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

EPPLER 625 AIRFOIL (e625-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5


Details Polar file
Airfoil: EPPLER 625 AIRFOIL (e625-il)
Reynolds number: 1,000,000
Max Cl/Cd: 93.2 at α=9.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-e625-il-1000000-n5.txt
Download as CSV file: xf-e625-il-1000000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 625 AIRFOIL                              
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     1.000 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -13.000  -0.8404   0.04929   0.04711  -0.0471   1.0000   0.0086
 -12.750  -0.9130   0.03951   0.03681  -0.0449   1.0000   0.0086
 -12.500  -0.9307   0.03634   0.03342  -0.0413   1.0000   0.0086
 -12.250  -0.9316   0.03474   0.03170  -0.0383   1.0000   0.0086
 -12.000  -0.9318   0.03328   0.03012  -0.0351   1.0000   0.0087
 -11.750  -0.9385   0.03137   0.02803  -0.0306   1.0000   0.0087
 -11.500  -0.9383   0.03006   0.02659  -0.0266   1.0000   0.0088
 -11.250  -0.9364   0.02848   0.02484  -0.0229   1.0000   0.0088
 -11.000  -0.9279   0.02731   0.02355  -0.0201   1.0000   0.0089
 -10.750  -0.9170   0.02630   0.02243  -0.0175   1.0000   0.0090
 -10.250  -0.8615   0.02306   0.01880  -0.0198   0.9534   0.0091
 -10.000  -0.8190   0.02158   0.01702  -0.0238   0.9025   0.0093
  -9.750  -0.8031   0.02079   0.01596  -0.0217   0.8614   0.0094
  -9.500  -0.7871   0.02002   0.01498  -0.0195   0.8344   0.0095
  -9.250  -0.7693   0.01937   0.01415  -0.0177   0.8120   0.0096
  -9.000  -0.7512   0.01855   0.01314  -0.0159   0.7926   0.0097
  -8.750  -0.7311   0.01797   0.01241  -0.0144   0.7749   0.0098
  -8.500  -0.7103   0.01740   0.01169  -0.0131   0.7589   0.0100
  -8.250  -0.6886   0.01692   0.01109  -0.0119   0.7434   0.0101
  -8.000  -0.6662   0.01647   0.01052  -0.0108   0.7289   0.0103
  -7.750  -0.6435   0.01603   0.00996  -0.0097   0.7152   0.0104
  -7.500  -0.6206   0.01557   0.00939  -0.0087   0.7015   0.0105
  -7.250  -0.5984   0.01492   0.00863  -0.0076   0.6891   0.0106
  -7.000  -0.5760   0.01432   0.00793  -0.0065   0.6762   0.0108
  -6.750  -0.5527   0.01389   0.00742  -0.0055   0.6627   0.0109
  -6.500  -0.5289   0.01351   0.00698  -0.0047   0.6505   0.0111
  -6.250  -0.5047   0.01322   0.00662  -0.0039   0.6380   0.0113
  -6.000  -0.4805   0.01290   0.00623  -0.0030   0.6263   0.0114
  -5.750  -0.4560   0.01262   0.00589  -0.0022   0.6143   0.0116
  -5.500  -0.4315   0.01232   0.00553  -0.0015   0.6031   0.0117
  -5.250  -0.4069   0.01208   0.00523  -0.0007   0.5917   0.0120
  -5.000  -0.3822   0.01183   0.00492   0.0001   0.5803   0.0122
  -4.750  -0.3575   0.01158   0.00461   0.0008   0.5681   0.0124
  -4.500  -0.3328   0.01136   0.00433   0.0016   0.5563   0.0127
  -4.250  -0.3080   0.01115   0.00406   0.0023   0.5447   0.0129
  -4.000  -0.2829   0.01098   0.00383   0.0030   0.5334   0.0131
  -3.750  -0.2576   0.01081   0.00362   0.0037   0.5232   0.0133
  -3.500  -0.2334   0.01056   0.00332   0.0045   0.5123   0.0138
  -3.250  -0.2082   0.01040   0.00312   0.0052   0.5023   0.0143
  -3.000  -0.1830   0.01026   0.00295   0.0059   0.4928   0.0148
  -2.750  -0.1577   0.01014   0.00279   0.0065   0.4826   0.0154
  -2.500  -0.1322   0.01003   0.00263   0.0072   0.4720   0.0160
  -2.250  -0.1067   0.00993   0.00250   0.0078   0.4628   0.0164
  -2.000  -0.0814   0.00982   0.00235   0.0084   0.4534   0.0173
  -1.750  -0.0559   0.00973   0.00224   0.0090   0.4432   0.0184
  -1.250  -0.0049   0.00958   0.00204   0.0102   0.4245   0.0242
  -1.000   0.0192   0.00942   0.00195   0.0111   0.4156   0.0481
  -0.750   0.0437   0.00929   0.00188   0.0118   0.4051   0.0734
  -0.500   0.0681   0.00916   0.00183   0.0126   0.3968   0.1045
  -0.250   0.0923   0.00904   0.00179   0.0134   0.3879   0.1424
   0.000   0.1159   0.00888   0.00176   0.0142   0.3794   0.1925
   0.250   0.1361   0.00855   0.00171   0.0157   0.3702   0.2944
   0.500   0.1221   0.00684   0.00144   0.0241   0.3663   0.7194
   0.750   0.1346   0.00650   0.00153   0.0278   0.3594   0.8424
   1.000   0.1599   0.00656   0.00162   0.0286   0.3519   0.8692
   1.250   0.1860   0.00664   0.00170   0.0292   0.3435   0.8864
   1.500   0.2121   0.00675   0.00175   0.0297   0.3352   0.8957
   1.750   0.2389   0.00685   0.00183   0.0301   0.3276   0.9038
   2.000   0.2653   0.00696   0.00190   0.0305   0.3206   0.9116
   2.250   0.2924   0.00709   0.00201   0.0309   0.3134   0.9196
   2.500   0.3172   0.00720   0.00206   0.0316   0.3058   0.9233
   2.750   0.3444   0.00727   0.00210   0.0317   0.2997   0.9251
   3.000   0.3727   0.00739   0.00216   0.0316   0.2920   0.9267
   3.250   0.4011   0.00748   0.00223   0.0315   0.2855   0.9283
   3.500   0.4287   0.00759   0.00230   0.0315   0.2792   0.9301
   3.750   0.4557   0.00770   0.00237   0.0316   0.2729   0.9319
   4.000   0.4825   0.00780   0.00245   0.0318   0.2667   0.9339
   4.250   0.5084   0.00793   0.00253   0.0322   0.2612   0.9360
   4.500   0.5340   0.00802   0.00260   0.0326   0.2569   0.9383
   4.750   0.5594   0.00812   0.00269   0.0331   0.2511   0.9405
   5.000   0.5878   0.00828   0.00281   0.0328   0.2450   0.9420
   5.250   0.6165   0.00838   0.00291   0.0325   0.2409   0.9435
   5.500   0.6444   0.00852   0.00303   0.0324   0.2358   0.9451
   5.750   0.6712   0.00868   0.00317   0.0324   0.2299   0.9470
   6.000   0.6981   0.00880   0.00328   0.0325   0.2256   0.9491
   6.250   0.7238   0.00894   0.00341   0.0328   0.2208   0.9516
   6.500   0.7474   0.00910   0.00355   0.0335   0.2154   0.9543
   6.750   0.7740   0.00923   0.00368   0.0336   0.2110   0.9562
   7.000   0.8032   0.00942   0.00385   0.0331   0.2056   0.9574
   7.250   0.8322   0.00962   0.00404   0.0325   0.2006   0.9588
   7.500   0.8613   0.00979   0.00422   0.0320   0.1958   0.9603
   7.750   0.8889   0.01001   0.00442   0.0317   0.1902   0.9621
   8.000   0.9159   0.01021   0.00462   0.0316   0.1859   0.9641
   8.250   0.9422   0.01039   0.00481   0.0316   0.1818   0.9663
   8.500   0.9647   0.01061   0.00502   0.0324   0.1760   0.9693
   8.750   0.9929   0.01085   0.00526   0.0319   0.1715   0.9706
   9.000   1.0232   0.01108   0.00550   0.0310   0.1672   0.9715
   9.250   1.0527   0.01138   0.00578   0.0301   0.1614   0.9725
   9.500   1.0822   0.01164   0.00606   0.0292   0.1571   0.9734
   9.750   1.1109   0.01192   0.00634   0.0285   0.1521   0.9745
  10.000   1.1381   0.01227   0.00667   0.0281   0.1466   0.9758
  10.250   1.1651   0.01254   0.00697   0.0277   0.1426   0.9774
  10.500   1.1903   0.01286   0.00730   0.0276   0.1381   0.9794
  10.750   1.2107   0.01324   0.00767   0.0285   0.1324   0.9825
  11.000   1.2378   0.01360   0.00804   0.0279   0.1268   0.9836
  11.250   1.2644   0.01407   0.00849   0.0273   0.1209   0.9848
  11.500   1.2911   0.01449   0.00892   0.0267   0.1155   0.9861
  11.750   1.3153   0.01503   0.00944   0.0264   0.1092   0.9879
  12.000   1.3374   0.01550   0.00993   0.0265   0.1032   0.9903
  12.250   1.3561   0.01603   0.01045   0.0273   0.0984   0.9932
  12.500   1.3765   0.01662   0.01104   0.0275   0.0920   0.9956
  12.750   1.3964   0.01735   0.01175   0.0274   0.0857   0.9980
  13.250   1.4259   0.01873   0.01318   0.0292   0.0774   1.0000
  13.500   1.4238   0.01958   0.01405   0.0328   0.0737   1.0000
  13.750   1.4256   0.02053   0.01502   0.0356   0.0698   1.0000
  14.000   1.4246   0.02185   0.01634   0.0380   0.0649   1.0000
  14.250   1.4300   0.02297   0.01752   0.0396   0.0625   1.0000
  14.500   1.4317   0.02451   0.01907   0.0411   0.0583   1.0000
  14.750   1.4334   0.02620   0.02080   0.0422   0.0552   1.0000
  15.000   1.4334   0.02818   0.02281   0.0430   0.0505   1.0000
  15.250   1.4330   0.03034   0.02501   0.0436   0.0477   1.0000
  15.500   1.4305   0.03280   0.02750   0.0439   0.0436   1.0000
  15.750   1.4265   0.03553   0.03028   0.0440   0.0407   1.0000
  16.000   1.4183   0.03879   0.03358   0.0439   0.0363   1.0000
  16.250   1.4139   0.04177   0.03662   0.0437   0.0347   1.0000
  16.500   1.3986   0.04601   0.04091   0.0430   0.0303   1.0000
  16.750   1.3928   0.04938   0.04436   0.0424   0.0294   1.0000
  17.000   1.3818   0.05346   0.04852   0.0414   0.0279   1.0000
  17.250   1.3715   0.05757   0.05271   0.0403   0.0262   1.0000
  17.500   1.3614   0.06177   0.05700   0.0391   0.0252   1.0000
  17.750   1.3387   0.06768   0.06298   0.0372   0.0220   1.0000
  18.000   1.3265   0.07238   0.06777   0.0357   0.0212   1.0000
  18.250   1.3084   0.07795   0.07342   0.0338   0.0195   1.0000
  18.500   1.3010   0.08212   0.07769   0.0323   0.0201   1.0000
  18.750   1.2828   0.08793   0.08358   0.0302   0.0186   1.0000
  19.000   1.2711   0.09286   0.08859   0.0283   0.0183   1.0000
  19.250   1.2539   0.09867   0.09449   0.0260   0.0169   1.0000
<< Back to EPPLER 625 AIRFOIL (e625-il)

Polar data table (+)

Polar graphs


<< Back to EPPLER 625 AIRFOIL (e625-il)