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

EPPLER E852 AIRFOIL (e852-il) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: EPPLER E852 AIRFOIL (e852-il)
Reynolds number: 500,000
Max Cl/Cd: 98.95 at α=3°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-e852-il-500000-n5.txt
Download as CSV file: xf-e852-il-500000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER E852 AIRFOIL                             
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.750  -0.4462   0.09248   0.09027  -0.0405   1.0000   0.0026
  -9.500  -0.4527   0.08793   0.08576  -0.0416   1.0000   0.0025
  -9.250  -0.4595   0.08345   0.08133  -0.0429   0.9999   0.0026
  -9.000  -0.4597   0.07500   0.07291  -0.0507   0.9975   0.0025
  -8.750  -0.4611   0.06403   0.06196  -0.0628   0.9931   0.0024
  -8.500  -0.4692   0.04974   0.04746  -0.0829   0.9819   0.0024
  -8.250  -0.4701   0.04205   0.03946  -0.0887   0.9702   0.0025
  -7.750  -0.4543   0.02882   0.02532  -0.0933   0.9492   0.0027
  -7.500  -0.4299   0.02488   0.02091  -0.0955   0.9443   0.0029
  -7.250  -0.4036   0.02207   0.01771  -0.0969   0.9385   0.0031
  -7.000  -0.3730   0.01956   0.01482  -0.0988   0.9338   0.0033
  -6.750  -0.3386   0.01848   0.01353  -0.1009   0.9302   0.0039
  -6.500  -0.3074   0.01733   0.01217  -0.1022   0.9234   0.0041
  -6.250  -0.2746   0.01581   0.01035  -0.1038   0.9177   0.0044
  -6.000  -0.2461   0.01412   0.00840  -0.1045   0.9097   0.0045
  -5.750  -0.2167   0.01285   0.00692  -0.1053   0.9021   0.0046
  -5.500  -0.1892   0.01178   0.00567  -0.1057   0.8934   0.0048
  -5.250  -0.1618   0.01099   0.00471  -0.1059   0.8850   0.0051
  -4.750  -0.1063   0.01003   0.00348  -0.1062   0.8698   0.0078
  -4.500  -0.0785   0.00965   0.00304  -0.1063   0.8630   0.0183
  -4.250  -0.0517   0.00930   0.00275  -0.1065   0.8559   0.0405
  -4.000  -0.0243   0.00894   0.00245  -0.1067   0.8499   0.0748
  -3.750   0.0024   0.00856   0.00220  -0.1068   0.8436   0.1241
  -3.500   0.0294   0.00814   0.00196  -0.1071   0.8379   0.1932
  -3.250   0.0554   0.00751   0.00173  -0.1074   0.8321   0.3168
  -3.000   0.0826   0.00723   0.00163  -0.1075   0.8268   0.3927
  -2.750   0.1102   0.00710   0.00157  -0.1076   0.8219   0.4388
  -2.500   0.1379   0.00703   0.00151  -0.1077   0.8167   0.4652
  -2.250   0.1659   0.00701   0.00145  -0.1078   0.8121   0.4835
  -2.000   0.1937   0.00697   0.00143  -0.1079   0.8077   0.5050
  -1.750   0.2214   0.00694   0.00143  -0.1079   0.8029   0.5265
  -1.500   0.2494   0.00694   0.00140  -0.1080   0.7987   0.5401
  -1.250   0.2773   0.00693   0.00137  -0.1081   0.7946   0.5498
  -1.000   0.3052   0.00692   0.00137  -0.1082   0.7901   0.5573
  -0.500   0.3612   0.00694   0.00137  -0.1084   0.7824   0.5723
  -0.250   0.3891   0.00694   0.00139  -0.1085   0.7780   0.5799
   0.000   0.4170   0.00695   0.00142  -0.1085   0.7739   0.5878
   0.250   0.4449   0.00697   0.00144  -0.1086   0.7702   0.5956
   0.500   0.4726   0.00698   0.00149  -0.1086   0.7652   0.6043
   0.750   0.5001   0.00699   0.00153  -0.1086   0.7600   0.6129
   1.000   0.5275   0.00701   0.00160  -0.1086   0.7540   0.6220
   1.250   0.5547   0.00703   0.00164  -0.1085   0.7464   0.6319
   1.500   0.5815   0.00704   0.00170  -0.1083   0.7372   0.6420
   1.750   0.6080   0.00707   0.00176  -0.1080   0.7269   0.6527
   2.000   0.6342   0.00710   0.00185  -0.1077   0.7143   0.6640
   2.250   0.6597   0.00714   0.00191  -0.1071   0.6967   0.6763
   2.500   0.6845   0.00720   0.00198  -0.1065   0.6744   0.6895
   2.750   0.7092   0.00729   0.00207  -0.1058   0.6510   0.7035
   3.000   0.7332   0.00741   0.00219  -0.1050   0.6208   0.7190
   3.250   0.7523   0.00777   0.00235  -0.1032   0.5544   0.7359
   3.500   0.7662   0.00854   0.00273  -0.1006   0.4504   0.7540
   3.750   0.7805   0.00940   0.00319  -0.0982   0.3461   0.7747
   4.000   0.7958   0.01021   0.00365  -0.0961   0.2504   0.7988
   4.250   0.8102   0.01106   0.00414  -0.0938   0.1565   0.8278
   4.500   0.8231   0.01186   0.00463  -0.0911   0.0767   0.8676
   4.750   0.8408   0.01222   0.00496  -0.0891   0.0433   0.9645
   5.000   0.8625   0.01279   0.00538  -0.0882   0.0191   1.0000
   5.250   0.8838   0.01336   0.00586  -0.0872   0.0050   1.0000
   5.500   0.9058   0.01386   0.00638  -0.0862   0.0025   1.0000
   5.750   0.9276   0.01438   0.00698  -0.0851   0.0020   1.0000
   6.000   0.9489   0.01495   0.00771  -0.0840   0.0018   1.0000
   6.250   0.9681   0.01574   0.00863  -0.0824   0.0017   1.0000
   6.500   0.9873   0.01650   0.00949  -0.0809   0.0016   1.0000
   6.750   1.0058   0.01733   0.01043  -0.0794   0.0016   1.0000
   7.000   1.0235   0.01828   0.01148  -0.0776   0.0016   1.0000
   7.250   1.0403   0.01938   0.01272  -0.0758   0.0016   1.0000
   7.500   1.0573   0.02059   0.01406  -0.0740   0.0015   1.0000
   7.750   1.0746   0.02197   0.01559  -0.0724   0.0015   1.0000
   8.000   1.0924   0.02360   0.01741  -0.0708   0.0015   1.0000
   8.250   1.1105   0.02541   0.01944  -0.0694   0.0015   1.0000
   8.500   1.1274   0.02753   0.02182  -0.0678   0.0015   1.0000
   8.750   1.1420   0.02986   0.02444  -0.0659   0.0016   1.0000
   9.000   1.1525   0.03255   0.02745  -0.0635   0.0016   1.0000
   9.250   1.1569   0.03549   0.03074  -0.0603   0.0016   1.0000
   9.500   1.1556   0.03843   0.03406  -0.0564   0.0016   1.0000
   9.750   1.1483   0.04162   0.03757  -0.0520   0.0016   1.0000
  10.000   1.1372   0.04493   0.04116  -0.0477   0.0016   1.0000
  10.250   1.1211   0.04864   0.04515  -0.0437   0.0016   1.0000
  10.750   1.0866   0.05645   0.05342  -0.0383   0.0017   1.0000
  11.000   1.0724   0.06039   0.05754  -0.0371   0.0017   1.0000
  11.250   1.0571   0.06477   0.06211  -0.0367   0.0017   1.0000
  11.500   1.0415   0.06964   0.06713  -0.0373   0.0017   1.0000
  11.750   1.0234   0.07546   0.07312  -0.0390   0.0017   1.0000
  12.000   1.0052   0.08211   0.07992  -0.0421   0.0017   1.0000
<< Back to EPPLER E852 AIRFOIL (e852-il)

Polar data table (+)

Polar graphs


<< Back to EPPLER E852 AIRFOIL (e852-il)