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=1,000,000 Ncrit=5


Details Polar file
Airfoil: DAE-21 AIRFOIL (dae21-il)
Reynolds number: 1,000,000
Max Cl/Cd: 158.11 at α=4.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-dae21-il-1000000-n5.txt
Download as CSV file: xf-dae21-il-1000000-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 =     1.000 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.0770   0.09562   0.09262  -0.0712   0.7023   0.0081
  -9.250  -0.0713   0.09235   0.08934  -0.0726   0.7001   0.0079
  -9.000  -0.0664   0.08898   0.08596  -0.0741   0.6980   0.0078
  -8.750  -0.0586   0.08640   0.08339  -0.0754   0.6962   0.0080
  -8.500  -0.0518   0.08359   0.08059  -0.0767   0.6943   0.0081
  -8.250  -0.0457   0.08070   0.07772  -0.0780   0.6920   0.0082
  -8.000  -0.0399   0.07786   0.07489  -0.0794   0.6895   0.0083
  -7.500  -0.1763   0.01960   0.01545  -0.1266   0.6843   0.0098
  -7.250  -0.1505   0.01822   0.01387  -0.1273   0.6820   0.0099
  -7.000  -0.1240   0.01712   0.01259  -0.1278   0.6799   0.0101
  -6.750  -0.0968   0.01621   0.01154  -0.1282   0.6779   0.0102
  -6.500  -0.0694   0.01538   0.01057  -0.1285   0.6756   0.0104
  -6.250  -0.0419   0.01459   0.00963  -0.1288   0.6731   0.0105
  -6.000  -0.0142   0.01383   0.00872  -0.1291   0.6706   0.0107
  -5.750   0.0137   0.01313   0.00787  -0.1293   0.6680   0.0109
  -5.500   0.0417   0.01253   0.00711  -0.1295   0.6655   0.0111
  -5.250   0.0699   0.01197   0.00642  -0.1297   0.6632   0.0114
  -5.000   0.0984   0.01148   0.00583  -0.1298   0.6608   0.0116
  -4.750   0.1269   0.01108   0.00534  -0.1300   0.6582   0.0118
  -4.500   0.1556   0.01080   0.00498  -0.1301   0.6555   0.0120
  -4.250   0.1840   0.01035   0.00447  -0.1303   0.6527   0.0125
  -4.000   0.2126   0.01012   0.00419  -0.1304   0.6499   0.0128
  -3.750   0.2414   0.00988   0.00391  -0.1305   0.6474   0.0131
  -3.500   0.2703   0.00963   0.00363  -0.1307   0.6447   0.0135
  -3.250   0.2991   0.00940   0.00335  -0.1308   0.6417   0.0139
  -3.000   0.3280   0.00919   0.00310  -0.1310   0.6386   0.0143
  -2.750   0.3568   0.00902   0.00288  -0.1311   0.6356   0.0146
  -2.500   0.3855   0.00882   0.00263  -0.1312   0.6325   0.0151
  -2.250   0.4145   0.00864   0.00245  -0.1314   0.6298   0.0157
  -2.000   0.4434   0.00850   0.00231  -0.1316   0.6265   0.0163
  -1.750   0.4723   0.00840   0.00218  -0.1317   0.6230   0.0171
  -1.500   0.5011   0.00832   0.00207  -0.1318   0.6196   0.0178
  -1.250   0.5299   0.00820   0.00193  -0.1320   0.6163   0.0189
  -1.000   0.5589   0.00810   0.00184  -0.1321   0.6130   0.0200
  -0.750   0.5878   0.00803   0.00176  -0.1323   0.6090   0.0211
  -0.500   0.6165   0.00796   0.00168  -0.1324   0.6051   0.0231
  -0.250   0.6451   0.00792   0.00162  -0.1325   0.6013   0.0260
   0.000   0.6741   0.00785   0.00158  -0.1327   0.5976   0.0307
   0.250   0.7028   0.00780   0.00155  -0.1329   0.5933   0.0395
   0.500   0.7314   0.00775   0.00154  -0.1330   0.5888   0.0574
   0.750   0.7598   0.00768   0.00154  -0.1332   0.5846   0.0883
   1.000   0.7885   0.00756   0.00156  -0.1334   0.5799   0.1420
   1.250   0.8167   0.00745   0.00159  -0.1336   0.5745   0.2141
   1.500   0.8447   0.00729   0.00166  -0.1338   0.5695   0.3230
   1.750   0.8727   0.00709   0.00176  -0.1340   0.5640   0.4510
   2.000   0.9004   0.00698   0.00184  -0.1341   0.5577   0.5458
   2.250   0.9238   0.00645   0.00200  -0.1333   0.5522   0.8251
   2.750   0.9832   0.00640   0.00210  -0.1341   0.5382   1.0000
   3.000   1.0113   0.00650   0.00217  -0.1342   0.5305   1.0000
   3.250   1.0388   0.00665   0.00225  -0.1342   0.5224   1.0000
   3.500   1.0665   0.00677   0.00234  -0.1342   0.5138   1.0000
   3.750   1.0937   0.00693   0.00245  -0.1342   0.5051   1.0000
   4.000   1.1209   0.00709   0.00256  -0.1342   0.4957   1.0000
   4.250   1.1479   0.00726   0.00269  -0.1342   0.4866   1.0000
   4.500   1.1744   0.00746   0.00283  -0.1340   0.4764   1.0000
   4.750   1.2011   0.00764   0.00298  -0.1340   0.4659   1.0000
   5.000   1.2272   0.00786   0.00315  -0.1338   0.4546   1.0000
   5.250   1.2528   0.00811   0.00334  -0.1336   0.4423   1.0000
   5.500   1.2779   0.00838   0.00356  -0.1333   0.4290   1.0000
   5.750   1.3028   0.00866   0.00377  -0.1330   0.4153   1.0000
   6.000   1.3273   0.00895   0.00401  -0.1326   0.4016   1.0000
   6.250   1.3511   0.00928   0.00427  -0.1322   0.3865   1.0000
   6.500   1.3739   0.00965   0.00457  -0.1316   0.3702   1.0000
   6.750   1.3957   0.01006   0.00490  -0.1308   0.3530   1.0000
   7.000   1.4166   0.01050   0.00525  -0.1299   0.3355   1.0000
   7.250   1.4365   0.01097   0.00564  -0.1289   0.3172   1.0000
   7.500   1.4555   0.01144   0.00604  -0.1277   0.3004   1.0000
   7.750   1.4727   0.01197   0.00648  -0.1263   0.2830   1.0000
   8.000   1.4887   0.01250   0.00694  -0.1246   0.2679   1.0000
   8.250   1.4996   0.01308   0.00746  -0.1220   0.2525   1.0000
   8.500   1.5067   0.01374   0.00808  -0.1190   0.2392   1.0000
   8.750   1.5133   0.01457   0.00884  -0.1162   0.2247   1.0000
   9.000   1.5207   0.01550   0.00971  -0.1137   0.2105   1.0000
   9.250   1.5289   0.01648   0.01064  -0.1115   0.1971   1.0000
   9.500   1.5374   0.01751   0.01162  -0.1095   0.1844   1.0000
   9.750   1.5462   0.01858   0.01264  -0.1076   0.1721   1.0000
  10.000   1.5538   0.01976   0.01377  -0.1057   0.1589   1.0000
  10.250   1.5620   0.02095   0.01492  -0.1040   0.1472   1.0000
  10.500   1.5699   0.02219   0.01613  -0.1022   0.1361   1.0000
  10.750   1.5773   0.02349   0.01739  -0.1005   0.1255   1.0000
  11.000   1.5843   0.02486   0.01872  -0.0989   0.1160   1.0000
  11.250   1.5909   0.02629   0.02012  -0.0973   0.1061   1.0000
  11.500   1.5984   0.02768   0.02149  -0.0958   0.0976   1.0000
  11.750   1.6054   0.02914   0.02294  -0.0943   0.0902   1.0000
  12.000   1.6111   0.03073   0.02449  -0.0929   0.0822   1.0000
  12.250   1.6182   0.03223   0.02601  -0.0916   0.0761   1.0000
  12.500   1.6254   0.03379   0.02756  -0.0903   0.0706   1.0000
  12.750   1.6311   0.03552   0.02929  -0.0891   0.0647   1.0000
  13.000   1.6390   0.03710   0.03088  -0.0881   0.0604   1.0000
  13.250   1.6439   0.03901   0.03278  -0.0870   0.0551   1.0000
  13.500   1.6520   0.04065   0.03446  -0.0861   0.0517   1.0000
  13.750   1.6577   0.04256   0.03638  -0.0852   0.0480   1.0000
  14.000   1.6644   0.04442   0.03826  -0.0843   0.0447   1.0000
  14.250   1.6701   0.04641   0.04028  -0.0836   0.0417   1.0000
  14.500   1.6755   0.04848   0.04238  -0.0828   0.0386   1.0000
  14.750   1.6814   0.05054   0.04447  -0.0822   0.0363   1.0000
  15.000   1.6856   0.05281   0.04677  -0.0816   0.0339   1.0000
  15.250   1.6913   0.05497   0.04897  -0.0811   0.0319   1.0000
  15.500   1.6950   0.05739   0.05143  -0.0807   0.0299   1.0000
  15.750   1.6997   0.05973   0.05381  -0.0803   0.0283   1.0000
  16.000   1.7037   0.06220   0.05635  -0.0800   0.0269   1.0000
  16.250   1.7068   0.06482   0.05901  -0.0797   0.0253   1.0000
  16.500   1.7094   0.06755   0.06179  -0.0796   0.0239   1.0000
  16.750   1.7130   0.07019   0.06449  -0.0795   0.0228   1.0000
  17.000   1.7139   0.07321   0.06757  -0.0795   0.0215   1.0000
  17.250   1.7154   0.07623   0.07064  -0.0796   0.0206   1.0000
  17.500   1.7179   0.07913   0.07362  -0.0798   0.0198   1.0000
  17.750   1.7187   0.08234   0.07690  -0.0801   0.0190   1.0000
  18.000   1.7191   0.08562   0.08025  -0.0804   0.0181   1.0000
  18.250   1.7177   0.08921   0.08390  -0.0809   0.0173   1.0000
  18.500   1.7179   0.09262   0.08739  -0.0815   0.0168   1.0000
  18.750   1.7180   0.09605   0.09090  -0.0821   0.0161   1.0000
  19.000   1.7163   0.09985   0.09477  -0.0829   0.0155   1.0000
  19.250   1.7138   0.10377   0.09877  -0.0838   0.0150   1.0000
<< Back to DAE-21 AIRFOIL (dae21-il)

Polar data table (+)

Polar graphs


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