Airfoil Tools
Search 1636 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

GOE 329 (PFALZ 58) AIRFOIL (goe329-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: GOE 329 (PFALZ 58) AIRFOIL (goe329-il)
Reynolds number: 50,000
Max Cl/Cd: 37.88 at α=6°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe329-il-50000-n5.txt
Download as CSV file: xf-goe329-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 329 (PFALZ 58) 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
  ------ -------- --------- --------- -------- -------- --------
  -7.500  -0.3182   0.10568   0.09960  -0.0253   1.0000   0.0994
  -7.250  -0.3345   0.10524   0.09929  -0.0245   1.0000   0.1012
  -7.000  -0.3499   0.10509   0.09923  -0.0261   1.0000   0.1024
  -6.750  -0.3438   0.10140   0.09559  -0.0279   0.9971   0.1041
  -6.500  -0.3240   0.09663   0.09082  -0.0266   0.9929   0.1085
  -6.250  -0.2996   0.09285   0.08698  -0.0336   0.9846   0.1151
  -6.000  -0.2720   0.08891   0.08293  -0.0450   0.9743   0.1199
  -5.750  -0.2531   0.08470   0.07873  -0.0444   0.9673   0.1243
  -5.250  -0.1989   0.07705   0.07095  -0.0558   0.9498   0.1403
  -5.000  -0.1692   0.07358   0.06735  -0.0629   0.9398   0.1533
  -4.000  -0.0508   0.06010   0.05353  -0.0789   0.9048   0.2175
  -3.750  -0.0152   0.05675   0.05010  -0.0829   0.8990   0.2490
  -3.500   0.0026   0.05427   0.04764  -0.0824   0.8876   0.2821
  -3.000   0.1396   0.04531   0.03693  -0.1040   0.8709   0.1049
  -2.750   0.1805   0.04278   0.03380  -0.1064   0.8614   0.0877
  -2.500   0.2160   0.04009   0.03099  -0.1086   0.8530   0.0839
  -2.250   0.2484   0.03810   0.02865  -0.1097   0.8415   0.0797
  -2.000   0.2875   0.03628   0.02623  -0.1113   0.8321   0.0761
  -1.750   0.3211   0.03466   0.02434  -0.1122   0.8216   0.0765
  -1.500   0.3513   0.03336   0.02277  -0.1125   0.8090   0.0767
  -1.250   0.3839   0.03232   0.02130  -0.1126   0.7972   0.0754
  -1.000   0.4178   0.03101   0.01971  -0.1131   0.7869   0.0747
  -0.750   0.4463   0.03000   0.01850  -0.1127   0.7734   0.0744
  -0.500   0.4742   0.02914   0.01745  -0.1122   0.7594   0.0743
  -0.250   0.5017   0.02836   0.01648  -0.1114   0.7454   0.0744
   0.000   0.5298   0.02770   0.01562  -0.1107   0.7314   0.0753
   0.250   0.5575   0.02696   0.01479  -0.1101   0.7173   0.0781
   0.500   0.5849   0.02641   0.01412  -0.1094   0.7030   0.0812
   0.750   0.6124   0.02594   0.01351  -0.1088   0.6889   0.0832
   1.000   0.6400   0.02556   0.01296  -0.1082   0.6748   0.0853
   1.250   0.6678   0.02525   0.01247  -0.1077   0.6608   0.0879
   1.500   0.6953   0.02498   0.01211  -0.1073   0.6467   0.0920
   1.750   0.7221   0.02486   0.01191  -0.1068   0.6324   0.1009
   2.250   0.7687   0.02291   0.01184  -0.1047   0.6062   1.0000
   2.500   0.7957   0.02319   0.01175  -0.1041   0.5940   1.0000
   2.750   0.8235   0.02344   0.01170  -0.1037   0.5831   1.0000
   3.000   0.8490   0.02385   0.01191  -0.1033   0.5710   1.0000
   3.250   0.8746   0.02428   0.01217  -0.1029   0.5599   1.0000
   3.500   0.9019   0.02463   0.01233  -0.1027   0.5505   1.0000
   3.750   0.9270   0.02514   0.01276  -0.1024   0.5403   1.0000
   4.000   0.9526   0.02564   0.01317  -0.1021   0.5311   1.0000
   4.250   0.9789   0.02611   0.01354  -0.1019   0.5227   1.0000
   4.500   1.0033   0.02672   0.01415  -0.1016   0.5140   1.0000
   4.750   1.0298   0.02723   0.01460  -0.1014   0.5066   1.0000
   5.000   1.0535   0.02791   0.01531  -0.1010   0.4987   1.0000
   5.250   1.0791   0.02849   0.01590  -0.1008   0.4917   1.0000
   5.500   1.1033   0.02920   0.01664  -0.1005   0.4849   1.0000
   5.750   1.1267   0.02995   0.01746  -0.1001   0.4780   1.0000
   6.000   1.1545   0.03048   0.01798  -0.1001   0.4729   1.0000
   6.250   1.1733   0.03153   0.01923  -0.0993   0.4656   1.0000
   6.500   1.1980   0.03224   0.02000  -0.0991   0.4600   1.0000
   6.750   1.2216   0.03307   0.02092  -0.0987   0.4548   1.0000
   7.000   1.2399   0.03422   0.02231  -0.0979   0.4487   1.0000
   7.250   1.2642   0.03502   0.02320  -0.0976   0.4438   1.0000
   7.500   1.2860   0.03600   0.02432  -0.0971   0.4389   1.0000
   7.750   1.3006   0.03741   0.02602  -0.0959   0.4330   1.0000
   8.000   1.3225   0.03839   0.02716  -0.0954   0.4285   1.0000
   8.250   1.3482   0.03921   0.02810  -0.0953   0.4247   1.0000
   8.500   1.3523   0.04131   0.03055  -0.0932   0.4187   1.0000
   8.750   1.3687   0.04265   0.03213  -0.0922   0.4139   1.0000
   9.000   1.3940   0.04350   0.03315  -0.0920   0.4104   1.0000
   9.250   1.3924   0.04606   0.03604  -0.0894   0.4051   1.0000
   9.500   1.3934   0.04842   0.03867  -0.0872   0.3998   1.0000
   9.750   1.4133   0.04957   0.04004  -0.0865   0.3959   1.0000
  10.000   1.4110   0.05213   0.04288  -0.0840   0.3907   1.0000
  10.250   1.3606   0.05818   0.04906  -0.0787   0.3844   1.0000
  10.500   1.4078   0.05556   0.04669  -0.0782   0.3743   1.0000
  10.750   1.2550   0.07567   0.06664  -0.0761   0.3667   1.0000
  11.000   1.2764   0.07639   0.06758  -0.0751   0.3638   1.0000
  11.250   1.2027   0.09083   0.08194  -0.0784   0.3500   1.0000
  11.500   1.2264   0.09099   0.08235  -0.0771   0.3474   1.0000
  11.750   1.3188   0.07499   0.06676  -0.0691   0.3269   1.0000
  12.000   1.2455   0.09351   0.08528  -0.0749   0.3306   1.0000
  12.250   1.1703   0.10992   0.10151  -0.0804   0.3207   1.0000
<< Back to GOE 329 (PFALZ 58) AIRFOIL (goe329-il)

Polar data table (+)

Polar graphs


<< Back to GOE 329 (PFALZ 58) AIRFOIL (goe329-il)