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=9


Details Polar file
Airfoil: EPPLER 342 AIRFOIL (e342-il)
Reynolds number: 1,000,000
Max Cl/Cd: 108.28 at α=10.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e342-il-1000000.txt
Download as CSV file: xf-e342-il-1000000.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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.750  -0.3130   0.08186   0.07955  -0.0481   0.7689   0.0110
  -9.500  -0.3208   0.07714   0.07473  -0.0507   0.7385   0.0110
  -9.250  -0.3301   0.07281   0.07032  -0.0541   0.7143   0.0109
  -9.000  -0.3588   0.06546   0.06286  -0.0572   0.7017   0.0111
  -8.750  -0.3644   0.06414   0.06146  -0.0557   0.6820   0.0109
  -8.500  -0.3868   0.05954   0.05672  -0.0531   0.6698   0.0111
  -8.250  -0.4001   0.05643   0.05349  -0.0497   0.6570   0.0111
  -7.500  -0.4123   0.04406   0.04050  -0.0405   0.6250   0.0113
  -7.250  -0.4239   0.03964   0.03585  -0.0363   0.6162   0.0115
  -7.000  -0.4162   0.03798   0.03407  -0.0341   0.6044   0.0116
  -6.750  -0.4062   0.03612   0.03205  -0.0320   0.5930   0.0117
  -6.500  -0.3945   0.03441   0.03017  -0.0298   0.5822   0.0118
  -6.250  -0.3815   0.03264   0.02825  -0.0278   0.5718   0.0120
  -6.000  -0.3677   0.03072   0.02613  -0.0256   0.5625   0.0123
  -5.750  -0.3529   0.02864   0.02384  -0.0233   0.5534   0.0127
  -4.750  -0.2796   0.01595   0.00942  -0.0130   0.5221   0.0086
  -4.500  -0.2534   0.01435   0.00761  -0.0125   0.5132   0.0082
  -4.250  -0.2274   0.01343   0.00657  -0.0119   0.5047   0.0079
  -4.000  -0.2028   0.01269   0.00570  -0.0110   0.4959   0.0078
  -3.750  -0.1783   0.01218   0.00512  -0.0102   0.4886   0.0077
  -3.500  -0.1544   0.01177   0.00463  -0.0093   0.4807   0.0077
  -3.250  -0.1306   0.01136   0.00417  -0.0084   0.4740   0.0078
  -3.000  -0.1064   0.01107   0.00382  -0.0075   0.4668   0.0079
  -2.750  -0.0822   0.01080   0.00349  -0.0067   0.4601   0.0082
  -2.500  -0.0574   0.01057   0.00321  -0.0059   0.4535   0.0087
  -2.250  -0.0336   0.01032   0.00291  -0.0050   0.4466   0.0094
  -2.000  -0.0083   0.01016   0.00272  -0.0044   0.4412   0.0103
  -1.750   0.0166   0.00999   0.00253  -0.0036   0.4356   0.0133
  -1.500   0.0400   0.00975   0.00236  -0.0027   0.4298   0.0362
  -1.250   0.0648   0.00957   0.00228  -0.0020   0.4253   0.0621
  -1.000   0.0891   0.00938   0.00220  -0.0013   0.4204   0.1003
  -0.750   0.1065   0.00879   0.00209   0.0005   0.4155   0.2549
  -0.500   0.0971   0.00707   0.00180   0.0075   0.4120   0.6677
  -0.250   0.1056   0.00669   0.00214   0.0123   0.4090   0.8722
   0.000   0.1299   0.00708   0.00253   0.0137   0.4049   0.8945
   0.250   0.1556   0.00752   0.00294   0.0148   0.4006   0.9063
   0.500   0.1829   0.00785   0.00322   0.0154   0.3957   0.9121
   0.750   0.2132   0.00831   0.00366   0.0156   0.3927   0.9198
   1.000   0.2401   0.00852   0.00384   0.0162   0.3891   0.9252
   1.250   0.2823   0.00895   0.00421   0.0137   0.3846   0.9281
   1.750   0.3448   0.00918   0.00434   0.0123   0.3775   0.9307
   2.000   0.3708   0.00918   0.00432   0.0127   0.3746   0.9321
   2.250   0.3933   0.00917   0.00429   0.0137   0.3713   0.9342
   2.500   0.4080   0.00913   0.00421   0.0164   0.3685   0.9374
   2.750   0.4280   0.00911   0.00415   0.0178   0.3651   0.9391
   3.000   0.4550   0.00911   0.00413   0.0179   0.3622   0.9398
   3.250   0.4829   0.00910   0.00412   0.0179   0.3598   0.9405
   3.500   0.5101   0.00911   0.00411   0.0179   0.3571   0.9411
   3.750   0.5365   0.00913   0.00411   0.0181   0.3542   0.9418
   4.000   0.5629   0.00919   0.00415   0.0182   0.3511   0.9426
   4.250   0.5878   0.00929   0.00421   0.0187   0.3473   0.9435
   4.500   0.6139   0.00928   0.00421   0.0189   0.3455   0.9443
   4.750   0.6396   0.00928   0.00422   0.0192   0.3431   0.9451
   5.000   0.6649   0.00929   0.00424   0.0195   0.3404   0.9460
   5.250   0.6897   0.00933   0.00427   0.0199   0.3377   0.9471
   5.500   0.7136   0.00940   0.00432   0.0205   0.3345   0.9483
   5.750   0.7372   0.00948   0.00438   0.0212   0.3312   0.9493
   6.000   0.7627   0.00948   0.00441   0.0214   0.3292   0.9502
   6.250   0.7884   0.00950   0.00445   0.0216   0.3266   0.9510
   6.500   0.8140   0.00954   0.00450   0.0218   0.3236   0.9517
   6.750   0.8393   0.00961   0.00457   0.0220   0.3205   0.9523
   7.000   0.8634   0.00972   0.00466   0.0224   0.3166   0.9531
   7.250   0.8897   0.00974   0.00472   0.0225   0.3141   0.9538
   7.500   0.9158   0.00978   0.00480   0.0226   0.3113   0.9546
   7.750   0.9414   0.00984   0.00488   0.0228   0.3077   0.9553
   8.000   0.9661   0.00995   0.00499   0.0231   0.3042   0.9560
   8.250   0.9903   0.01008   0.00512   0.0234   0.3002   0.9569
   8.500   1.0165   0.01012   0.00522   0.0235   0.2969   0.9576
   8.750   1.0416   0.01020   0.00532   0.0237   0.2930   0.9585
   9.000   1.0653   0.01034   0.00546   0.0241   0.2884   0.9595
   9.250   1.0894   0.01046   0.00561   0.0244   0.2842   0.9606
   9.500   1.1145   0.01054   0.00573   0.0246   0.2796   0.9615
   9.750   1.1378   0.01069   0.00589   0.0250   0.2744   0.9626
  10.000   1.1613   0.01084   0.00606   0.0254   0.2690   0.9636
  10.250   1.1850   0.01098   0.00623   0.0257   0.2629   0.9645
  10.500   1.2056   0.01119   0.00643   0.0266   0.2565   0.9660
  10.750   1.2279   0.01134   0.00663   0.0272   0.2504   0.9674
  11.000   1.2464   0.01162   0.00689   0.0284   0.2426   0.9692
  11.250   1.2664   0.01180   0.00712   0.0294   0.2354   0.9709
  11.500   1.2821   0.01211   0.00742   0.0311   0.2264   0.9732
  11.750   1.2977   0.01245   0.00776   0.0327   0.2171   0.9756
  12.000   1.3159   0.01288   0.00819   0.0337   0.2066   0.9777
  12.250   1.3326   0.01343   0.00872   0.0347   0.1946   0.9801
  12.500   1.3489   0.01403   0.00931   0.0356   0.1834   0.9829
  12.750   1.3641   0.01477   0.01001   0.0364   0.1710   0.9858
  13.000   1.3809   0.01573   0.01093   0.0364   0.1558   0.9882
  13.250   1.3965   0.01677   0.01196   0.0363   0.1435   0.9912
  13.500   1.4092   0.01804   0.01320   0.0361   0.1314   0.9945
  13.750   1.4232   0.01965   0.01479   0.0350   0.1189   0.9965
  14.000   1.4354   0.02151   0.01665   0.0336   0.1088   0.9992
  14.250   1.4268   0.02307   0.01825   0.0364   0.1035   1.0000
  14.500   1.4176   0.02524   0.02043   0.0383   0.0978   1.0000
  14.750   1.4144   0.02746   0.02269   0.0392   0.0926   1.0000
  15.000   1.4056   0.03036   0.02561   0.0397   0.0867   1.0000
  15.250   1.3981   0.03331   0.02860   0.0400   0.0815   1.0000
  15.500   1.3869   0.03670   0.03203   0.0402   0.0774   1.0000
  15.750   1.3807   0.03968   0.03508   0.0402   0.0742   1.0000
  16.000   1.3677   0.04341   0.03886   0.0400   0.0703   1.0000
  16.250   1.3552   0.04719   0.04269   0.0397   0.0672   1.0000
  16.500   1.3447   0.05091   0.04647   0.0392   0.0636   1.0000
  16.750   1.3327   0.05490   0.05052   0.0385   0.0612   1.0000
  17.000   1.3204   0.05906   0.05473   0.0377   0.0580   1.0000
  17.250   1.3119   0.06284   0.05858   0.0368   0.0555   1.0000
  17.500   1.3001   0.06712   0.06290   0.0357   0.0527   1.0000
  17.750   1.2903   0.07125   0.06707   0.0346   0.0496   1.0000
  18.000   1.2825   0.07521   0.07110   0.0334   0.0475   1.0000
  18.250   1.2693   0.07995   0.07585   0.0319   0.0443   1.0000
  18.500   1.2632   0.08383   0.07979   0.0306   0.0415   1.0000
<< Back to EPPLER 342 AIRFOIL (e342-il)

Polar data table (+)

Polar graphs


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