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

GOE 287 AIRFOIL (goe287-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: GOE 287 AIRFOIL (goe287-il)
Reynolds number: 100,000
Max Cl/Cd: 58.26 at α=7.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe287-il-100000.txt
Download as CSV file: xf-goe287-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 287 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.3359   0.09949   0.09521  -0.0161   1.0000   0.0625
  -7.750  -0.3331   0.09730   0.09311  -0.0205   1.0000   0.0646
  -7.500  -0.3274   0.09584   0.09173  -0.0313   1.0000   0.0657
  -7.250  -0.3164   0.09402   0.08990  -0.0414   1.0000   0.0662
  -7.000  -0.3111   0.08660   0.08262  -0.0298   1.0000   0.0687
  -6.750  -0.3121   0.08447   0.08060  -0.0283   1.0000   0.0707
  -6.500  -0.3146   0.08281   0.07902  -0.0288   0.9951   0.0725
  -6.250  -0.2518   0.07827   0.07414  -0.0544   0.9783   0.0802
  -6.000  -0.2273   0.07126   0.06722  -0.0562   0.9711   0.0823
  -5.750  -0.1986   0.06698   0.06292  -0.0587   0.9590   0.0869
  -5.500  -0.1537   0.06260   0.05820  -0.0723   0.9412   0.0961
  -5.250  -0.1383   0.05906   0.05474  -0.0704   0.9253   0.1005
  -5.000  -0.1047   0.05599   0.05128  -0.0779   0.9070   0.1106
  -4.750  -0.0916   0.05276   0.04812  -0.0759   0.8920   0.1142
  -4.500  -0.0640   0.05009   0.04512  -0.0795   0.8777   0.1258
  -4.250  -0.0481   0.04746   0.04251  -0.0782   0.8653   0.1321
  -4.000  -0.0260   0.04500   0.03986  -0.0791   0.8546   0.1455
  -3.750  -0.0035   0.04279   0.03750  -0.0798   0.8438   0.1616
  -3.500   0.0229   0.04090   0.03533  -0.0816   0.8335   0.1859
  -3.250   0.0428   0.03869   0.03309  -0.0809   0.8255   0.2038
  -3.000   0.0654   0.03691   0.03122  -0.0811   0.8160   0.2339
  -2.250   0.1203   0.03192   0.02635  -0.0769   0.7945   0.3841
  -2.000   0.1405   0.03026   0.02474  -0.0755   0.7877   0.4350
  -1.750   0.2374   0.02716   0.01912  -0.0871   0.7831   0.1418
  -1.500   0.2709   0.02571   0.01725  -0.0874   0.7764   0.1164
  -1.250   0.3021   0.02539   0.01638  -0.0870   0.7708   0.1096
  -1.000   0.3305   0.02428   0.01511  -0.0870   0.7660   0.1084
  -0.750   0.3598   0.02345   0.01419  -0.0874   0.7598   0.1060
  -0.500   0.3874   0.02276   0.01335  -0.0868   0.7544   0.1046
  -0.250   0.4153   0.02228   0.01280  -0.0867   0.7467   0.1046
   0.000   0.4415   0.02172   0.01214  -0.0856   0.7395   0.1060
   0.250   0.4683   0.02142   0.01180  -0.0852   0.7308   0.1094
   0.500   0.4931   0.02075   0.01119  -0.0840   0.7239   0.1185
   0.750   0.5197   0.02052   0.01097  -0.0837   0.7152   0.1275
   1.000   0.5465   0.02009   0.01049  -0.0827   0.7089   0.1478
   1.250   0.5725   0.01802   0.01031  -0.0819   0.7005   1.0000
   1.500   0.5990   0.01807   0.01005  -0.0809   0.6943   1.0000
   1.750   0.6261   0.01843   0.01031  -0.0811   0.6864   1.0000
   2.000   0.6528   0.01854   0.01027  -0.0805   0.6804   1.0000
   2.250   0.6795   0.01886   0.01054  -0.0804   0.6727   1.0000
   2.500   0.7062   0.01896   0.01053  -0.0798   0.6660   1.0000
   2.750   0.7327   0.01925   0.01080  -0.0797   0.6579   1.0000
   3.000   0.7594   0.01932   0.01081  -0.0790   0.6510   1.0000
   3.250   0.7857   0.01962   0.01111  -0.0789   0.6422   1.0000
   3.500   0.8126   0.01960   0.01101  -0.0781   0.6352   1.0000
   3.750   0.8386   0.01987   0.01135  -0.0778   0.6247   1.0000
   4.000   0.8650   0.01996   0.01142  -0.0772   0.6153   1.0000
   4.250   0.8917   0.01990   0.01130  -0.0764   0.6057   1.0000
   4.500   0.9176   0.02000   0.01145  -0.0758   0.5929   1.0000
   4.750   0.9435   0.02004   0.01154  -0.0751   0.5796   1.0000
   5.000   0.9696   0.02004   0.01155  -0.0743   0.5653   1.0000
   5.250   0.9956   0.02004   0.01155  -0.0735   0.5497   1.0000
   5.500   1.0217   0.02002   0.01154  -0.0727   0.5328   1.0000
   5.750   1.0474   0.02008   0.01159  -0.0719   0.5142   1.0000
   6.000   1.0724   0.02029   0.01187  -0.0713   0.4927   1.0000
   6.250   1.0980   0.02046   0.01204  -0.0707   0.4743   1.0000
   6.500   1.1231   0.02047   0.01208  -0.0700   0.4547   1.0000
   6.750   1.1467   0.02016   0.01176  -0.0690   0.4289   1.0000
   7.000   1.1697   0.02020   0.01182  -0.0682   0.4029   1.0000
   7.250   1.1920   0.02046   0.01204  -0.0674   0.3781   1.0000
   7.500   1.2124   0.02095   0.01244  -0.0664   0.3512   1.0000
   7.750   1.2322   0.02166   0.01304  -0.0654   0.3286   1.0000
   8.000   1.2489   0.02255   0.01386  -0.0642   0.2969   1.0000
   8.250   1.2657   0.02345   0.01475  -0.0630   0.2692   1.0000
   8.500   1.2818   0.02439   0.01572  -0.0617   0.2365   1.0000
   8.750   1.2955   0.02557   0.01683  -0.0603   0.1824   1.0000
   9.000   1.2873   0.02876   0.01903  -0.0572   0.0728   1.0000
   9.250   1.2903   0.03095   0.02104  -0.0548   0.0594   1.0000
   9.500   1.2939   0.03274   0.02295  -0.0524   0.0554   1.0000
   9.750   1.2945   0.03484   0.02518  -0.0501   0.0530   1.0000
  10.000   1.2921   0.03737   0.02785  -0.0483   0.0513   1.0000
  10.250   1.2892   0.04013   0.03077  -0.0470   0.0502   1.0000
  10.500   1.2879   0.04289   0.03373  -0.0461   0.0491   1.0000
  10.750   1.2857   0.04584   0.03684  -0.0453   0.0479   1.0000
  11.000   1.2833   0.04889   0.04002  -0.0446   0.0465   1.0000
  11.250   1.2816   0.05186   0.04313  -0.0439   0.0454   1.0000
  11.500   1.2826   0.05451   0.04584  -0.0428   0.0445   1.0000
  11.750   1.2893   0.05649   0.04785  -0.0411   0.0439   1.0000
  12.000   1.3056   0.05758   0.04891  -0.0386   0.0432   1.0000
  12.250   1.3350   0.05809   0.04941  -0.0355   0.0429   1.0000
  12.500   1.3691   0.05930   0.05072  -0.0330   0.0424   1.0000
  12.750   1.4195   0.06246   0.05394  -0.0315   0.0411   1.0000
  13.000   1.4278   0.06536   0.05712  -0.0303   0.0413   1.0000
  13.250   1.4253   0.06814   0.06022  -0.0292   0.0419   1.0000
  13.500   1.4185   0.07148   0.06395  -0.0283   0.0427   1.0000
  13.750   1.4073   0.07554   0.06840  -0.0278   0.0436   1.0000
  14.000   1.3896   0.08032   0.07356  -0.0280   0.0444   1.0000
  14.250   1.3711   0.08557   0.07915  -0.0288   0.0452   1.0000
  14.500   1.3516   0.09121   0.08509  -0.0301   0.0459   1.0000
  14.750   1.3312   0.09724   0.09140  -0.0321   0.0466   1.0000
  15.000   1.3099   0.10367   0.09806  -0.0346   0.0472   1.0000
  15.250   1.2878   0.11053   0.10514  -0.0378   0.0477   1.0000
  15.500   1.2652   0.11800   0.11281  -0.0418   0.0482   1.0000
  15.750   1.2419   0.12618   0.12118  -0.0467   0.0486   1.0000
  16.000   1.2179   0.13525   0.13041  -0.0525   0.0491   1.0000
  16.250   1.1961   0.14478   0.14006  -0.0586   0.0497   1.0000
  16.500   1.1176   0.17455   0.16999  -0.0813   0.0531   1.0000
  16.750   1.0991   0.18951   0.18487  -0.0889   0.0597   1.0000
<< Back to GOE 287 AIRFOIL (goe287-il)

Polar data table (+)

Polar graphs


<< Back to GOE 287 AIRFOIL (goe287-il)