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GOE 741 AIRFOIL (goe741-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: GOE 741 AIRFOIL (goe741-il)
Reynolds number: 50,000
Max Cl/Cd: 14.06 at α=2.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe741-il-50000-n5.txt
Download as CSV file: xf-goe741-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 741 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
  ------ -------- --------- --------- -------- -------- --------
 -10.250  -0.2664   0.12320   0.11618  -0.0133   0.7312   0.1402
 -10.000  -0.2742   0.12109   0.11407  -0.0160   0.7266   0.1414
  -9.750  -0.2760   0.11825   0.11125  -0.0189   0.7208   0.1418
  -9.250  -0.2713   0.10381   0.09667  -0.0264   0.7117   0.0902
  -9.000  -0.2571   0.10101   0.09381  -0.0258   0.7066   0.0887
  -8.750  -0.2493   0.09757   0.09040  -0.0276   0.7005   0.0870
  -8.500  -0.2476   0.09380   0.08665  -0.0298   0.6952   0.0852
  -8.000  -0.2815   0.08281   0.07560  -0.0391   0.6882   0.0806
  -7.750  -0.2844   0.07942   0.07220  -0.0405   0.6832   0.0804
  -7.500  -0.2881   0.07555   0.06826  -0.0422   0.6782   0.0804
  -7.250  -0.2904   0.07179   0.06439  -0.0432   0.6738   0.0804
  -7.000  -0.2923   0.06799   0.06042  -0.0435   0.6702   0.0804
  -6.750  -0.2910   0.06437   0.05660  -0.0437   0.6664   0.0804
  -6.500  -0.2862   0.06087   0.05291  -0.0441   0.6611   0.0803
  -6.250  -0.2810   0.05750   0.04927  -0.0437   0.6564   0.0802
  -6.000  -0.2754   0.05418   0.04558  -0.0426   0.6525   0.0804
  -5.750  -0.2682   0.05100   0.04195  -0.0411   0.6492   0.0806
  -5.500  -0.2505   0.04947   0.04035  -0.0409   0.6436   0.0817
  -5.250  -0.2329   0.04813   0.03888  -0.0404   0.6385   0.0833
  -5.000  -0.2169   0.04642   0.03690  -0.0392   0.6344   0.0854
  -4.750  -0.2013   0.04436   0.03441  -0.0377   0.6311   0.0875
  -4.500  -0.1838   0.04237   0.03193  -0.0367   0.6266   0.0893
  -4.250  -0.1634   0.04091   0.03016  -0.0362   0.6211   0.0910
  -4.000  -0.1415   0.04002   0.02918  -0.0356   0.6167   0.0933
  -3.750  -0.1190   0.03906   0.02800  -0.0347   0.6132   0.0969
  -3.500  -0.0950   0.03778   0.02626  -0.0336   0.6104   0.1017
  -3.250  -0.0718   0.03755   0.02613  -0.0340   0.6041   0.1055
  -3.000  -0.0466   0.03693   0.02532  -0.0339   0.5992   0.1111
  -2.750  -0.0206   0.03632   0.02462  -0.0335   0.5955   0.1178
  -2.500   0.0078   0.03565   0.02372  -0.0331   0.5925   0.1274
  -2.250   0.0334   0.03572   0.02392  -0.0336   0.5876   0.1366
  -2.000   0.0599   0.03586   0.02409  -0.0342   0.5819   0.1489
  -1.750   0.0885   0.03577   0.02395  -0.0344   0.5778   0.1647
  -1.500   0.1177   0.03558   0.02367  -0.0344   0.5746   0.1845
  -1.250   0.1477   0.03529   0.02330  -0.0343   0.5720   0.2073
  -1.000   0.1672   0.03627   0.02441  -0.0352   0.5647   0.2270
  -0.750   0.1910   0.03637   0.02460  -0.0351   0.5600   0.2517
  -0.500   0.2167   0.03617   0.02443  -0.0346   0.5566   0.2818
  -0.250   0.2422   0.03577   0.02414  -0.0338   0.5539   0.3154
   0.000   0.2547   0.03679   0.02537  -0.0335   0.5473   0.3476
   0.250   0.2684   0.03709   0.02605  -0.0325   0.5416   0.4033
   0.750   0.4831   0.03576   0.02555  -0.0575   0.5349   1.0000
   1.000   0.4936   0.03751   0.02726  -0.0571   0.5280   1.0000
   1.250   0.5061   0.03878   0.02843  -0.0561   0.5218   1.0000
   1.500   0.5257   0.03917   0.02866  -0.0546   0.5181   1.0000
   1.750   0.5482   0.03928   0.02860  -0.0531   0.5154   1.0000
   2.000   0.5428   0.04216   0.03153  -0.0520   0.5058   1.0000
   2.250   0.5567   0.04300   0.03227  -0.0504   0.5006   1.0000
   2.500   0.5788   0.04312   0.03224  -0.0489   0.4975   1.0000
   2.750   0.6042   0.04296   0.03192  -0.0474   0.4952   1.0000
   3.000   0.5771   0.04720   0.03626  -0.0452   0.4820   1.0000
   3.250   0.5993   0.04726   0.03621  -0.0436   0.4788   1.0000
   3.500   0.6264   0.04694   0.03575  -0.0421   0.4767   1.0000
   4.000   0.6126   0.05170   0.04048  -0.0380   0.4598   1.0000
   4.250   0.6399   0.05140   0.04006  -0.0365   0.4577   1.0000
   4.750   0.6162   0.05614   0.04474  -0.0312   0.4412   1.0000
   5.250   0.6011   0.06059   0.04911  -0.0268   0.4256   1.0000
   5.500   0.6277   0.06050   0.04895  -0.0256   0.4232   1.0000
   6.000   0.6162   0.06540   0.05379  -0.0225   0.4081   1.0000
   6.250   0.6384   0.06572   0.05405  -0.0213   0.4055   1.0000
   6.500   0.6133   0.07001   0.05835  -0.0202   0.3943   1.0000
   6.750   0.6339   0.07046   0.05875  -0.0190   0.3911   1.0000
   7.250   0.6278   0.07591   0.06418  -0.0174   0.3779   1.0000
   7.500   0.6460   0.07673   0.06497  -0.0165   0.3750   1.0000
   8.000   0.6412   0.08241   0.07066  -0.0155   0.3625   1.0000
   8.250   0.6588   0.08331   0.07153  -0.0146   0.3594   1.0000
   8.500   0.6808   0.08381   0.07201  -0.0137   0.3572   1.0000
   8.750   0.6521   0.08979   0.07805  -0.0143   0.3488   1.0000
   9.000   0.6636   0.09147   0.07972  -0.0138   0.3451   1.0000
   9.250   0.6828   0.09232   0.08056  -0.0130   0.3426   1.0000
   9.500   0.6700   0.09664   0.08492  -0.0134   0.3361   1.0000
   9.750   0.6722   0.09950   0.08780  -0.0134   0.3318   1.0000
  10.000   0.6853   0.10113   0.08945  -0.0129   0.3287   1.0000
  10.250   0.7053   0.10196   0.09028  -0.0122   0.3263   1.0000
  10.500   0.6876   0.10700   0.09537  -0.0131   0.3195   1.0000
  10.750   0.6941   0.10934   0.09774  -0.0130   0.3152   1.0000
  11.000   0.7107   0.11057   0.09898  -0.0125   0.3121   1.0000
  11.250   0.7151   0.11312   0.10157  -0.0125   0.3076   1.0000
  11.500   0.7097   0.11671   0.10521  -0.0131   0.3011   1.0000
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