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

EPPLER 583 AIRFOIL (e583-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: EPPLER 583 AIRFOIL (e583-il)
Reynolds number: 500,000
Max Cl/Cd: 111.51 at α=7.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e583-il-500000.txt
Download as CSV file: xf-e583-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 583 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -11.500  -0.0489   0.09714   0.09391  -0.1160   0.8147   0.0168
 -11.250  -0.0632   0.08949   0.08626  -0.1203   0.8051   0.0178
 -11.000  -0.0676   0.08446   0.08119  -0.1226   0.7964   0.0179
 -10.750  -0.0631   0.08222   0.07892  -0.1227   0.7878   0.0181
 -10.500  -0.0596   0.07962   0.07627  -0.1233   0.7804   0.0182
 -10.250  -0.0584   0.07641   0.07305  -0.1245   0.7734   0.0183
 -10.000  -0.0556   0.07379   0.07040  -0.1254   0.7672   0.0187
  -9.750  -0.0597   0.06948   0.06608  -0.1276   0.7618   0.0188
  -9.500  -0.0663   0.06475   0.06138  -0.1305   0.7563   0.0191
  -9.250  -0.0905   0.05725   0.05381  -0.1360   0.7510   0.0189
  -9.000  -0.1113   0.05288   0.04934  -0.1367   0.7460   0.0189
  -7.500  -0.2113   0.02576   0.02004  -0.1145   0.7214   0.0120
  -7.250  -0.1931   0.02401   0.01814  -0.1134   0.7183   0.0112
  -7.000  -0.1776   0.02155   0.01530  -0.1117   0.7154   0.0107
  -6.750  -0.1582   0.01953   0.01296  -0.1103   0.7126   0.0104
  -6.500  -0.1360   0.01825   0.01146  -0.1094   0.7096   0.0108
  -6.250  -0.1125   0.01726   0.01027  -0.1087   0.7065   0.0112
  -6.000  -0.0883   0.01648   0.00934  -0.1081   0.7036   0.0115
  -5.750  -0.0641   0.01578   0.00850  -0.1076   0.7009   0.0118
  -5.500  -0.0412   0.01500   0.00769  -0.1070   0.6983   0.0124
  -5.250  -0.0181   0.01447   0.00715  -0.1063   0.6958   0.0128
  -5.000   0.0056   0.01403   0.00669  -0.1057   0.6930   0.0136
  -4.750   0.0299   0.01366   0.00627  -0.1052   0.6902   0.0148
  -4.500   0.0536   0.01323   0.00584  -0.1047   0.6877   0.0163
  -4.250   0.0784   0.01290   0.00544  -0.1043   0.6853   0.0179
  -4.000   0.1031   0.01252   0.00502  -0.1039   0.6827   0.0210
  -3.750   0.1267   0.01207   0.00462  -0.1033   0.6804   0.0316
  -3.500   0.1499   0.01159   0.00435  -0.1027   0.6779   0.0687
  -3.250   0.1739   0.01119   0.00417  -0.1024   0.6753   0.1216
  -3.000   0.1971   0.01067   0.00400  -0.1021   0.6728   0.2135
  -2.750   0.2181   0.00984   0.00378  -0.1017   0.6705   0.3875
  -2.500   0.2351   0.00875   0.00378  -0.1002   0.6682   0.6861
  -2.250   0.2624   0.00897   0.00402  -0.0999   0.6657   0.7438
  -2.000   0.2892   0.00912   0.00417  -0.0995   0.6635   0.7640
  -1.750   0.3163   0.00928   0.00430  -0.0992   0.6611   0.7780
  -1.500   0.3438   0.00946   0.00443  -0.0990   0.6585   0.7907
  -1.250   0.3701   0.00966   0.00461  -0.0984   0.6560   0.7992
  -1.000   0.3981   0.00981   0.00468  -0.0984   0.6536   0.8079
  -0.750   0.4247   0.01000   0.00483  -0.0980   0.6513   0.8134
  -0.500   0.4522   0.01024   0.00502  -0.0979   0.6489   0.8227
  -0.250   0.4728   0.01052   0.00538  -0.0956   0.6464   0.8332
   0.000   0.4959   0.01072   0.00558  -0.0943   0.6437   0.8427
   0.250   0.5210   0.01081   0.00565  -0.0936   0.6409   0.8469
   0.500   0.5491   0.01083   0.00563  -0.0939   0.6384   0.8498
   0.750   0.5785   0.01085   0.00559  -0.0945   0.6359   0.8524
   1.000   0.6089   0.01095   0.00561  -0.0954   0.6333   0.8547
   1.250   0.6355   0.01090   0.00558  -0.0954   0.6306   0.8564
   1.500   0.6616   0.01088   0.00557  -0.0953   0.6276   0.8580
   1.750   0.6888   0.01087   0.00555  -0.0953   0.6246   0.8595
   2.000   0.7170   0.01087   0.00552  -0.0957   0.6218   0.8610
   2.250   0.7457   0.01091   0.00552  -0.0961   0.6191   0.8626
   2.500   0.7734   0.01097   0.00558  -0.0964   0.6161   0.8645
   2.750   0.8000   0.01095   0.00559  -0.0965   0.6127   0.8663
   3.000   0.8278   0.01094   0.00559  -0.0968   0.6090   0.8680
   3.250   0.8567   0.01094   0.00556  -0.0973   0.6056   0.8696
   3.500   0.8869   0.01101   0.00557  -0.0982   0.6022   0.8713
   3.750   0.9114   0.01098   0.00561  -0.0978   0.5985   0.8728
   4.000   0.9365   0.01096   0.00562  -0.0975   0.5942   0.8743
   4.250   0.9630   0.01095   0.00562  -0.0975   0.5903   0.8757
   4.500   0.9910   0.01099   0.00562  -0.0978   0.5866   0.8772
   4.750   1.0155   0.01100   0.00571  -0.0975   0.5822   0.8790
   5.000   1.0407   0.01100   0.00575  -0.0973   0.5773   0.8809
   5.250   1.0672   0.01101   0.00575  -0.0973   0.5728   0.8830
   5.500   1.0935   0.01107   0.00583  -0.0974   0.5680   0.8851
   5.750   1.1181   0.01109   0.00591  -0.0971   0.5623   0.8869
   6.000   1.1421   0.01108   0.00591  -0.0966   0.5569   0.8886
   6.250   1.1647   0.01111   0.00600  -0.0959   0.5510   0.8904
   6.500   1.1865   0.01113   0.00607  -0.0950   0.5444   0.8925
   6.750   1.2095   0.01120   0.00614  -0.0943   0.5382   0.8947
   7.000   1.2302   0.01126   0.00627  -0.0933   0.5302   0.8972
   7.250   1.2514   0.01136   0.00637  -0.0924   0.5225   0.9000
   7.500   1.2713   0.01146   0.00652  -0.0912   0.5134   0.9026
   7.750   1.2868   0.01154   0.00664  -0.0891   0.5049   0.9054
   8.000   1.3003   0.01168   0.00680  -0.0866   0.4953   0.9086
   8.250   1.3151   0.01188   0.00703  -0.0845   0.4851   0.9121
   8.500   1.3290   0.01215   0.00731  -0.0823   0.4741   0.9159
   8.750   1.3395   0.01246   0.00762  -0.0795   0.4622   0.9198
   9.000   1.3487   0.01281   0.00798  -0.0765   0.4492   0.9247
   9.250   1.3583   0.01323   0.00841  -0.0738   0.4359   0.9304
   9.500   1.3648   0.01368   0.00889  -0.0705   0.4230   0.9370
   9.750   1.3713   0.01423   0.00945  -0.0674   0.4099   0.9449
  10.000   1.3749   0.01486   0.01011  -0.0640   0.3960   0.9571
  10.250   1.3875   0.01568   0.01094  -0.0630   0.3794   0.9891
  10.500   1.3955   0.01674   0.01196  -0.0613   0.3631   1.0000
  10.750   1.4015   0.01796   0.01315  -0.0595   0.3462   1.0000
  11.000   1.4044   0.01940   0.01453  -0.0574   0.3283   1.0000
  11.250   1.4070   0.02092   0.01601  -0.0554   0.3116   1.0000
  11.500   1.4073   0.02264   0.01767  -0.0533   0.2947   1.0000
  12.000   1.4071   0.02633   0.02126  -0.0495   0.2612   1.0000
  12.250   1.4076   0.02823   0.02312  -0.0478   0.2457   1.0000
  12.500   1.4079   0.03020   0.02505  -0.0462   0.2306   1.0000
  12.750   1.4082   0.03224   0.02705  -0.0447   0.2160   1.0000
  13.000   1.4088   0.03431   0.02909  -0.0433   0.2027   1.0000
  13.250   1.4095   0.03643   0.03117  -0.0421   0.1888   1.0000
  13.500   1.4108   0.03857   0.03329  -0.0410   0.1755   1.0000
  13.750   1.4113   0.04084   0.03553  -0.0399   0.1622   1.0000
  14.000   1.4124   0.04311   0.03778  -0.0390   0.1497   1.0000
  14.250   1.4130   0.04549   0.04013  -0.0382   0.1373   1.0000
  14.500   1.4126   0.04802   0.04263  -0.0374   0.1251   1.0000
  14.750   1.4125   0.05058   0.04517  -0.0368   0.1136   1.0000
  15.000   1.4123   0.05322   0.04779  -0.0363   0.1027   1.0000
  15.250   1.4108   0.05608   0.05061  -0.0358   0.0916   1.0000
  15.500   1.4100   0.05892   0.05343  -0.0355   0.0823   1.0000
  15.750   1.4101   0.06172   0.05624  -0.0353   0.0738   1.0000
  16.000   1.4099   0.06464   0.05917  -0.0352   0.0659   1.0000
  16.250   1.4093   0.06766   0.06219  -0.0352   0.0588   1.0000
  16.500   1.4064   0.07103   0.06556  -0.0353   0.0519   1.0000
  16.750   1.4050   0.07429   0.06883  -0.0355   0.0459   1.0000
  17.000   1.4044   0.07750   0.07208  -0.0358   0.0407   1.0000
  17.250   1.4015   0.08110   0.07570  -0.0363   0.0356   1.0000
  17.500   1.3979   0.08486   0.07947  -0.0369   0.0315   1.0000
  17.750   1.3973   0.08827   0.08294  -0.0376   0.0281   1.0000
  18.000   1.3941   0.09210   0.08683  -0.0384   0.0255   1.0000
  18.250   1.3910   0.09598   0.09075  -0.0393   0.0226   1.0000
  18.500   1.3893   0.09971   0.09456  -0.0404   0.0203   1.0000
  18.750   1.3853   0.10384   0.09874  -0.0416   0.0185   1.0000
  19.000   1.3833   0.10771   0.10269  -0.0428   0.0168   1.0000
<< Back to EPPLER 583 AIRFOIL (e583-il)

Polar data table (+)

Polar graphs


<< Back to EPPLER 583 AIRFOIL (e583-il)