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

GOE 274 (DAIMLER V) AIRFOIL (goe274-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: GOE 274 (DAIMLER V) AIRFOIL (goe274-il)
Reynolds number: 50,000
Max Cl/Cd: 40.73 at α=5.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe274-il-50000-n5.txt
Download as CSV file: xf-goe274-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 274 (DAIMLER V) 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
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.3296   0.11013   0.10335  -0.0263   1.0000   0.0829
  -8.000  -0.3438   0.10972   0.10310  -0.0258   1.0000   0.0844
  -7.750  -0.3543   0.10920   0.10270  -0.0283   1.0000   0.0852
  -7.500  -0.3504   0.10484   0.09842  -0.0266   1.0000   0.0866
  -7.250  -0.3419   0.10070   0.09430  -0.0232   1.0000   0.0898
  -7.000  -0.3411   0.09816   0.09183  -0.0225   1.0000   0.0929
  -6.750  -0.3418   0.09593   0.08967  -0.0231   1.0000   0.0965
  -6.500  -0.3419   0.09473   0.08850  -0.0294   1.0000   0.1001
  -6.000  -0.3341   0.08761   0.08152  -0.0254   1.0000   0.1056
  -5.750  -0.3276   0.08512   0.07904  -0.0270   1.0000   0.1129
  -5.500  -0.3178   0.08201   0.07593  -0.0309   1.0000   0.1180
  -5.250  -0.3124   0.07908   0.07304  -0.0284   1.0000   0.1238
  -5.000  -0.2949   0.07607   0.06995  -0.0345   1.0000   0.1329
  -4.750  -0.2883   0.07306   0.06701  -0.0320   1.0000   0.1383
  -4.500  -0.2701   0.06997   0.06384  -0.0360   1.0000   0.1492
  -4.250  -0.2509   0.06717   0.06096  -0.0392   1.0000   0.1630
  -3.750  -0.2232   0.06162   0.05544  -0.0396   1.0000   0.1958
  -3.500  -0.1999   0.05876   0.05256  -0.0415   0.9968   0.2278
  -3.250  -0.1702   0.05579   0.04960  -0.0438   0.9898   0.2752
  -3.000  -0.1443   0.05302   0.04691  -0.0440   0.9828   0.3267
  -2.500   0.0397   0.04358   0.03524  -0.0770   0.9689   0.1148
  -2.250   0.0906   0.04056   0.03180  -0.0820   0.9626   0.0976
  -2.000   0.1356   0.03816   0.02894  -0.0857   0.9533   0.0884
  -1.750   0.1809   0.03641   0.02673  -0.0891   0.9443   0.0830
  -1.500   0.2261   0.03453   0.02456  -0.0926   0.9354   0.0801
  -1.250   0.2662   0.03311   0.02283  -0.0949   0.9238   0.0795
  -1.000   0.3076   0.03184   0.02125  -0.0973   0.9126   0.0807
  -0.750   0.3514   0.03053   0.01968  -0.1000   0.9022   0.0802
  -0.500   0.3913   0.02938   0.01831  -0.1017   0.8895   0.0791
  -0.250   0.4290   0.02833   0.01710  -0.1029   0.8755   0.0786
   0.000   0.4664   0.02737   0.01596  -0.1038   0.8610   0.0785
   0.250   0.5028   0.02647   0.01492  -0.1044   0.8455   0.0789
   0.500   0.5368   0.02570   0.01405  -0.1047   0.8275   0.0808
   0.750   0.5694   0.02508   0.01332  -0.1048   0.8068   0.0846
   1.000   0.6051   0.02431   0.01251  -0.1056   0.7865   0.0882
   1.250   0.6390   0.02375   0.01186  -0.1060   0.7629   0.0911
   1.500   0.6745   0.02323   0.01122  -0.1065   0.7389   0.0953
   1.750   0.7073   0.02292   0.01077  -0.1067   0.7121   0.1012
   2.000   0.7400   0.02264   0.01040  -0.1069   0.6856   0.1153
   2.250   0.7618   0.02067   0.01020  -0.1050   0.6619   1.0000
   2.500   0.7909   0.02093   0.01008  -0.1046   0.6364   1.0000
   2.750   0.8194   0.02122   0.01001  -0.1042   0.6123   1.0000
   3.000   0.8462   0.02160   0.01007  -0.1035   0.5877   1.0000
   3.250   0.8723   0.02203   0.01021  -0.1029   0.5646   1.0000
   3.500   0.8979   0.02251   0.01045  -0.1022   0.5427   1.0000
   3.750   0.9237   0.02301   0.01073  -0.1017   0.5234   1.0000
   4.000   0.9496   0.02354   0.01111  -0.1013   0.5061   1.0000
   4.250   0.9755   0.02409   0.01155  -0.1009   0.4903   1.0000
   4.500   1.0010   0.02465   0.01205  -0.1005   0.4752   1.0000
   4.750   1.0263   0.02522   0.01262  -0.1001   0.4605   1.0000
   5.000   1.0511   0.02581   0.01319  -0.0996   0.4462   1.0000
   5.250   1.0758   0.02641   0.01381  -0.0992   0.4326   1.0000
   5.500   1.1005   0.02704   0.01449  -0.0987   0.4198   1.0000
   5.750   1.1250   0.02768   0.01515  -0.0982   0.4073   1.0000
   6.000   1.1494   0.02833   0.01580  -0.0977   0.3953   1.0000
   6.250   1.1725   0.02902   0.01657  -0.0970   0.3826   1.0000
   6.500   1.1953   0.02975   0.01744  -0.0963   0.3702   1.0000
   6.750   1.2182   0.03051   0.01829  -0.0956   0.3589   1.0000
   7.000   1.2424   0.03125   0.01904  -0.0950   0.3494   1.0000
   7.250   1.2643   0.03218   0.02017  -0.0943   0.3389   1.0000
   7.500   1.2873   0.03308   0.02122  -0.0937   0.3302   1.0000
   7.750   1.3095   0.03399   0.02227  -0.0930   0.3214   1.0000
   8.000   1.3307   0.03502   0.02349  -0.0921   0.3130   1.0000
   8.250   1.3526   0.03593   0.02455  -0.0914   0.3053   1.0000
   8.500   1.3718   0.03708   0.02595  -0.0903   0.2974   1.0000
   8.750   1.3927   0.03798   0.02698  -0.0894   0.2900   1.0000
   9.000   1.4089   0.03916   0.02844  -0.0880   0.2817   1.0000
   9.250   1.4255   0.03976   0.02912  -0.0863   0.2718   1.0000
   9.500   1.4368   0.04028   0.02978  -0.0840   0.2595   1.0000
   9.750   1.4435   0.04107   0.03077  -0.0812   0.2467   1.0000
  10.000   1.4491   0.04198   0.03188  -0.0783   0.2349   1.0000
  10.250   1.4540   0.04298   0.03306  -0.0755   0.2240   1.0000
  10.500   1.4589   0.04397   0.03417  -0.0727   0.2148   1.0000
  10.750   1.4572   0.04527   0.03567  -0.0692   0.2049   1.0000
  11.000   1.4545   0.04687   0.03752  -0.0660   0.1948   1.0000
  11.250   1.4546   0.04846   0.03923  -0.0634   0.1870   1.0000
  11.500   1.4506   0.05069   0.04175  -0.0610   0.1776   1.0000
  11.750   1.4433   0.05325   0.04450  -0.0589   0.1670   1.0000
  12.000   1.4352   0.05614   0.04751  -0.0573   0.1574   1.0000
  12.250   1.4256   0.05983   0.05154  -0.0565   0.1460   1.0000
  12.500   1.4127   0.06430   0.05631  -0.0565   0.1325   1.0000
  12.750   1.3978   0.06955   0.06186  -0.0574   0.1173   1.0000
  13.000   1.3815   0.07531   0.06779  -0.0588   0.1023   1.0000
  13.250   1.3633   0.08155   0.07401  -0.0608   0.0925   1.0000
  13.500   1.3443   0.08824   0.08070  -0.0631   0.0859   1.0000
  13.750   1.3256   0.09510   0.08761  -0.0656   0.0807   1.0000
<< Back to GOE 274 (DAIMLER V) AIRFOIL (goe274-il)

Polar data table (+)

Polar graphs


<< Back to GOE 274 (DAIMLER V) AIRFOIL (goe274-il)