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 S802 Airfoil (s802-nr) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: NREL's S802 Airfoil (s802-nr)
Reynolds number: 500,000
Max Cl/Cd: 124.33 at α=5.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-s802-nr-500000.txt
Download as CSV file: xf-s802-nr-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NREL's S802 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.250  -0.3050   0.09008   0.08786  -0.0614   0.9913   0.0225
  -9.000  -0.2923   0.08520   0.08298  -0.0668   0.9880   0.0236
  -8.250  -0.3002   0.05053   0.04808  -0.1104   0.9528   0.0264
  -8.000  -0.2777   0.04854   0.04610  -0.1131   0.9482   0.0269
  -7.750  -0.2588   0.04532   0.04277  -0.1170   0.9385   0.0275
  -7.500  -0.2284   0.04144   0.03872  -0.1236   0.9339   0.0290
  -7.250  -0.2079   0.03253   0.02892  -0.1321   0.9211   0.0335
  -7.000  -0.1760   0.03104   0.02749  -0.1346   0.9122   0.0347
  -6.750  -0.1568   0.02121   0.01642  -0.1366   0.8980   0.0247
  -6.500  -0.1293   0.01843   0.01320  -0.1376   0.8873   0.0240
  -6.250  -0.1021   0.01666   0.01110  -0.1379   0.8767   0.0238
  -6.000  -0.0756   0.01564   0.00985  -0.1379   0.8664   0.0245
  -5.750  -0.0481   0.01478   0.00879  -0.1379   0.8576   0.0251
  -5.500  -0.0220   0.01394   0.00779  -0.1377   0.8485   0.0255
  -5.250   0.0048   0.01332   0.00703  -0.1376   0.8406   0.0258
  -5.000   0.0305   0.01245   0.00606  -0.1373   0.8330   0.0262
  -4.750   0.0566   0.01164   0.00518  -0.1373   0.8264   0.0271
  -4.500   0.0831   0.01117   0.00468  -0.1372   0.8197   0.0283
  -4.000   0.1381   0.01053   0.00392  -0.1374   0.8081   0.0313
  -3.750   0.1660   0.01027   0.00358  -0.1375   0.8026   0.0325
  -3.500   0.1944   0.00991   0.00317  -0.1378   0.7978   0.0360
  -3.250   0.2221   0.00969   0.00293  -0.1378   0.7927   0.0417
  -3.000   0.2503   0.00912   0.00266  -0.1383   0.7882   0.1119
  -2.750   0.2792   0.00871   0.00252  -0.1390   0.7843   0.2050
  -2.500   0.3074   0.00832   0.00244  -0.1396   0.7801   0.3023
  -2.250   0.3359   0.00784   0.00237  -0.1403   0.7758   0.4336
  -2.000   0.3644   0.00737   0.00238  -0.1408   0.7720   0.5887
  -1.750   0.3922   0.00739   0.00259  -0.1407   0.7686   0.6865
  -1.500   0.4197   0.00748   0.00270  -0.1405   0.7651   0.7166
  -1.250   0.4474   0.00760   0.00280  -0.1403   0.7614   0.7388
  -1.000   0.4748   0.00772   0.00292  -0.1400   0.7579   0.7558
  -0.750   0.5027   0.00784   0.00300  -0.1398   0.7548   0.7662
  -0.500   0.5304   0.00795   0.00309  -0.1397   0.7517   0.7749
  -0.250   0.5578   0.00804   0.00318  -0.1396   0.7482   0.7840
   0.000   0.5848   0.00810   0.00325  -0.1393   0.7446   0.7908
   0.250   0.6128   0.00820   0.00331  -0.1393   0.7412   0.7986
   0.500   0.6406   0.00830   0.00339  -0.1391   0.7377   0.8047
   0.750   0.6668   0.00835   0.00347  -0.1387   0.7334   0.8114
   1.000   0.6941   0.00840   0.00352  -0.1386   0.7289   0.8175
   1.250   0.7211   0.00846   0.00356  -0.1383   0.7244   0.8229
   1.500   0.7481   0.00852   0.00363  -0.1380   0.7190   0.8289
   1.750   0.7741   0.00852   0.00365  -0.1376   0.7129   0.8338
   2.000   0.8013   0.00859   0.00368  -0.1373   0.7072   0.8384
   2.250   0.8274   0.00859   0.00372  -0.1370   0.7006   0.8430
   2.500   0.8549   0.00861   0.00373  -0.1369   0.6943   0.8464
   2.750   0.8806   0.00862   0.00376  -0.1364   0.6872   0.8494
   3.000   0.9070   0.00863   0.00378  -0.1361   0.6800   0.8526
   3.250   0.9337   0.00866   0.00383  -0.1358   0.6728   0.8558
   3.500   0.9605   0.00869   0.00387  -0.1357   0.6650   0.8590
   3.750   0.9864   0.00870   0.00392  -0.1353   0.6565   0.8618
   4.000   1.0119   0.00874   0.00395  -0.1347   0.6471   0.8646
   4.250   1.0369   0.00875   0.00403  -0.1342   0.6364   0.8680
   4.500   1.0621   0.00880   0.00410  -0.1336   0.6236   0.8714
   5.000   1.1093   0.00894   0.00423  -0.1319   0.5871   0.8780
   5.250   1.1302   0.00909   0.00434  -0.1305   0.5543   0.8821
   5.500   1.1447   0.00958   0.00455  -0.1279   0.4925   0.8867
   5.750   1.1550   0.01037   0.00500  -0.1247   0.4247   0.8914
   6.000   1.1659   0.01117   0.00550  -0.1217   0.3639   0.8969
   6.250   1.1774   0.01199   0.00604  -0.1189   0.3061   0.9034
   6.500   1.1883   0.01269   0.00654  -0.1159   0.2598   0.9106
   6.750   1.1996   0.01334   0.00701  -0.1131   0.2209   0.9193
   7.000   1.2083   0.01389   0.00745  -0.1097   0.1896   0.9337
   7.250   1.2202   0.01443   0.00790  -0.1070   0.1622   1.0000
   7.500   1.2364   0.01514   0.00849  -0.1054   0.1385   1.0000
   7.750   1.2525   0.01583   0.00908  -0.1038   0.1194   1.0000
   8.000   1.2680   0.01654   0.00970  -0.1020   0.1039   1.0000
   8.250   1.2834   0.01723   0.01034  -0.1003   0.0917   1.0000
   8.500   1.2988   0.01791   0.01100  -0.0986   0.0821   1.0000
   8.750   1.3123   0.01871   0.01177  -0.0966   0.0741   1.0000
   9.000   1.3275   0.01940   0.01247  -0.0950   0.0681   1.0000
   9.250   1.3393   0.02031   0.01338  -0.0928   0.0624   1.0000
   9.500   1.3548   0.02100   0.01412  -0.0913   0.0584   1.0000
   9.750   1.3658   0.02200   0.01510  -0.0893   0.0543   1.0000
  10.000   1.3769   0.02303   0.01619  -0.0873   0.0512   1.0000
  10.250   1.3915   0.02383   0.01705  -0.0858   0.0485   1.0000
  10.500   1.4037   0.02482   0.01807  -0.0842   0.0458   1.0000
  10.750   1.4088   0.02638   0.01965  -0.0818   0.0431   1.0000
  11.000   1.4211   0.02744   0.02080  -0.0803   0.0416   1.0000
  11.250   1.4345   0.02842   0.02187  -0.0790   0.0397   1.0000
  11.500   1.4473   0.02948   0.02298  -0.0778   0.0378   1.0000
  11.750   1.4560   0.03089   0.02441  -0.0762   0.0358   1.0000
  12.000   1.4604   0.03275   0.02633  -0.0744   0.0339   1.0000
  12.250   1.4755   0.03367   0.02734  -0.0735   0.0324   1.0000
  12.500   1.4874   0.03489   0.02864  -0.0725   0.0309   1.0000
  12.750   1.4980   0.03624   0.03003  -0.0714   0.0294   1.0000
  13.000   1.5029   0.03818   0.03201  -0.0701   0.0280   1.0000
  13.250   1.5058   0.04038   0.03431  -0.0686   0.0268   1.0000
  13.500   1.5170   0.04178   0.03583  -0.0678   0.0259   1.0000
  13.750   1.5264   0.04338   0.03752  -0.0670   0.0247   1.0000
  14.000   1.5354   0.04505   0.03926  -0.0663   0.0235   1.0000
  14.250   1.5408   0.04713   0.04138  -0.0655   0.0223   1.0000
  14.500   1.5372   0.05028   0.04462  -0.0644   0.0211   1.0000
  14.750   1.5472   0.05198   0.04644  -0.0640   0.0202   1.0000
  15.000   1.5549   0.05396   0.04853  -0.0636   0.0191   1.0000
  15.250   1.5612   0.05615   0.05079  -0.0633   0.0181   1.0000
  15.500   1.5612   0.05917   0.05387  -0.0630   0.0171   1.0000
  15.750   1.5565   0.06288   0.05771  -0.0627   0.0163   1.0000
  16.000   1.5611   0.06551   0.06048  -0.0627   0.0154   1.0000
  16.250   1.5632   0.06853   0.06362  -0.0629   0.0146   1.0000
  16.500   1.5636   0.07185   0.06703  -0.0633   0.0139   1.0000
  16.750   1.5587   0.07603   0.07129  -0.0639   0.0133   1.0000
  17.000   1.5457   0.08158   0.07695  -0.0650   0.0127   1.0000
  17.250   1.5439   0.08564   0.08118  -0.0660   0.0123   1.0000
  17.500   1.5391   0.09027   0.08598  -0.0674   0.0118   1.0000
  17.750   1.5332   0.09523   0.09108  -0.0691   0.0114   1.0000
  18.000   1.5265   0.10045   0.09642  -0.0711   0.0111   1.0000
  18.250   1.5186   0.10604   0.10214  -0.0735   0.0107   1.0000
  18.500   1.5093   0.11202   0.10824  -0.0762   0.0105   1.0000
  18.750   1.4976   0.11857   0.11492  -0.0795   0.0103   1.0000
  19.000   1.4842   0.12554   0.12201  -0.0832   0.0101   1.0000
<< Back to NREL's S802 Airfoil (s802-nr)

Polar data table (+)

Polar graphs


<< Back to NREL's S802 Airfoil (s802-nr)