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

NREL's S810 Airfoil (s810-nr) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: NREL's S810 Airfoil (s810-nr)
Reynolds number: 100,000
Max Cl/Cd: 24.86 at α=9.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-s810-nr-100000.txt
Download as CSV file: xf-s810-nr-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NREL's S810 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
  ------ -------- --------- --------- -------- -------- --------
 -12.500  -0.3442   0.10846   0.10450  -0.0749   0.9778   0.1076
 -12.250  -0.3732   0.10133   0.09739  -0.0802   0.9756   0.1079
 -12.000  -0.3954   0.09516   0.09117  -0.0841   0.9734   0.1079
 -11.750  -0.4196   0.08916   0.08509  -0.0874   0.9713   0.1079
 -11.500  -0.4455   0.08389   0.07973  -0.0895   0.9679   0.1079
 -11.250  -0.4735   0.07939   0.07511  -0.0904   0.9633   0.1080
 -11.000  -0.4957   0.07525   0.07084  -0.0912   0.9595   0.1080
 -10.750  -0.5235   0.07210   0.06752  -0.0906   0.9548   0.1081
 -10.500  -0.5503   0.07025   0.06555  -0.0876   0.9480   0.1082
 -10.250  -0.5746   0.06841   0.06350  -0.0847   0.9432   0.1083
  -7.250  -0.7333   0.05003   0.04394  -0.0433   0.9143   0.2149
  -7.000  -0.7385   0.04793   0.04175  -0.0392   0.9142   0.2262
  -6.750  -0.7453   0.04721   0.04071  -0.0343   0.9139   0.2320
  -6.500  -0.7400   0.04446   0.03796  -0.0315   0.9134   0.2412
  -6.250  -0.7314   0.04304   0.03625  -0.0287   0.9138   0.2445
  -6.000  -0.7137   0.04119   0.03415  -0.0268   0.9143   0.2409
  -5.750  -0.6921   0.04070   0.03312  -0.0245   0.9145   0.2215
  -5.000  -0.5580   0.04111   0.03172  -0.0205   0.9134   0.0787
  -4.750  -0.5269   0.03952   0.03012  -0.0201   0.9132   0.0734
  -4.500   0.1153   0.04940   0.04197  -0.0747   0.9135   0.8144
  -4.250   0.0818   0.05097   0.04361  -0.0650   0.9095   0.8194
  -4.000  -0.7083   0.03949   0.02991   0.0219   1.0000   0.0811
  -3.750  -0.6821   0.03698   0.02756   0.0228   1.0000   0.0765
  -3.500  -0.6580   0.03600   0.02641   0.0244   1.0000   0.0702
  -3.250  -0.6357   0.03572   0.02596   0.0259   1.0000   0.0657
  -3.000  -0.6137   0.03455   0.02492   0.0275   1.0000   0.0630
  -2.750  -0.5943   0.03382   0.02421   0.0293   1.0000   0.0607
  -2.500  -0.5772   0.03321   0.02362   0.0313   1.0000   0.0596
  -2.250  -0.5624   0.03267   0.02310   0.0334   1.0000   0.0592
  -2.000  -0.5484   0.03213   0.02256   0.0353   1.0000   0.0595
  -1.750  -0.5331   0.03152   0.02185   0.0365   1.0000   0.0609
  -1.500  -0.5155   0.03104   0.02122   0.0371   1.0000   0.0655
  -1.250  -0.4969   0.03033   0.02033   0.0373   1.0000   0.0719
  -1.000  -0.4765   0.03001   0.01982   0.0374   1.0000   0.0788
  -0.750  -0.2612   0.05403   0.04697   0.0466   1.0000   0.8896
  -0.500  -0.2230   0.05415   0.04694   0.0433   1.0000   0.9046
  -0.250  -0.1804   0.05434   0.04703   0.0390   1.0000   0.9213
   0.000  -0.4446   0.03846   0.03116   0.0635   0.9893   0.7529
   0.250  -0.4221   0.04075   0.03334   0.0659   0.9833   0.7756
   0.500  -0.4164   0.04246   0.03508   0.0738   0.9753   0.8038
   0.750  -0.3881   0.04419   0.03672   0.0744   0.9700   0.8161
   1.000  -0.3668   0.04357   0.03596   0.0729   0.9605   0.8174
   1.250  -0.3396   0.04390   0.03618   0.0706   0.9543   0.8183
   1.500  -0.3099   0.04412   0.03630   0.0682   0.9448   0.8190
   1.750  -0.2870   0.04412   0.03624   0.0670   0.9358   0.8197
   2.000  -0.2509   0.04518   0.03722   0.0634   0.9290   0.8208
   2.250  -0.2316   0.04491   0.03691   0.0631   0.9182   0.8223
   2.500  -0.2067   0.04531   0.03727   0.0616   0.9097   0.8237
   2.750  -0.1707   0.04623   0.03815   0.0581   0.9011   0.8248
   3.000  -0.1497   0.04620   0.03810   0.0572   0.8897   0.8259
   3.250  -0.1240   0.04665   0.03854   0.0554   0.8798   0.8272
   3.500  -0.0823   0.04809   0.03997   0.0509   0.8718   0.8290
   3.750  -0.0593   0.04816   0.04003   0.0496   0.8591   0.8307
   4.000  -0.0342   0.04854   0.04043   0.0480   0.8467   0.8321
   4.250  -0.0098   0.04903   0.04096   0.0470   0.8341   0.8333
   4.500   0.0176   0.04964   0.04162   0.0456   0.8207   0.8346
   4.750   0.0510   0.05028   0.04231   0.0435   0.8046   0.8361
   5.000   0.1289   0.04924   0.04131   0.0391   0.7494   0.8379
   5.250   0.1669   0.04925   0.04142   0.0372   0.7338   0.8400
   5.500   0.1994   0.04933   0.04157   0.0357   0.7200   0.8421
   5.750   0.2319   0.04938   0.04171   0.0340   0.7068   0.8439
   6.000   0.2648   0.04944   0.04190   0.0323   0.6942   0.8455
   6.250   0.2971   0.04931   0.04190   0.0313   0.6826   0.8471
   6.500   0.3421   0.04884   0.04160   0.0293   0.6743   0.8489
   6.750   0.3689   0.04852   0.04141   0.0292   0.6612   0.8512
   7.000   0.3988   0.04805   0.04113   0.0288   0.6483   0.8537
   7.250   0.4327   0.04732   0.04058   0.0280   0.6355   0.8562
   7.500   0.4719   0.04613   0.03959   0.0269   0.6228   0.8584
   7.750   0.5134   0.04441   0.03813   0.0261   0.6099   0.8605
   8.000   0.5542   0.04212   0.03610   0.0263   0.5965   0.8627
   8.750   0.6724   0.03355   0.02839   0.0285   0.5370   0.8719
   9.000   0.7092   0.03075   0.02566   0.0300   0.4756   0.8749
   9.250   0.7301   0.02937   0.02313   0.0340   0.3107   0.8775
   9.500   0.7194   0.03152   0.02447   0.0374   0.2169   0.8804
   9.750   0.7204   0.03339   0.02577   0.0394   0.1622   0.8833
  10.000   0.7322   0.03484   0.02690   0.0404   0.1322   0.8864
  10.250   0.7540   0.03588   0.02778   0.0412   0.1133   0.8895
  10.500   0.7801   0.03687   0.02874   0.0417   0.0987   0.8931
  10.750   0.8065   0.03816   0.03009   0.0417   0.0848   0.8969
  11.000   0.8220   0.03967   0.03164   0.0419   0.0715   0.9011
  11.250   0.8378   0.04132   0.03337   0.0426   0.0594   0.9050
  11.500   0.8646   0.04357   0.03570   0.0425   0.0498   0.9089
  11.750   0.8896   0.04602   0.03843   0.0426   0.0442   0.9133
  12.000   0.9264   0.05095   0.04359   0.0414   0.0409   0.9167
  12.250   0.9249   0.05339   0.04648   0.0435   0.0402   0.9226
  12.500   0.9204   0.05615   0.04966   0.0453   0.0395   0.9295
  12.750   0.9120   0.05905   0.05298   0.0472   0.0388   0.9372
  13.000   0.9035   0.06260   0.05690   0.0483   0.0388   0.9462
  13.250   0.8919   0.06652   0.06118   0.0487   0.0387   0.9574
  13.500   0.8779   0.07076   0.06577   0.0480   0.0385   0.9758
  13.750   0.8586   0.07539   0.07070   0.0472   0.0383   1.0000
  14.000   0.8421   0.08056   0.07610   0.0456   0.0387   1.0000
  14.250   0.8197   0.08644   0.08221   0.0433   0.0388   1.0000
  14.500   0.8019   0.09252   0.08845   0.0404   0.0395   1.0000
  14.750   0.7813   0.09947   0.09557   0.0367   0.0399   1.0000
  15.000   0.7633   0.10684   0.10304   0.0323   0.0404   1.0000
  15.250   0.7548   0.11417   0.11040   0.0286   0.0411   1.0000
  15.500   0.6573   0.14749   0.14386   0.0076   0.0620   1.0000
<< Back to NREL's S810 Airfoil (s810-nr)

Polar data table (+)

Polar graphs


<< Back to NREL's S810 Airfoil (s810-nr)