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GOE 564 AIRFOIL (goe564-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: GOE 564 AIRFOIL (goe564-il)
Reynolds number: 50,000
Max Cl/Cd: 37.13 at α=5.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe564-il-50000.txt
Download as CSV file: xf-goe564-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 564 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.250  -0.4543   0.11508   0.10785  -0.0224   1.0000   0.1997
  -9.000  -0.4273   0.10885   0.10155  -0.0201   1.0000   0.2116
  -8.750  -0.4546   0.10919   0.10209  -0.0218   1.0000   0.2148
  -8.500  -0.4259   0.10305   0.09586  -0.0195   1.0000   0.2266
  -8.250  -0.4322   0.10047   0.09340  -0.0195   1.0000   0.2324
  -8.000  -0.4331   0.09806   0.09106  -0.0186   1.0000   0.2431
  -7.750  -0.4245   0.09441   0.08744  -0.0171   1.0000   0.2542
  -7.500  -0.4295   0.09191   0.08504  -0.0159   1.0000   0.2649
  -7.250  -0.4517   0.09110   0.08441  -0.0144   1.0000   0.2743
  -7.000  -0.4533   0.08840   0.08179  -0.0127   1.0000   0.2894
  -6.750  -0.4564   0.08585   0.07933  -0.0109   1.0000   0.3047
  -6.500  -0.4632   0.08360   0.07717  -0.0094   1.0000   0.3204
  -6.250  -0.4471   0.07960   0.07319  -0.0056   1.0000   0.3394
  -6.000  -0.4442   0.07703   0.07068  -0.0027   1.0000   0.3608
  -5.750  -0.4519   0.07486   0.06861   0.0001   1.0000   0.3834
  -5.500  -0.4405   0.07211   0.06587   0.0050   1.0000   0.4181
  -5.250  -0.4516   0.07077   0.06464   0.0098   1.0000   0.4561
  -5.000  -0.4234   0.06742   0.06126   0.0161   1.0000   0.5125
  -4.250  -0.3891   0.04497   0.03713  -0.0304   1.0000   0.1937
  -4.000  -0.3623   0.03993   0.03109  -0.0327   1.0000   0.1707
  -3.750  -0.3401   0.03720   0.02780  -0.0324   1.0000   0.1694
  -3.500  -0.3206   0.03483   0.02539  -0.0316   1.0000   0.1745
  -3.250  -0.2971   0.03250   0.02261  -0.0311   1.0000   0.1754
  -3.000  -0.2731   0.03049   0.02019  -0.0304   1.0000   0.1783
  -2.750  -0.2488   0.02878   0.01802  -0.0298   1.0000   0.1867
  -2.500  -0.2250   0.02746   0.01640  -0.0290   1.0000   0.2000
  -2.250  -0.2005   0.02603   0.01479  -0.0282   1.0000   0.2149
  -2.000  -0.1768   0.02477   0.01359  -0.0275   1.0000   0.2415
  -1.750  -0.1536   0.02377   0.01257  -0.0265   1.0000   0.2757
  -1.500  -0.1283   0.02290   0.01181  -0.0261   1.0000   0.3184
  -1.250  -0.1030   0.02207   0.01117  -0.0256   1.0000   0.3656
  -1.000  -0.0788   0.02104   0.01069  -0.0249   1.0000   0.4387
  -0.750  -0.0352   0.01897   0.00994  -0.0264   1.0000   1.0000
  -0.500  -0.0152   0.01925   0.00982  -0.0254   1.0000   1.0000
  -0.250   0.0041   0.01957   0.00985  -0.0244   1.0000   1.0000
   0.000   0.0230   0.01993   0.00998  -0.0235   1.0000   1.0000
   0.250   0.0416   0.02033   0.01017  -0.0226   1.0000   1.0000
   0.500   0.0598   0.02077   0.01045  -0.0218   1.0000   1.0000
   0.750   0.0778   0.02125   0.01080  -0.0210   1.0000   1.0000
   1.000   0.0956   0.02178   0.01121  -0.0203   1.0000   1.0000
   1.250   0.1132   0.02235   0.01169  -0.0196   1.0000   1.0000
   1.500   0.1305   0.02297   0.01222  -0.0190   1.0000   1.0000
   1.750   0.1477   0.02363   0.01282  -0.0185   1.0000   1.0000
   2.000   0.1647   0.02433   0.01348  -0.0180   1.0000   1.0000
   2.250   0.1815   0.02509   0.01420  -0.0176   1.0000   1.0000
   2.500   0.1979   0.02590   0.01500  -0.0172   1.0000   1.0000
   2.750   0.2515   0.02757   0.01671  -0.0239   0.9836   1.0000
   3.000   0.3276   0.02892   0.01816  -0.0339   0.9481   1.0000
   3.250   0.3947   0.02967   0.01904  -0.0412   0.9177   1.0000
   3.500   0.4476   0.03009   0.01964  -0.0457   0.8914   1.0000
   3.750   0.5017   0.03020   0.01995  -0.0498   0.8650   1.0000
   4.000   0.5592   0.02990   0.01990  -0.0537   0.8379   1.0000
   4.250   0.6198   0.02899   0.01933  -0.0572   0.8081   1.0000
   4.500   0.6841   0.02736   0.01807  -0.0600   0.7740   1.0000
   4.750   0.7451   0.02523   0.01630  -0.0611   0.7304   1.0000
   5.000   0.8028   0.02300   0.01417  -0.0607   0.6671   1.0000
   5.250   0.8385   0.02259   0.01346  -0.0585   0.5888   1.0000
   5.500   0.8596   0.02315   0.01370  -0.0556   0.5243   1.0000
   5.750   0.8817   0.02390   0.01420  -0.0536   0.4762   1.0000
   6.000   0.9054   0.02493   0.01508  -0.0522   0.4384   1.0000
   6.250   0.9293   0.02611   0.01617  -0.0509   0.4044   1.0000
   6.500   0.9529   0.02742   0.01733  -0.0497   0.3717   1.0000
   6.750   0.9756   0.02895   0.01878  -0.0483   0.3403   1.0000
   7.000   0.9905   0.03021   0.02002  -0.0459   0.3043   1.0000
   7.250   1.0057   0.03139   0.02092  -0.0436   0.2669   1.0000
   7.500   1.0196   0.03313   0.02272  -0.0410   0.2319   1.0000
   7.750   1.0361   0.03529   0.02456  -0.0390   0.1945   1.0000
   8.000   1.0502   0.03745   0.02670  -0.0365   0.1657   1.0000
   8.250   1.0643   0.04002   0.02961  -0.0341   0.1479   1.0000
   8.500   1.0794   0.04239   0.03215  -0.0320   0.1344   1.0000
   8.750   1.0983   0.04494   0.03478  -0.0307   0.1246   1.0000
   9.000   1.1035   0.04864   0.03915  -0.0276   0.1199   1.0000
   9.250   1.1176   0.05120   0.04183  -0.0259   0.1128   1.0000
   9.500   1.1205   0.05498   0.04599  -0.0233   0.1096   1.0000
   9.750   1.1146   0.05929   0.05082  -0.0201   0.1085   1.0000
  10.000   1.1057   0.06372   0.05567  -0.0172   0.1083   1.0000
  10.250   1.0919   0.06829   0.06057  -0.0144   0.1086   1.0000
  10.500   1.0745   0.07283   0.06536  -0.0119   0.1091   1.0000
  10.750   1.0548   0.07718   0.06987  -0.0095   0.1098   1.0000
  11.000   1.0354   0.08190   0.07471  -0.0081   0.1105   1.0000
  11.250   1.0210   0.08705   0.07993  -0.0078   0.1111   1.0000
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