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 S830 Airfoil (s830-nr) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: NREL's S830 Airfoil (s830-nr)
Reynolds number: 500,000
Max Cl/Cd: 94.01 at α=7°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-s830-nr-500000-n5.txt
Download as CSV file: xf-s830-nr-500000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NREL's S830 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -13.000  -0.1647   0.08684   0.08425  -0.1363   0.9636   0.0056
 -12.750  -0.1801   0.07801   0.07536  -0.1428   0.9622   0.0054
 -12.500  -0.2023   0.06934   0.06655  -0.1489   0.9608   0.0053
 -12.250  -0.2320   0.06330   0.06039  -0.1494   0.9547   0.0054
 -12.000  -0.2482   0.05824   0.05518  -0.1515   0.9505   0.0053
 -11.500  -0.2932   0.04670   0.04318  -0.1517   0.9378   0.0054
 -11.000  -0.3225   0.03569   0.03131  -0.1479   0.9216   0.0054
 -10.750  -0.3116   0.03353   0.02886  -0.1471   0.9141   0.0056
 -10.500  -0.2899   0.03153   0.02658  -0.1477   0.9086   0.0056
 -10.250  -0.2585   0.02998   0.02479  -0.1496   0.9044   0.0057
 -10.000  -0.2266   0.02888   0.02350  -0.1510   0.8978   0.0057
  -9.750  -0.1895   0.02796   0.02242  -0.1537   0.8922   0.0057
  -9.500  -0.1540   0.02727   0.02158  -0.1555   0.8836   0.0058
  -9.250  -0.1203   0.02666   0.02084  -0.1575   0.8732   0.0059
  -9.000  -0.0927   0.02614   0.02021  -0.1585   0.8610   0.0061
  -8.750  -0.0712   0.02569   0.01965  -0.1586   0.8477   0.0061
  -8.500  -0.0547   0.02524   0.01910  -0.1577   0.8345   0.0062
  -8.250  -0.0401   0.02480   0.01858  -0.1566   0.8220   0.0064
  -8.000  -0.0271   0.02435   0.01804  -0.1552   0.8102   0.0065
  -7.750  -0.0152   0.02389   0.01750  -0.1537   0.7992   0.0067
  -7.500  -0.0051   0.02340   0.01696  -0.1519   0.7886   0.0070
  -7.250   0.0035   0.02289   0.01637  -0.1500   0.7785   0.0070
  -7.000   0.0114   0.02237   0.01578  -0.1479   0.7691   0.0072
  -6.750   0.0162   0.02177   0.01513  -0.1453   0.7593   0.0078
  -6.500   0.0205   0.02121   0.01451  -0.1426   0.7505   0.0081
  -6.250   0.0174   0.02053   0.01379  -0.1388   0.7418   0.0089
  -6.000   0.0124   0.01994   0.01316  -0.1344   0.7338   0.0089
  -5.750   0.0150   0.01924   0.01242  -0.1317   0.7258   0.0095
  -5.500   0.0215   0.01846   0.01157  -0.1300   0.7185   0.0108
  -5.250   0.0320   0.01756   0.01064  -0.1292   0.7111   0.0134
  -5.000   0.0483   0.01667   0.00970  -0.1294   0.7042   0.0165
  -4.750   0.0698   0.01578   0.00878  -0.1307   0.6974   0.0211
  -4.250   0.1266   0.01386   0.00684  -0.1361   0.6853   0.0453
  -3.750   0.2197   0.00953   0.00346  -0.1533   0.6733   0.4446
  -3.500   0.2483   0.00953   0.00347  -0.1539   0.6670   0.4681
  -3.250   0.2757   0.00963   0.00351  -0.1541   0.6610   0.4821
  -3.000   0.3034   0.00972   0.00352  -0.1543   0.6552   0.4913
  -2.750   0.3302   0.00982   0.00359  -0.1544   0.6494   0.4962
  -2.250   0.3850   0.00997   0.00355  -0.1548   0.6390   0.5043
  -2.000   0.4113   0.01004   0.00361  -0.1548   0.6335   0.5070
  -1.500   0.4643   0.01012   0.00358  -0.1549   0.6236   0.5092
  -1.250   0.4905   0.01015   0.00357  -0.1549   0.6183   0.5101
  -1.000   0.5160   0.01021   0.00356  -0.1547   0.6133   0.5111
  -0.750   0.5423   0.01024   0.00355  -0.1548   0.6083   0.5121
  -0.500   0.5681   0.01028   0.00356  -0.1547   0.6031   0.5133
  -0.250   0.5935   0.01035   0.00357  -0.1546   0.5983   0.5145
   0.000   0.6198   0.01040   0.00359  -0.1546   0.5940   0.5158
   0.250   0.6462   0.01046   0.00361  -0.1547   0.5895   0.5171
   0.500   0.6720   0.01053   0.00364  -0.1547   0.5850   0.5182
   0.750   0.6975   0.01062   0.00368  -0.1546   0.5810   0.5193
   1.000   0.7238   0.01067   0.00375  -0.1546   0.5770   0.5203
   1.250   0.7496   0.01074   0.00383  -0.1546   0.5729   0.5214
   1.500   0.7750   0.01084   0.00392  -0.1544   0.5688   0.5225
   1.750   0.8003   0.01096   0.00402  -0.1543   0.5650   0.5236
   2.000   0.8265   0.01104   0.00413  -0.1544   0.5610   0.5248
   2.250   0.8521   0.01115   0.00424  -0.1543   0.5569   0.5260
   2.500   0.8775   0.01127   0.00436  -0.1542   0.5532   0.5273
   2.750   0.9027   0.01141   0.00448  -0.1541   0.5499   0.5287
   3.000   0.9289   0.01151   0.00460  -0.1542   0.5467   0.5301
   3.250   0.9546   0.01162   0.00473  -0.1542   0.5433   0.5314
   3.500   0.9800   0.01174   0.00487  -0.1541   0.5400   0.5327
   3.750   1.0051   0.01189   0.00500  -0.1540   0.5368   0.5339
   4.000   1.0299   0.01204   0.00517  -0.1538   0.5338   0.5351
   4.250   1.0555   0.01215   0.00535  -0.1538   0.5308   0.5364
   4.500   1.0805   0.01228   0.00553  -0.1537   0.5275   0.5377
   4.750   1.1053   0.01243   0.00573  -0.1535   0.5245   0.5391
   5.000   1.1299   0.01260   0.00593  -0.1533   0.5217   0.5406
   5.250   1.1544   0.01278   0.00613  -0.1531   0.5192   0.5422
   5.500   1.1794   0.01294   0.00635  -0.1530   0.5168   0.5437
   5.750   1.2040   0.01309   0.00656  -0.1528   0.5140   0.5452
   6.000   1.2283   0.01326   0.00678  -0.1526   0.5110   0.5466
   6.250   1.2523   0.01344   0.00701  -0.1523   0.5082   0.5480
   6.500   1.2764   0.01364   0.00724  -0.1521   0.5058   0.5493
   6.750   1.3001   0.01385   0.00749  -0.1518   0.5031   0.5506
   7.000   1.3199   0.01404   0.00780  -0.1507   0.4966   0.5518
   7.250   1.3335   0.01440   0.00815  -0.1485   0.4857   0.5531
   7.500   1.3461   0.01480   0.00855  -0.1462   0.4724   0.5544
   7.750   1.3629   0.01513   0.00894  -0.1447   0.4621   0.5557
   8.000   1.3788   0.01554   0.00937  -0.1431   0.4528   0.5572
   8.250   1.3925   0.01603   0.00988  -0.1411   0.4411   0.5587
   8.500   1.4087   0.01647   0.01036  -0.1396   0.4299   0.5603
   8.750   1.4221   0.01703   0.01093  -0.1377   0.4160   0.5619
   9.000   1.4319   0.01776   0.01164  -0.1353   0.3979   0.5633
   9.250   1.4328   0.01893   0.01269  -0.1317   0.3671   0.5645
   9.500   1.4231   0.02075   0.01431  -0.1268   0.3255   0.5655
   9.750   1.4108   0.02293   0.01631  -0.1219   0.2865   0.5665
  10.000   1.4019   0.02513   0.01835  -0.1179   0.2524   0.5674
  10.250   1.3937   0.02745   0.02055  -0.1143   0.2210   0.5684
  10.500   1.3876   0.02978   0.02278  -0.1112   0.1937   0.5694
  10.750   1.3813   0.03229   0.02519  -0.1084   0.1676   0.5705
  11.000   1.3760   0.03485   0.02766  -0.1060   0.1443   0.5716
  11.250   1.3710   0.03752   0.03025  -0.1038   0.1222   0.5727
  11.500   1.3673   0.04021   0.03288  -0.1019   0.1036   0.5738
  11.750   1.3654   0.04284   0.03546  -0.1003   0.0884   0.5750
  12.000   1.3654   0.04538   0.03799  -0.0990   0.0759   0.5764
  12.500   1.3691   0.05034   0.04296  -0.0970   0.0572   0.5790
  12.750   1.3722   0.05279   0.04542  -0.0962   0.0504   0.5801
  13.000   1.3743   0.05539   0.04804  -0.0955   0.0442   0.5812
  13.250   1.3784   0.05784   0.05054  -0.0949   0.0390   0.5824
  13.500   1.3821   0.06037   0.05312  -0.0944   0.0347   0.5835
  13.750   1.3857   0.06297   0.05577  -0.0940   0.0307   0.5846
  14.000   1.3899   0.06554   0.05840  -0.0937   0.0275   0.5858
  14.250   1.3935   0.06820   0.06110  -0.0934   0.0244   0.5871
  14.500   1.3975   0.07087   0.06383  -0.0933   0.0216   0.5884
  14.750   1.4007   0.07365   0.06665  -0.0931   0.0187   0.5899
  15.000   1.4046   0.07641   0.06946  -0.0931   0.0163   0.5914
  15.500   1.4123   0.08198   0.07514  -0.0932   0.0126   0.5941
  15.750   1.4154   0.08491   0.07811  -0.0933   0.0110   0.5953
  16.000   1.4195   0.08774   0.08101  -0.0935   0.0095   0.5964
<< Back to NREL's S830 Airfoil (s830-nr)

Polar data table (+)

Polar graphs


<< Back to NREL's S830 Airfoil (s830-nr)