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NREL's S806A Airfoil (modified line 35) (s806a-nr) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: NREL's S806A Airfoil (modified line 35) (s806a-nr)
Reynolds number: 500,000
Max Cl/Cd: 107.63 at α=4.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-s806a-nr-500000.txt
Download as CSV file: xf-s806a-nr-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NREL's S806A Airfoil                            
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -3.500  -0.1227   0.02084   0.01488  -0.0555   0.7681   0.0525
  -3.250  -0.0944   0.02056   0.01462  -0.0557   0.7652   0.0572
  -3.000  -0.0674   0.01848   0.01245  -0.0562   0.7622   0.0667
  -2.750  -0.0388   0.01774   0.01165  -0.0564   0.7594   0.0746
  -2.500  -0.0112   0.01659   0.01055  -0.0570   0.7561   0.0905
  -2.250   0.0252   0.01390   0.00703  -0.0542   0.7525   0.0283
  -2.000   0.0536   0.01294   0.00600  -0.0542   0.7497   0.0227
  -1.750   0.0828   0.01260   0.00565  -0.0545   0.7492   0.0208
  -1.500   0.1120   0.01224   0.00527  -0.0548   0.7485   0.0200
  -1.250   0.1414   0.01175   0.00471  -0.0551   0.7477   0.0198
  -1.000   0.1711   0.01146   0.00430  -0.0555   0.7467   0.0217
  -0.750   0.1893   0.00868   0.00434  -0.0546   0.7457   0.8126
  -0.500   0.2093   0.00886   0.00467  -0.0521   0.7443   0.8770
  -0.250   0.2254   0.00896   0.00485  -0.0486   0.7428   0.9143
   0.000   0.2472   0.00888   0.00479  -0.0470   0.7411   0.9287
   0.250   0.2719   0.00883   0.00474  -0.0462   0.7388   0.9391
   0.500   0.2978   0.00873   0.00463  -0.0457   0.7363   0.9465
   0.750   0.3250   0.00871   0.00460  -0.0456   0.7339   0.9530
   1.000   0.3533   0.00871   0.00453  -0.0459   0.7317   0.9588
   1.250   0.3828   0.00868   0.00452  -0.0464   0.7290   0.9642
   1.500   0.4132   0.00862   0.00445  -0.0471   0.7259   0.9688
   1.750   0.4442   0.00855   0.00436  -0.0478   0.7225   0.9735
   2.000   0.4750   0.00844   0.00421  -0.0485   0.7184   0.9787
   2.250   0.5083   0.00833   0.00407  -0.0497   0.7136   0.9826
   2.500   0.5419   0.00839   0.00409  -0.0512   0.7070   0.9859
   2.750   0.5752   0.00859   0.00430  -0.0527   0.6992   0.9898
   3.000   0.6090   0.00835   0.00407  -0.0541   0.6966   0.9942
   3.250   0.6424   0.00819   0.00393  -0.0554   0.6928   0.9996
   3.500   0.6694   0.00806   0.00381  -0.0555   0.6875   1.0000
   3.750   0.6977   0.00792   0.00367  -0.0557   0.6803   1.0000
   4.000   0.7267   0.00780   0.00355  -0.0561   0.6725   1.0000
   4.250   0.7556   0.00773   0.00352  -0.0564   0.6638   1.0000
   4.500   0.7845   0.00770   0.00343  -0.0567   0.6477   1.0000
   4.750   0.8137   0.00756   0.00327  -0.0571   0.6218   1.0000
   5.000   0.8323   0.00872   0.00335  -0.0560   0.4269   1.0000
   5.250   0.8574   0.00939   0.00376  -0.0562   0.3692   1.0000
   5.500   0.8743   0.01103   0.00462  -0.0556   0.2257   1.0000
   5.750   0.8955   0.01199   0.00518  -0.0553   0.1580   1.0000
   6.000   0.9195   0.01253   0.00560  -0.0553   0.1297   1.0000
   6.250   0.9420   0.01318   0.00606  -0.0549   0.1042   1.0000
   6.500   0.9650   0.01372   0.00651  -0.0547   0.0872   1.0000
   6.750   0.9883   0.01420   0.00693  -0.0544   0.0741   1.0000
   7.000   1.0104   0.01476   0.00742  -0.0539   0.0604   1.0000
   7.250   1.0320   0.01531   0.00791  -0.0533   0.0508   1.0000
   7.500   1.0450   0.01600   0.00867  -0.0508   0.0365   1.0000
   7.750   1.0575   0.01685   0.00932  -0.0486   0.0170   1.0000
   8.000   1.0722   0.01759   0.00997  -0.0467   0.0096   1.0000
   8.250   1.0888   0.01811   0.01051  -0.0451   0.0079   1.0000
   8.500   1.1032   0.01872   0.01114  -0.0432   0.0070   1.0000
   8.750   1.1158   0.01942   0.01188  -0.0411   0.0063   1.0000
   9.000   1.1286   0.02021   0.01272  -0.0392   0.0060   1.0000
   9.250   1.1401   0.02112   0.01370  -0.0373   0.0057   1.0000
   9.500   1.1521   0.02208   0.01473  -0.0355   0.0055   1.0000
   9.750   1.1632   0.02314   0.01587  -0.0336   0.0053   1.0000
  10.000   1.1719   0.02440   0.01722  -0.0316   0.0052   1.0000
  10.250   1.1805   0.02569   0.01861  -0.0296   0.0052   1.0000
  10.500   1.1984   0.02654   0.01956  -0.0293   0.0052   1.0000
  10.750   1.2058   0.02799   0.02111  -0.0273   0.0052   1.0000
  11.000   1.2120   0.02959   0.02282  -0.0253   0.0052   1.0000
  11.250   1.2144   0.03153   0.02490  -0.0228   0.0052   1.0000
  11.500   1.2165   0.03362   0.02714  -0.0202   0.0053   1.0000
  11.750   1.2181   0.03587   0.02957  -0.0175   0.0054   1.0000
  12.000   1.2110   0.03955   0.03357  -0.0134   0.0058   1.0000
  12.250   1.2064   0.04313   0.03743  -0.0104   0.0060   1.0000
  12.500   1.1721   0.05109   0.04595  -0.0052   0.0070   1.0000
  12.750   1.1427   0.05756   0.05280  -0.0026   0.0076   1.0000
  13.000   1.0969   0.06635   0.06199  -0.0012   0.0084   1.0000
  13.250   1.0725   0.07244   0.06831  -0.0017   0.0085   1.0000
  13.500   1.0463   0.07912   0.07521  -0.0032   0.0087   1.0000
  13.750   1.0267   0.08562   0.08188  -0.0056   0.0087   1.0000
  14.000   0.9983   0.09475   0.09123  -0.0101   0.0089   1.0000
  14.250   0.9840   0.10217   0.09879  -0.0147   0.0087   1.0000
  14.500   0.9709   0.11024   0.10700  -0.0200   0.0086   1.0000
  14.750   0.9482   0.12237   0.11931  -0.0282   0.0085   1.0000
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