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EPPLER 49 AIRFOIL (e49-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: EPPLER 49 AIRFOIL (e49-il)
Reynolds number: 50,000
Max Cl/Cd: 34.28 at α=6.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-e49-il-50000-n5.txt
Download as CSV file: xf-e49-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 49 AIRFOIL                               
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.3909   0.12709   0.12090  -0.0148   1.0000   0.0528
  -7.750  -0.3969   0.12540   0.11929  -0.0140   1.0000   0.0533
  -7.500  -0.4046   0.12384   0.11781  -0.0129   1.0000   0.0536
  -7.250  -0.4027   0.12083   0.11485  -0.0114   1.0000   0.0552
  -7.000  -0.4058   0.11868   0.11277  -0.0108   1.0000   0.0568
  -6.750  -0.4080   0.11652   0.11068  -0.0114   1.0000   0.0579
  -6.500  -0.4064   0.11382   0.10804  -0.0114   1.0000   0.0593
  -6.250  -0.4042   0.11164   0.10592  -0.0144   1.0000   0.0611
  -6.000  -0.4006   0.10835   0.10267  -0.0128   1.0000   0.0631
  -5.750  -0.3945   0.10562   0.09998  -0.0149   1.0000   0.0661
  -5.250  -0.3793   0.09935   0.09362  -0.0178   1.0000   0.0707
  -5.000  -0.3491   0.09590   0.09013  -0.0287   0.9980   0.0760
  -4.500  -0.2942   0.08821   0.08238  -0.0390   0.9859   0.0968
  -4.250  -0.2669   0.08464   0.07878  -0.0437   0.9808   0.1085
  -4.000  -0.2340   0.08105   0.07513  -0.0497   0.9761   0.1211
  -3.750  -0.2013   0.07743   0.07143  -0.0556   0.9704   0.1349
  -3.500  -0.1610   0.07373   0.06763  -0.0634   0.9662   0.1529
  -3.250  -0.1219   0.07041   0.06422  -0.0699   0.9624   0.1747
  -2.750   0.0061   0.06042   0.05343  -0.0937   0.9555   0.1100
  -2.500   0.0866   0.05429   0.04641  -0.1065   0.9544   0.0576
  -2.250   0.1370   0.05149   0.04312  -0.1126   0.9519   0.0521
  -2.000   0.1856   0.04936   0.04045  -0.1180   0.9496   0.0517
  -1.750   0.2221   0.04779   0.03840  -0.1209   0.9443   0.0591
  -1.500   0.2625   0.04655   0.03654  -0.1240   0.9405   0.0612
  -1.250   0.3020   0.04541   0.03476  -0.1265   0.9376   0.0598
  -1.000   0.3337   0.04442   0.03332  -0.1274   0.9321   0.0588
  -0.750   0.3671   0.04377   0.03229  -0.1286   0.9274   0.0587
  -0.500   0.4028   0.04346   0.03162  -0.1300   0.9238   0.0593
  -0.250   0.4291   0.04312   0.03104  -0.1300   0.9172   0.0602
   0.000   0.4615   0.04304   0.03072  -0.1312   0.9121   0.0620
   0.250   0.4951   0.04308   0.03056  -0.1326   0.9073   0.0650
   0.500   0.5219   0.04305   0.03042  -0.1330   0.8997   0.0703
   0.750   0.5586   0.04312   0.03047  -0.1351   0.8951   0.0906
   1.000   0.5790   0.04135   0.03068  -0.1345   0.8872   1.0000
   1.250   0.6108   0.04179   0.03074  -0.1356   0.8807   1.0000
   1.500   0.6358   0.04215   0.03087  -0.1357   0.8721   1.0000
   1.750   0.6670   0.04251   0.03103  -0.1367   0.8650   1.0000
   2.000   0.6921   0.04283   0.03124  -0.1368   0.8558   1.0000
   2.250   0.7248   0.04309   0.03139  -0.1381   0.8485   1.0000
   2.500   0.7488   0.04334   0.03161  -0.1380   0.8380   1.0000
   2.750   0.7851   0.04340   0.03163  -0.1397   0.8313   1.0000
   3.250   0.8333   0.04368   0.03197  -0.1391   0.8080   1.0000
   3.500   0.8703   0.04344   0.03190  -0.1407   0.8004   1.0000
   3.750   0.8931   0.04355   0.03208  -0.1401   0.7872   1.0000
   4.000   0.9178   0.04362   0.03225  -0.1397   0.7746   1.0000
   4.250   0.9465   0.04351   0.03225  -0.1399   0.7632   1.0000
   4.500   0.9817   0.04302   0.03192  -0.1408   0.7538   1.0000
   4.750   1.0069   0.04294   0.03201  -0.1403   0.7398   1.0000
   5.000   1.0343   0.04268   0.03193  -0.1399   0.7259   1.0000
   5.250   1.0640   0.04216   0.03183  -0.1396   0.7118   1.0000
   5.500   1.0955   0.04137   0.03129  -0.1393   0.6974   1.0000
   5.750   1.1270   0.04052   0.03073  -0.1389   0.6811   1.0000
   6.000   1.1578   0.03973   0.03023  -0.1383   0.6619   1.0000
   6.250   1.1906   0.03886   0.02970  -0.1378   0.6410   1.0000
   6.500   1.2168   0.03848   0.02965  -0.1365   0.6132   1.0000
   6.750   1.2418   0.03623   0.02617  -0.1311   0.4317   1.0000
   7.000   1.2330   0.03893   0.02777  -0.1257   0.2834   1.0000
   7.250   1.2263   0.04215   0.02998  -0.1219   0.1570   1.0000
   7.500   1.2196   0.04641   0.03286  -0.1189   0.0253   1.0000
   7.750   1.2309   0.04846   0.03501  -0.1171   0.0198   1.0000
   8.000   1.2426   0.05035   0.03720  -0.1153   0.0180   1.0000
   8.250   1.2531   0.05237   0.03951  -0.1135   0.0169   1.0000
   8.500   1.2617   0.05455   0.04200  -0.1116   0.0161   1.0000
   8.750   1.2683   0.05694   0.04468  -0.1097   0.0156   1.0000
   9.000   1.2732   0.05951   0.04754  -0.1077   0.0151   1.0000
   9.250   1.2770   0.06223   0.05055  -0.1059   0.0148   1.0000
   9.500   1.2804   0.06508   0.05367  -0.1040   0.0146   1.0000
   9.750   1.2842   0.06797   0.05684  -0.1024   0.0144   1.0000
  10.000   1.2894   0.07088   0.06002  -0.1008   0.0142   1.0000
  10.250   1.2973   0.07373   0.06313  -0.0995   0.0141   1.0000
  10.500   1.3082   0.07654   0.06623  -0.0983   0.0140   1.0000
  10.750   1.3204   0.07949   0.06949  -0.0972   0.0140   1.0000
  11.000   1.3306   0.08276   0.07310  -0.0962   0.0140   1.0000
  11.250   1.3363   0.08638   0.07708  -0.0952   0.0140   1.0000
  11.500   1.3371   0.09036   0.08142  -0.0944   0.0141   1.0000
  11.750   1.3339   0.09464   0.08606  -0.0938   0.0141   1.0000
  12.000   1.3275   0.09924   0.09101  -0.0936   0.0142   1.0000
  12.250   1.3185   0.10420   0.09630  -0.0939   0.0143   1.0000
  12.500   1.3073   0.10954   0.10196  -0.0949   0.0144   1.0000
  12.750   1.2946   0.11531   0.10803  -0.0966   0.0146   1.0000
  13.000   1.2808   0.12156   0.11455  -0.0991   0.0147   1.0000
  13.250   1.2662   0.12835   0.12158  -0.1024   0.0148   1.0000
  13.500   1.2515   0.13573   0.12919  -0.1067   0.0150   1.0000
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