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

EPPLER 342 AIRFOIL (e342-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5


Details Polar file
Airfoil: EPPLER 342 AIRFOIL (e342-il)
Reynolds number: 1,000,000
Max Cl/Cd: 106.32 at α=9.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-e342-il-1000000-n5.txt
Download as CSV file: xf-e342-il-1000000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 342 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  -0.2902   0.06738   0.06440  -0.0424   0.5543   0.0063
  -9.250  -0.3047   0.06101   0.05803  -0.0468   0.5506   0.0063
  -9.000  -0.3343   0.05402   0.05100  -0.0515   0.5492   0.0062
  -8.750  -0.3553   0.04962   0.04653  -0.0516   0.5458   0.0062
  -7.750  -0.5458   0.02510   0.02010  -0.0249   0.5537   0.0043
  -7.500  -0.5366   0.02206   0.01661  -0.0218   0.5462   0.0043
  -7.250  -0.5201   0.02013   0.01437  -0.0198   0.5378   0.0043
  -7.000  -0.5010   0.01851   0.01243  -0.0182   0.5296   0.0043
  -6.750  -0.4788   0.01753   0.01127  -0.0172   0.5213   0.0043
  -6.500  -0.4560   0.01657   0.01011  -0.0162   0.5129   0.0043
  -6.250  -0.4323   0.01575   0.00914  -0.0154   0.5042   0.0043
  -6.000  -0.4083   0.01507   0.00830  -0.0146   0.4954   0.0043
  -5.750  -0.3837   0.01451   0.00764  -0.0139   0.4874   0.0043
  -5.250  -0.3349   0.01345   0.00637  -0.0124   0.4734   0.0043
  -5.000  -0.3104   0.01303   0.00586  -0.0117   0.4663   0.0043
  -4.750  -0.2857   0.01268   0.00544  -0.0110   0.4600   0.0043
  -4.500  -0.2613   0.01229   0.00499  -0.0102   0.4536   0.0044
  -4.250  -0.2368   0.01200   0.00462  -0.0095   0.4463   0.0044
  -4.000  -0.2121   0.01170   0.00427  -0.0087   0.4401   0.0045
  -3.750  -0.1873   0.01146   0.00398  -0.0081   0.4340   0.0045
  -3.500  -0.1624   0.01126   0.00374  -0.0074   0.4285   0.0045
  -3.250  -0.1381   0.01094   0.00337  -0.0066   0.4239   0.0047
  -3.000  -0.1130   0.01076   0.00315  -0.0059   0.4186   0.0049
  -2.750  -0.0879   0.01061   0.00296  -0.0053   0.4133   0.0053
  -2.500  -0.0623   0.01047   0.00280  -0.0048   0.4088   0.0061
  -2.250  -0.0367   0.01035   0.00265  -0.0042   0.4036   0.0066
  -2.000  -0.0112   0.01024   0.00251  -0.0037   0.3987   0.0075
  -1.750   0.0145   0.01015   0.00240  -0.0032   0.3948   0.0089
  -1.500   0.0400   0.01002   0.00230  -0.0027   0.3914   0.0146
  -1.250   0.0657   0.00992   0.00222  -0.0022   0.3875   0.0219
  -1.000   0.0914   0.00984   0.00215  -0.0017   0.3834   0.0303
  -0.750   0.1158   0.00968   0.00209  -0.0010   0.3792   0.0607
  -0.500   0.1398   0.00946   0.00203  -0.0003   0.3763   0.1114
  -0.250   0.1510   0.00848   0.00189   0.0026   0.3731   0.3760
   0.000   0.1348   0.00669   0.00182   0.0117   0.3713   0.8392
   0.250   0.1603   0.00687   0.00199   0.0125   0.3679   0.8724
   0.500   0.1869   0.00703   0.00212   0.0129   0.3644   0.8827
   0.750   0.2129   0.00721   0.00226   0.0136   0.3618   0.8933
   1.000   0.2403   0.00739   0.00243   0.0140   0.3590   0.8988
   1.250   0.2669   0.00755   0.00256   0.0144   0.3555   0.9044
   1.500   0.2934   0.00766   0.00262   0.0147   0.3524   0.9079
   1.750   0.3211   0.00770   0.00261   0.0147   0.3495   0.9087
   2.000   0.3488   0.00775   0.00262   0.0147   0.3467   0.9094
   2.250   0.3769   0.00777   0.00263   0.0146   0.3441   0.9100
   2.500   0.4049   0.00781   0.00264   0.0146   0.3414   0.9105
   2.750   0.4327   0.00786   0.00266   0.0145   0.3387   0.9111
   3.000   0.4602   0.00793   0.00270   0.0145   0.3357   0.9118
   3.250   0.4875   0.00801   0.00275   0.0145   0.3327   0.9125
   3.500   0.5153   0.00807   0.00279   0.0144   0.3304   0.9130
   3.750   0.5431   0.00811   0.00283   0.0143   0.3282   0.9136
   4.000   0.5708   0.00817   0.00288   0.0142   0.3256   0.9142
   4.250   0.5983   0.00824   0.00293   0.0142   0.3227   0.9147
   4.500   0.6256   0.00832   0.00300   0.0141   0.3198   0.9153
   4.750   0.6526   0.00842   0.00307   0.0141   0.3165   0.9160
   5.000   0.6802   0.00848   0.00314   0.0140   0.3145   0.9166
   5.250   0.7077   0.00855   0.00321   0.0139   0.3122   0.9172
   5.500   0.7350   0.00862   0.00329   0.0138   0.3092   0.9179
   5.750   0.7620   0.00872   0.00338   0.0138   0.3061   0.9187
   6.000   0.7887   0.00883   0.00348   0.0138   0.3027   0.9194
   6.250   0.8157   0.00893   0.00358   0.0138   0.2998   0.9200
   6.500   0.8430   0.00900   0.00368   0.0136   0.2970   0.9206
   6.750   0.8699   0.00910   0.00378   0.0136   0.2936   0.9211
   7.000   0.8965   0.00922   0.00390   0.0136   0.2898   0.9217
   7.250   0.9225   0.00936   0.00404   0.0136   0.2857   0.9223
   7.500   0.9495   0.00946   0.00416   0.0135   0.2825   0.9228
   7.750   0.9759   0.00958   0.00429   0.0135   0.2779   0.9232
   8.000   1.0015   0.00973   0.00444   0.0136   0.2732   0.9238
   8.250   1.0271   0.00987   0.00459   0.0137   0.2689   0.9245
   8.500   1.0524   0.01001   0.00474   0.0139   0.2631   0.9252
   8.750   1.0765   0.01021   0.00494   0.0142   0.2573   0.9259
   9.000   1.1014   0.01036   0.00512   0.0144   0.2520   0.9266
   9.250   1.1247   0.01059   0.00533   0.0149   0.2442   0.9274
   9.500   1.1482   0.01080   0.00556   0.0153   0.2369   0.9284
   9.750   1.1698   0.01108   0.00583   0.0160   0.2285   0.9295
  10.000   1.1918   0.01135   0.00609   0.0167   0.2190   0.9305
  10.250   1.2125   0.01167   0.00640   0.0175   0.2099   0.9316
  10.500   1.2317   0.01204   0.00675   0.0186   0.1999   0.9327
  10.750   1.2506   0.01240   0.00710   0.0196   0.1900   0.9340
  11.000   1.2678   0.01280   0.00749   0.0210   0.1799   0.9353
  11.250   1.2814   0.01322   0.00789   0.0230   0.1704   0.9368
  11.500   1.2920   0.01376   0.00839   0.0254   0.1591   0.9384
  11.750   1.3020   0.01428   0.00890   0.0279   0.1495   0.9405
  12.000   1.3104   0.01485   0.00948   0.0306   0.1398   0.9429
  12.250   1.3161   0.01556   0.01017   0.0334   0.1297   0.9458
  12.500   1.3210   0.01635   0.01096   0.0361   0.1211   0.9491
  12.750   1.3239   0.01729   0.01189   0.0387   0.1119   0.9528
  13.000   1.3244   0.01836   0.01299   0.0414   0.1036   0.9579
  13.250   1.3253   0.01965   0.01431   0.0435   0.0970   0.9639
  13.500   1.3293   0.02124   0.01594   0.0443   0.0895   0.9704
  13.750   1.3377   0.02321   0.01794   0.0434   0.0815   0.9759
  14.000   1.3460   0.02566   0.02041   0.0416   0.0735   0.9838
  14.250   1.3502   0.02837   0.02315   0.0401   0.0674   0.9977
  14.500   1.3489   0.03102   0.02584   0.0397   0.0624   1.0000
  14.750   1.3431   0.03381   0.02868   0.0400   0.0595   1.0000
  15.000   1.3380   0.03664   0.03157   0.0402   0.0566   1.0000
  15.250   1.3293   0.03987   0.03485   0.0402   0.0533   1.0000
  15.500   1.3188   0.04338   0.03842   0.0400   0.0506   1.0000
  15.750   1.3132   0.04644   0.04154   0.0398   0.0492   1.0000
  16.000   1.3034   0.05006   0.04522   0.0394   0.0463   1.0000
  16.250   1.2937   0.05378   0.04900   0.0388   0.0444   1.0000
  16.500   1.2860   0.05739   0.05268   0.0381   0.0424   1.0000
  16.750   1.2772   0.06120   0.05654   0.0373   0.0399   1.0000
  17.000   1.2660   0.06539   0.06077   0.0363   0.0374   1.0000
  17.250   1.2608   0.06896   0.06441   0.0353   0.0363   1.0000
  17.500   1.2536   0.07282   0.06833   0.0342   0.0343   1.0000
  17.750   1.2441   0.07702   0.07257   0.0329   0.0318   1.0000
  18.000   1.2394   0.08072   0.07634   0.0317   0.0304   1.0000
  18.250   1.2325   0.08476   0.08043   0.0303   0.0286   1.0000
<< Back to EPPLER 342 AIRFOIL (e342-il)

Polar data table (+)

Polar graphs


<< Back to EPPLER 342 AIRFOIL (e342-il)