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

GOE 723 AIRFOIL (goe723-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: GOE 723 AIRFOIL (goe723-il)
Reynolds number: 100,000
Max Cl/Cd: 58.42 at α=8°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe723-il-100000.txt
Download as CSV file: xf-goe723-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 723 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
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.2885   0.10459   0.09947  -0.0383   1.0000   0.1226
  -8.750  -0.2957   0.10291   0.09788  -0.0387   1.0000   0.1272
  -8.500  -0.3367   0.10391   0.09910  -0.0402   1.0000   0.1294
  -8.250  -0.3277   0.09956   0.09480  -0.0382   1.0000   0.1309
  -8.000  -0.3082   0.09634   0.09160  -0.0352   1.0000   0.1337
  -7.750  -0.3141   0.09511   0.09047  -0.0321   1.0000   0.1362
  -7.500  -0.3337   0.09486   0.09036  -0.0279   1.0000   0.1382
  -7.250  -0.3610   0.09518   0.09081  -0.0232   1.0000   0.1399
  -7.000  -0.3939   0.09353   0.08927  -0.0383   0.9878   0.1452
  -6.750  -0.3632   0.08996   0.08569  -0.0300   0.9879   0.1480
  -6.500  -0.3346   0.08650   0.08220  -0.0351   0.9803   0.1572
  -6.250  -0.3100   0.08143   0.07711  -0.0439   0.9712   0.1639
  -6.000  -0.2854   0.07836   0.07402  -0.0464   0.9621   0.1707
  -5.750  -0.2649   0.07370   0.06933  -0.0552   0.9510   0.1799
  -5.500  -0.2356   0.06903   0.06455  -0.0667   0.9421   0.1942
  -5.250  -0.2136   0.06653   0.06208  -0.0645   0.9342   0.1975
  -5.000  -0.1779   0.06201   0.05745  -0.0737   0.9285   0.2121
  -4.750  -0.1577   0.05981   0.05522  -0.0750   0.9178   0.2227
  -4.500  -0.1269   0.03455   0.02739  -0.1003   0.9062   0.0995
  -4.250  -0.0919   0.03157   0.02406  -0.1024   0.9004   0.0962
  -4.000  -0.0487   0.02878   0.02069  -0.1054   0.8970   0.0928
  -3.750  -0.0265   0.02750   0.01901  -0.1044   0.8870   0.0914
  -3.500   0.0163   0.02590   0.01711  -0.1069   0.8827   0.0912
  -3.250   0.0617   0.02452   0.01558  -0.1098   0.8796   0.0933
  -3.000   0.0809   0.02409   0.01508  -0.1081   0.8688   0.0958
  -2.750   0.1225   0.02308   0.01395  -0.1102   0.8645   0.0987
  -2.500   0.1461   0.02268   0.01351  -0.1092   0.8554   0.1006
  -2.250   0.1800   0.02181   0.01273  -0.1100   0.8495   0.1050
  -2.000   0.2128   0.02129   0.01220  -0.1105   0.8436   0.1122
  -1.750   0.2353   0.02102   0.01200  -0.1094   0.8346   0.1229
  -1.500   0.2719   0.01983   0.01143  -0.1108   0.8301   0.2287
  -1.250   0.2881   0.01990   0.01198  -0.1086   0.8201   0.3924
  -1.000   0.3197   0.01983   0.01202  -0.1086   0.8145   0.4664
  -0.750   0.3390   0.02010   0.01237  -0.1068   0.8056   0.5135
  -0.500   0.3674   0.02000   0.01235  -0.1063   0.7991   0.5619
  -0.250   0.3956   0.01991   0.01236  -0.1059   0.7926   0.6111
   0.000   0.4200   0.01978   0.01244  -0.1048   0.7841   0.6694
   0.250   0.4603   0.01897   0.01196  -0.1062   0.7795   0.7660
   0.500   0.5343   0.01860   0.01172  -0.1159   0.7707   1.0000
   0.750   0.5683   0.01859   0.01147  -0.1167   0.7640   1.0000
   1.000   0.5903   0.01900   0.01175  -0.1157   0.7545   1.0000
   1.250   0.6215   0.01905   0.01164  -0.1159   0.7476   1.0000
   1.500   0.6435   0.01947   0.01198  -0.1148   0.7387   1.0000
   1.750   0.6728   0.01958   0.01197  -0.1147   0.7314   1.0000
   2.000   0.6960   0.01998   0.01232  -0.1138   0.7231   1.0000
   2.250   0.7232   0.02016   0.01243  -0.1133   0.7153   1.0000
   2.500   0.7481   0.02049   0.01270  -0.1126   0.7074   1.0000
   2.750   0.7734   0.02075   0.01293  -0.1119   0.6991   1.0000
   3.000   0.7998   0.02101   0.01316  -0.1113   0.6915   1.0000
   3.250   0.8235   0.02135   0.01348  -0.1104   0.6826   1.0000
   3.500   0.8505   0.02157   0.01368  -0.1099   0.6751   1.0000
   3.750   0.8738   0.02190   0.01403  -0.1089   0.6657   1.0000
   4.000   0.8997   0.02218   0.01430  -0.1083   0.6579   1.0000
   4.250   0.9239   0.02248   0.01463  -0.1074   0.6487   1.0000
   4.500   0.9479   0.02281   0.01499  -0.1064   0.6399   1.0000
   4.750   0.9763   0.02281   0.01498  -0.1059   0.6308   1.0000
   5.000   0.9974   0.02315   0.01537  -0.1045   0.6197   1.0000
   5.250   1.0290   0.02276   0.01494  -0.1041   0.6091   1.0000
   5.500   1.0580   0.02237   0.01451  -0.1033   0.5964   1.0000
   5.750   1.0801   0.02232   0.01450  -0.1017   0.5825   1.0000
   6.000   1.1034   0.02227   0.01449  -0.1003   0.5693   1.0000
   6.250   1.1284   0.02211   0.01434  -0.0991   0.5561   1.0000
   6.500   1.1541   0.02190   0.01412  -0.0980   0.5427   1.0000
   6.750   1.1800   0.02171   0.01393  -0.0969   0.5293   1.0000
   7.000   1.2014   0.02168   0.01395  -0.0952   0.5146   1.0000
   7.250   1.2211   0.02167   0.01400  -0.0933   0.4989   1.0000
   7.500   1.2402   0.02166   0.01405  -0.0912   0.4823   1.0000
   7.750   1.2541   0.02170   0.01421  -0.0883   0.4612   1.0000
   8.000   1.2666   0.02168   0.01420  -0.0850   0.4346   1.0000
   8.250   1.2731   0.02192   0.01443  -0.0809   0.3995   1.0000
   8.500   1.2701   0.02258   0.01486  -0.0753   0.3441   1.0000
   8.750   1.2539   0.02403   0.01582  -0.0682   0.2689   1.0000
   9.000   1.2338   0.02618   0.01738  -0.0617   0.1970   1.0000
   9.250   1.2205   0.02847   0.01928  -0.0569   0.1655   1.0000
   9.500   1.2141   0.03058   0.02124  -0.0533   0.1508   1.0000
   9.750   1.2109   0.03265   0.02320  -0.0503   0.1420   1.0000
  10.000   1.2127   0.03450   0.02503  -0.0479   0.1347   1.0000
  10.250   1.2178   0.03629   0.02672  -0.0457   0.1285   1.0000
  10.500   1.2289   0.03774   0.02817  -0.0440   0.1223   1.0000
  10.750   1.2519   0.03903   0.02921  -0.0428   0.1159   1.0000
  11.000   1.2696   0.04032   0.03064  -0.0415   0.1108   1.0000
  11.250   1.3038   0.04160   0.03173  -0.0416   0.1043   1.0000
  11.500   1.3308   0.04324   0.03348  -0.0412   0.0994   1.0000
  11.750   1.3503   0.04485   0.03524  -0.0403   0.0949   1.0000
  12.000   1.3994   0.04744   0.03769  -0.0427   0.0887   1.0000
  12.250   1.4044   0.04920   0.03975  -0.0403   0.0862   1.0000
  12.500   1.4155   0.05121   0.04198  -0.0388   0.0831   1.0000
  12.750   1.4318   0.05333   0.04418  -0.0380   0.0798   1.0000
  13.000   1.4577   0.05709   0.04806  -0.0385   0.0767   1.0000
  13.250   1.4486   0.05939   0.05071  -0.0352   0.0757   1.0000
  13.500   1.4392   0.06205   0.05370  -0.0325   0.0745   1.0000
  13.750   1.4298   0.06504   0.05699  -0.0302   0.0733   1.0000
  14.000   1.4196   0.06828   0.06050  -0.0284   0.0723   1.0000
  14.250   1.4089   0.07165   0.06412  -0.0269   0.0712   1.0000
  14.500   1.3955   0.07551   0.06824  -0.0258   0.0707   1.0000
  14.750   1.3777   0.07990   0.07289  -0.0250   0.0704   1.0000
  15.000   1.3552   0.08484   0.07811  -0.0248   0.0703   1.0000
  15.250   1.3247   0.09084   0.08441  -0.0256   0.0709   1.0000
  15.500   1.2900   0.09777   0.09162  -0.0275   0.0716   1.0000
  15.750   1.2514   0.10588   0.10000  -0.0307   0.0725   1.0000
  16.000   1.2114   0.11527   0.10962  -0.0354   0.0736   1.0000
  16.250   1.1716   0.12604   0.12057  -0.0414   0.0749   1.0000
  16.500   1.1389   0.13702   0.13164  -0.0476   0.0762   1.0000
  16.750   1.1214   0.14621   0.14088  -0.0521   0.0771   1.0000
<< Back to GOE 723 AIRFOIL (goe723-il)

Polar data table (+)

Polar graphs


<< Back to GOE 723 AIRFOIL (goe723-il)