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

EPPLER 562 AIRFOIL (e562-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: EPPLER 562 AIRFOIL (e562-il)
Reynolds number: 500,000
Max Cl/Cd: 104.91 at α=6.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e562-il-500000.txt
Download as CSV file: xf-e562-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 562 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
  ------ -------- --------- --------- -------- -------- --------
 -15.000  -0.7025   0.08821   0.08500  -0.0594   1.0000   0.0139
 -14.750  -0.7282   0.08021   0.07682  -0.0638   1.0000   0.0137
 -14.500  -0.7325   0.07625   0.07283  -0.0658   1.0000   0.0140
 -14.250  -0.7450   0.07113   0.06761  -0.0685   1.0000   0.0142
 -14.000  -0.7604   0.06598   0.06232  -0.0709   1.0000   0.0141
 -13.750  -0.7718   0.06172   0.05797  -0.0727   1.0000   0.0141
 -13.500  -0.7805   0.05799   0.05416  -0.0742   1.0000   0.0143
 -13.250  -0.7909   0.05425   0.05032  -0.0754   1.0000   0.0143
 -13.000  -0.7971   0.05118   0.04720  -0.0763   1.0000   0.0146
 -12.750  -0.8078   0.04783   0.04373  -0.0768   1.0000   0.0145
 -12.500  -0.8152   0.04505   0.04089  -0.0770   1.0000   0.0147
 -12.250  -0.8256   0.04226   0.03801  -0.0767   1.0000   0.0148
 -12.000  -0.8366   0.03978   0.03547  -0.0761   1.0000   0.0151
 -11.750  -0.8531   0.03750   0.03312  -0.0741   1.0000   0.0150
 -11.500  -0.8774   0.03537   0.03092  -0.0713   1.0000   0.0151
 -11.250  -0.8657   0.03268   0.02807  -0.0752   0.9982   0.0154
 -11.000  -0.8347   0.03040   0.02560  -0.0804   0.9948   0.0158
 -10.750  -0.8042   0.02868   0.02370  -0.0841   0.9912   0.0162
 -10.500  -0.7774   0.02634   0.02124  -0.0872   0.9873   0.0167
 -10.250  -0.7445   0.02473   0.01959  -0.0907   0.9848   0.0174
 -10.000  -0.7134   0.02351   0.01832  -0.0932   0.9814   0.0181
  -9.750  -0.6827   0.02234   0.01706  -0.0953   0.9772   0.0189
  -9.500  -0.6476   0.02137   0.01598  -0.0979   0.9748   0.0198
  -9.250  -0.6125   0.01978   0.01433  -0.1013   0.9729   0.0210
  -9.000  -0.5851   0.01887   0.01339  -0.1022   0.9674   0.0222
  -8.750  -0.5506   0.01806   0.01251  -0.1043   0.9642   0.0237
  -8.500  -0.5145   0.01691   0.01128  -0.1071   0.9620   0.0253
  -8.250  -0.4765   0.01603   0.01036  -0.1099   0.9604   0.0274
  -8.000  -0.4509   0.01539   0.00965  -0.1098   0.9531   0.0291
  -7.750  -0.4147   0.01446   0.00870  -0.1122   0.9503   0.0326
  -7.500  -0.3769   0.01372   0.00792  -0.1146   0.9484   0.0372
  -7.250  -0.3486   0.01302   0.00722  -0.1150   0.9417   0.0440
  -7.000  -0.3137   0.01207   0.00640  -0.1170   0.9380   0.0670
  -6.750  -0.2770   0.01120   0.00571  -0.1193   0.9354   0.1043
  -6.500  -0.2466   0.01060   0.00521  -0.1200   0.9291   0.1328
  -6.250  -0.2124   0.01006   0.00477  -0.1215   0.9238   0.1637
  -6.000  -0.1727   0.00959   0.00438  -0.1240   0.9204   0.1908
  -5.750  -0.1390   0.00929   0.00410  -0.1252   0.9116   0.2120
  -5.500  -0.0952   0.00898   0.00380  -0.1285   0.9063   0.2317
  -5.250  -0.0538   0.00876   0.00359  -0.1313   0.8973   0.2501
  -4.750   0.0323   0.00846   0.00322  -0.1375   0.8750   0.2835
  -4.500   0.0689   0.00839   0.00307  -0.1393   0.8595   0.2955
  -4.250   0.1024   0.00839   0.00296  -0.1403   0.8429   0.3073
  -4.000   0.1330   0.00835   0.00286  -0.1407   0.8259   0.3169
  -3.750   0.1621   0.00836   0.00276  -0.1408   0.8090   0.3260
  -3.500   0.1899   0.00837   0.00270  -0.1407   0.7921   0.3349
  -3.250   0.2173   0.00839   0.00263  -0.1404   0.7759   0.3430
  -2.750   0.2711   0.00845   0.00256  -0.1397   0.7444   0.3617
  -2.500   0.2979   0.00852   0.00253  -0.1394   0.7294   0.3713
  -2.250   0.3246   0.00853   0.00251  -0.1390   0.7147   0.3791
  -2.000   0.3514   0.00859   0.00249  -0.1386   0.7005   0.3872
  -1.750   0.3780   0.00861   0.00248  -0.1383   0.6868   0.3952
  -1.500   0.4048   0.00869   0.00247  -0.1379   0.6735   0.4036
  -1.250   0.4314   0.00873   0.00248  -0.1375   0.6605   0.4110
  -1.000   0.4581   0.00880   0.00249  -0.1372   0.6476   0.4193
  -0.750   0.4848   0.00884   0.00250  -0.1368   0.6352   0.4272
  -0.500   0.5116   0.00892   0.00253  -0.1365   0.6231   0.4357
  -0.250   0.5381   0.00899   0.00256  -0.1361   0.6115   0.4433
   0.000   0.5649   0.00906   0.00260  -0.1358   0.5999   0.4519
   0.250   0.5917   0.00912   0.00264  -0.1355   0.5886   0.4599
   0.500   0.6182   0.00922   0.00269  -0.1352   0.5778   0.4685
   0.750   0.6448   0.00929   0.00275  -0.1348   0.5670   0.4769
   1.000   0.6716   0.00937   0.00281  -0.1345   0.5566   0.4858
   1.250   0.6980   0.00947   0.00288  -0.1341   0.5465   0.4942
   1.500   0.7247   0.00955   0.00296  -0.1338   0.5364   0.5038
   1.750   0.7512   0.00963   0.00304  -0.1335   0.5269   0.5127
   2.000   0.7775   0.00975   0.00313  -0.1331   0.5173   0.5227
   2.250   0.8042   0.00982   0.00324  -0.1328   0.5080   0.5328
   2.500   0.8301   0.00995   0.00334  -0.1324   0.4992   0.5433
   2.750   0.8568   0.01003   0.00345  -0.1321   0.4902   0.5550
   3.000   0.8828   0.01015   0.00358  -0.1317   0.4820   0.5671
   3.250   0.9092   0.01023   0.00371  -0.1314   0.4732   0.5802
   3.500   0.9350   0.01036   0.00385  -0.1310   0.4654   0.5946
   3.750   0.9612   0.01045   0.00400  -0.1306   0.4571   0.6112
   4.000   0.9867   0.01059   0.00416  -0.1301   0.4495   0.6297
   4.250   1.0128   0.01067   0.00433  -0.1297   0.4416   0.6515
   4.500   1.0378   0.01082   0.00452  -0.1292   0.4341   0.6764
   4.750   1.0634   0.01088   0.00470  -0.1287   0.4263   0.7066
   5.000   1.0873   0.01103   0.00491  -0.1279   0.4190   0.7434
   5.250   1.1114   0.01104   0.00511  -0.1270   0.4116   0.7906
   5.500   1.1305   0.01109   0.00531  -0.1251   0.4045   0.8574
   5.750   1.1492   0.01101   0.00539  -0.1229   0.3975   1.0000
   6.000   1.1739   0.01124   0.00558  -0.1224   0.3896   1.0000
   6.250   1.1990   0.01144   0.00579  -0.1219   0.3819   1.0000
   6.500   1.2230   0.01168   0.00600  -0.1212   0.3739   1.0000
   6.750   1.2474   0.01189   0.00622  -0.1206   0.3661   1.0000
   7.000   1.2707   0.01214   0.00646  -0.1198   0.3579   1.0000
   7.250   1.2943   0.01237   0.00670  -0.1190   0.3499   1.0000
   7.500   1.3157   0.01265   0.00694  -0.1178   0.3413   1.0000
   7.750   1.3381   0.01287   0.00720  -0.1168   0.3326   1.0000
   8.000   1.3579   0.01320   0.00749  -0.1154   0.3240   1.0000
   8.250   1.3799   0.01343   0.00776  -0.1143   0.3149   1.0000
   8.500   1.3995   0.01376   0.00807  -0.1129   0.3059   1.0000
   8.750   1.4194   0.01409   0.00840  -0.1115   0.2964   1.0000
   9.000   1.4393   0.01442   0.00875  -0.1102   0.2867   1.0000
   9.250   1.4569   0.01483   0.00914  -0.1085   0.2763   1.0000
   9.500   1.4747   0.01525   0.00954  -0.1068   0.2646   1.0000
   9.750   1.4919   0.01569   0.00998  -0.1051   0.2516   1.0000
  10.000   1.5079   0.01620   0.01046  -0.1033   0.2382   1.0000
  10.250   1.5226   0.01678   0.01100  -0.1012   0.2247   1.0000
  10.500   1.5363   0.01741   0.01160  -0.0991   0.2110   1.0000
  10.750   1.5486   0.01812   0.01227  -0.0969   0.1972   1.0000
  11.000   1.5596   0.01891   0.01302  -0.0945   0.1828   1.0000
  11.250   1.5689   0.01980   0.01387  -0.0920   0.1679   1.0000
  11.500   1.5763   0.02082   0.01483  -0.0894   0.1526   1.0000
  11.750   1.5825   0.02195   0.01591  -0.0867   0.1384   1.0000
  12.000   1.5872   0.02321   0.01713  -0.0841   0.1248   1.0000
  12.250   1.5910   0.02459   0.01847  -0.0815   0.1126   1.0000
  12.500   1.5937   0.02610   0.01996  -0.0790   0.1015   1.0000
  12.750   1.5954   0.02777   0.02162  -0.0767   0.0920   1.0000
  13.000   1.5967   0.02957   0.02341  -0.0745   0.0829   1.0000
  13.250   1.5986   0.03143   0.02529  -0.0727   0.0751   1.0000
  13.500   1.5982   0.03358   0.02745  -0.0709   0.0683   1.0000
  13.750   1.5976   0.03587   0.02977  -0.0695   0.0620   1.0000
  14.000   1.5971   0.03828   0.03222  -0.0682   0.0566   1.0000
  14.250   1.5936   0.04112   0.03508  -0.0672   0.0514   1.0000
  14.500   1.5923   0.04387   0.03790  -0.0664   0.0471   1.0000
  14.750   1.5875   0.04713   0.04120  -0.0659   0.0432   1.0000
  15.000   1.5839   0.05040   0.04454  -0.0655   0.0397   1.0000
  15.250   1.5784   0.05402   0.04823  -0.0655   0.0365   1.0000
  15.500   1.5709   0.05806   0.05233  -0.0657   0.0339   1.0000
  15.750   1.5665   0.06182   0.05620  -0.0660   0.0315   1.0000
  16.000   1.5580   0.06628   0.06071  -0.0667   0.0295   1.0000
  16.250   1.5490   0.07095   0.06548  -0.0676   0.0277   1.0000
  16.500   1.5431   0.07531   0.06995  -0.0685   0.0261   1.0000
  16.750   1.5343   0.08021   0.07493  -0.0698   0.0246   1.0000
  17.000   1.5215   0.08586   0.08066  -0.0714   0.0234   1.0000
  17.250   1.5146   0.09071   0.08562  -0.0729   0.0223   1.0000
  17.500   1.5081   0.09556   0.09058  -0.0745   0.0212   1.0000
  17.750   1.4992   0.10089   0.09599  -0.0765   0.0202   1.0000
  18.000   1.4877   0.10676   0.10195  -0.0787   0.0194   1.0000
<< Back to EPPLER 562 AIRFOIL (e562-il)

Polar data table (+)

Polar graphs


<< Back to EPPLER 562 AIRFOIL (e562-il)