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E230 (9.96%) (e230-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: E230 (9.96%) (e230-il)
Reynolds number: 500,000
Max Cl/Cd: 50.38 at α=4.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e230-il-500000.txt
Download as CSV file: xf-e230-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: E230  (9.96%)                                   
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -1.0547   0.03144   0.02622   0.0169   1.0000   0.0121
  -9.250  -1.0382   0.02979   0.02436   0.0183   1.0000   0.0120
  -9.000  -1.0207   0.02799   0.02236   0.0195   1.0000   0.0119
  -8.750  -1.0020   0.02614   0.02031   0.0207   1.0000   0.0119
  -8.500  -0.9834   0.02385   0.01783   0.0218   1.0000   0.0120
  -8.250  -0.9650   0.02163   0.01547   0.0230   1.0000   0.0123
  -8.000  -0.9451   0.02014   0.01389   0.0241   1.0000   0.0126
  -7.750  -0.9243   0.01893   0.01258   0.0252   1.0000   0.0126
  -7.500  -0.9038   0.01776   0.01133   0.0263   1.0000   0.0128
  -7.250  -0.8828   0.01675   0.01025   0.0274   1.0000   0.0131
  -7.000  -0.8611   0.01588   0.00931   0.0284   1.0000   0.0134
  -6.750  -0.8386   0.01509   0.00846   0.0293   1.0000   0.0139
  -6.500  -0.8155   0.01437   0.00768   0.0301   1.0000   0.0146
  -6.250  -0.7916   0.01375   0.00698   0.0308   1.0000   0.0154
  -6.000  -0.7671   0.01319   0.00635   0.0313   1.0000   0.0160
  -5.750  -0.7433   0.01251   0.00561   0.0320   1.0000   0.0179
  -5.500  -0.7179   0.01205   0.00512   0.0325   1.0000   0.0212
  -5.250  -0.6966   0.01103   0.00440   0.0335   1.0000   0.0627
  -5.000  -0.6740   0.01027   0.00395   0.0341   1.0000   0.1189
  -4.750  -0.6501   0.00975   0.00364   0.0345   1.0000   0.1624
  -4.500  -0.6271   0.00929   0.00338   0.0352   1.0000   0.2048
  -4.250  -0.5942   0.00882   0.00313   0.0337   0.9870   0.2571
  -4.000  -0.5598   0.00841   0.00292   0.0320   0.9745   0.3102
  -3.750  -0.5310   0.00804   0.00276   0.0316   0.9601   0.3635
  -3.500  -0.5089   0.00777   0.00265   0.0329   0.9443   0.4099
  -3.250  -0.4885   0.00753   0.00253   0.0347   0.9291   0.4539
  -3.000  -0.4674   0.00732   0.00242   0.0363   0.9150   0.4913
  -2.750  -0.4455   0.00711   0.00231   0.0377   0.9017   0.5282
  -2.500  -0.4229   0.00691   0.00219   0.0390   0.8885   0.5655
  -2.250  -0.3998   0.00671   0.00208   0.0402   0.8752   0.6046
  -2.000  -0.3765   0.00651   0.00199   0.0414   0.8619   0.6430
  -1.750  -0.3526   0.00634   0.00191   0.0425   0.8485   0.6800
  -1.500  -0.3286   0.00618   0.00183   0.0435   0.8349   0.7151
  -1.250  -0.3040   0.00604   0.00177   0.0445   0.8209   0.7499
  -0.750  -0.2533   0.00582   0.00170   0.0462   0.7922   0.8150
  -0.500  -0.2265   0.00576   0.00168   0.0468   0.7772   0.8447
  -0.250  -0.1982   0.00574   0.00168   0.0470   0.7617   0.8718
   0.000  -0.1666   0.00576   0.00172   0.0466   0.7455   0.8936
   0.250  -0.1330   0.00584   0.00177   0.0456   0.7285   0.9114
   0.500  -0.0975   0.00596   0.00184   0.0443   0.7105   0.9245
   0.750  -0.0621   0.00609   0.00192   0.0429   0.6910   0.9350
   1.000  -0.0300   0.00624   0.00200   0.0423   0.6705   0.9452
   1.250   0.0089   0.00643   0.00211   0.0401   0.6482   0.9500
   1.500   0.0401   0.00661   0.00222   0.0396   0.6251   0.9585
   1.750   0.0800   0.00679   0.00232   0.0372   0.6003   0.9615
   2.000   0.1175   0.00698   0.00242   0.0352   0.5732   0.9661
   2.500   0.1900   0.00735   0.00262   0.0318   0.5153   0.9753
   2.750   0.2274   0.00754   0.00272   0.0297   0.4835   0.9796
   3.000   0.2617   0.00776   0.00284   0.0283   0.4491   0.9845
   3.250   0.3011   0.00796   0.00291   0.0257   0.4056   0.9871
   3.500   0.3387   0.00821   0.00299   0.0234   0.3577   0.9904
   3.750   0.3747   0.00851   0.00311   0.0214   0.3086   0.9940
   4.000   0.4120   0.00882   0.00320   0.0190   0.2521   0.9965
   4.250   0.4494   0.00914   0.00333   0.0166   0.2022   0.9993
   4.500   0.4776   0.00948   0.00351   0.0161   0.1589   1.0000
   5.250   0.5491   0.01110   0.00446   0.0172   0.0303   1.0000
   5.500   0.5730   0.01157   0.00492   0.0178   0.0227   1.0000
   5.750   0.5965   0.01209   0.00549   0.0185   0.0203   1.0000
   6.000   0.6200   0.01253   0.00600   0.0192   0.0192   1.0000
   6.250   0.6429   0.01304   0.00657   0.0200   0.0183   1.0000
   6.500   0.6654   0.01363   0.00724   0.0208   0.0175   1.0000
   6.750   0.6872   0.01433   0.00799   0.0217   0.0168   1.0000
   7.000   0.7080   0.01516   0.00889   0.0227   0.0164   1.0000
   7.250   0.7275   0.01622   0.01002   0.0239   0.0159   1.0000
   7.500   0.7447   0.01781   0.01172   0.0253   0.0154   1.0000
   7.750   0.7653   0.01857   0.01258   0.0265   0.0151   1.0000
   8.000   0.7843   0.01974   0.01388   0.0278   0.0150   1.0000
   8.250   0.8029   0.02097   0.01523   0.0291   0.0149   1.0000
   8.500   0.8208   0.02233   0.01675   0.0305   0.0147   1.0000
   8.750   0.8373   0.02393   0.01853   0.0320   0.0147   1.0000
   9.000   0.8521   0.02580   0.02061   0.0337   0.0146   1.0000
   9.250   0.8642   0.02802   0.02310   0.0355   0.0146   1.0000
   9.500   0.8720   0.03077   0.02614   0.0376   0.0148   1.0000
   9.750   0.8730   0.03545   0.03106   0.0397   0.0155   1.0000
  10.000   0.8866   0.03618   0.03195   0.0414   0.0158   1.0000
  11.750   0.6302   0.10286   0.10050   0.0180   0.0286   1.0000
  12.000   0.6297   0.10693   0.10456   0.0165   0.0283   1.0000
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