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 S834 Airfoil (s834-nr) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: NREL's S834 Airfoil (s834-nr)
Reynolds number: 50,000
Max Cl/Cd: 29.85 at α=8.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-s834-nr-50000.txt
Download as CSV file: xf-s834-nr-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NREL's S834 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -11.500  -0.5301   0.11104   0.10393  -0.0465   1.0000   0.1887
 -11.250  -0.5379   0.10623   0.09917  -0.0465   1.0000   0.1896
 -11.000  -0.7108   0.07960   0.07295  -0.0615   1.0000   0.1463
 -10.750  -0.7401   0.07562   0.06899  -0.0591   1.0000   0.1456
 -10.500  -0.7713   0.07236   0.06573  -0.0553   1.0000   0.1446
 -10.250  -0.8035   0.06935   0.06270  -0.0506   1.0000   0.1434
 -10.000  -0.8351   0.06591   0.05920  -0.0460   1.0000   0.1425
  -9.750  -0.8631   0.06247   0.05557  -0.0413   1.0000   0.1427
  -9.500  -0.8873   0.05905   0.05187  -0.0366   1.0000   0.1437
  -9.250  -0.9096   0.05559   0.04797  -0.0318   1.0000   0.1456
  -9.000  -0.8974   0.05381   0.04627  -0.0298   1.0000   0.1542
  -8.750  -0.9086   0.05044   0.04244  -0.0257   1.0000   0.1582
  -8.500  -0.9045   0.04780   0.03964  -0.0230   1.0000   0.1650
  -8.250  -0.9063   0.04534   0.03669  -0.0195   1.0000   0.1742
  -8.000  -0.8944   0.04365   0.03501  -0.0174   1.0000   0.1866
  -7.750  -0.8828   0.04172   0.03298  -0.0152   1.0000   0.1985
  -7.500  -0.8727   0.04009   0.03121  -0.0128   1.0000   0.2144
  -7.250  -0.8615   0.03841   0.02932  -0.0106   1.0000   0.2307
  -7.000  -0.8465   0.03700   0.02792  -0.0087   1.0000   0.2487
  -6.750  -0.8318   0.03594   0.02690  -0.0067   1.0000   0.2700
  -6.500  -0.8172   0.03471   0.02560  -0.0047   1.0000   0.2924
  -6.250  -0.8024   0.03371   0.02452  -0.0027   1.0000   0.3172
  -6.000  -0.7858   0.03288   0.02380  -0.0008   1.0000   0.3425
  -5.750  -0.7697   0.03209   0.02307   0.0012   1.0000   0.3699
  -5.500  -0.7533   0.03137   0.02240   0.0032   1.0000   0.3985
  -5.250  -0.7370   0.03076   0.02184   0.0053   1.0000   0.4288
  -5.000  -0.7205   0.03020   0.02134   0.0074   1.0000   0.4604
  -4.750  -0.7043   0.02968   0.02084   0.0096   1.0000   0.4936
  -4.500  -0.6876   0.02939   0.02064   0.0121   1.0000   0.5257
  -4.250  -0.6714   0.02918   0.02053   0.0147   1.0000   0.5588
  -4.000  -0.6558   0.02900   0.02040   0.0174   1.0000   0.5936
  -3.750  -0.6405   0.02902   0.02050   0.0206   1.0000   0.6272
  -3.500  -0.6257   0.02921   0.02076   0.0241   1.0000   0.6600
  -3.250  -0.6115   0.02949   0.02107   0.0277   1.0000   0.6929
  -3.000  -0.5982   0.02981   0.02136   0.0315   1.0000   0.7268
  -2.750  -0.5833   0.03044   0.02197   0.0354   1.0000   0.7575
  -2.500  -0.5663   0.03121   0.02270   0.0391   1.0000   0.7883
  -2.250  -0.5438   0.03212   0.02351   0.0417   1.0000   0.8197
  -2.000  -0.5144   0.03304   0.02427   0.0425   1.0000   0.8532
  -1.750  -0.4683   0.03413   0.02516   0.0399   1.0000   0.8869
  -1.500  -0.3845   0.03567   0.02639   0.0298   1.0000   0.9172
  -1.250  -0.2869   0.03685   0.02724   0.0160   1.0000   0.9459
  -1.000  -0.1916   0.03753   0.02765   0.0014   1.0000   0.9747
  -0.750  -0.1038   0.03787   0.02775  -0.0129   1.0000   1.0000
  -0.500  -0.1102   0.03716   0.02698  -0.0101   1.0000   1.0000
  -0.250  -0.1161   0.03652   0.02629  -0.0072   1.0000   1.0000
   0.000  -0.1211   0.03597   0.02567  -0.0043   1.0000   1.0000
   0.250  -0.1250   0.03550   0.02514  -0.0013   1.0000   1.0000
   0.500  -0.1275   0.03511   0.02469   0.0017   1.0000   1.0000
   0.750  -0.1286   0.03480   0.02432   0.0045   1.0000   1.0000
   1.000  -0.1284   0.03458   0.02403   0.0073   1.0000   1.0000
   1.250  -0.1267   0.03443   0.02382   0.0099   1.0000   1.0000
   1.500  -0.1237   0.03436   0.02370   0.0123   1.0000   1.0000
   1.750  -0.1192   0.03438   0.02367   0.0145   1.0000   1.0000
   2.000  -0.1135   0.03449   0.02372   0.0165   1.0000   1.0000
   2.250  -0.1063   0.03469   0.02387   0.0183   1.0000   1.0000
   2.500  -0.0974   0.03500   0.02414   0.0197   1.0000   1.0000
   2.750  -0.0869   0.03543   0.02453   0.0208   1.0000   1.0000
   3.000  -0.0753   0.03596   0.02504   0.0216   1.0000   1.0000
   3.250  -0.0628   0.03659   0.02565   0.0222   1.0000   1.0000
   3.500  -0.0498   0.03731   0.02636   0.0227   1.0000   1.0000
   3.750  -0.0364   0.03812   0.02717   0.0230   1.0000   1.0000
   4.000  -0.0229   0.03902   0.02807   0.0231   1.0000   1.0000
   4.250   0.1395   0.04498   0.03423  -0.0018   0.9022   1.0000
   4.500   0.1702   0.04582   0.03513  -0.0036   0.8795   1.0000
   4.750   0.2123   0.04700   0.03641  -0.0069   0.8579   1.0000
   5.000   0.2413   0.04777   0.03727  -0.0081   0.8366   1.0000
   5.250   0.2855   0.04876   0.03837  -0.0113   0.8162   1.0000
   5.500   0.3095   0.04939   0.03912  -0.0115   0.7948   1.0000
   5.750   0.3579   0.05003   0.03992  -0.0147   0.7747   1.0000
   6.000   0.3785   0.05056   0.04056  -0.0142   0.7523   1.0000
   6.250   0.4287   0.05066   0.04089  -0.0168   0.7320   1.0000
   6.500   0.4497   0.05106   0.04142  -0.0160   0.7088   1.0000
   6.750   0.5060   0.05029   0.04091  -0.0183   0.6884   1.0000
   7.000   0.5296   0.05029   0.04108  -0.0172   0.6638   1.0000
   7.250   0.5694   0.04943   0.04048  -0.0171   0.6405   1.0000
   7.500   0.6355   0.04640   0.03783  -0.0180   0.6183   1.0000
   7.750   0.6843   0.04360   0.03535  -0.0168   0.5929   1.0000
   8.000   0.8171   0.03431   0.02671  -0.0205   0.5555   1.0000
   8.250   0.8912   0.03049   0.02283  -0.0215   0.4852   1.0000
   8.500   0.9097   0.03048   0.02245  -0.0181   0.4245   1.0000
   8.750   0.9230   0.03139   0.02295  -0.0148   0.3677   1.0000
   9.000   0.9312   0.03278   0.02399  -0.0113   0.3193   1.0000
   9.250   0.9404   0.03439   0.02526  -0.0083   0.2760   1.0000
   9.500   0.9541   0.03625   0.02682  -0.0061   0.2371   1.0000
   9.750   0.9695   0.03831   0.02872  -0.0043   0.2045   1.0000
  10.000   0.9861   0.04050   0.03079  -0.0028   0.1778   1.0000
  10.250   1.0125   0.04325   0.03340  -0.0028   0.1542   1.0000
  10.500   1.0246   0.04567   0.03595  -0.0008   0.1391   1.0000
  10.750   1.0445   0.04883   0.03923  -0.0001   0.1265   1.0000
  11.000   1.0523   0.05120   0.04184   0.0023   0.1173   1.0000
  11.250   1.0527   0.05444   0.04541   0.0054   0.1120   1.0000
  11.500   1.0448   0.05755   0.04893   0.0094   0.1094   1.0000
  11.750   1.0388   0.06065   0.05229   0.0126   0.1063   1.0000
  12.000   1.0506   0.06431   0.05593   0.0134   0.1007   1.0000
  12.250   1.0288   0.06736   0.05931   0.0175   0.1001   1.0000
  12.500   1.0059   0.07088   0.06314   0.0209   0.0998   1.0000
  12.750   0.9811   0.07476   0.06728   0.0234   0.0996   1.0000
  13.000   0.9567   0.07918   0.07192   0.0249   0.0999   1.0000
  13.250   0.9336   0.08411   0.07703   0.0254   0.1004   1.0000
  13.500   0.9084   0.08960   0.08267   0.0250   0.1009   1.0000
  13.750   0.8878   0.09554   0.08872   0.0239   0.1015   1.0000
<< Back to NREL's S834 Airfoil (s834-nr)

Polar data table (+)

Polar graphs


<< Back to NREL's S834 Airfoil (s834-nr)