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

GOE 517 AIRFOIL (goe517-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5


Details Polar file
Airfoil: GOE 517 AIRFOIL (goe517-il)
Reynolds number: 1,000,000
Max Cl/Cd: 117.69 at α=4.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe517-il-1000000-n5.txt
Download as CSV file: xf-goe517-il-1000000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 517 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     1.000 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.000  -0.3355   0.11201   0.11036  -0.0255   1.0000   0.0045
  -9.750  -0.3315   0.10883   0.10720  -0.0264   1.0000   0.0045
  -9.500  -0.3283   0.10558   0.10396  -0.0272   1.0000   0.0045
  -9.250  -0.3247   0.10269   0.10110  -0.0277   1.0000   0.0045
  -9.000  -0.3202   0.10074   0.09916  -0.0275   1.0000   0.0041
  -8.750  -0.3118   0.09715   0.09558  -0.0298   0.9992   0.0038
  -8.500  -0.3024   0.09345   0.09189  -0.0324   0.9912   0.0049
  -8.250  -0.1595   0.06740   0.06580  -0.0569   0.9385   0.0043
  -8.000  -0.1477   0.06309   0.06140  -0.0605   0.9115   0.0042
  -7.750  -0.2504   0.07993   0.07835  -0.0487   0.9609   0.0044
  -7.500  -0.2174   0.07473   0.07310  -0.0589   0.9476   0.0042
  -7.250  -0.1841   0.06930   0.06756  -0.0698   0.9180   0.0039
  -7.000  -0.1724   0.06538   0.06349  -0.0740   0.8822   0.0037
  -6.750  -0.1652   0.06180   0.05978  -0.0765   0.8572   0.0038
  -5.750  -0.1465   0.01536   0.01137  -0.0936   0.8077   0.0040
  -5.500  -0.1237   0.01354   0.00916  -0.0930   0.8016   0.0043
  -5.000  -0.0741   0.01148   0.00656  -0.0919   0.7903   0.0046
  -4.750  -0.0490   0.01062   0.00549  -0.0914   0.7841   0.0049
  -4.500  -0.0230   0.01017   0.00492  -0.0910   0.7782   0.0052
  -4.250   0.0035   0.00974   0.00438  -0.0907   0.7715   0.0055
  -4.000   0.0296   0.00935   0.00386  -0.0903   0.7645   0.0059
  -3.750   0.0562   0.00899   0.00340  -0.0899   0.7569   0.0063
  -3.500   0.0822   0.00865   0.00294  -0.0895   0.7473   0.0067
  -3.250   0.1085   0.00841   0.00264  -0.0891   0.7349   0.0075
  -3.000   0.1348   0.00823   0.00238  -0.0887   0.7213   0.0084
  -2.750   0.1611   0.00803   0.00207  -0.0883   0.7076   0.0092
  -2.500   0.1869   0.00786   0.00182  -0.0878   0.6887   0.0106
  -2.250   0.2125   0.00780   0.00163  -0.0872   0.6644   0.0120
  -2.000   0.2376   0.00773   0.00145  -0.0866   0.6357   0.0149
  -1.750   0.2627   0.00773   0.00130  -0.0860   0.6052   0.0183
  -1.500   0.2879   0.00773   0.00119  -0.0854   0.5769   0.0246
  -0.750   0.3659   0.00767   0.00105  -0.0843   0.5316   0.0700
  -0.500   0.3925   0.00768   0.00103  -0.0840   0.5236   0.0787
  -0.250   0.4194   0.00767   0.00101  -0.0838   0.5164   0.0864
   0.000   0.4460   0.00768   0.00100  -0.0835   0.5090   0.0959
   0.250   0.4730   0.00768   0.00099  -0.0834   0.5019   0.1012
   0.500   0.4998   0.00769   0.00099  -0.0831   0.4955   0.1098
   0.750   0.5266   0.00769   0.00099  -0.0829   0.4899   0.1188
   1.000   0.5534   0.00770   0.00101  -0.0827   0.4824   0.1308
   1.250   0.5800   0.00771   0.00103  -0.0825   0.4746   0.1485
   1.500   0.6065   0.00770   0.00107  -0.0823   0.4663   0.1768
   1.750   0.6327   0.00767   0.00113  -0.0820   0.4583   0.2253
   2.000   0.6561   0.00726   0.00124  -0.0814   0.4482   0.4747
   2.500   0.7464   0.00644   0.00148  -0.0897   0.4067   1.0000
   2.750   0.7700   0.00666   0.00157  -0.0889   0.3822   1.0000
   3.000   0.7936   0.00687   0.00169  -0.0881   0.3599   1.0000
   3.250   0.8174   0.00709   0.00182  -0.0874   0.3416   1.0000
   3.500   0.8414   0.00728   0.00195  -0.0866   0.3272   1.0000
   3.750   0.8655   0.00747   0.00209  -0.0859   0.3156   1.0000
   4.000   0.8895   0.00767   0.00224  -0.0852   0.3028   1.0000
   4.250   0.9139   0.00784   0.00238  -0.0846   0.2924   1.0000
   4.500   0.9381   0.00802   0.00254  -0.0840   0.2828   1.0000
   4.750   0.9627   0.00818   0.00269  -0.0834   0.2749   1.0000
   5.000   0.9866   0.00840   0.00287  -0.0827   0.2621   1.0000
   5.250   1.0092   0.00872   0.00309  -0.0818   0.2391   1.0000
   5.500   1.0315   0.00907   0.00332  -0.0809   0.2137   1.0000
   5.750   1.0525   0.00952   0.00362  -0.0798   0.1805   1.0000
   6.000   1.0655   0.01068   0.00430  -0.0774   0.0935   1.0000
   6.250   1.0826   0.01150   0.00489  -0.0757   0.0501   1.0000
   6.500   1.1010   0.01221   0.00545  -0.0741   0.0167   1.0000
   6.750   1.1224   0.01265   0.00586  -0.0731   0.0109   1.0000
   7.000   1.1443   0.01303   0.00625  -0.0721   0.0084   1.0000
   7.250   1.1663   0.01339   0.00665  -0.0712   0.0071   1.0000
   7.500   1.1876   0.01381   0.00710  -0.0702   0.0060   1.0000
   7.750   1.2089   0.01421   0.00753  -0.0692   0.0052   1.0000
   8.000   1.2290   0.01470   0.00806  -0.0680   0.0045   1.0000
   8.250   1.2495   0.01512   0.00854  -0.0669   0.0041   1.0000
   8.500   1.2691   0.01561   0.00908  -0.0656   0.0037   1.0000
   8.750   1.2881   0.01612   0.00964  -0.0643   0.0034   1.0000
   9.000   1.3044   0.01681   0.01039  -0.0625   0.0031   1.0000
   9.250   1.3216   0.01739   0.01104  -0.0609   0.0030   1.0000
   9.500   1.3379   0.01800   0.01173  -0.0592   0.0028   1.0000
   9.750   1.3530   0.01859   0.01239  -0.0573   0.0026   1.0000
  10.000   1.3649   0.01923   0.01310  -0.0547   0.0024   1.0000
  10.250   1.3780   0.01978   0.01370  -0.0525   0.0023   1.0000
  10.500   1.3860   0.02065   0.01464  -0.0496   0.0022   1.0000
  10.750   1.3911   0.02172   0.01580  -0.0464   0.0021   1.0000
  11.000   1.3998   0.02261   0.01679  -0.0440   0.0020   1.0000
  11.250   1.4067   0.02366   0.01794  -0.0415   0.0019   1.0000
  11.500   1.4100   0.02501   0.01940  -0.0387   0.0019   1.0000
  11.750   1.4175   0.02614   0.02065  -0.0367   0.0017   1.0000
  12.000   1.4221   0.02755   0.02216  -0.0346   0.0017   1.0000
  12.250   1.4246   0.02921   0.02394  -0.0326   0.0016   1.0000
  12.500   1.4218   0.03144   0.02631  -0.0304   0.0017   1.0000
  12.750   1.4281   0.03299   0.02794  -0.0292   0.0015   1.0000
  13.000   1.4249   0.03551   0.03060  -0.0277   0.0016   1.0000
  13.250   1.4227   0.03806   0.03329  -0.0265   0.0016   1.0000
  13.500   1.4253   0.04027   0.03559  -0.0258   0.0015   1.0000
  13.750   1.4266   0.04268   0.03810  -0.0253   0.0014   1.0000
  14.000   1.4253   0.04550   0.04104  -0.0249   0.0014   1.0000
  14.250   1.4276   0.04803   0.04365  -0.0249   0.0013   1.0000
  14.500   1.4115   0.05296   0.04878  -0.0250   0.0014   1.0000
  14.750   1.4142   0.05574   0.05163  -0.0256   0.0013   1.0000
  15.000   1.4071   0.05996   0.05598  -0.0265   0.0013   1.0000
  15.250   1.3961   0.06485   0.06103  -0.0276   0.0013   1.0000
  15.500   1.3889   0.06941   0.06570  -0.0289   0.0013   1.0000
  15.750   1.3787   0.07454   0.07097  -0.0305   0.0013   1.0000
  16.000   1.3674   0.08000   0.07656  -0.0325   0.0012   1.0000
  16.250   1.3531   0.08617   0.08287  -0.0348   0.0012   1.0000
  16.500   1.3363   0.09300   0.08985  -0.0376   0.0012   1.0000
<< Back to GOE 517 AIRFOIL (goe517-il)

Polar data table (+)

Polar graphs


<< Back to GOE 517 AIRFOIL (goe517-il)