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GOE 553 AIRFOIL (goe553-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: GOE 553 AIRFOIL (goe553-il)
Reynolds number: 50,000
Max Cl/Cd: 30.12 at α=7.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe553-il-50000-n5.txt
Download as CSV file: xf-goe553-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 553 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.250  -0.2580   0.10894   0.10219  -0.0376   1.0000   0.1509
  -9.000  -0.2828   0.10820   0.10164  -0.0391   1.0000   0.1529
  -8.500  -0.3077   0.09594   0.08954  -0.0430   1.0000   0.0915
  -8.250  -0.3003   0.09390   0.08758  -0.0400   1.0000   0.0899
  -8.000  -0.3122   0.09252   0.08637  -0.0370   1.0000   0.0884
  -7.750  -0.3184   0.08303   0.07688  -0.0520   0.9777   0.0803
  -7.500  -0.2945   0.07836   0.07214  -0.0574   0.9635   0.0789
  -7.250  -0.2753   0.07234   0.06604  -0.0657   0.9477   0.0779
  -7.000  -0.2590   0.06599   0.05951  -0.0744   0.9317   0.0775
  -6.750  -0.2416   0.06005   0.05330  -0.0820   0.9167   0.0774
  -6.500  -0.2206   0.05477   0.04766  -0.0882   0.9033   0.0773
  -6.250  -0.1937   0.04995   0.04241  -0.0937   0.8929   0.0767
  -6.000  -0.1714   0.04623   0.03826  -0.0966   0.8790   0.0764
  -5.750  -0.1462   0.04298   0.03454  -0.0989   0.8667   0.0765
  -5.500  -0.1138   0.03997   0.03099  -0.1017   0.8579   0.0778
  -5.250  -0.0887   0.03779   0.02829  -0.1026   0.8450   0.0793
  -5.000  -0.0598   0.03583   0.02594  -0.1035   0.8345   0.0806
  -4.750  -0.0291   0.03414   0.02401  -0.1043   0.8249   0.0817
  -4.500  -0.0024   0.03282   0.02247  -0.1043   0.8136   0.0829
  -4.250   0.0303   0.03145   0.02085  -0.1050   0.8055   0.0846
  -4.000   0.0553   0.03056   0.01977  -0.1045   0.7939   0.0874
  -3.750   0.0856   0.02960   0.01852  -0.1045   0.7850   0.0910
  -3.500   0.1117   0.02877   0.01763  -0.1039   0.7748   0.0939
  -3.250   0.1387   0.02804   0.01682  -0.1034   0.7655   0.0971
  -3.000   0.1663   0.02738   0.01603  -0.1029   0.7562   0.1013
  -2.750   0.1926   0.02685   0.01538  -0.1024   0.7466   0.1074
  -2.500   0.2205   0.02626   0.01472  -0.1022   0.7378   0.1167
  -2.250   0.2457   0.02578   0.01422  -0.1018   0.7281   0.1278
  -2.000   0.2750   0.02507   0.01352  -0.1019   0.7198   0.1531
  -1.750   0.3000   0.02401   0.01314  -0.1023   0.7100   0.2577
  -1.500   0.3220   0.02339   0.01347  -0.1005   0.7021   0.5236
  -1.250   0.3391   0.02357   0.01385  -0.0976   0.6925   0.6266
  -1.000   0.3551   0.02353   0.01400  -0.0935   0.6848   0.7137
  -0.750   0.3673   0.02349   0.01412  -0.0892   0.6757   0.7784
  -0.500   0.3864   0.02326   0.01392  -0.0861   0.6678   0.8324
  -0.250   0.4149   0.02316   0.01379  -0.0856   0.6590   0.8806
   0.000   0.4619   0.02307   0.01353  -0.0890   0.6500   0.9352
   0.250   0.4977   0.02329   0.01353  -0.0913   0.6407   1.0000
   0.500   0.5263   0.02357   0.01356  -0.0920   0.6323   1.0000
   0.750   0.5537   0.02394   0.01370  -0.0925   0.6238   1.0000
   1.000   0.5808   0.02430   0.01388  -0.0927   0.6151   1.0000
   1.250   0.6093   0.02464   0.01401  -0.0930   0.6076   1.0000
   1.500   0.6334   0.02514   0.01439  -0.0928   0.5984   1.0000
   1.750   0.6647   0.02534   0.01436  -0.0932   0.5923   1.0000
   2.000   0.6845   0.02605   0.01505  -0.0925   0.5822   1.0000
   2.250   0.7138   0.02632   0.01515  -0.0926   0.5758   1.0000
   2.500   0.7344   0.02703   0.01583  -0.0920   0.5668   1.0000
   2.750   0.7613   0.02741   0.01609  -0.0918   0.5598   1.0000
   3.000   0.7842   0.02802   0.01665  -0.0913   0.5522   1.0000
   3.250   0.8073   0.02860   0.01719  -0.0908   0.5446   1.0000
   3.500   0.8374   0.02881   0.01727  -0.0909   0.5394   1.0000
   3.750   0.8523   0.02986   0.01840  -0.0896   0.5300   1.0000
   4.000   0.8802   0.03018   0.01863  -0.0895   0.5244   1.0000
   4.250   0.8975   0.03114   0.01964  -0.0885   0.5166   1.0000
   4.500   0.9205   0.03174   0.02022  -0.0879   0.5100   1.0000
   4.750   0.9487   0.03203   0.02045  -0.0878   0.5050   1.0000
   5.000   0.9599   0.03336   0.02190  -0.0863   0.4963   1.0000
   5.250   0.9877   0.03366   0.02215  -0.0860   0.4912   1.0000
   5.500   1.0005   0.03486   0.02343  -0.0846   0.4831   1.0000
   5.750   1.0243   0.03525   0.02384  -0.0839   0.4763   1.0000
   6.000   1.0431   0.03592   0.02452  -0.0828   0.4686   1.0000
   6.250   1.0636   0.03636   0.02499  -0.0818   0.4603   1.0000
   6.500   1.0826   0.03689   0.02555  -0.0806   0.4521   1.0000
   6.750   1.1025   0.03732   0.02601  -0.0794   0.4438   1.0000
   7.000   1.1209   0.03796   0.02668  -0.0783   0.4364   1.0000
   7.250   1.1331   0.03904   0.02784  -0.0767   0.4287   1.0000
   7.500   1.1679   0.03877   0.02756  -0.0769   0.4240   1.0000
   7.750   1.1562   0.04140   0.03038  -0.0736   0.4148   1.0000
   8.000   1.1854   0.04142   0.03041  -0.0733   0.4093   1.0000
   8.250   1.1767   0.04375   0.03286  -0.0702   0.4012   1.0000
   8.500   1.1913   0.04463   0.03382  -0.0688   0.3946   1.0000
   8.750   1.2116   0.04518   0.03442  -0.0678   0.3886   1.0000
   9.000   1.1874   0.04886   0.03824  -0.0646   0.3793   1.0000
   9.250   1.2310   0.04771   0.03711  -0.0648   0.3751   1.0000
   9.500   1.1721   0.05493   0.04449  -0.0618   0.3631   1.0000
   9.750   1.2093   0.05393   0.04356  -0.0611   0.3593   1.0000
  10.000   1.1507   0.06286   0.05260  -0.0608   0.3455   1.0000
  10.250   1.1861   0.06167   0.05149  -0.0596   0.3425   1.0000
  10.750   1.1665   0.07013   0.06012  -0.0596   0.3254   1.0000
  11.250   1.1325   0.08164   0.07176  -0.0613   0.3069   1.0000
  11.750   1.1058   0.09346   0.08372  -0.0638   0.2913   1.0000
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