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

E10(08%) (e228-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: E10(08%) (e228-il)
Reynolds number: 50,000
Max Cl/Cd: 27.01 at α=5.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e228-il-50000.txt
Download as CSV file: xf-e228-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: E10(08%)                                        
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.6618   0.09176   0.08524  -0.0040   1.0000   0.1379
  -9.250  -0.7056   0.08120   0.07477  -0.0141   1.0000   0.1257
  -9.000  -0.7125   0.07584   0.06941  -0.0151   1.0000   0.1217
  -8.750  -0.7707   0.06812   0.06122  -0.0184   1.0000   0.1136
  -8.500  -0.7640   0.06304   0.05607  -0.0181   1.0000   0.1118
  -8.250  -0.7645   0.05809   0.05088  -0.0176   1.0000   0.1099
  -8.000  -0.7652   0.05316   0.04555  -0.0167   1.0000   0.1084
  -7.750  -0.7613   0.04867   0.04059  -0.0154   1.0000   0.1081
  -7.500  -0.7537   0.04483   0.03619  -0.0138   1.0000   0.1108
  -7.250  -0.7441   0.04132   0.03194  -0.0119   1.0000   0.1139
  -7.000  -0.7234   0.03810   0.02871  -0.0110   1.0000   0.1193
  -6.750  -0.7042   0.03500   0.02509  -0.0095   1.0000   0.1252
  -6.500  -0.6828   0.03249   0.02239  -0.0083   1.0000   0.1371
  -6.250  -0.6617   0.03017   0.01978  -0.0068   1.0000   0.1549
  -6.000  -0.6396   0.02781   0.01763  -0.0054   1.0000   0.1843
  -5.750  -0.6230   0.02601   0.01621  -0.0033   1.0000   0.2345
  -5.500  -0.6096   0.02460   0.01520  -0.0005   1.0000   0.2996
  -5.250  -0.5957   0.02344   0.01435   0.0025   1.0000   0.3668
  -5.000  -0.5801   0.02251   0.01377   0.0056   1.0000   0.4297
  -4.750  -0.5641   0.02177   0.01329   0.0088   1.0000   0.4906
  -4.500  -0.5476   0.02118   0.01291   0.0122   1.0000   0.5494
  -4.250  -0.5304   0.02074   0.01266   0.0158   1.0000   0.6061
  -4.000  -0.5120   0.02046   0.01250   0.0194   1.0000   0.6606
  -3.750  -0.4926   0.02032   0.01242   0.0231   1.0000   0.7139
  -3.500  -0.4677   0.02044   0.01249   0.0263   1.0000   0.7651
  -3.250  -0.4303   0.02084   0.01276   0.0276   1.0000   0.8158
  -3.000  -0.3595   0.02163   0.01314   0.0231   1.0000   0.8655
  -2.750  -0.2553   0.02213   0.01307   0.0112   1.0000   0.9104
  -2.500  -0.1592   0.02169   0.01217  -0.0014   1.0000   0.9498
  -2.250  -0.0707   0.02055   0.01069  -0.0142   1.0000   0.9858
  -2.000  -0.0259   0.01949   0.00950  -0.0200   1.0000   1.0000
  -1.750  -0.0161   0.01894   0.00893  -0.0192   1.0000   1.0000
  -1.500  -0.0080   0.01851   0.00850  -0.0179   1.0000   1.0000
  -1.250  -0.0033   0.01821   0.00821  -0.0159   1.0000   1.0000
  -1.000  -0.0045   0.01804   0.00806  -0.0128   1.0000   1.0000
  -0.750  -0.0112   0.01799   0.00802  -0.0087   1.0000   1.0000
  -0.500  -0.0205   0.01804   0.00803  -0.0040   1.0000   1.0000
  -0.250  -0.0283   0.01816   0.00810   0.0004   1.0000   1.0000
   0.000  -0.0322   0.01834   0.00822   0.0042   1.0000   1.0000
   0.250  -0.0323   0.01859   0.00840   0.0073   1.0000   1.0000
   0.500  -0.0293   0.01890   0.00865   0.0099   1.0000   1.0000
   0.750  -0.0244   0.01928   0.00897   0.0122   1.0000   1.0000
   1.000  -0.0177   0.01973   0.00937   0.0140   1.0000   1.0000
   1.250  -0.0095   0.02024   0.00984   0.0155   1.0000   1.0000
   1.500   0.0647   0.02142   0.01113   0.0047   0.9762   1.0000
   1.750   0.1366   0.02240   0.01227  -0.0051   0.9515   1.0000
   2.000   0.2107   0.02314   0.01320  -0.0146   0.9256   1.0000
   2.250   0.2935   0.02351   0.01388  -0.0248   0.8987   1.0000
   2.500   0.3589   0.02357   0.01420  -0.0310   0.8698   1.0000
   2.750   0.4067   0.02356   0.01441  -0.0334   0.8398   1.0000
   3.000   0.4423   0.02355   0.01459  -0.0333   0.8097   1.0000
   3.250   0.4701   0.02354   0.01470  -0.0315   0.7796   1.0000
   3.500   0.4935   0.02351   0.01478  -0.0288   0.7493   1.0000
   3.750   0.5147   0.02346   0.01484  -0.0255   0.7186   1.0000
   4.000   0.5349   0.02335   0.01479  -0.0219   0.6872   1.0000
   4.250   0.5545   0.02324   0.01470  -0.0181   0.6548   1.0000
   4.500   0.5726   0.02331   0.01480  -0.0146   0.6180   1.0000
   4.750   0.5916   0.02330   0.01476  -0.0110   0.5803   1.0000
   5.000   0.6108   0.02336   0.01477  -0.0075   0.5398   1.0000
   5.250   0.6292   0.02363   0.01496  -0.0043   0.4945   1.0000
   5.500   0.6473   0.02403   0.01521  -0.0012   0.4461   1.0000
   5.750   0.6653   0.02463   0.01558   0.0017   0.3957   1.0000
   6.000   0.6826   0.02552   0.01623   0.0043   0.3442   1.0000
   6.250   0.7000   0.02671   0.01718   0.0067   0.2959   1.0000
   6.500   0.7173   0.02813   0.01845   0.0088   0.2528   1.0000
   6.750   0.7354   0.02975   0.01990   0.0107   0.2161   1.0000
   7.000   0.7545   0.03176   0.02183   0.0123   0.1874   1.0000
   7.250   0.7719   0.03397   0.02421   0.0141   0.1635   1.0000
   7.500   0.7915   0.03641   0.02654   0.0154   0.1451   1.0000
   7.750   0.8086   0.03921   0.02956   0.0169   0.1315   1.0000
   8.000   0.8193   0.04251   0.03348   0.0189   0.1223   1.0000
   8.250   0.8304   0.04584   0.03713   0.0204   0.1140   1.0000
   8.500   0.8344   0.05008   0.04192   0.0222   0.1112   1.0000
   8.750   0.8375   0.05402   0.04621   0.0236   0.1079   1.0000
   9.000   0.8472   0.05769   0.04989   0.0246   0.1029   1.0000
   9.250   0.8377   0.06231   0.05492   0.0258   0.1025   1.0000
   9.500   0.8271   0.06703   0.05990   0.0265   0.1028   1.0000
   9.750   0.8154   0.07192   0.06495   0.0268   0.1033   1.0000
  10.000   0.7173   0.08439   0.07761   0.0175   0.1244   1.0000
  10.250   0.7021   0.09237   0.08551   0.0128   0.1286   1.0000
<< Back to E10(08%) (e228-il)

Polar data table (+)

Polar graphs


<< Back to E10(08%) (e228-il)