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 S825 Airfoil (s825-nr) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: NREL's S825 Airfoil (s825-nr)
Reynolds number: 100,000
Max Cl/Cd: 50.33 at α=8.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-s825-nr-100000.txt
Download as CSV file: xf-s825-nr-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NREL's S825 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
  ------ -------- --------- --------- -------- -------- --------
 -11.000  -0.4150   0.09128   0.08685  -0.0544   1.0000   0.0677
 -10.750  -0.4714   0.08061   0.07613  -0.0595   1.0000   0.0656
 -10.500  -0.5140   0.07387   0.06929  -0.0620   1.0000   0.0649
 -10.250  -0.5362   0.06896   0.06427  -0.0623   1.0000   0.0607
 -10.000  -0.5539   0.06453   0.05981  -0.0620   1.0000   0.0594
  -9.750  -0.5793   0.06052   0.05574  -0.0614   1.0000   0.0585
  -9.500  -0.6062   0.05735   0.05253  -0.0600   1.0000   0.0579
  -9.250  -0.7063   0.06529   0.05996  -0.0639   1.0000   0.0607
  -9.000  -0.7096   0.06099   0.05555  -0.0637   1.0000   0.0578
  -8.750  -0.7190   0.05612   0.04984  -0.0662   1.0000   0.0504
  -8.500  -0.7110   0.05179   0.04543  -0.0655   1.0000   0.0491
  -8.250  -0.7010   0.04827   0.04169  -0.0652   1.0000   0.0480
  -8.000  -0.6880   0.04509   0.03826  -0.0648   1.0000   0.0472
  -7.750  -0.6729   0.04241   0.03533  -0.0638   1.0000   0.0470
  -7.500  -0.6570   0.04021   0.03293  -0.0623   1.0000   0.0474
  -7.250  -0.6403   0.03858   0.03108  -0.0610   1.0000   0.0492
  -7.000  -0.6256   0.03702   0.02940  -0.0587   1.0000   0.0520
  -6.750  -0.6146   0.03603   0.02858  -0.0552   1.0000   0.0550
  -6.500  -0.6031   0.03544   0.02802  -0.0515   1.0000   0.0577
  -6.250  -0.5887   0.03486   0.02734  -0.0491   1.0000   0.0624
  -6.000  -0.5885   0.03574   0.02857  -0.0408   1.0000   0.0674
  -5.750  -0.5808   0.03673   0.02965  -0.0349   1.0000   0.0745
  -5.500  -0.5584   0.03873   0.03184  -0.0305   0.9959   0.0896
  -5.250  -0.4560   0.03238   0.02697  -0.0616   0.9903   0.3647
  -5.000  -0.4418   0.04742   0.04223  -0.0377   0.9736   0.4070
  -4.750  -0.3883   0.06646   0.06133  -0.0042   0.9709   0.5400
  -4.500  -0.3134   0.06520   0.05974  -0.0047   0.9709   0.5630
  -4.250  -0.2922   0.06451   0.05891  -0.0067   0.9587   0.5729
  -4.000  -0.2503   0.06334   0.05756  -0.0096   0.9501   0.5814
  -3.750  -0.2282   0.06235   0.05646  -0.0116   0.9387   0.5880
  -3.500  -0.1790   0.06076   0.05469  -0.0159   0.9325   0.5927
  -3.250  -0.1699   0.06043   0.05429  -0.0172   0.9184   0.6015
  -3.000  -0.1338   0.05861   0.05238  -0.0189   0.9090   0.6034
  -2.750  -0.0876   0.05694   0.05057  -0.0234   0.9038   0.6057
  -2.500  -0.0781   0.05697   0.05054  -0.0252   0.8893   0.6158
  -2.250  -0.0517   0.05519   0.04870  -0.0256   0.8779   0.6166
  -2.000  -0.0196   0.05354   0.04699  -0.0272   0.8687   0.6177
  -1.750   0.0219   0.05183   0.04519  -0.0305   0.8624   0.6195
  -1.500   0.0503   0.05058   0.04390  -0.0315   0.8520   0.6226
  -1.250   0.0837   0.04945   0.04269  -0.0357   0.8456   0.6317
  -1.000   0.1100   0.04794   0.04116  -0.0355   0.8346   0.6335
  -0.750   0.1597   0.04602   0.03918  -0.0400   0.8312   0.6360
  -0.500   0.1768   0.04565   0.03878  -0.0409   0.8186   0.6465
  -0.250   0.2328   0.04345   0.03652  -0.0462   0.8158   0.6481
   0.000   0.2638   0.04221   0.03527  -0.0471   0.8048   0.6513
   0.250   0.3231   0.04103   0.03402  -0.0554   0.8001   0.6619
   0.500   0.3547   0.03967   0.03265  -0.0561   0.7883   0.6644
   0.750   0.3947   0.03858   0.03153  -0.0590   0.7784   0.6694
   1.000   0.4410   0.03754   0.03044  -0.0640   0.7699   0.6788
   1.250   0.4667   0.03679   0.02969  -0.0639   0.7575   0.6840
   1.500   0.4881   0.03631   0.02920  -0.0639   0.7458   0.6945
   1.750   0.5204   0.03587   0.02872  -0.0658   0.7362   0.7074
   2.000   0.5474   0.03495   0.02781  -0.0656   0.7258   0.7123
   2.250   0.5609   0.03466   0.02756  -0.0635   0.7155   0.7255
   2.500   0.5974   0.03399   0.02684  -0.0655   0.7078   0.7396
   2.750   0.7741   0.02943   0.02161  -0.1247   0.6927   0.5391
   3.000   0.8012   0.02928   0.02152  -0.1261   0.6826   0.5390
   3.250   0.8405   0.02893   0.02115  -0.1297   0.6748   0.5391
   3.500   0.8714   0.02878   0.02104  -0.1316   0.6658   0.5391
   3.750   0.9079   0.02862   0.02088  -0.1346   0.6584   0.5396
   4.000   0.9382   0.02858   0.02092  -0.1362   0.6501   0.5405
   4.250   0.9668   0.02865   0.02106  -0.1368   0.6429   0.5418
   4.500   0.9931   0.02876   0.02128  -0.1371   0.6346   0.5433
   4.750   1.0212   0.02891   0.02152  -0.1376   0.6268   0.5455
   5.000   1.0510   0.02898   0.02167  -0.1385   0.6182   0.5484
   5.250   1.0795   0.02913   0.02190  -0.1394   0.6090   0.5514
   5.500   1.1239   0.02884   0.02152  -0.1431   0.5997   0.5550
   5.750   1.1423   0.02905   0.02194  -0.1417   0.5887   0.5568
   6.000   1.1674   0.02914   0.02215  -0.1412   0.5785   0.5592
   6.250   1.2024   0.02895   0.02195  -0.1424   0.5684   0.5624
   6.500   1.2260   0.02905   0.02221  -0.1420   0.5564   0.5657
   6.750   1.2542   0.02909   0.02231  -0.1425   0.5440   0.5698
   7.000   1.2781   0.02905   0.02241  -0.1415   0.5319   0.5732
   7.250   1.3040   0.02894   0.02239  -0.1408   0.5192   0.5774
   7.500   1.3309   0.02879   0.02231  -0.1406   0.5050   0.5825
   7.750   1.3546   0.02864   0.02224  -0.1397   0.4894   0.5869
   8.000   1.3728   0.02852   0.02227  -0.1377   0.4730   0.5907
   8.250   1.3897   0.02841   0.02230  -0.1356   0.4550   0.5950
   8.500   1.4071   0.02828   0.02224  -0.1336   0.4348   0.6000
   8.750   1.4184   0.02818   0.02224  -0.1304   0.4132   0.6039
   9.000   1.4226   0.02829   0.02253  -0.1262   0.3878   0.6082
   9.250   1.4224   0.02857   0.02285  -0.1213   0.3573   0.6130
   9.500   1.4170   0.02921   0.02346  -0.1160   0.3186   0.6172
   9.750   1.4076   0.03051   0.02450  -0.1105   0.2753   0.6208
  10.000   1.3964   0.03250   0.02617  -0.1055   0.2390   0.6248
  10.250   1.3872   0.03486   0.02827  -0.1015   0.2094   0.6293
  10.500   1.3801   0.03727   0.03054  -0.0979   0.1854   0.6331
  10.750   1.3745   0.03975   0.03290  -0.0949   0.1664   0.6373
  11.000   1.3720   0.04232   0.03531  -0.0925   0.1499   0.6424
  11.250   1.3725   0.04484   0.03772  -0.0907   0.1349   0.6476
  11.500   1.3742   0.04726   0.04015  -0.0889   0.1222   0.6531
  11.750   1.3773   0.04976   0.04270  -0.0876   0.1107   0.6601
  12.000   1.3820   0.05221   0.04525  -0.0862   0.1006   0.6665
  12.250   1.3899   0.05469   0.04768  -0.0852   0.0915   0.6750
  12.500   1.3951   0.05710   0.05015  -0.0841   0.0838   0.6826
  12.750   1.4033   0.05981   0.05300  -0.0832   0.0765   0.6922
  13.000   1.4198   0.06214   0.05518  -0.0825   0.0687   0.7033
  13.250   1.4176   0.06524   0.05868  -0.0814   0.0646   0.7122
  13.500   1.4211   0.06798   0.06150  -0.0809   0.0599   0.7242
  13.750   1.4300   0.07125   0.06495  -0.0802   0.0553   0.7384
  14.000   1.4262   0.07480   0.06883  -0.0797   0.0521   0.7523
  14.250   1.4267   0.07783   0.07200  -0.0794   0.0490   0.7707
  14.500   1.4342   0.08162   0.07600  -0.0787   0.0459   0.7986
  14.750   1.4201   0.08588   0.08073  -0.0784   0.0447   0.8343
  15.000   1.4054   0.09004   0.08529  -0.0784   0.0436   1.0000
  15.250   1.3954   0.09545   0.09094  -0.0804   0.0421   1.0000
  15.500   1.3882   0.10046   0.09609  -0.0823   0.0406   1.0000
  15.750   1.3796   0.10561   0.10137  -0.0844   0.0394   1.0000
  16.000   1.3752   0.11049   0.10632  -0.0862   0.0382   1.0000
  16.250   1.3642   0.11663   0.11263  -0.0887   0.0377   1.0000
  16.500   1.3454   0.12383   0.12008  -0.0925   0.0377   1.0000
  16.750   1.3189   0.13251   0.12909  -0.0981   0.0381   1.0000
<< Back to NREL's S825 Airfoil (s825-nr)

Polar data table (+)

Polar graphs


<< Back to NREL's S825 Airfoil (s825-nr)