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GOE 511 AIRFOIL (goe511-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: GOE 511 AIRFOIL (goe511-il)
Reynolds number: 100,000
Max Cl/Cd: 41.42 at α=2.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe511-il-100000.txt
Download as CSV file: xf-goe511-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 511 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  -0.0269   0.10403   0.09973  -0.0763   0.9093   0.1154
  -7.750  -0.0232   0.10131   0.09702  -0.0785   0.8967   0.1198
  -7.500  -0.0593   0.10066   0.09640  -0.0820   0.8789   0.1223
  -7.250  -0.0270   0.09466   0.09039  -0.0852   0.8749   0.1246
  -7.000   0.0879   0.07778   0.07366  -0.0943   0.8397   0.1392
  -6.750   0.1117   0.07435   0.07023  -0.0926   0.8272   0.1431
  -6.500   0.1120   0.07102   0.06688  -0.0934   0.8178   0.1480
  -6.250   0.0122   0.08138   0.07705  -0.0876   0.8278   0.1409
  -6.000   0.0409   0.07737   0.07298  -0.0906   0.8231   0.1482
  -5.500   0.0477   0.07142   0.06693  -0.0895   0.7967   0.1562
  -5.250   0.0762   0.06789   0.06330  -0.0918   0.7908   0.1639
  -5.000   0.0799   0.06518   0.06043  -0.0921   0.7779   0.1702
  -4.750   0.1207   0.06094   0.05615  -0.0947   0.7732   0.1761
  -4.500   0.1314   0.05887   0.05384  -0.0955   0.7605   0.1867
  -4.250   0.1814   0.05461   0.04951  -0.0999   0.7551   0.1960
  -4.000   0.1988   0.05233   0.04701  -0.1005   0.7421   0.2056
  -3.500   0.2299   0.03901   0.03201  -0.1012   0.7235   0.1228
  -3.250   0.2856   0.03521   0.02803  -0.1069   0.7156   0.1164
  -3.000   0.2860   0.03156   0.02349  -0.1014   0.7032   0.1081
  -2.750   0.3213   0.02946   0.02089  -0.1025   0.6929   0.1079
  -2.500   0.3515   0.02790   0.01892  -0.1025   0.6812   0.1074
  -2.250   0.3754   0.02690   0.01758  -0.1013   0.6698   0.1077
  -2.000   0.4189   0.02578   0.01598  -0.1037   0.6597   0.1092
  -1.750   0.4352   0.02520   0.01543  -0.1015   0.6482   0.1123
  -1.500   0.4807   0.02446   0.01446  -0.1045   0.6387   0.1184
  -1.250   0.4946   0.02423   0.01414  -0.1015   0.6274   0.1219
  -1.000   0.5339   0.02360   0.01347  -0.1034   0.6182   0.1293
  -0.750   0.5521   0.02344   0.01330  -0.1013   0.6079   0.1385
  -0.500   0.5890   0.02323   0.01318  -0.1028   0.5995   0.1621
  -0.250   0.6032   0.02340   0.01353  -0.0998   0.5900   0.2055
   0.000   0.6302   0.02371   0.01374  -0.0991   0.5828   0.2918
   0.250   0.6342   0.02406   0.01415  -0.0945   0.5738   0.3169
   0.750   0.6879   0.02431   0.01433  -0.0942   0.5577   0.3780
   1.000   0.7215   0.02433   0.01435  -0.0955   0.5494   0.4117
   1.250   0.7629   0.02436   0.01433  -0.0984   0.5424   0.4451
   1.500   0.7748   0.02459   0.01469  -0.0956   0.5350   0.4642
   1.750   0.8024   0.02450   0.01467  -0.0957   0.5290   0.5003
   2.000   1.0410   0.02548   0.01632  -0.1383   0.5129   1.0000
   2.250   1.0710   0.02586   0.01643  -0.1388   0.5077   1.0000
   2.500   1.0839   0.02650   0.01703  -0.1362   0.5026   1.0000
   2.750   1.0898   0.02715   0.01769  -0.1323   0.4972   1.0000
   3.000   1.1062   0.02768   0.01814  -0.1303   0.4926   1.0000
   3.250   1.1327   0.02815   0.01845  -0.1302   0.4887   1.0000
   3.500   1.1505   0.02887   0.01909  -0.1287   0.4848   1.0000
   3.750   1.1481   0.02974   0.02008  -0.1234   0.4807   1.0000
   4.000   1.1543   0.03051   0.02088  -0.1197   0.4765   1.0000
   4.250   1.1700   0.03112   0.02144  -0.1177   0.4728   1.0000
   4.500   1.1964   0.03161   0.02180  -0.1177   0.4694   1.0000
   4.750   1.2133   0.03243   0.02257  -0.1161   0.4661   1.0000
   5.000   1.1959   0.03367   0.02399  -0.1083   0.4624   1.0000
   5.250   1.1896   0.03483   0.02524  -0.1027   0.4590   1.0000
   5.500   1.1938   0.03583   0.02627  -0.0990   0.4559   1.0000
   5.750   1.2081   0.03660   0.02703  -0.0970   0.4530   1.0000
   6.000   1.2344   0.03724   0.02758  -0.0971   0.4505   1.0000
   6.250   1.2571   0.03818   0.02847  -0.0967   0.4481   1.0000
   6.500   1.1763   0.04105   0.03166  -0.0792   0.4452   1.0000
   6.750   0.6726   0.08501   0.07690  -0.0430   0.4291   1.0000
   7.000   0.6477   0.09182   0.08378  -0.0428   0.4308   1.0000
   7.250   0.6494   0.09551   0.08745  -0.0425   0.4304   1.0000
   7.500   0.6520   0.09921   0.09114  -0.0422   0.4301   1.0000
   7.750   0.5547   0.11604   0.10833  -0.0450   0.4684   1.0000
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