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GOE 456 AIRFOIL (goe456-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: GOE 456 AIRFOIL (goe456-il)
Reynolds number: 200,000
Max Cl/Cd: 87.49 at α=4.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe456-il-200000.txt
Download as CSV file: xf-goe456-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 456 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.750  -0.3342   0.09970   0.09660  -0.0288   1.0000   0.0304
  -7.500  -0.3503   0.09928   0.09627  -0.0257   1.0000   0.0306
  -7.250  -0.3639   0.09821   0.09526  -0.0225   1.0000   0.0305
  -7.000  -0.3642   0.09642   0.09352  -0.0255   0.9980   0.0309
  -6.500  -0.3077   0.08565   0.08273  -0.0441   0.9868   0.0314
  -6.250  -0.2985   0.07959   0.07667  -0.0437   0.9841   0.0330
  -6.000  -0.2773   0.07592   0.07298  -0.0466   0.9792   0.0345
  -5.750  -0.2507   0.07173   0.06876  -0.0527   0.9735   0.0366
  -5.500  -0.2138   0.06670   0.06363  -0.0623   0.9697   0.0396
  -5.250  -0.1613   0.06084   0.05758  -0.0788   0.9607   0.0439
  -5.000  -0.1315   0.05385   0.05052  -0.0860   0.9572   0.0455
  -4.750  -0.1124   0.05140   0.04807  -0.0864   0.9507   0.0478
  -4.500  -0.0738   0.04722   0.04374  -0.0928   0.9461   0.0536
  -4.250  -0.0281   0.04160   0.03790  -0.1014   0.9433   0.0611
  -4.000   0.0054   0.03740   0.03339  -0.1060   0.9357   0.0725
  -3.750   0.0409   0.03526   0.03119  -0.1090   0.9322   0.0874
  -3.500   0.0859   0.03228   0.02785  -0.1143   0.9300   0.1114
  -3.250   0.1270   0.02093   0.01488  -0.1160   0.9242   0.0403
  -3.000   0.1641   0.01829   0.01159  -0.1172   0.9199   0.0365
  -2.750   0.2038   0.01652   0.00947  -0.1189   0.9170   0.0365
  -2.500   0.2345   0.01546   0.00828  -0.1190   0.9103   0.0401
  -2.250   0.2692   0.01441   0.00711  -0.1199   0.9050   0.0453
  -2.000   0.3026   0.01340   0.00603  -0.1207   0.8995   0.0549
  -1.750   0.3324   0.01255   0.00556  -0.1207   0.8918   0.1744
  -1.500   0.3637   0.01251   0.00545  -0.1212   0.8847   0.2203
  -1.250   0.3941   0.01235   0.00529  -0.1215   0.8768   0.2505
  -1.000   0.4219   0.01217   0.00510  -0.1213   0.8674   0.2737
  -0.750   0.4532   0.01181   0.00473  -0.1217   0.8606   0.2925
  -0.500   0.4789   0.01154   0.00454  -0.1211   0.8499   0.3200
  -0.250   0.5046   0.01113   0.00438  -0.1206   0.8397   0.3862
   0.000   0.5493   0.00958   0.00409  -0.1237   0.8336   1.0000
   0.250   0.5749   0.00964   0.00400  -0.1230   0.8222   1.0000
   0.500   0.6011   0.00971   0.00394  -0.1225   0.8112   1.0000
   0.750   0.6277   0.00977   0.00389  -0.1220   0.8008   1.0000
   1.000   0.6547   0.00982   0.00382  -0.1216   0.7902   1.0000
   1.250   0.6814   0.00986   0.00376  -0.1211   0.7786   1.0000
   1.500   0.7070   0.00990   0.00371  -0.1204   0.7657   1.0000
   1.750   0.7325   0.00997   0.00371  -0.1197   0.7529   1.0000
   2.000   0.7579   0.01005   0.00373  -0.1190   0.7402   1.0000
   2.250   0.7834   0.01014   0.00377  -0.1184   0.7275   1.0000
   2.500   0.8087   0.01023   0.00384  -0.1177   0.7142   1.0000
   2.750   0.8341   0.01034   0.00392  -0.1170   0.7014   1.0000
   3.000   0.8597   0.01048   0.00404  -0.1165   0.6903   1.0000
   3.250   0.8856   0.01063   0.00419  -0.1160   0.6797   1.0000
   3.500   0.9116   0.01080   0.00438  -0.1155   0.6694   1.0000
   3.750   0.9359   0.01095   0.00453  -0.1147   0.6542   1.0000
   4.000   0.9578   0.01106   0.00456  -0.1131   0.6275   1.0000
   4.250   0.9808   0.01123   0.00475  -0.1120   0.6073   1.0000
   4.500   1.0000   0.01143   0.00482  -0.1099   0.5677   1.0000
   4.750   1.0189   0.01170   0.00502  -0.1079   0.5239   1.0000
   5.000   1.0309   0.01233   0.00522  -0.1047   0.4242   1.0000
   5.250   1.0277   0.01445   0.00614  -0.0996   0.2368   1.0000
   5.500   1.0261   0.01694   0.00745  -0.0953   0.0525   1.0000
   5.750   1.0416   0.01802   0.00847  -0.0930   0.0336   1.0000
   6.000   1.0594   0.01890   0.00955  -0.0910   0.0300   1.0000
   6.250   1.0771   0.01974   0.01059  -0.0891   0.0290   1.0000
   6.500   1.0924   0.02072   0.01172  -0.0869   0.0264   1.0000
   6.750   1.1046   0.02185   0.01301  -0.0843   0.0248   1.0000
   7.000   1.1141   0.02315   0.01441  -0.0813   0.0242   1.0000
   7.250   1.1231   0.02451   0.01583  -0.0781   0.0241   1.0000
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