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GOE 492 AIRFOIL (goe492-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: GOE 492 AIRFOIL (goe492-il)
Reynolds number: 50,000
Max Cl/Cd: 31.13 at α=5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe492-il-50000.txt
Download as CSV file: xf-goe492-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 492 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
  ------ -------- --------- --------- -------- -------- --------
  -7.500  -0.4151   0.09792   0.09135  -0.0226   1.0000   0.2107
  -7.250  -0.4077   0.09411   0.08759  -0.0212   1.0000   0.2186
  -7.000  -0.4258   0.09391   0.08758  -0.0217   1.0000   0.2252
  -6.750  -0.4149   0.08964   0.08327  -0.0193   1.0000   0.2360
  -6.500  -0.4158   0.08688   0.08061  -0.0185   1.0000   0.2440
  -6.250  -0.4274   0.08583   0.07969  -0.0204   1.0000   0.2537
  -6.000  -0.4239   0.08279   0.07671  -0.0188   1.0000   0.2662
  -5.750  -0.4192   0.07940   0.07337  -0.0165   1.0000   0.2783
  -5.500  -0.4169   0.07641   0.07044  -0.0151   1.0000   0.2905
  -5.000  -0.4154   0.07104   0.06519  -0.0145   1.0000   0.3225
  -4.750  -0.4134   0.06852   0.06271  -0.0134   1.0000   0.3487
  -4.500  -0.4105   0.06553   0.05978  -0.0095   1.0000   0.3767
  -3.750  -0.2639   0.04182   0.03410  -0.0555   1.0000   0.1570
  -3.500  -0.2305   0.03658   0.02824  -0.0584   1.0000   0.1412
  -3.250  -0.1997   0.03283   0.02386  -0.0601   1.0000   0.1402
  -3.000  -0.1705   0.02999   0.02043  -0.0609   1.0000   0.1474
  -2.750  -0.1415   0.02749   0.01737  -0.0611   1.0000   0.1512
  -2.500  -0.1147   0.02585   0.01540  -0.0609   1.0000   0.1645
  -2.250  -0.0881   0.02451   0.01379  -0.0605   1.0000   0.1796
  -2.000  -0.0623   0.02335   0.01241  -0.0598   1.0000   0.1943
  -1.750  -0.0371   0.02266   0.01163  -0.0591   1.0000   0.2225
  -1.500  -0.0106   0.02192   0.01090  -0.0587   1.0000   0.2786
  -1.250   0.0155   0.01834   0.00950  -0.0571   1.0000   0.6570
  -1.000   0.0372   0.01775   0.00877  -0.0559   1.0000   1.0000
  -0.750   0.0600   0.01805   0.00860  -0.0555   1.0000   1.0000
  -0.500   0.0817   0.01840   0.00858  -0.0551   1.0000   1.0000
  -0.250   0.1028   0.01877   0.00867  -0.0546   1.0000   1.0000
   0.000   0.1234   0.01919   0.00885  -0.0542   1.0000   1.0000
   0.250   0.1435   0.01966   0.00907  -0.0537   1.0000   1.0000
   0.500   0.1631   0.02017   0.00941  -0.0532   1.0000   1.0000
   0.750   0.1824   0.02072   0.00982  -0.0528   1.0000   1.0000
   1.000   0.2013   0.02131   0.01030  -0.0524   1.0000   1.0000
   1.250   0.2199   0.02196   0.01085  -0.0520   1.0000   1.0000
   1.500   0.2383   0.02264   0.01147  -0.0517   1.0000   1.0000
   1.750   0.2563   0.02338   0.01215  -0.0514   1.0000   1.0000
   2.000   0.2741   0.02415   0.01289  -0.0511   1.0000   1.0000
   2.250   0.2917   0.02498   0.01370  -0.0509   1.0000   1.0000
   2.500   0.3091   0.02585   0.01458  -0.0507   1.0000   1.0000
   2.750   0.3413   0.02706   0.01584  -0.0535   0.9927   1.0000
   3.000   0.3843   0.02845   0.01736  -0.0582   0.9788   1.0000
   3.250   0.4245   0.02974   0.01878  -0.0622   0.9650   1.0000
   3.500   0.4626   0.03098   0.02017  -0.0657   0.9510   1.0000
   3.750   0.4990   0.03218   0.02156  -0.0687   0.9369   1.0000
   4.000   0.5341   0.03336   0.02304  -0.0715   0.9226   1.0000
   4.250   0.5679   0.03454   0.02448  -0.0739   0.9083   1.0000
   4.500   0.6009   0.03572   0.02596  -0.0761   0.8937   1.0000
   4.750   0.6352   0.03687   0.02747  -0.0783   0.8783   1.0000
   5.000   0.8019   0.02576   0.01380  -0.0662   0.1364   1.0000
   5.250   0.8182   0.02788   0.01590  -0.0638   0.1191   1.0000
   5.500   0.8479   0.03008   0.01811  -0.0629   0.1073   1.0000
   5.750   0.8868   0.03256   0.02067  -0.0633   0.0947   1.0000
   6.000   0.9277   0.03549   0.02371  -0.0638   0.0910   1.0000
   6.250   0.9617   0.03859   0.02719  -0.0633   0.0914   1.0000
   6.500   0.9906   0.04191   0.03106  -0.0620   0.0943   1.0000
   6.750   1.0156   0.04565   0.03525  -0.0606   0.0974   1.0000
   7.000   1.0361   0.04952   0.03954  -0.0590   0.0990   1.0000
   7.250   1.0557   0.05415   0.04443  -0.0577   0.1005   1.0000
   7.500   1.0648   0.05698   0.04834  -0.0542   0.1102   1.0000
   7.750   1.0777   0.06182   0.05380  -0.0521   0.1250   1.0000
   8.000   1.0956   0.06858   0.06111  -0.0512   0.1542   1.0000
   8.250   1.0725   0.07318   0.06667  -0.0493   0.1760   1.0000
   8.500   1.0559   0.08076   0.07471  -0.0508   0.2098   1.0000
   8.750   1.0172   0.08654   0.08062  -0.0518   0.2172   1.0000
   9.000   0.9701   0.09475   0.08881  -0.0573   0.2386   1.0000
   9.250   0.9168   0.10673   0.10061  -0.0702   0.2801   1.0000
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