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

EPPLER 521 AIRFOIL (e521-il) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: EPPLER 521 AIRFOIL (e521-il)
Reynolds number: 500,000
Max Cl/Cd: 56.41 at α=7.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-e521-il-500000-n5.txt
Download as CSV file: xf-e521-il-500000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 521 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
  ------ -------- --------- --------- -------- -------- --------
 -17.750  -1.0080   0.11060   0.10720  -0.0029   1.0000   0.0032
 -17.500  -1.0293   0.10197   0.09843  -0.0077   1.0000   0.0031
 -17.250  -1.0516   0.09336   0.08967  -0.0125   1.0000   0.0032
 -17.000  -1.0689   0.08596   0.08212  -0.0166   1.0000   0.0032
 -16.750  -1.0845   0.07910   0.07512  -0.0204   1.0000   0.0031
 -16.500  -1.1004   0.07242   0.06829  -0.0240   1.0000   0.0032
 -16.250  -1.1130   0.06660   0.06231  -0.0270   1.0000   0.0031
 -16.000  -1.1225   0.06154   0.05711  -0.0294   1.0000   0.0031
 -15.750  -1.1287   0.05722   0.05266  -0.0312   1.0000   0.0033
 -15.500  -1.1358   0.05302   0.04833  -0.0327   1.0000   0.0034
 -15.250  -1.1407   0.04936   0.04454  -0.0338   1.0000   0.0034
 -15.000  -1.1451   0.04601   0.04105  -0.0344   1.0000   0.0035
 -14.750  -1.1470   0.04314   0.03806  -0.0347   1.0000   0.0034
 -14.500  -1.1471   0.04059   0.03538  -0.0346   1.0000   0.0036
 -14.250  -1.1452   0.03835   0.03303  -0.0344   1.0000   0.0037
 -14.000  -1.1436   0.03618   0.03075  -0.0339   1.0000   0.0038
 -13.750  -1.1404   0.03423   0.02869  -0.0333   1.0000   0.0038
 -13.500  -1.1388   0.03221   0.02658  -0.0324   1.0000   0.0042
 -13.250  -1.1332   0.03060   0.02490  -0.0315   1.0000   0.0045
 -13.000  -1.1265   0.02911   0.02333  -0.0305   1.0000   0.0047
 -12.750  -1.1179   0.02782   0.02195  -0.0294   1.0000   0.0049
 -12.500  -1.1086   0.02660   0.02065  -0.0284   1.0000   0.0052
 -12.250  -1.0987   0.02545   0.01942  -0.0272   1.0000   0.0056
 -12.000  -1.0886   0.02433   0.01823  -0.0260   1.0000   0.0060
 -11.750  -1.0775   0.02331   0.01718  -0.0248   1.0000   0.0068
 -11.500  -1.0504   0.02231   0.01603  -0.0267   0.9233   0.0075
 -11.250  -1.0375   0.02162   0.01518  -0.0252   0.8959   0.0082
 -11.000  -1.0269   0.02089   0.01434  -0.0233   0.8810   0.0092
 -10.750  -1.0149   0.02021   0.01358  -0.0215   0.8708   0.0103
 -10.500  -1.0014   0.01962   0.01290  -0.0198   0.8633   0.0115
 -10.250  -0.9881   0.01902   0.01223  -0.0180   0.8565   0.0127
 -10.000  -0.9751   0.01844   0.01161  -0.0160   0.8507   0.0149
  -9.750  -0.9600   0.01798   0.01108  -0.0143   0.8454   0.0164
  -9.500  -0.9448   0.01744   0.01050  -0.0126   0.8410   0.0191
  -9.250  -0.9264   0.01697   0.00998  -0.0114   0.8374   0.0218
  -9.000  -0.9072   0.01648   0.00946  -0.0104   0.8337   0.0249
  -8.750  -0.8876   0.01600   0.00896  -0.0093   0.8299   0.0288
  -8.500  -0.8672   0.01555   0.00849  -0.0084   0.8266   0.0334
  -8.250  -0.8464   0.01512   0.00804  -0.0075   0.8239   0.0393
  -8.000  -0.8251   0.01465   0.00760  -0.0068   0.8210   0.0470
  -7.750  -0.8029   0.01423   0.00717  -0.0061   0.8179   0.0553
  -7.500  -0.7802   0.01386   0.00680  -0.0055   0.8149   0.0640
  -7.250  -0.7576   0.01346   0.00642  -0.0049   0.8123   0.0751
  -7.000  -0.7347   0.01309   0.00607  -0.0043   0.8101   0.0872
  -6.750  -0.7110   0.01272   0.00574  -0.0038   0.8075   0.1001
  -6.500  -0.6870   0.01237   0.00542  -0.0034   0.8047   0.1146
  -6.250  -0.6633   0.01198   0.00511  -0.0030   0.8021   0.1334
  -6.000  -0.6393   0.01162   0.00482  -0.0026   0.7996   0.1541
  -5.750  -0.6150   0.01130   0.00455  -0.0022   0.7972   0.1742
  -5.500  -0.5905   0.01097   0.00429  -0.0018   0.7950   0.1971
  -5.250  -0.5655   0.01064   0.00404  -0.0016   0.7925   0.2205
  -5.000  -0.5403   0.01033   0.00382  -0.0014   0.7899   0.2450
  -4.750  -0.5150   0.01003   0.00361  -0.0012   0.7872   0.2709
  -4.500  -0.4896   0.00974   0.00340  -0.0009   0.7847   0.2983
  -4.250  -0.4641   0.00946   0.00320  -0.0007   0.7823   0.3261
  -4.000  -0.4384   0.00920   0.00302  -0.0005   0.7800   0.3550
  -3.750  -0.4125   0.00891   0.00285  -0.0004   0.7770   0.3878
  -3.500  -0.3865   0.00863   0.00269  -0.0002   0.7738   0.4197
  -3.250  -0.3605   0.00837   0.00253   0.0000   0.7706   0.4523
  -3.000  -0.3345   0.00813   0.00239   0.0002   0.7677   0.4859
  -2.750  -0.3081   0.00790   0.00227   0.0003   0.7645   0.5191
  -2.500  -0.2814   0.00769   0.00218   0.0004   0.7606   0.5517
  -2.250  -0.2546   0.00751   0.00211   0.0006   0.7567   0.5838
  -2.000  -0.2273   0.00739   0.00206   0.0007   0.7532   0.6109
  -1.750  -0.1996   0.00731   0.00204   0.0007   0.7494   0.6332
  -1.500  -0.1713   0.00725   0.00203   0.0006   0.7446   0.6512
  -1.250  -0.1431   0.00722   0.00202   0.0006   0.7402   0.6659
  -1.000  -0.1147   0.00721   0.00200   0.0005   0.7359   0.6781
  -0.750  -0.0860   0.00719   0.00202   0.0004   0.7304   0.6885
  -0.500  -0.0574   0.00718   0.00200   0.0003   0.7252   0.6975
  -0.250  -0.0287   0.00718   0.00200   0.0001   0.7197   0.7060
   0.000   0.0000   0.00718   0.00201   0.0000   0.7129   0.7128
   0.250   0.0287   0.00718   0.00200  -0.0001   0.7060   0.7197
   0.500   0.0574   0.00718   0.00201  -0.0003   0.6975   0.7252
   0.750   0.0860   0.00719   0.00202  -0.0004   0.6884   0.7304
   1.000   0.1147   0.00721   0.00200  -0.0005   0.6782   0.7360
   1.250   0.1431   0.00722   0.00201  -0.0006   0.6658   0.7402
   1.500   0.1714   0.00725   0.00203  -0.0007   0.6512   0.7446
   1.750   0.1996   0.00731   0.00204  -0.0007   0.6334   0.7494
   2.000   0.2274   0.00739   0.00206  -0.0007   0.6108   0.7532
   2.250   0.2546   0.00751   0.00211  -0.0006   0.5838   0.7567
   2.500   0.2814   0.00769   0.00218  -0.0004   0.5520   0.7606
   2.750   0.3081   0.00790   0.00227  -0.0003   0.5192   0.7645
   3.000   0.3345   0.00812   0.00239  -0.0002   0.4862   0.7677
   3.250   0.3605   0.00837   0.00253   0.0000   0.4524   0.7706
   3.500   0.3865   0.00863   0.00269   0.0002   0.4197   0.7738
   3.750   0.4125   0.00891   0.00285   0.0004   0.3876   0.7769
   4.000   0.4383   0.00920   0.00302   0.0005   0.3543   0.7800
   4.250   0.4641   0.00946   0.00320   0.0007   0.3261   0.7823
   4.500   0.4896   0.00973   0.00340   0.0009   0.2986   0.7847
   4.750   0.5149   0.01003   0.00361   0.0012   0.2709   0.7872
   5.000   0.5402   0.01034   0.00382   0.0014   0.2448   0.7899
   5.250   0.5654   0.01064   0.00404   0.0016   0.2205   0.7925
   5.500   0.5904   0.01097   0.00429   0.0019   0.1972   0.7950
   5.750   0.6148   0.01130   0.00455   0.0022   0.1742   0.7972
   6.000   0.6392   0.01162   0.00482   0.0026   0.1540   0.7996
   6.250   0.6631   0.01199   0.00511   0.0030   0.1332   0.8021
   6.500   0.6869   0.01237   0.00542   0.0035   0.1147   0.8047
   6.750   0.7108   0.01272   0.00574   0.0039   0.1001   0.8075
   7.000   0.7345   0.01309   0.00607   0.0043   0.0871   0.8101
   7.250   0.7574   0.01346   0.00642   0.0049   0.0750   0.8124
   7.750   0.8027   0.01423   0.00717   0.0061   0.0554   0.8179
   8.000   0.8248   0.01465   0.00759   0.0068   0.0470   0.8210
   8.250   0.8461   0.01512   0.00804   0.0076   0.0394   0.8240
   8.500   0.8669   0.01555   0.00848   0.0085   0.0334   0.8267
   8.750   0.8872   0.01599   0.00896   0.0094   0.0288   0.8300
   9.000   0.9068   0.01648   0.00945   0.0104   0.0249   0.8337
   9.250   0.9260   0.01696   0.00997   0.0115   0.0217   0.8375
   9.500   0.9444   0.01744   0.01050   0.0127   0.0191   0.8411
   9.750   0.9596   0.01797   0.01107   0.0144   0.0164   0.8455
  10.000   0.9747   0.01843   0.01160   0.0161   0.0149   0.8509
  10.250   0.9881   0.01899   0.01222   0.0180   0.0132   0.8566
  10.500   1.0008   0.01963   0.01291   0.0199   0.0117   0.8635
  10.750   1.0145   0.02020   0.01358   0.0216   0.0104   0.8711
  11.000   1.0265   0.02088   0.01433   0.0234   0.0092   0.8813
  11.500   1.0500   0.02232   0.01605   0.0267   0.0077   0.9248
  11.750   1.0772   0.02330   0.01717   0.0249   0.0067   1.0000
  12.000   1.0885   0.02432   0.01823   0.0261   0.0061   1.0000
  12.250   1.0985   0.02545   0.01942   0.0273   0.0056   1.0000
  12.500   1.1087   0.02657   0.02062   0.0284   0.0054   1.0000
  12.750   1.1183   0.02776   0.02189   0.0294   0.0048   1.0000
  13.000   1.1265   0.02909   0.02330   0.0305   0.0047   1.0000
  13.250   1.1333   0.03056   0.02484   0.0315   0.0043   1.0000
  13.500   1.1390   0.03217   0.02653   0.0324   0.0041   1.0000
  13.750   1.1412   0.03413   0.02859   0.0333   0.0039   1.0000
  14.000   1.1435   0.03615   0.03072   0.0340   0.0038   1.0000
  14.250   1.1463   0.03821   0.03288   0.0344   0.0037   1.0000
  14.500   1.1472   0.04055   0.03534   0.0347   0.0036   1.0000
  14.750   1.1474   0.04307   0.03799   0.0347   0.0035   1.0000
  15.000   1.1459   0.04591   0.04095   0.0344   0.0034   1.0000
  15.250   1.1408   0.04933   0.04451   0.0338   0.0033   1.0000
  15.500   1.1378   0.05274   0.04803   0.0328   0.0034   1.0000
  15.750   1.1324   0.05668   0.05211   0.0314   0.0033   1.0000
  16.000   1.1237   0.06134   0.05690   0.0294   0.0032   1.0000
  16.250   1.1125   0.06668   0.06239   0.0269   0.0033   1.0000
  16.500   1.1018   0.07220   0.06805   0.0240   0.0033   1.0000
  16.750   1.0875   0.07857   0.07458   0.0206   0.0031   1.0000
  17.000   1.0709   0.08563   0.08178   0.0168   0.0032   1.0000
  17.250   1.0523   0.09327   0.08958   0.0125   0.0032   1.0000
  17.500   1.0323   0.10141   0.09786   0.0080   0.0032   1.0000
  17.750   1.0091   0.11042   0.10702   0.0029   0.0032   1.0000
<< Back to EPPLER 521 AIRFOIL (e521-il)

Polar data table (+)

Polar graphs


<< Back to EPPLER 521 AIRFOIL (e521-il)