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 S826 Airfoil (s826-nr) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: NREL's S826 Airfoil (s826-nr)
Reynolds number: 50,000
Max Cl/Cd: 34.41 at α=8°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-s826-nr-50000-n5.txt
Download as CSV file: xf-s826-nr-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NREL's S826 Airfoil                             
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.500  -0.3494   0.12362   0.11659  -0.0457   1.0000   0.0501
 -10.250  -0.3496   0.12003   0.11307  -0.0460   1.0000   0.0478
  -9.750  -0.3668   0.11024   0.10347  -0.0499   1.0000   0.0416
  -9.500  -0.3666   0.10714   0.10043  -0.0493   1.0000   0.0409
  -9.250  -0.3693   0.10377   0.09714  -0.0492   1.0000   0.0403
  -9.000  -0.3739   0.10027   0.09373  -0.0493   1.0000   0.0396
  -8.750  -0.3811   0.09649   0.09004  -0.0496   1.0000   0.0389
  -8.500  -0.3908   0.09263   0.08628  -0.0499   1.0000   0.0383
  -8.250  -0.4042   0.08877   0.08252  -0.0502   1.0000   0.0376
  -8.000  -0.4218   0.08517   0.07902  -0.0501   1.0000   0.0370
  -7.750  -0.4447   0.08197   0.07593  -0.0494   1.0000   0.0366
  -7.500  -0.4690   0.07864   0.07268  -0.0493   1.0000   0.0361
  -7.250  -0.4889   0.07454   0.06858  -0.0514   1.0000   0.0356
  -7.000  -0.4995   0.06937   0.06327  -0.0564   0.9984   0.0350
  -6.750  -0.4792   0.06172   0.05506  -0.0679   0.9884   0.0339
  -6.500  -0.4515   0.05574   0.04847  -0.0761   0.9802   0.0335
  -6.250  -0.4191   0.05097   0.04314  -0.0822   0.9730   0.0336
  -6.000  -0.3882   0.04760   0.03946  -0.0861   0.9656   0.0351
  -5.750  -0.3516   0.04450   0.03592  -0.0904   0.9594   0.0378
  -5.500  -0.3180   0.04166   0.03253  -0.0929   0.9518   0.0401
  -5.250  -0.2811   0.03921   0.02955  -0.0946   0.9461   0.0419
  -5.000  -0.2553   0.03743   0.02779  -0.0942   0.9379   0.0449
  -4.750  -0.2224   0.03619   0.02637  -0.0945   0.9318   0.0515
  -4.500  -0.2001   0.03511   0.02524  -0.0925   0.9224   0.0559
  -4.250  -0.1720   0.03416   0.02416  -0.0925   0.9138   0.0654
  -4.000  -0.1377   0.03288   0.02287  -0.0943   0.9067   0.0789
  -3.750  -0.1014   0.03120   0.02126  -0.0981   0.8982   0.1017
  -3.250  -0.0232   0.03052   0.02192  -0.1040   0.8824   0.4920
  -3.000  -0.0071   0.03201   0.02332  -0.0990   0.8726   0.5261
  -2.750   0.0128   0.03303   0.02420  -0.0950   0.8646   0.5541
  -2.500   0.0319   0.03344   0.02446  -0.0924   0.8545   0.5709
  -2.250   0.0739   0.03298   0.02361  -0.0959   0.8494   0.5794
  -2.000   0.1024   0.03268   0.02308  -0.0969   0.8400   0.5812
  -1.750   0.1426   0.03215   0.02227  -0.0998   0.8350   0.5830
  -1.500   0.1721   0.03191   0.02182  -0.1012   0.8260   0.5854
  -1.250   0.2123   0.03143   0.02111  -0.1042   0.8211   0.5885
  -1.000   0.2454   0.03122   0.02070  -0.1064   0.8130   0.5922
  -0.750   0.2830   0.03088   0.02018  -0.1089   0.8076   0.5951
  -0.500   0.3141   0.03074   0.01993  -0.1101   0.8007   0.5977
  -0.250   0.3468   0.03061   0.01969  -0.1115   0.7940   0.6010
   0.000   0.3894   0.03026   0.01918  -0.1148   0.7903   0.6054
   0.250   0.4157   0.03044   0.01929  -0.1157   0.7813   0.6094
   0.500   0.4521   0.03025   0.01905  -0.1175   0.7769   0.6126
   0.750   0.4781   0.03046   0.01921  -0.1180   0.7691   0.6164
   1.000   0.5147   0.03041   0.01910  -0.1202   0.7638   0.6214
   1.250   0.5552   0.03022   0.01887  -0.1228   0.7603   0.6258
   1.500   0.5726   0.03078   0.01946  -0.1219   0.7512   0.6296
   1.750   0.6112   0.03070   0.01936  -0.1243   0.7471   0.6347
   2.000   0.6353   0.03117   0.01984  -0.1246   0.7396   0.6395
   2.250   0.6648   0.03135   0.02007  -0.1253   0.7342   0.6441
   2.500   0.7040   0.03128   0.02004  -0.1275   0.7308   0.6502
   2.750   0.7190   0.03214   0.02096  -0.1265   0.7219   0.6550
   3.000   0.7529   0.03220   0.02109  -0.1278   0.7176   0.6603
   3.250   0.7741   0.03291   0.02187  -0.1277   0.7104   0.6663
   3.500   0.8014   0.03325   0.02235  -0.1280   0.7045   0.6720
   3.750   0.8400   0.03316   0.02236  -0.1299   0.7010   0.6792
   4.000   0.8499   0.03432   0.02364  -0.1281   0.6913   0.6851
   4.250   0.8863   0.03425   0.02374  -0.1295   0.6871   0.6923
   4.500   0.8983   0.03535   0.02497  -0.1280   0.6777   0.6992
   4.750   0.9332   0.03531   0.02510  -0.1290   0.6728   0.7073
   5.000   0.9457   0.03635   0.02629  -0.1274   0.6634   0.7150
   5.250   0.9822   0.03620   0.02637  -0.1285   0.6578   0.7246
   5.500   0.9930   0.03718   0.02754  -0.1265   0.6476   0.7329
   5.750   1.0367   0.03657   0.02715  -0.1283   0.6419   0.7448
   6.000   1.0467   0.03750   0.02831  -0.1260   0.6304   0.7554
   6.250   1.0683   0.03773   0.02878  -0.1248   0.6202   0.7676
   6.500   1.1140   0.03661   0.02793  -0.1262   0.6116   0.7839
   6.750   1.1278   0.03702   0.02859  -0.1239   0.5987   0.7999
   7.000   1.1434   0.03712   0.02900  -0.1215   0.5856   0.8196
   7.250   1.1604   0.03699   0.02914  -0.1192   0.5722   0.8473
   7.750   1.2006   0.03639   0.02903  -0.1155   0.5412   1.0000
   8.000   1.2323   0.03581   0.02863  -0.1152   0.5224   1.0000
   8.250   1.2376   0.03668   0.02965  -0.1123   0.5020   1.0000
   8.500   1.2492   0.03708   0.03020  -0.1098   0.4793   1.0000
   8.750   1.2494   0.03833   0.03161  -0.1066   0.4541   1.0000
   9.000   1.2521   0.03949   0.03290  -0.1037   0.4244   1.0000
   9.250   1.2605   0.04020   0.03363  -0.1012   0.3873   1.0000
   9.500   1.2720   0.04073   0.03392  -0.0986   0.3408   1.0000
   9.750   1.2747   0.04238   0.03527  -0.0960   0.2976   1.0000
  10.000   1.2719   0.04485   0.03749  -0.0936   0.2626   1.0000
  10.250   1.2679   0.04768   0.04014  -0.0916   0.2337   1.0000
  10.500   1.2642   0.05070   0.04303  -0.0900   0.2088   1.0000
  10.750   1.2610   0.05381   0.04603  -0.0887   0.1882   1.0000
  11.000   1.2596   0.05690   0.04907  -0.0876   0.1693   1.0000
  11.250   1.2590   0.06001   0.05215  -0.0867   0.1523   1.0000
  11.500   1.2598   0.06307   0.05520  -0.0859   0.1376   1.0000
  11.750   1.2613   0.06613   0.05827  -0.0852   0.1242   1.0000
  12.000   1.2641   0.06913   0.06130  -0.0846   0.1125   1.0000
  12.250   1.2668   0.07217   0.06434  -0.0840   0.1019   1.0000
  12.500   1.2700   0.07524   0.06751  -0.0836   0.0921   1.0000
  12.750   1.2746   0.07834   0.07079  -0.0832   0.0828   1.0000
  13.000   1.2790   0.08146   0.07398  -0.0829   0.0750   1.0000
  13.250   1.2819   0.08477   0.07740  -0.0828   0.0680   1.0000
  13.500   1.2860   0.08822   0.08108  -0.0826   0.0616   1.0000
  13.750   1.2874   0.09179   0.08476  -0.0829   0.0565   1.0000
  14.000   1.2889   0.09567   0.08883  -0.0831   0.0520   1.0000
  14.250   1.2871   0.10016   0.09364  -0.0839   0.0483   1.0000
  14.500   1.2858   0.10418   0.09773  -0.0849   0.0452   1.0000
  14.750   1.2828   0.10888   0.10260  -0.0860   0.0427   1.0000
  15.000   1.2732   0.11504   0.10914  -0.0883   0.0412   1.0000
  15.250   1.2617   0.12167   0.11609  -0.0912   0.0401   1.0000
  15.500   1.2481   0.12891   0.12361  -0.0950   0.0394   1.0000
  15.750   1.2323   0.13704   0.13199  -0.0997   0.0391   1.0000
  16.000   1.2130   0.14658   0.14176  -0.1057   0.0394   1.0000
  16.250   1.1906   0.15799   0.15336  -0.1134   0.0402   1.0000
<< Back to NREL's S826 Airfoil (s826-nr)

Polar data table (+)

Polar graphs


<< Back to NREL's S826 Airfoil (s826-nr)