Airfoil Tools
Search 1636 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

DAE-21 AIRFOIL (dae21-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: DAE-21 AIRFOIL (dae21-il)
Reynolds number: 100,000
Max Cl/Cd: 50.14 at α=7.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-dae21-il-100000-n5.txt
Download as CSV file: xf-dae21-il-100000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: DAE-21 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.250  -0.0789   0.11438   0.10880  -0.0738   0.8334   0.0452
 -10.000  -0.0758   0.11244   0.10687  -0.0761   0.8297   0.0459
  -9.750  -0.0751   0.11070   0.10517  -0.0787   0.8253   0.0462
  -9.500  -0.0690   0.10794   0.10242  -0.0803   0.8215   0.0464
  -9.250  -0.0508   0.10380   0.09825  -0.0795   0.8189   0.0472
  -9.000  -0.0403   0.10105   0.09548  -0.0799   0.8162   0.0482
  -8.750  -0.0312   0.09855   0.09300  -0.0811   0.8127   0.0492
  -8.500  -0.0234   0.09612   0.09058  -0.0824   0.8087   0.0505
  -8.250  -0.0172   0.09381   0.08829  -0.0837   0.8049   0.0518
  -8.000  -0.0159   0.09190   0.08638  -0.0859   0.8012   0.0533
  -7.750  -0.0175   0.09013   0.08464  -0.0888   0.7976   0.0538
  -7.500  -0.0183   0.08826   0.08284  -0.0914   0.7926   0.0540
  -7.250  -0.0025   0.08450   0.07909  -0.0906   0.7896   0.0547
  -7.000   0.0100   0.08178   0.07637  -0.0902   0.7864   0.0556
  -6.750   0.0173   0.07958   0.07416  -0.0903   0.7834   0.0565
  -6.500   0.0239   0.07747   0.07206  -0.0912   0.7803   0.0578
  -6.250   0.0310   0.07532   0.06996  -0.0941   0.7755   0.0609
  -6.000   0.0402   0.07227   0.06686  -0.1068   0.7701   0.0632
  -5.750   0.0505   0.06933   0.06394  -0.1037   0.7675   0.0643
  -5.500   0.0636   0.06727   0.06188  -0.1017   0.7653   0.0666
  -5.250   0.0780   0.06479   0.05940  -0.1045   0.7613   0.0687
  -4.750   0.1232   0.05164   0.04582  -0.1214   0.7526   0.0400
  -4.500   0.1436   0.04908   0.04316  -0.1228   0.7499   0.0385
  -4.250   0.1684   0.04528   0.03914  -0.1260   0.7476   0.0371
  -4.000   0.1944   0.04083   0.03433  -0.1300   0.7429   0.0355
  -3.750   0.2213   0.03710   0.03014  -0.1326   0.7387   0.0351
  -3.500   0.2474   0.03516   0.02792  -0.1337   0.7355   0.0366
  -3.250   0.2760   0.03270   0.02498  -0.1348   0.7329   0.0377
  -3.000   0.3054   0.03043   0.02218  -0.1355   0.7309   0.0380
  -2.750   0.3307   0.02902   0.02038  -0.1357   0.7262   0.0384
  -2.500   0.3567   0.02782   0.01880  -0.1357   0.7218   0.0391
  -2.250   0.3837   0.02687   0.01765  -0.1357   0.7186   0.0409
  -2.000   0.4115   0.02612   0.01676  -0.1356   0.7161   0.0436
  -1.750   0.4405   0.02519   0.01555  -0.1354   0.7140   0.0456
  -1.500   0.4628   0.02490   0.01526  -0.1349   0.7085   0.0472
  -1.250   0.4874   0.02457   0.01492  -0.1346   0.7041   0.0503
  -1.000   0.5145   0.02409   0.01437  -0.1343   0.7011   0.0552
  -0.750   0.5426   0.02358   0.01382  -0.1341   0.6987   0.0605
  -0.500   0.5717   0.02307   0.01327  -0.1340   0.6968   0.0698
  -0.250   0.5902   0.02343   0.01370  -0.1333   0.6892   0.0819
   0.000   0.6167   0.02307   0.01345  -0.1331   0.6857   0.1182
   0.250   0.6459   0.02225   0.01327  -0.1335   0.6831   0.3137
   0.500   0.6705   0.02114   0.01317  -0.1324   0.6811   0.6587
   1.000   0.7187   0.02148   0.01368  -0.1314   0.6700   1.0000
   1.250   0.7477   0.02147   0.01348  -0.1313   0.6672   1.0000
   1.500   0.7778   0.02139   0.01321  -0.1313   0.6650   1.0000
   1.750   0.7941   0.02224   0.01403  -0.1303   0.6570   1.0000
   2.000   0.8209   0.02235   0.01403  -0.1300   0.6531   1.0000
   2.250   0.8505   0.02227   0.01382  -0.1299   0.6505   1.0000
   2.750   0.8937   0.02316   0.01461  -0.1285   0.6391   1.0000
   3.000   0.9226   0.02309   0.01446  -0.1284   0.6358   1.0000
   3.250   0.9540   0.02287   0.01415  -0.1285   0.6335   1.0000
   3.500   0.9663   0.02385   0.01517  -0.1269   0.6246   1.0000
   3.750   0.9945   0.02380   0.01507  -0.1267   0.6208   1.0000
   4.000   1.0264   0.02352   0.01473  -0.1268   0.6181   1.0000
   4.250   1.0381   0.02447   0.01573  -0.1251   0.6092   1.0000
   4.500   1.0664   0.02437   0.01561  -0.1249   0.6051   1.0000
   4.750   1.0994   0.02400   0.01519  -0.1251   0.6022   1.0000
   5.000   1.1085   0.02502   0.01629  -0.1230   0.5926   1.0000
   5.250   1.1388   0.02477   0.01604  -0.1230   0.5885   1.0000
   5.500   1.1584   0.02514   0.01644  -0.1219   0.5818   1.0000
   5.750   1.1789   0.02543   0.01675  -0.1209   0.5749   1.0000
   6.000   1.2136   0.02492   0.01622  -0.1213   0.5711   1.0000
   6.250   1.2198   0.02596   0.01737  -0.1188   0.5609   1.0000
   6.750   1.2592   0.02646   0.01794  -0.1164   0.5462   1.0000
   7.000   1.2923   0.02604   0.01755  -0.1166   0.5407   1.0000
   7.250   1.2954   0.02707   0.01864  -0.1136   0.5305   1.0000
   7.500   1.3301   0.02653   0.01811  -0.1140   0.5244   1.0000
   7.750   1.3315   0.02773   0.01939  -0.1110   0.5134   1.0000
   8.000   1.3551   0.02776   0.01946  -0.1102   0.5051   1.0000
   8.250   1.3706   0.02825   0.01999  -0.1087   0.4950   1.0000
   8.500   1.3805   0.02913   0.02093  -0.1068   0.4840   1.0000
   8.750   1.4028   0.02923   0.02105  -0.1059   0.4741   1.0000
   9.000   1.4188   0.02973   0.02160  -0.1045   0.4627   1.0000
   9.250   1.4270   0.03080   0.02272  -0.1027   0.4502   1.0000
   9.500   1.4393   0.03161   0.02356  -0.1011   0.4378   1.0000
   9.750   1.4536   0.03229   0.02426  -0.0997   0.4253   1.0000
  10.000   1.4664   0.03307   0.02506  -0.0982   0.4120   1.0000
  10.250   1.4781   0.03396   0.02594  -0.0967   0.3983   1.0000
  10.500   1.4844   0.03530   0.02731  -0.0949   0.3839   1.0000
  10.750   1.4905   0.03670   0.02874  -0.0932   0.3693   1.0000
  11.000   1.4963   0.03816   0.03023  -0.0916   0.3548   1.0000
  11.250   1.5015   0.03973   0.03181  -0.0900   0.3402   1.0000
  11.500   1.5058   0.04140   0.03349  -0.0885   0.3254   1.0000
  11.750   1.5095   0.04318   0.03528  -0.0870   0.3108   1.0000
  12.000   1.5125   0.04507   0.03718  -0.0855   0.2966   1.0000
  12.250   1.5145   0.04710   0.03923  -0.0842   0.2823   1.0000
  12.500   1.5161   0.04924   0.04137  -0.0829   0.2683   1.0000
  12.750   1.5170   0.05153   0.04367  -0.0817   0.2545   1.0000
  13.000   1.5172   0.05398   0.04613  -0.0806   0.2410   1.0000
  13.250   1.5169   0.05658   0.04873  -0.0797   0.2277   1.0000
  13.500   1.5162   0.05932   0.05149  -0.0789   0.2149   1.0000
  13.750   1.5149   0.06223   0.05444  -0.0782   0.2022   1.0000
  14.000   1.5135   0.06524   0.05750  -0.0777   0.1902   1.0000
  14.250   1.5117   0.06837   0.06066  -0.0773   0.1787   1.0000
  14.500   1.5093   0.07166   0.06396  -0.0770   0.1680   1.0000
  14.750   1.5062   0.07511   0.06743  -0.0768   0.1578   1.0000
  15.000   1.5039   0.07860   0.07100  -0.0768   0.1478   1.0000
  15.250   1.5011   0.08218   0.07464  -0.0769   0.1389   1.0000
  15.500   1.4969   0.08600   0.07847  -0.0771   0.1307   1.0000
  15.750   1.4946   0.08970   0.08228  -0.0775   0.1224   1.0000
  16.000   1.4909   0.09362   0.08623  -0.0779   0.1155   1.0000
  16.250   1.4877   0.09756   0.09025  -0.0786   0.1085   1.0000
<< Back to DAE-21 AIRFOIL (dae21-il)

Polar data table (+)

Polar graphs


<< Back to DAE-21 AIRFOIL (dae21-il)