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 S815 Airfoil (s815-nr) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: NREL's S815 Airfoil (s815-nr)
Reynolds number: 200,000
Max Cl/Cd: 54.38 at α=10°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-s815-nr-200000.txt
Download as CSV file: xf-s815-nr-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NREL's S815 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
  ------ -------- --------- --------- -------- -------- --------
  -7.500  -0.0147   0.11256   0.10719  -0.0479   0.9707   0.3131
  -7.250   0.0180   0.11008   0.10471  -0.0519   0.9672   0.3147
  -7.000  -0.0417   0.11124   0.10577  -0.0561   0.9566   0.3248
  -6.750   0.0021   0.10689   0.10143  -0.0595   0.9529   0.3252
  -6.500   0.0474   0.10300   0.09755  -0.0636   0.9502   0.3256
  -6.250   0.0782   0.10012   0.09471  -0.0645   0.9393   0.3260
  -6.000   0.1177   0.09712   0.09174  -0.0681   0.9346   0.3267
  -5.750   0.1420   0.09495   0.08960  -0.0684   0.9194   0.3274
  -5.500   0.1636   0.09302   0.08769  -0.0685   0.9007   0.3284
  -5.250   0.1938   0.09069   0.08537  -0.0708   0.8857   0.3297
  -5.000   0.2314   0.08800   0.08264  -0.0754   0.8704   0.3315
  -4.750   0.2010   0.08756   0.08202  -0.0837   0.8543   0.3414
  -4.500   0.2899   0.08194   0.07628  -0.0968   0.8370   0.3419
  -4.250   0.3785   0.07710   0.07119  -0.1107   0.8047   0.3424
  -4.000   0.4278   0.07456   0.06835  -0.1164   0.7660   0.3430
  -3.750   0.4562   0.07314   0.06669  -0.1176   0.7357   0.3437
  -3.500   0.4773   0.07210   0.06549  -0.1174   0.7122   0.3446
  -3.250   0.4959   0.07118   0.06444  -0.1169   0.6937   0.3457
  -3.000   0.5129   0.07034   0.06351  -0.1164   0.6776   0.3471
  -2.750   0.5280   0.06953   0.06265  -0.1157   0.6640   0.3488
  -2.500   0.4568   0.07147   0.06449  -0.1125   0.6576   0.3583
  -2.250   0.4835   0.06945   0.06245  -0.1124   0.6463   0.3587
  -2.000   0.5104   0.06782   0.06073  -0.1125   0.6367   0.3591
  -1.750   0.5350   0.06640   0.05933  -0.1123   0.6272   0.3595
  -1.500   0.5593   0.06516   0.05803  -0.1122   0.6189   0.3601
  -1.250   0.5823   0.06407   0.05693  -0.1120   0.6115   0.3608
  -1.000   0.6026   0.06310   0.05598  -0.1115   0.6041   0.3617
  -0.750   0.6236   0.06223   0.05506  -0.1114   0.5982   0.3629
  -0.500   0.6420   0.06144   0.05429  -0.1110   0.5927   0.3643
  -0.250   0.6574   0.06069   0.05358  -0.1104   0.5872   0.3661
   0.000   0.6707   0.05999   0.05289  -0.1097   0.5825   0.3688
   0.250   0.6123   0.05972   0.05261  -0.1061   0.5801   0.3763
   0.500   0.6444   0.05845   0.05130  -0.1068   0.5757   0.3768
   0.750   0.6689   0.05741   0.05035  -0.1063   0.5712   0.3775
   1.000   0.6923   0.05649   0.04948  -0.1058   0.5668   0.3782
   1.250   0.7138   0.05565   0.04866  -0.1054   0.5624   0.3791
   1.500   0.7359   0.05488   0.04789  -0.1053   0.5588   0.3802
   1.750   0.7565   0.05424   0.04726  -0.1051   0.5556   0.3816
   2.000   0.7699   0.05364   0.04676  -0.1038   0.5524   0.3835
   2.250   0.7805   0.05308   0.04629  -0.1025   0.5491   0.3863
   2.500   0.5613   0.02800   0.02044  -0.1150   0.5512   0.3619
   2.750   0.5866   0.02779   0.02028  -0.1152   0.5477   0.3630
   3.000   0.6141   0.02748   0.02000  -0.1159   0.5445   0.3643
   3.250   0.6446   0.02699   0.01949  -0.1176   0.5415   0.3660
   3.500   0.6786   0.02654   0.01895  -0.1200   0.5385   0.3685
   3.750   0.7112   0.02591   0.01824  -0.1226   0.5353   0.3716
   4.000   0.7423   0.02528   0.01756  -0.1248   0.5321   0.3742
   4.250   0.7727   0.02489   0.01714  -0.1263   0.5288   0.3760
   4.500   0.7998   0.02475   0.01710  -0.1265   0.5254   0.3776
   4.750   0.8290   0.02478   0.01719  -0.1268   0.5221   0.3791
   5.000   0.8613   0.02490   0.01731  -0.1278   0.5190   0.3807
   5.250   0.8889   0.02508   0.01757  -0.1280   0.5157   0.3826
   5.500   0.9119   0.02512   0.01774  -0.1274   0.5119   0.3846
   5.750   0.9384   0.02513   0.01781  -0.1275   0.5079   0.3870
   6.000   0.9697   0.02503   0.01770  -0.1285   0.5038   0.3898
   6.250   1.0074   0.02496   0.01749  -0.1306   0.4996   0.3927
   6.500   1.0327   0.02495   0.01758  -0.1304   0.4950   0.3950
   6.750   1.0519   0.02499   0.01781  -0.1289   0.4895   0.3970
   7.000   1.0782   0.02497   0.01785  -0.1284   0.4839   0.3995
   7.250   1.1149   0.02504   0.01783  -0.1298   0.4784   0.4028
   7.500   1.1289   0.02504   0.01802  -0.1274   0.4723   0.4057
   7.750   1.1547   0.02488   0.01786  -0.1271   0.4656   0.4092
   8.000   1.1929   0.02471   0.01753  -0.1289   0.4595   0.4127
   8.250   1.2037   0.02472   0.01780  -0.1259   0.4533   0.4149
   8.500   1.2243   0.02465   0.01783  -0.1244   0.4465   0.4176
   8.750   1.2560   0.02459   0.01771  -0.1248   0.4400   0.4212
   9.000   1.2646   0.02463   0.01794  -0.1215   0.4328   0.4244
   9.250   1.2861   0.02447   0.01775  -0.1203   0.4252   0.4284
   9.500   1.3029   0.02441   0.01774  -0.1183   0.4176   0.4317
   9.750   1.3080   0.02435   0.01784  -0.1143   0.4096   0.4342
  10.000   1.3236   0.02434   0.01784  -0.1119   0.4014   0.4374
  10.250   1.3259   0.02462   0.01830  -0.1078   0.3921   0.4406
  10.500   1.3382   0.02478   0.01844  -0.1053   0.3824   0.4449
  10.750   1.3433   0.02518   0.01893  -0.1019   0.3705   0.4486
  11.000   1.3490   0.02556   0.01944  -0.0988   0.3573   0.4518
  11.250   1.3522   0.02614   0.02012  -0.0954   0.3415   0.4548
  11.500   1.3526   0.02698   0.02099  -0.0919   0.3212   0.4579
  12.000   1.3383   0.03005   0.02384  -0.0847   0.2720   0.4640
  12.250   1.3257   0.03239   0.02603  -0.0815   0.2494   0.4667
  12.750   1.2993   0.03808   0.03158  -0.0768   0.2100   0.4712
  13.000   1.2861   0.04138   0.03486  -0.0751   0.1929   0.4733
  13.250   1.2729   0.04497   0.03844  -0.0740   0.1777   0.4755
  13.750   1.2503   0.05259   0.04601  -0.0731   0.1509   0.4808
  14.000   1.2419   0.05642   0.04983  -0.0731   0.1393   0.4839
  14.250   1.2339   0.06038   0.05374  -0.0734   0.1292   0.4869
  14.500   1.2245   0.06455   0.05789  -0.0738   0.1205   0.4895
  14.750   1.2215   0.06810   0.06151  -0.0742   0.1117   0.4924
  15.000   1.2166   0.07193   0.06536  -0.0747   0.1044   0.4956
  15.250   1.2127   0.07575   0.06917  -0.0754   0.0976   0.4992
  15.500   1.2123   0.07926   0.07270  -0.0760   0.0911   0.5033
  15.750   1.2100   0.08308   0.07643  -0.0769   0.0858   0.5071
  16.000   1.2121   0.08634   0.07981  -0.0777   0.0803   0.5108
  16.250   1.2114   0.08990   0.08336  -0.0785   0.0758   0.5144
  16.500   1.2149   0.09306   0.08659  -0.0792   0.0713   0.5189
  16.750   1.2169   0.09646   0.08996  -0.0802   0.0676   0.5239
  17.000   1.2215   0.09950   0.09304  -0.0811   0.0639   0.5285
  17.250   1.2253   0.10262   0.09624  -0.0820   0.0608   0.5330
<< Back to NREL's S815 Airfoil (s815-nr)

Polar data table (+)

Polar graphs


<< Back to NREL's S815 Airfoil (s815-nr)