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

EPPLER 544 AIRFOIL (e544-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: EPPLER 544 AIRFOIL (e544-il)
Reynolds number: 200,000
Max Cl/Cd: 61.74 at α=7.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-e544-il-200000-n5.txt
Download as CSV file: xf-e544-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 544 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
  ------ -------- --------- --------- -------- -------- --------
 -13.750  -0.5013   0.09507   0.09127  -0.0614   1.0000   0.0076
 -13.500  -0.5232   0.08530   0.08140  -0.0673   1.0000   0.0075
 -13.250  -0.5473   0.07718   0.07315  -0.0719   1.0000   0.0074
 -13.000  -0.5760   0.06993   0.06572  -0.0749   1.0000   0.0074
 -12.750  -0.5949   0.06517   0.06082  -0.0760   1.0000   0.0073
 -12.500  -0.6200   0.06008   0.05554  -0.0761   1.0000   0.0074
 -12.250  -0.6372   0.05640   0.05169  -0.0754   1.0000   0.0073
 -12.000  -0.6567   0.05268   0.04777  -0.0738   1.0000   0.0073
 -11.750  -0.6728   0.04958   0.04447  -0.0717   1.0000   0.0073
 -11.500  -0.6861   0.04705   0.04177  -0.0690   1.0000   0.0073
 -10.750  -0.6732   0.03735   0.03103  -0.0691   0.9709   0.0075
 -10.500  -0.6537   0.03484   0.02816  -0.0699   0.9599   0.0077
 -10.250  -0.6281   0.03273   0.02581  -0.0715   0.9508   0.0079
 -10.000  -0.5976   0.03098   0.02390  -0.0740   0.9433   0.0082
  -9.750  -0.5662   0.02948   0.02225  -0.0763   0.9342   0.0085
  -9.500  -0.5315   0.02799   0.02059  -0.0789   0.9260   0.0090
  -9.250  -0.4977   0.02673   0.01914  -0.0812   0.9153   0.0098
  -9.000  -0.4653   0.02561   0.01791  -0.0836   0.9038   0.0108
  -8.750  -0.4343   0.02461   0.01680  -0.0855   0.8917   0.0118
  -8.500  -0.4064   0.02370   0.01571  -0.0866   0.8792   0.0130
  -8.250  -0.3862   0.02274   0.01470  -0.0866   0.8663   0.0137
  -8.000  -0.3675   0.02194   0.01380  -0.0861   0.8544   0.0147
  -7.750  -0.3504   0.02118   0.01290  -0.0851   0.8433   0.0163
  -7.500  -0.3350   0.02050   0.01216  -0.0839   0.8322   0.0183
  -7.250  -0.3202   0.01988   0.01149  -0.0824   0.8221   0.0214
  -7.000  -0.3060   0.01924   0.01078  -0.0808   0.8132   0.0262
  -6.750  -0.2943   0.01859   0.01012  -0.0788   0.8039   0.0332
  -6.500  -0.2828   0.01794   0.00946  -0.0767   0.7963   0.0453
  -6.250  -0.2746   0.01729   0.00889  -0.0741   0.7877   0.0642
  -6.000  -0.2699   0.01666   0.00836  -0.0709   0.7804   0.0928
  -5.750  -0.2661   0.01604   0.00791  -0.0674   0.7722   0.1315
  -5.250  -0.2651   0.01414   0.00672  -0.0597   0.7585   0.2905
  -5.000  -0.2709   0.01275   0.00606  -0.0549   0.7521   0.4424
  -4.750  -0.2531   0.01309   0.00710  -0.0519   0.7473   0.6002
  -4.500  -0.2297   0.01325   0.00714  -0.0511   0.7413   0.6304
  -4.250  -0.2032   0.01359   0.00733  -0.0506   0.7362   0.6478
  -4.000  -0.1761   0.01407   0.00768  -0.0500   0.7316   0.6634
  -3.750  -0.1497   0.01471   0.00827  -0.0491   0.7262   0.6782
  -3.500  -0.1212   0.01554   0.00904  -0.0482   0.7216   0.6913
  -3.250  -0.0907   0.01613   0.00954  -0.0479   0.7177   0.6966
  -3.000  -0.0642   0.01615   0.00948  -0.0477   0.7130   0.6991
  -2.750  -0.0389   0.01605   0.00928  -0.0475   0.7080   0.7018
  -2.500  -0.0135   0.01586   0.00896  -0.0475   0.7038   0.7050
  -2.250   0.0115   0.01555   0.00848  -0.0478   0.7002   0.7088
  -2.000   0.0377   0.01550   0.00838  -0.0477   0.6956   0.7099
  -1.750   0.0643   0.01545   0.00827  -0.0477   0.6913   0.7110
  -1.500   0.0916   0.01541   0.00815  -0.0478   0.6875   0.7123
  -1.250   0.1190   0.01535   0.00801  -0.0480   0.6840   0.7137
  -1.000   0.1449   0.01528   0.00790  -0.0480   0.6797   0.7150
  -0.750   0.1713   0.01522   0.00779  -0.0480   0.6757   0.7168
  -0.500   0.1984   0.01514   0.00764  -0.0483   0.6720   0.7186
   0.000   0.2521   0.01495   0.00734  -0.0488   0.6647   0.7220
   0.250   0.2787   0.01486   0.00720  -0.0491   0.6608   0.7241
   0.500   0.3059   0.01482   0.00712  -0.0493   0.6571   0.7254
   0.750   0.3339   0.01481   0.00708  -0.0496   0.6540   0.7263
   1.000   0.3605   0.01481   0.00709  -0.0497   0.6503   0.7272
   1.250   0.3866   0.01481   0.00711  -0.0497   0.6462   0.7282
   1.500   0.4136   0.01481   0.00711  -0.0498   0.6423   0.7293
   1.750   0.4414   0.01481   0.00709  -0.0501   0.6390   0.7304
   2.000   0.4685   0.01483   0.00711  -0.0503   0.6353   0.7317
   2.250   0.4940   0.01485   0.00717  -0.0502   0.6306   0.7333
   2.500   0.5207   0.01485   0.00718  -0.0503   0.6262   0.7349
   2.750   0.5490   0.01485   0.00715  -0.0507   0.6222   0.7364
   3.000   0.5745   0.01486   0.00720  -0.0506   0.6170   0.7378
   3.250   0.6006   0.01486   0.00722  -0.0507   0.6118   0.7393
   3.500   0.6285   0.01486   0.00721  -0.0511   0.6073   0.7407
   3.750   0.6540   0.01490   0.00730  -0.0509   0.6023   0.7416
   4.000   0.6786   0.01493   0.00741  -0.0505   0.5963   0.7426
   4.250   0.7055   0.01495   0.00745  -0.0506   0.5910   0.7437
   4.500   0.7295   0.01501   0.00759  -0.0501   0.5846   0.7450
   4.750   0.7542   0.01505   0.00769  -0.0497   0.5781   0.7465
   5.000   0.7800   0.01509   0.00778  -0.0496   0.5720   0.7480
   5.250   0.8031   0.01514   0.00792  -0.0490   0.5640   0.7495
   5.500   0.8284   0.01518   0.00797  -0.0488   0.5569   0.7510
   5.750   0.8509   0.01523   0.00812  -0.0481   0.5477   0.7526
   6.000   0.8744   0.01529   0.00823  -0.0475   0.5388   0.7542
   6.250   0.8972   0.01535   0.00832  -0.0469   0.5287   0.7558
   6.500   0.9179   0.01543   0.00850  -0.0458   0.5171   0.7572
   6.750   0.9378   0.01552   0.00866  -0.0446   0.5047   0.7584
   7.000   0.9562   0.01564   0.00883  -0.0430   0.4898   0.7599
   7.250   0.9729   0.01579   0.00903  -0.0412   0.4723   0.7616
   7.500   0.9872   0.01599   0.00922  -0.0389   0.4517   0.7635
   7.750   0.9973   0.01624   0.00945  -0.0359   0.4287   0.7656
   8.250   1.0103   0.01718   0.01021  -0.0291   0.3774   0.7698
   8.500   1.0148   0.01787   0.01081  -0.0257   0.3513   0.7719
   9.000   1.0208   0.01951   0.01232  -0.0190   0.3032   0.7756
   9.250   1.0227   0.02049   0.01324  -0.0158   0.2808   0.7777
   9.500   1.0256   0.02154   0.01424  -0.0131   0.2581   0.7799
   9.750   1.0280   0.02269   0.01534  -0.0105   0.2374   0.7823
  10.000   1.0307   0.02394   0.01655  -0.0081   0.2175   0.7851
  10.250   1.0345   0.02522   0.01778  -0.0061   0.1985   0.7878
  10.500   1.0376   0.02655   0.01910  -0.0041   0.1807   0.7902
  10.750   1.0408   0.02796   0.02049  -0.0023   0.1641   0.7925
  11.000   1.0446   0.02941   0.02193  -0.0006   0.1488   0.7951
  11.250   1.0484   0.03094   0.02344   0.0009   0.1340   0.7978
  11.500   1.0530   0.03248   0.02499   0.0021   0.1208   0.8007
  11.750   1.0578   0.03407   0.02658   0.0032   0.1085   0.8036
  12.000   1.0623   0.03569   0.02822   0.0044   0.0975   0.8062
  12.250   1.0667   0.03737   0.02993   0.0054   0.0878   0.8090
  12.500   1.0705   0.03918   0.03176   0.0063   0.0790   0.8122
  12.750   1.0761   0.04093   0.03356   0.0070   0.0706   0.8156
  13.000   1.0812   0.04281   0.03548   0.0075   0.0635   0.8194
  13.250   1.0842   0.04488   0.03759   0.0081   0.0569   0.8230
  13.500   1.0897   0.04679   0.03959   0.0086   0.0511   0.8272
  13.750   1.0931   0.04897   0.04182   0.0089   0.0460   0.8319
  14.000   1.0972   0.05115   0.04407   0.0091   0.0415   0.8366
  14.250   1.1001   0.05347   0.04648   0.0093   0.0380   0.8418
  14.500   1.1035   0.05581   0.04892   0.0094   0.0344   0.8476
  14.750   1.1050   0.05842   0.05162   0.0094   0.0318   0.8539
  15.000   1.1078   0.06091   0.05425   0.0093   0.0291   0.8617
  15.250   1.1080   0.06375   0.05719   0.0092   0.0270   0.8711
  15.500   1.1095   0.06648   0.06008   0.0090   0.0249   0.8844
  15.750   1.1108   0.06935   0.06313   0.0085   0.0229   0.9098
  16.250   1.1085   0.07584   0.06987   0.0067   0.0201   1.0000
  16.500   1.1085   0.07928   0.07341   0.0057   0.0188   1.0000
  16.750   1.1049   0.08330   0.07749   0.0043   0.0176   1.0000
  17.000   1.1040   0.08701   0.08133   0.0031   0.0164   1.0000
  17.250   1.1014   0.09103   0.08548   0.0016   0.0155   1.0000
  17.500   1.0982   0.09527   0.08980  -0.0002   0.0145   1.0000
  17.750   1.0920   0.10001   0.09464  -0.0021   0.0141   1.0000
  18.000   1.0885   0.10442   0.09920  -0.0040   0.0132   1.0000
  18.250   1.0845   0.10900   0.10391  -0.0061   0.0125   1.0000
  18.500   1.0794   0.11384   0.10887  -0.0085   0.0118   1.0000
  18.750   1.0732   0.11895   0.11407  -0.0111   0.0115   1.0000
<< Back to EPPLER 544 AIRFOIL (e544-il)

Polar data table (+)

Polar graphs


<< Back to EPPLER 544 AIRFOIL (e544-il)