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

EPPLER 472 AIRFOIL (e472-il) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: EPPLER 472 AIRFOIL (e472-il)
Reynolds number: 500,000
Max Cl/Cd: 72.64 at α=9.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-e472-il-500000-n5.txt
Download as CSV file: xf-e472-il-500000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 472 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
  ------ -------- --------- --------- -------- -------- --------
 -17.500  -1.1811   0.08823   0.08424  -0.0202   1.0000   0.0314
 -17.250  -1.2246   0.07833   0.07420  -0.0255   1.0000   0.0311
 -17.000  -1.2797   0.06667   0.06235  -0.0319   1.0000   0.0304
 -16.750  -1.3267   0.05594   0.05140  -0.0381   1.0000   0.0297
 -16.500  -1.3595   0.04710   0.04233  -0.0439   1.0000   0.0297
 -16.250  -1.3784   0.04091   0.03597  -0.0479   1.0000   0.0303
 -16.000  -1.3893   0.03667   0.03156  -0.0499   1.0000   0.0310
 -15.750  -1.3983   0.03337   0.02814  -0.0505   1.0000   0.0321
 -15.500  -1.4029   0.03101   0.02568  -0.0497   1.0000   0.0331
 -15.250  -1.4042   0.02923   0.02380  -0.0482   1.0000   0.0342
 -15.000  -1.4023   0.02785   0.02234  -0.0462   1.0000   0.0357
 -14.750  -1.3996   0.02665   0.02106  -0.0438   1.0000   0.0373
 -14.500  -1.3968   0.02556   0.01991  -0.0410   1.0000   0.0390
 -14.250  -1.3917   0.02469   0.01897  -0.0381   1.0000   0.0407
 -14.000  -1.3852   0.02397   0.01818  -0.0351   1.0000   0.0424
 -13.750  -1.3793   0.02317   0.01734  -0.0319   1.0000   0.0442
 -13.500  -1.3681   0.02243   0.01655  -0.0296   1.0000   0.0462
 -13.250  -1.3546   0.02180   0.01586  -0.0274   1.0000   0.0483
 -13.000  -1.3402   0.02119   0.01519  -0.0254   1.0000   0.0501
 -12.750  -1.3261   0.02053   0.01450  -0.0234   1.0000   0.0525
 -12.500  -1.3099   0.01997   0.01389  -0.0215   1.0000   0.0549
 -12.250  -1.2923   0.01949   0.01335  -0.0198   1.0000   0.0570
 -12.000  -1.2756   0.01892   0.01275  -0.0180   1.0000   0.0594
 -11.750  -1.2578   0.01841   0.01222  -0.0163   1.0000   0.0619
 -11.500  -1.2388   0.01797   0.01172  -0.0148   1.0000   0.0645
 -11.250  -1.2197   0.01753   0.01124  -0.0132   1.0000   0.0669
 -11.000  -1.2007   0.01706   0.01076  -0.0117   1.0000   0.0697
 -10.750  -1.1806   0.01666   0.01033  -0.0102   1.0000   0.0728
 -10.500  -1.1598   0.01632   0.00993  -0.0089   1.0000   0.0756
 -10.250  -1.1400   0.01587   0.00950  -0.0074   1.0000   0.0792
 -10.000  -1.1190   0.01552   0.00913  -0.0060   1.0000   0.0825
  -9.750  -1.0973   0.01522   0.00878  -0.0048   1.0000   0.0851
  -9.500  -1.0762   0.01486   0.00840  -0.0034   1.0000   0.0877
  -9.250  -1.0548   0.01452   0.00806  -0.0021   1.0000   0.0909
  -9.000  -1.0329   0.01423   0.00774  -0.0008   1.0000   0.0938
  -8.750  -1.0104   0.01397   0.00744   0.0004   1.0000   0.0964
  -8.500  -0.9885   0.01366   0.00712   0.0017   1.0000   0.0991
  -8.250  -0.9663   0.01338   0.00684   0.0029   1.0000   0.1024
  -8.000  -0.9437   0.01313   0.00658   0.0041   1.0000   0.1056
  -7.750  -0.9207   0.01291   0.00632   0.0052   1.0000   0.1081
  -7.500  -0.8981   0.01266   0.00606   0.0064   1.0000   0.1110
  -7.250  -0.8754   0.01241   0.00583   0.0076   1.0000   0.1145
  -7.000  -0.8524   0.01220   0.00561   0.0088   1.0000   0.1177
  -6.750  -0.8292   0.01202   0.00540   0.0099   1.0000   0.1206
  -6.500  -0.8061   0.01182   0.00519   0.0111   1.0000   0.1235
  -6.250  -0.7832   0.01161   0.00500   0.0122   1.0000   0.1269
  -6.000  -0.7600   0.01143   0.00483   0.0134   1.0000   0.1301
  -5.750  -0.7366   0.01128   0.00467   0.0145   1.0000   0.1334
  -5.500  -0.6998   0.01113   0.00451   0.0128   0.9976   0.1367
  -5.250  -0.6544   0.01095   0.00438   0.0092   0.9932   0.1412
  -5.000  -0.6118   0.01075   0.00419   0.0063   0.9868   0.1458
  -4.750  -0.5686   0.01058   0.00401   0.0033   0.9801   0.1494
  -4.500  -0.5372   0.01034   0.00380   0.0028   0.9717   0.1531
  -4.250  -0.5101   0.01014   0.00364   0.0033   0.9605   0.1571
  -4.000  -0.4814   0.00995   0.00346   0.0034   0.9455   0.1610
  -3.750  -0.4423   0.00977   0.00326   0.0014   0.9284   0.1644
  -3.500  -0.3582   0.00943   0.00292  -0.0103   0.8800   0.1717
  -3.250  -0.3217   0.00959   0.00271  -0.0116   0.7736   0.1767
  -3.000  -0.2998   0.00979   0.00261  -0.0100   0.6960   0.1796
  -2.750  -0.2777   0.00992   0.00252  -0.0085   0.6288   0.1835
  -2.500  -0.2548   0.01004   0.00244  -0.0072   0.5655   0.1880
  -2.250  -0.2310   0.01018   0.00237  -0.0062   0.5053   0.1928
  -2.000  -0.2062   0.01027   0.00230  -0.0053   0.4581   0.1971
  -1.750  -0.1811   0.01032   0.00224  -0.0045   0.4192   0.2023
  -1.500  -0.1556   0.01037   0.00219  -0.0038   0.3850   0.2079
  -1.250  -0.1299   0.01043   0.00215  -0.0031   0.3539   0.2140
  -1.000  -0.1043   0.01046   0.00212  -0.0024   0.3264   0.2215
  -0.750  -0.0782   0.01050   0.00209  -0.0018   0.3044   0.2282
  -0.500  -0.0522   0.01050   0.00207  -0.0012   0.2871   0.2367
  -0.250  -0.0260   0.01052   0.00206  -0.0006   0.2722   0.2461
   0.000   0.0000   0.01052   0.00206   0.0000   0.2587   0.2585
   0.250   0.0260   0.01052   0.00206   0.0006   0.2460   0.2722
   0.500   0.0522   0.01050   0.00207   0.0012   0.2368   0.2871
   0.750   0.0782   0.01050   0.00209   0.0018   0.2282   0.3044
   1.000   0.1043   0.01046   0.00212   0.0024   0.2215   0.3267
   1.250   0.1299   0.01043   0.00215   0.0031   0.2139   0.3539
   1.500   0.1556   0.01037   0.00219   0.0038   0.2079   0.3851
   1.750   0.1811   0.01032   0.00224   0.0045   0.2023   0.4189
   2.000   0.2062   0.01028   0.00230   0.0053   0.1971   0.4580
   2.250   0.2310   0.01018   0.00237   0.0062   0.1929   0.5053
   2.500   0.2548   0.01004   0.00244   0.0072   0.1881   0.5653
   2.750   0.2777   0.00992   0.00252   0.0085   0.1835   0.6291
   3.000   0.2999   0.00979   0.00261   0.0100   0.1796   0.6956
   3.250   0.3216   0.00959   0.00271   0.0116   0.1767   0.7740
   3.500   0.3598   0.00943   0.00293   0.0100   0.1716   0.8827
   3.750   0.4424   0.00977   0.00326  -0.0015   0.1644   0.9284
   4.000   0.4816   0.00995   0.00346  -0.0035   0.1610   0.9458
   4.250   0.5101   0.01014   0.00363  -0.0033   0.1570   0.9604
   4.500   0.5371   0.01034   0.00380  -0.0028   0.1531   0.9716
   4.750   0.5687   0.01058   0.00401  -0.0033   0.1495   0.9803
   5.000   0.6121   0.01075   0.00419  -0.0064   0.1458   0.9869
   5.250   0.6546   0.01095   0.00438  -0.0092   0.1412   0.9933
   5.500   0.6998   0.01113   0.00451  -0.0128   0.1367   0.9976
   5.750   0.7366   0.01128   0.00467  -0.0145   0.1334   1.0000
   6.000   0.7600   0.01143   0.00483  -0.0134   0.1301   1.0000
   6.250   0.7832   0.01161   0.00500  -0.0122   0.1269   1.0000
   6.500   0.8061   0.01182   0.00519  -0.0111   0.1235   1.0000
   6.750   0.8292   0.01202   0.00540  -0.0099   0.1206   1.0000
   7.000   0.8524   0.01220   0.00561  -0.0088   0.1177   1.0000
   7.250   0.8754   0.01241   0.00583  -0.0076   0.1144   1.0000
   7.500   0.8981   0.01266   0.00605  -0.0065   0.1109   1.0000
   7.750   0.9208   0.01291   0.00632  -0.0053   0.1082   1.0000
   8.000   0.9437   0.01313   0.00658  -0.0041   0.1056   1.0000
   8.250   0.9663   0.01338   0.00684  -0.0029   0.1024   1.0000
   8.500   0.9885   0.01366   0.00712  -0.0017   0.0991   1.0000
   8.750   1.0104   0.01397   0.00744  -0.0004   0.0964   1.0000
   9.000   1.0329   0.01423   0.00774   0.0008   0.0939   1.0000
   9.250   1.0548   0.01452   0.00806   0.0021   0.0909   1.0000
   9.500   1.0762   0.01486   0.00840   0.0034   0.0878   1.0000
   9.750   1.0973   0.01522   0.00878   0.0047   0.0851   1.0000
  10.000   1.1190   0.01552   0.00913   0.0060   0.0825   1.0000
  10.250   1.1400   0.01587   0.00950   0.0074   0.0792   1.0000
  10.500   1.1598   0.01632   0.00993   0.0088   0.0756   1.0000
  10.750   1.1807   0.01666   0.01033   0.0102   0.0728   1.0000
  11.000   1.2007   0.01706   0.01076   0.0117   0.0697   1.0000
  11.250   1.2196   0.01753   0.01124   0.0132   0.0669   1.0000
  11.500   1.2389   0.01797   0.01172   0.0148   0.0644   1.0000
  11.750   1.2578   0.01841   0.01221   0.0163   0.0620   1.0000
  12.000   1.2757   0.01891   0.01275   0.0180   0.0594   1.0000
  12.250   1.2923   0.01949   0.01335   0.0198   0.0570   1.0000
  12.500   1.3099   0.01997   0.01389   0.0215   0.0549   1.0000
  12.750   1.3261   0.02053   0.01449   0.0233   0.0525   1.0000
  13.000   1.3404   0.02119   0.01519   0.0254   0.0502   1.0000
  13.250   1.3547   0.02180   0.01586   0.0274   0.0482   1.0000
  13.500   1.3682   0.02243   0.01655   0.0295   0.0462   1.0000
  13.750   1.3793   0.02318   0.01734   0.0319   0.0441   1.0000
  14.000   1.3856   0.02397   0.01818   0.0350   0.0424   1.0000
  14.250   1.3920   0.02469   0.01897   0.0381   0.0406   1.0000
  14.500   1.3971   0.02556   0.01991   0.0410   0.0389   1.0000
  14.750   1.4001   0.02665   0.02106   0.0437   0.0373   1.0000
  15.000   1.4029   0.02784   0.02233   0.0461   0.0360   1.0000
  15.250   1.4051   0.02919   0.02377   0.0481   0.0346   1.0000
  15.500   1.4040   0.03097   0.02564   0.0496   0.0331   1.0000
  15.750   1.3983   0.03343   0.02820   0.0503   0.0319   1.0000
  16.000   1.3906   0.03661   0.03150   0.0498   0.0311   1.0000
  16.250   1.3799   0.04084   0.03589   0.0478   0.0300   1.0000
  16.500   1.3611   0.04703   0.04226   0.0438   0.0297   1.0000
  16.750   1.3296   0.05566   0.05112   0.0381   0.0298   1.0000
  17.000   1.2815   0.06663   0.06231   0.0317   0.0302   1.0000
  17.250   1.2278   0.07808   0.07395   0.0254   0.0308   1.0000
  17.500   1.1839   0.08805   0.08407   0.0201   0.0314   1.0000
<< Back to EPPLER 472 AIRFOIL (e472-il)

Polar data table (+)

Polar graphs


<< Back to EPPLER 472 AIRFOIL (e472-il)