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EH 1.5/9.0 (eh1590-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: EH 1.5/9.0 (eh1590-il)
Reynolds number: 100,000
Max Cl/Cd: 45.6 at α=7°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-eh1590-il-100000.txt
Download as CSV file: xf-eh1590-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EH 1.5/9.0                                      
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.5552   0.08876   0.08429  -0.0023   1.0000   0.1017
  -8.000  -0.5732   0.08445   0.08007  -0.0081   1.0000   0.1046
  -7.750  -0.6075   0.08139   0.07675  -0.0155   1.0000   0.1064
  -7.500  -0.5758   0.07558   0.07122  -0.0113   1.0000   0.1098
  -7.250  -0.5703   0.07204   0.06766  -0.0117   1.0000   0.1148
  -7.000  -0.5823   0.06752   0.06290  -0.0157   1.0000   0.1213
  -6.750  -0.5641   0.06423   0.05974  -0.0137   1.0000   0.1279
  -6.500  -0.5604   0.06033   0.05573  -0.0145   1.0000   0.1378
  -6.250  -0.5536   0.05693   0.05221  -0.0147   1.0000   0.1505
  -6.000  -0.5441   0.05368   0.04884  -0.0144   1.0000   0.1638
  -5.750  -0.5327   0.05047   0.04548  -0.0138   1.0000   0.1775
  -5.500  -0.5082   0.03639   0.02940  -0.0133   1.0000   0.0649
  -5.250  -0.4902   0.03245   0.02518  -0.0120   1.0000   0.0618
  -5.000  -0.4701   0.02902   0.02125  -0.0103   1.0000   0.0595
  -4.750  -0.4482   0.02619   0.01791  -0.0086   1.0000   0.0592
  -4.500  -0.4249   0.02394   0.01528  -0.0072   1.0000   0.0611
  -4.250  -0.4014   0.02235   0.01332  -0.0058   1.0000   0.0668
  -4.000  -0.3789   0.02062   0.01162  -0.0047   1.0000   0.0754
  -3.750  -0.3565   0.01905   0.01005  -0.0033   1.0000   0.0868
  -3.500  -0.3356   0.01780   0.00880  -0.0018   1.0000   0.1073
  -3.250  -0.3177   0.01659   0.00785  -0.0001   1.0000   0.1378
  -3.000  -0.3019   0.01549   0.00700   0.0019   1.0000   0.1799
  -2.750  -0.2892   0.01428   0.00634   0.0042   1.0000   0.2551
  -2.500  -0.2873   0.01186   0.00655   0.0104   1.0000   0.7737
  -2.250  -0.1005   0.01311   0.00700  -0.0125   1.0000   0.9793
  -2.000  -0.0194   0.01269   0.00621  -0.0233   1.0000   1.0000
  -1.750  -0.0042   0.01255   0.00603  -0.0221   1.0000   1.0000
  -1.500   0.0017   0.01259   0.00606  -0.0196   1.0000   1.0000
  -1.250   0.0122   0.01287   0.00632  -0.0190   0.9898   1.0000
  -1.000   0.0798   0.01275   0.00604  -0.0273   0.9663   1.0000
  -0.750   0.1394   0.01258   0.00572  -0.0339   0.9365   1.0000
  -0.500   0.1826   0.01246   0.00546  -0.0368   0.9020   1.0000
  -0.250   0.2079   0.01247   0.00531  -0.0360   0.8679   1.0000
   0.000   0.2276   0.01254   0.00524  -0.0341   0.8378   1.0000
   0.250   0.2467   0.01263   0.00520  -0.0322   0.8114   1.0000
   0.500   0.2663   0.01274   0.00516  -0.0303   0.7883   1.0000
   0.750   0.2863   0.01287   0.00519  -0.0287   0.7650   1.0000
   1.000   0.3064   0.01302   0.00521  -0.0270   0.7447   1.0000
   1.250   0.3272   0.01319   0.00531  -0.0255   0.7243   1.0000
   1.500   0.3481   0.01337   0.00537  -0.0240   0.7061   1.0000
   1.750   0.3695   0.01356   0.00550  -0.0227   0.6877   1.0000
   2.000   0.3912   0.01377   0.00565  -0.0215   0.6698   1.0000
   2.250   0.4131   0.01398   0.00580  -0.0202   0.6530   1.0000
   2.500   0.4351   0.01421   0.00599  -0.0190   0.6369   1.0000
   2.750   0.4574   0.01445   0.00618  -0.0178   0.6214   1.0000
   3.000   0.4798   0.01471   0.00643  -0.0167   0.6050   1.0000
   3.250   0.5024   0.01498   0.00670  -0.0156   0.5889   1.0000
   3.500   0.5250   0.01526   0.00698  -0.0146   0.5732   1.0000
   3.750   0.5478   0.01555   0.00731  -0.0135   0.5579   1.0000
   4.000   0.5707   0.01585   0.00762  -0.0125   0.5426   1.0000
   4.250   0.5937   0.01615   0.00794  -0.0114   0.5275   1.0000
   4.500   0.6168   0.01647   0.00827  -0.0104   0.5124   1.0000
   4.750   0.6399   0.01679   0.00861  -0.0094   0.4973   1.0000
   5.000   0.6630   0.01712   0.00902  -0.0083   0.4819   1.0000
   5.250   0.6858   0.01747   0.00948  -0.0073   0.4650   1.0000
   5.500   0.7087   0.01780   0.00990  -0.0063   0.4480   1.0000
   5.750   0.7318   0.01811   0.01028  -0.0052   0.4307   1.0000
   6.000   0.7551   0.01840   0.01058  -0.0041   0.4134   1.0000
   6.250   0.7773   0.01869   0.01109  -0.0030   0.3926   1.0000
   6.500   0.7989   0.01872   0.01117  -0.0016   0.3676   1.0000
   6.750   0.8176   0.01834   0.01078   0.0002   0.3281   1.0000
   7.000   0.8368   0.01835   0.01088   0.0016   0.2847   1.0000
   7.250   0.8509   0.01915   0.01136   0.0034   0.1789   1.0000
   7.500   0.8509   0.02276   0.01389   0.0063   0.0753   1.0000
   7.750   0.8623   0.02476   0.01587   0.0085   0.0621   1.0000
   8.000   0.8771   0.02648   0.01765   0.0105   0.0551   1.0000
   8.250   0.8923   0.02869   0.01985   0.0122   0.0501   1.0000
   8.500   0.9110   0.03033   0.02166   0.0135   0.0450   1.0000
   8.750   0.9310   0.03263   0.02400   0.0147   0.0424   1.0000
   9.000   0.9521   0.03579   0.02725   0.0156   0.0410   1.0000
   9.250   0.9709   0.03953   0.03125   0.0165   0.0404   1.0000
   9.500   0.9852   0.04306   0.03515   0.0179   0.0404   1.0000
   9.750   0.9949   0.04661   0.03910   0.0193   0.0404   1.0000
  10.000   0.9986   0.05077   0.04365   0.0207   0.0402   1.0000
  10.250   0.9963   0.05529   0.04855   0.0221   0.0400   1.0000
  10.500   0.9910   0.05872   0.05229   0.0236   0.0399   1.0000
  10.750   0.9790   0.06166   0.05555   0.0252   0.0396   1.0000
  11.000   0.9709   0.06561   0.05971   0.0264   0.0401   1.0000
  11.250   0.9563   0.06834   0.06264   0.0272   0.0403   1.0000
  11.500   0.9367   0.07109   0.06566   0.0263   0.0412   1.0000
  11.750   0.9113   0.07639   0.07116   0.0229   0.0416   1.0000
  12.000   0.8382   0.09158   0.08672   0.0093   0.0478   1.0000
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