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E184 (8.33%) (e184-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: E184 (8.33%) (e184-il)
Reynolds number: 50,000
Max Cl/Cd: 30.86 at α=6°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e184-il-50000.txt
Download as CSV file: xf-e184-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: E184  (8.33%)                                   
  
 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
  ------ -------- --------- --------- -------- -------- --------
 -10.750  -0.6625   0.14376   0.13753   0.0399   1.0000   0.1802
 -10.500  -0.6333   0.13638   0.13008   0.0419   1.0000   0.1899
 -10.250  -0.6529   0.13563   0.12946   0.0374   1.0000   0.1942
 -10.000  -0.6299   0.12946   0.12325   0.0388   1.0000   0.2063
  -9.750  -0.6189   0.12486   0.11866   0.0387   1.0000   0.2164
  -9.500  -0.6390   0.12373   0.11766   0.0346   1.0000   0.2222
  -9.250  -0.6254   0.11896   0.11288   0.0351   1.0000   0.2357
  -9.000  -0.6101   0.11404   0.10796   0.0355   1.0000   0.2476
  -8.750  -0.6018   0.10981   0.10376   0.0351   1.0000   0.2576
  -8.500  -0.6008   0.10620   0.10021   0.0340   1.0000   0.2686
  -8.000  -0.6002   0.09939   0.09354   0.0323   1.0000   0.2946
  -7.500  -0.5838   0.09168   0.08586   0.0325   1.0000   0.3247
  -7.250  -0.5750   0.08793   0.08214   0.0327   1.0000   0.3413
  -7.000  -0.5778   0.08504   0.07934   0.0326   1.0000   0.3616
  -6.750  -0.5702   0.08121   0.07557   0.0332   1.0000   0.3803
  -6.000  -0.5829   0.05184   0.04496  -0.0079   1.0000   0.1351
  -5.750  -0.5656   0.04628   0.03878  -0.0087   1.0000   0.1213
  -5.500  -0.5468   0.04143   0.03323  -0.0087   1.0000   0.1132
  -5.250  -0.5246   0.03794   0.02898  -0.0081   1.0000   0.1102
  -5.000  -0.5015   0.03473   0.02552  -0.0079   1.0000   0.1136
  -4.750  -0.4766   0.03205   0.02254  -0.0076   1.0000   0.1173
  -4.500  -0.4500   0.02940   0.01944  -0.0071   1.0000   0.1198
  -4.250  -0.4224   0.02715   0.01666  -0.0065   1.0000   0.1246
  -4.000  -0.3954   0.02497   0.01450  -0.0062   1.0000   0.1362
  -3.750  -0.3714   0.02319   0.01283  -0.0055   1.0000   0.1623
  -3.500  -0.3473   0.02094   0.01097  -0.0046   1.0000   0.2331
  -3.250  -0.1651   0.01699   0.00929  -0.0211   1.0000   1.0000
  -3.000  -0.1450   0.01650   0.00843  -0.0222   1.0000   1.0000
  -2.750  -0.1238   0.01621   0.00786  -0.0230   1.0000   1.0000
  -2.500  -0.1048   0.01609   0.00748  -0.0233   1.0000   1.0000
  -2.250  -0.0930   0.01612   0.00732  -0.0223   1.0000   1.0000
  -2.000  -0.0932   0.01634   0.00736  -0.0191   1.0000   1.0000
  -1.750  -0.0941   0.01660   0.00740  -0.0157   1.0000   1.0000
  -1.500  -0.0921   0.01686   0.00747  -0.0126   1.0000   1.0000
  -1.250  -0.0873   0.01713   0.00756  -0.0099   1.0000   1.0000
  -1.000  -0.0801   0.01742   0.00766  -0.0076   1.0000   1.0000
  -0.750  -0.0713   0.01772   0.00779  -0.0055   1.0000   1.0000
  -0.500  -0.0612   0.01804   0.00796  -0.0037   1.0000   1.0000
  -0.250  -0.0498   0.01839   0.00815  -0.0021   1.0000   1.0000
   0.000  -0.0353   0.01877   0.00841  -0.0011   0.9992   1.0000
   0.250   0.0087   0.01942   0.00891  -0.0057   0.9883   1.0000
   0.500   0.0518   0.02011   0.00951  -0.0101   0.9775   1.0000
   0.750   0.0950   0.02083   0.01018  -0.0144   0.9668   1.0000
   1.000   0.1398   0.02159   0.01092  -0.0190   0.9559   1.0000
   1.250   0.1783   0.02231   0.01164  -0.0222   0.9445   1.0000
   1.500   0.2146   0.02306   0.01241  -0.0250   0.9326   1.0000
   1.750   0.2515   0.02384   0.01326  -0.0279   0.9203   1.0000
   2.000   0.2892   0.02463   0.01415  -0.0307   0.9074   1.0000
   2.250   0.3272   0.02544   0.01506  -0.0334   0.8936   1.0000
   2.500   0.3660   0.02623   0.01604  -0.0362   0.8788   1.0000
   2.750   0.4038   0.02700   0.01698  -0.0383   0.8630   1.0000
   3.000   0.4432   0.02774   0.01792  -0.0403   0.8461   1.0000
   3.250   0.4648   0.02855   0.01888  -0.0394   0.8271   1.0000
   3.500   0.4990   0.02914   0.01974  -0.0397   0.8068   1.0000
   3.750   0.5231   0.02978   0.02058  -0.0384   0.7849   1.0000
   4.000   0.5549   0.02998   0.02103  -0.0367   0.7619   1.0000
   4.250   0.5770   0.03020   0.02148  -0.0338   0.7362   1.0000
   4.500   0.5985   0.03012   0.02167  -0.0299   0.7089   1.0000
   4.750   0.6194   0.02962   0.02141  -0.0251   0.6801   1.0000
   5.000   0.6395   0.02858   0.02058  -0.0192   0.6501   1.0000
   5.500   0.6745   0.02557   0.01796  -0.0064   0.5687   1.0000
   5.750   0.6908   0.02327   0.01576   0.0010   0.4992   1.0000
   6.000   0.7006   0.02270   0.01417   0.0077   0.3448   1.0000
   6.250   0.7074   0.02598   0.01611   0.0106   0.2130   1.0000
   6.500   0.7207   0.02906   0.01857   0.0126   0.1426   1.0000
   6.750   0.7424   0.03171   0.02103   0.0144   0.1139   1.0000
   7.000   0.7662   0.03441   0.02378   0.0158   0.0995   1.0000
   7.250   0.7884   0.03737   0.02703   0.0170   0.0907   1.0000
   7.500   0.8100   0.04067   0.03077   0.0181   0.0871   1.0000
   7.750   0.8289   0.04436   0.03487   0.0191   0.0857   1.0000
   8.000   0.8425   0.04857   0.03967   0.0200   0.0862   1.0000
   8.250   0.8499   0.05326   0.04497   0.0207   0.0881   1.0000
   8.500   0.8522   0.05811   0.05029   0.0210   0.0901   1.0000
   8.750   0.8504   0.06298   0.05552   0.0210   0.0919   1.0000
   9.000   0.8469   0.06787   0.06062   0.0209   0.0935   1.0000
   9.250   0.8484   0.07290   0.06575   0.0208   0.0952   1.0000
   9.500   0.7948   0.07902   0.07221   0.0166   0.1014   1.0000
   9.750   0.7706   0.08563   0.07882   0.0123   0.1048   1.0000
  10.000   0.7833   0.09065   0.08382   0.0131   0.1087   1.0000
  10.250   0.7309   0.10302   0.09601  -0.0005   0.1193   1.0000
  10.500   0.7090   0.11340   0.10627  -0.0088   0.1433   1.0000
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