0000
  -0.750  -0.0087   0.01060   0.00536  -0.0131   0.9962   1.0000
  -0.500   0.0368   0.01062   0.00529  -0.0172   0.9904   1.0000
  -0.250   0.0844   0.01061   0.00521  -0.0217   0.9848   1.0000
   0.000   0.1315   0.01055   0.00512  -0.0260   0.9779   1.0000
   0.250   0.1824   0.01045   0.00499  -0.0310   0.9723   1.0000
   0.500   0.2284   0.01033   0.00487  -0.0349   0.9645   1.0000
   0.750   0.2745   0.01017   0.00472  -0.0387   0.9564   1.0000
   1.000   0.3097   0.01005   0.00463  -0.0400   0.9437   1.0000
   1.250   0.3392   0.00995   0.00455  -0.0401   0.9297   1.0000
   1.500   0.3640   0.00987   0.00448  -0.0389   0.9144   1.0000
   1.750   0.3858   0.00977   0.00439  -0.0371   0.8979   1.0000
   2.000   0.4065   0.00969   0.00434  -0.0351   0.8791   1.0000
   2.250   0.4276   0.00961   0.00427  -0.0331   0.8601   1.0000
   2.500   0.4489   0.00951   0.00415  -0.0310   0.8397   1.0000
   2.750   0.4702   0.00943   0.00406  -0.0291   0.8128   1.0000
   3.000   0.4919   0.00937   0.00397  -0.0271   0.7800   1.0000
   3.250   0.5140   0.00937   0.00390  -0.0254   0.7405   1.0000
   3.500   0.5365   0.00946   0.00387  -0.0238   0.6921   1.0000
   3.750   0.5585   0.00968   0.00387  -0.0222   0.6246   1.0000
   4.000   0.5791   0.01016   0.00396  -0.0205   0.5274   1.0000
   4.250   0.5983   0.01094   0.00420  -0.0189   0.4122   1.0000
   4.750   0.6333   0.01332   0.00521  -0.0162   0.1485   1.0000
   5.000   0.6511   0.01460   0.00605  -0.0147   0.0913   1.0000
   5.250   0.6709   0.01548   0.00693  -0.0133   0.0751   1.0000
   5.500   0.6895   0.01654   0.00791  -0.0117   0.0655   1.0000
   5.750   0.7107   0.01720   0.00865  -0.0104   0.0592   1.0000
   6.000   0.7301   0.01823   0.00965  -0.0090   0.0547   1.0000
   6.250   0.7501   0.01970   0.01116  -0.0075   0.0518   1.0000
   6.500   0.7720   0.02067   0.01227  -0.0062   0.0489   1.0000
   6.750   0.7933   0.02162   0.01330  -0.0050   0.0456   1.0000
   7.000   0.8146   0.02303   0.01476  -0.0039   0.0435   1.0000
   7.250   0.8356   0.02574   0.01762  -0.0028   0.0419   1.0000
   7.500   0.8547   0.02763   0.01984  -0.0012   0.0405   1.0000
   7.750   0.8733   0.02902   0.02159   0.0006   0.0387   1.0000
   8.000   0.8893   0.03154   0.02449   0.0027   0.0380   1.0000
   8.250   0.9019   0.03468   0.02806   0.0051   0.0379   1.0000
   8.500   0.9105   0.03814   0.03197   0.0079   0.0379   1.0000
   8.750   0.9163   0.04149   0.03574   0.0107   0.0373   1.0000
   9.000   0.9171   0.04558   0.04025   0.0137   0.0375   1.0000
   9.250   0.9126   0.05063   0.04563   0.0165   0.0394   1.0000
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