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EH 0.0/9.0 (eh0009-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: EH 0.0/9.0 (eh0009-il)
Reynolds number: 100,000
Max Cl/Cd: 37.58 at α=4.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-eh0009-il-100000.txt
Download as CSV file: xf-eh0009-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EH 0.0/9.0                                      
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.500  -0.6404   0.11534   0.11047   0.0106   1.0000   0.1209
 -10.250  -0.6368   0.11140   0.10655   0.0096   1.0000   0.1253
 -10.000  -0.6668   0.10662   0.10191   0.0017   1.0000   0.1300
  -9.750  -0.6502   0.10184   0.09711   0.0039   1.0000   0.1331
  -9.500  -0.6516   0.06832   0.06390  -0.0217   1.0000   0.0636
  -9.250  -0.6634   0.06284   0.05840  -0.0232   1.0000   0.0623
  -8.500  -0.7828   0.05459   0.04896  -0.0211   1.0000   0.0506
  -8.250  -0.7814   0.04974   0.04384  -0.0201   1.0000   0.0491
  -8.000  -0.7777   0.04543   0.03910  -0.0187   1.0000   0.0490
  -7.750  -0.7706   0.04175   0.03486  -0.0170   1.0000   0.0504
  -7.500  -0.7590   0.03821   0.03077  -0.0152   1.0000   0.0510
  -7.250  -0.7437   0.03453   0.02658  -0.0135   1.0000   0.0510
  -7.000  -0.7250   0.03145   0.02290  -0.0119   1.0000   0.0516
  -6.750  -0.7041   0.02801   0.01912  -0.0106   1.0000   0.0536
  -6.500  -0.6822   0.02591   0.01697  -0.0097   1.0000   0.0594
  -6.250  -0.6596   0.02390   0.01477  -0.0085   1.0000   0.0673
  -6.000  -0.6366   0.02214   0.01294  -0.0072   1.0000   0.0764
  -5.750  -0.6167   0.02044   0.01137  -0.0059   1.0000   0.0934
  -5.500  -0.5985   0.01868   0.00983  -0.0040   1.0000   0.1214
  -5.250  -0.5831   0.01693   0.00859  -0.0021   1.0000   0.1897
  -5.000  -0.5671   0.01561   0.00776  -0.0003   1.0000   0.2792
  -4.750  -0.5502   0.01477   0.00728   0.0015   1.0000   0.3649
  -4.500  -0.5323   0.01416   0.00695   0.0035   1.0000   0.4377
  -4.250  -0.5135   0.01369   0.00666   0.0055   1.0000   0.5006
  -4.000  -0.4943   0.01331   0.00645   0.0076   1.0000   0.5571
  -3.750  -0.4751   0.01301   0.00628   0.0098   1.0000   0.6097
  -3.500  -0.4557   0.01278   0.00617   0.0122   1.0000   0.6567
  -3.250  -0.4366   0.01261   0.00605   0.0146   1.0000   0.7011
  -3.000  -0.4178   0.01250   0.00600   0.0173   1.0000   0.7425
  -2.750  -0.3993   0.01244   0.00599   0.0200   1.0000   0.7812
  -2.500  -0.3810   0.01243   0.00599   0.0228   1.0000   0.8185
  -2.250  -0.3601   0.01249   0.00605   0.0253   1.0000   0.8536
  -2.000  -0.3333   0.01263   0.00611   0.0265   1.0000   0.8879
  -1.750  -0.2905   0.01289   0.00626   0.0246   1.0000   0.9175
  -1.500  -0.2349   0.01314   0.00636   0.0197   1.0000   0.9417
  -1.250  -0.1638   0.01329   0.00632   0.0115   1.0000   0.9580
  -1.000  -0.0956   0.01322   0.00614   0.0033   1.0000   0.9740
  -0.750  -0.0299   0.01299   0.00582  -0.0046   1.0000   0.9895
  -0.500   0.0181   0.01268   0.00548  -0.0098   1.0000   1.0000
  -0.250   0.0136   0.01252   0.00536  -0.0057   1.0000   1.0000
   0.000   0.0000   0.01247   0.00533   0.0000   1.0000   1.0000
   0.250  -0.0136   0.01252   0.00536   0.0057   1.0000   1.0000
   0.500  -0.0181   0.01268   0.00548   0.0098   1.0000   1.0000
   0.750   0.0298   0.01299   0.00582   0.0046   0.9895   1.0000
   1.000   0.0955   0.01322   0.00614  -0.0033   0.9740   1.0000
   1.250   0.1636   0.01328   0.00632  -0.0114   0.9581   1.0000
   1.500   0.2347   0.01314   0.00636  -0.0196   0.9418   1.0000
   1.750   0.2904   0.01288   0.00625  -0.0245   0.9175   1.0000
   2.000   0.3331   0.01262   0.00610  -0.0265   0.8879   1.0000
   2.250   0.3599   0.01249   0.00605  -0.0253   0.8537   1.0000
   2.500   0.3807   0.01243   0.00599  -0.0228   0.8185   1.0000
   2.750   0.3991   0.01244   0.00599  -0.0200   0.7813   1.0000
   3.000   0.4176   0.01250   0.00600  -0.0172   0.7426   1.0000
   3.250   0.4363   0.01261   0.00604  -0.0146   0.7013   1.0000
   3.500   0.4555   0.01278   0.00617  -0.0121   0.6568   1.0000
   3.750   0.4748   0.01301   0.00627  -0.0098   0.6100   1.0000
   4.000   0.4941   0.01331   0.00645  -0.0076   0.5573   1.0000
   4.250   0.5132   0.01369   0.00666  -0.0055   0.5008   1.0000
   4.500   0.5321   0.01416   0.00695  -0.0035   0.4378   1.0000
   4.750   0.5500   0.01477   0.00728  -0.0015   0.3652   1.0000
   5.000   0.5669   0.01561   0.00776   0.0003   0.2797   1.0000
   5.250   0.5830   0.01694   0.00859   0.0022   0.1896   1.0000
   5.500   0.5985   0.01868   0.00983   0.0040   0.1215   1.0000
   5.750   0.6167   0.02044   0.01137   0.0059   0.0935   1.0000
   6.000   0.6366   0.02214   0.01294   0.0072   0.0764   1.0000
   6.250   0.6596   0.02389   0.01476   0.0085   0.0673   1.0000
   6.500   0.6823   0.02592   0.01698   0.0097   0.0595   1.0000
   6.750   0.7041   0.02805   0.01914   0.0106   0.0535   1.0000
   7.000   0.7251   0.03146   0.02291   0.0119   0.0516   1.0000
   7.250   0.7438   0.03452   0.02657   0.0135   0.0509   1.0000
   7.500   0.7591   0.03821   0.03077   0.0152   0.0510   1.0000
   7.750   0.7706   0.04171   0.03484   0.0170   0.0501   1.0000
   8.000   0.7781   0.04546   0.03912   0.0186   0.0491   1.0000
   8.250   0.7815   0.04977   0.04387   0.0201   0.0490   1.0000
   8.500   0.7831   0.05462   0.04899   0.0211   0.0506   1.0000
   8.750   0.7615   0.06175   0.05681   0.0217   0.0580   1.0000
   9.000   0.7490   0.06741   0.06262   0.0214   0.0620   1.0000
   9.250   0.5599   0.08630   0.08190   0.0016   0.1376   1.0000
   9.500   0.5514   0.09214   0.08770  -0.0008   0.1388   1.0000
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