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

EPPLER 340 AIRFOIL (e340-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: EPPLER 340 AIRFOIL (e340-il)
Reynolds number: 200,000
Max Cl/Cd: 57.05 at α=10.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-e340-il-200000-n5.txt
Download as CSV file: xf-e340-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 340 AIRFOIL                              
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.750  -0.5350   0.09256   0.08949  -0.0023   1.0000   0.0264
  -9.250  -0.5909   0.07206   0.06871  -0.0136   1.0000   0.0176
  -8.750  -0.6233   0.06198   0.05835  -0.0097   1.0000   0.0141
  -8.500  -0.6272   0.05939   0.05568  -0.0076   1.0000   0.0140
  -8.250  -0.6244   0.05623   0.05232  -0.0066   0.8920   0.0138
  -8.000  -0.6262   0.05333   0.04907  -0.0034   0.8252   0.0136
  -7.750  -0.6257   0.05013   0.04554  -0.0004   0.7894   0.0133
  -7.500  -0.6235   0.04655   0.04159   0.0027   0.7632   0.0129
  -7.250  -0.6198   0.04249   0.03709   0.0060   0.7424   0.0125
  -6.750  -0.6073   0.03171   0.02470   0.0143   0.7104   0.0112
  -6.500  -0.5894   0.02952   0.02209   0.0161   0.6927   0.0112
  -6.250  -0.5688   0.02751   0.01966   0.0176   0.6759   0.0112
  -6.000  -0.5457   0.02574   0.01750   0.0187   0.6599   0.0112
  -5.750  -0.5207   0.02419   0.01565   0.0194   0.6445   0.0113
  -5.500  -0.4950   0.02287   0.01413   0.0199   0.6299   0.0115
  -5.250  -0.4688   0.02178   0.01286   0.0203   0.6159   0.0117
  -5.000  -0.4424   0.02081   0.01173   0.0207   0.6027   0.0121
  -4.500  -0.3906   0.01913   0.00977   0.0217   0.5788   0.0130
  -4.250  -0.3658   0.01840   0.00890   0.0224   0.5677   0.0137
  -4.000  -0.3419   0.01776   0.00815   0.0232   0.5570   0.0145
  -3.750  -0.3176   0.01725   0.00760   0.0239   0.5462   0.0159
  -3.500  -0.2935   0.01682   0.00707   0.0247   0.5365   0.0183
  -3.250  -0.2694   0.01642   0.00659   0.0255   0.5270   0.0213
  -3.000  -0.2455   0.01598   0.00608   0.0263   0.5187   0.0253
  -2.750  -0.2218   0.01556   0.00567   0.0272   0.5100   0.0349
  -2.500  -0.1992   0.01508   0.00527   0.0282   0.5021   0.0609
  -2.250  -0.1782   0.01451   0.00496   0.0293   0.4938   0.1254
  -2.000  -0.1220   0.01335   0.00682   0.0259   0.4842   0.8255
  -1.750  -0.1078   0.01382   0.00716   0.0294   0.4775   0.8622
  -1.500  -0.0892   0.01439   0.00757   0.0321   0.4716   0.8838
  -1.250  -0.0506   0.01515   0.00819   0.0310   0.4636   0.8992
  -1.000  -0.0104   0.01588   0.00871   0.0295   0.4564   0.9135
  -0.750   0.0390   0.01618   0.00886   0.0257   0.4481   0.9188
  -0.500   0.0693   0.01620   0.00873   0.0252   0.4421   0.9222
  -0.250   0.0913   0.01622   0.00864   0.0262   0.4372   0.9269
   0.000   0.1208   0.01618   0.00851   0.0258   0.4313   0.9292
   0.250   0.1532   0.01616   0.00837   0.0247   0.4256   0.9308
   0.500   0.1844   0.01615   0.00825   0.0239   0.4204   0.9326
   0.750   0.2142   0.01614   0.00817   0.0234   0.4149   0.9348
   1.000   0.2416   0.01615   0.00810   0.0233   0.4103   0.9373
   1.250   0.2643   0.01621   0.00805   0.0241   0.4063   0.9406
   1.500   0.2919   0.01619   0.00802   0.0240   0.4013   0.9424
   1.750   0.3233   0.01617   0.00795   0.0231   0.3964   0.9436
   2.000   0.3537   0.01619   0.00789   0.0224   0.3923   0.9450
   2.250   0.3828   0.01622   0.00787   0.0219   0.3886   0.9464
   2.500   0.4111   0.01624   0.00789   0.0216   0.3843   0.9482
   2.750   0.4377   0.01628   0.00791   0.0216   0.3801   0.9502
   3.000   0.4618   0.01634   0.00792   0.0221   0.3763   0.9521
   3.250   0.4864   0.01642   0.00797   0.0225   0.3728   0.9540
   3.500   0.5169   0.01644   0.00802   0.0217   0.3684   0.9550
   3.750   0.5460   0.01647   0.00806   0.0212   0.3643   0.9559
   4.000   0.5745   0.01654   0.00811   0.0208   0.3607   0.9570
   4.250   0.6026   0.01664   0.00817   0.0205   0.3577   0.9584
   4.500   0.6295   0.01671   0.00830   0.0204   0.3538   0.9597
   4.750   0.6549   0.01679   0.00842   0.0205   0.3498   0.9609
   5.000   0.6790   0.01690   0.00855   0.0210   0.3461   0.9625
   5.250   0.7029   0.01704   0.00866   0.0215   0.3429   0.9641
   5.500   0.7312   0.01713   0.00882   0.0211   0.3390   0.9648
   5.750   0.7588   0.01723   0.00900   0.0208   0.3350   0.9655
   6.000   0.7860   0.01733   0.00915   0.0206   0.3312   0.9663
   6.250   0.8133   0.01747   0.00928   0.0203   0.3276   0.9673
   6.500   0.8399   0.01762   0.00951   0.0202   0.3237   0.9685
   6.750   0.8648   0.01776   0.00976   0.0204   0.3193   0.9695
   7.000   0.8887   0.01790   0.00996   0.0208   0.3150   0.9705
   7.250   0.9116   0.01808   0.01014   0.0214   0.3113   0.9715
   7.500   0.9340   0.01827   0.01045   0.0221   0.3068   0.9730
   7.750   0.9612   0.01842   0.01071   0.0217   0.3018   0.9737
   8.000   0.9866   0.01858   0.01091   0.0217   0.2971   0.9744
   8.250   1.0113   0.01878   0.01121   0.0218   0.2920   0.9752
   8.500   1.0355   0.01897   0.01152   0.0221   0.2862   0.9760
   8.750   1.0587   0.01917   0.01177   0.0225   0.2812   0.9769
   9.000   1.0817   0.01941   0.01215   0.0228   0.2753   0.9779
   9.250   1.1039   0.01964   0.01247   0.0234   0.2686   0.9794
   9.500   1.1235   0.01991   0.01283   0.0244   0.2623   0.9805
   9.750   1.1449   0.02018   0.01322   0.0250   0.2549   0.9814
  10.000   1.1669   0.02049   0.01360   0.0254   0.2475   0.9822
  10.250   1.1877   0.02082   0.01404   0.0259   0.2387   0.9832
  10.500   1.2072   0.02122   0.01454   0.0267   0.2298   0.9843
  10.750   1.2240   0.02169   0.01506   0.0278   0.2210   0.9856
  11.000   1.2403   0.02222   0.01569   0.0289   0.2106   0.9874
  11.250   1.2565   0.02284   0.01638   0.0299   0.1997   0.9889
  11.500   1.2707   0.02357   0.01717   0.0309   0.1886   0.9904
  11.750   1.2808   0.02444   0.01809   0.0324   0.1775   0.9923
  12.000   1.2861   0.02547   0.01916   0.0344   0.1666   0.9945
  12.250   1.2891   0.02658   0.02034   0.0363   0.1559   0.9966
  12.500   1.2878   0.02812   0.02195   0.0377   0.1465   0.9992
  12.750   1.2724   0.03000   0.02387   0.0409   0.1405   1.0000
  13.000   1.2551   0.03204   0.02599   0.0442   0.1356   1.0000
  13.250   1.2378   0.03438   0.02839   0.0469   0.1304   1.0000
  13.500   1.2177   0.03713   0.03118   0.0492   0.1261   1.0000
  13.750   1.2018   0.03975   0.03387   0.0512   0.1212   1.0000
  14.000   1.1838   0.04276   0.03693   0.0527   0.1166   1.0000
  14.250   1.1647   0.04614   0.04034   0.0538   0.1126   1.0000
  14.500   1.1516   0.04932   0.04359   0.0543   0.1075   1.0000
  14.750   1.1366   0.05306   0.04736   0.0541   0.1029   1.0000
  15.000   1.1235   0.05687   0.05122   0.0536   0.0985   1.0000
  15.250   1.1128   0.06069   0.05511   0.0527   0.0935   1.0000
  15.500   1.0996   0.06503   0.05946   0.0515   0.0893   1.0000
  15.750   1.0922   0.06883   0.06335   0.0503   0.0846   1.0000
  16.000   1.0824   0.07305   0.06761   0.0489   0.0802   1.0000
  16.250   1.0731   0.07728   0.07189   0.0475   0.0764   1.0000
  16.500   1.0666   0.08126   0.07594   0.0461   0.0722   1.0000
  16.750   1.0574   0.08568   0.08039   0.0444   0.0686   1.0000
  17.000   1.0506   0.08981   0.08458   0.0429   0.0651   1.0000
<< Back to EPPLER 340 AIRFOIL (e340-il)

Polar data table (+)

Polar graphs


<< Back to EPPLER 340 AIRFOIL (e340-il)