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GOE 506 AIRFOIL (goe506-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: GOE 506 AIRFOIL (goe506-il)
Reynolds number: 50,000
Max Cl/Cd: 10.2 at α=3.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe506-il-50000.txt
Download as CSV file: xf-goe506-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 506 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
  ------ -------- --------- --------- -------- -------- --------
  -7.750  -0.4343   0.11874   0.11333  -0.0097   1.0000   0.2779
  -7.500  -0.4363   0.11543   0.11005  -0.0085   1.0000   0.2786
  -7.250  -0.5680   0.09776   0.09249  -0.0228   1.0000   0.1655
  -7.000  -0.5486   0.09534   0.09011  -0.0191   1.0000   0.1613
  -6.750  -0.6209   0.07845   0.07286  -0.0280   1.0000   0.1368
  -6.500  -0.6220   0.07416   0.06849  -0.0270   1.0000   0.1353
  -6.250  -0.6251   0.06894   0.06308  -0.0266   1.0000   0.1324
  -6.000  -0.6464   0.05725   0.05003  -0.0280   1.0000   0.1238
  -5.750  -0.6371   0.05351   0.04595  -0.0266   1.0000   0.1233
  -5.500  -0.6260   0.05004   0.04204  -0.0252   1.0000   0.1229
  -5.250  -0.6136   0.04698   0.03837  -0.0237   1.0000   0.1237
  -5.000  -0.5980   0.04501   0.03635  -0.0223   1.0000   0.1266
  -4.750  -0.5816   0.04337   0.03453  -0.0209   1.0000   0.1297
  -4.500  -0.5641   0.04154   0.03236  -0.0195   1.0000   0.1324
  -4.250  -0.5454   0.03988   0.03029  -0.0181   1.0000   0.1360
  -4.000  -0.5273   0.03887   0.02925  -0.0168   1.0000   0.1414
  -3.750  -0.5089   0.03807   0.02832  -0.0155   1.0000   0.1488
  -3.500  -0.4904   0.03736   0.02763  -0.0141   1.0000   0.1568
  -3.250  -0.4714   0.03676   0.02698  -0.0128   1.0000   0.1678
  -3.000  -0.4538   0.03648   0.02680  -0.0116   1.0000   0.1862
  -2.750  -0.4355   0.03611   0.02658  -0.0103   1.0000   0.2186
  -2.500  -0.4216   0.03573   0.02664  -0.0081   1.0000   0.3002
  -2.250  -0.4135   0.03619   0.02725  -0.0051   1.0000   0.3655
  -2.000  -0.4033   0.03671   0.02783  -0.0027   1.0000   0.4049
  -1.750  -0.3926   0.03719   0.02836  -0.0007   1.0000   0.4358
  -1.500  -0.3816   0.03769   0.02883   0.0010   1.0000   0.4636
  -1.250  -0.2720   0.04181   0.03304  -0.0127   0.9366   0.5440
  -1.000  -0.2157   0.04262   0.03376  -0.0178   0.9078   0.5788
  -0.750  -0.1632   0.04295   0.03400  -0.0226   0.8834   0.6061
  -0.500  -0.1116   0.04312   0.03424  -0.0268   0.8621   0.6499
  -0.250   0.0553   0.04310   0.03520  -0.0531   0.8381   1.0000
   0.000   0.1040   0.04335   0.03477  -0.0576   0.8188   1.0000
   0.250   0.1369   0.04354   0.03461  -0.0588   0.7973   1.0000
   0.500   0.1724   0.04366   0.03442  -0.0601   0.7770   1.0000
   0.750   0.2114   0.04367   0.03417  -0.0617   0.7592   1.0000
   1.000   0.2497   0.04360   0.03386  -0.0630   0.7423   1.0000
   1.250   0.2848   0.04351   0.03357  -0.0638   0.7260   1.0000
   1.500   0.3177   0.04339   0.03328  -0.0642   0.7101   1.0000
   1.750   0.3453   0.04348   0.03321  -0.0639   0.6947   1.0000
   2.000   0.3687   0.04378   0.03337  -0.0632   0.6797   1.0000
   2.250   0.3861   0.04447   0.03394  -0.0621   0.6655   1.0000
   2.500   0.4012   0.04542   0.03477  -0.0608   0.6527   1.0000
   2.750   0.4470   0.04472   0.03394  -0.0624   0.6466   1.0000
   3.000   0.4488   0.04651   0.03565  -0.0601   0.6334   1.0000
   3.250   0.4520   0.04839   0.03745  -0.0580   0.6220   1.0000
   3.500   0.4904   0.04809   0.03706  -0.0588   0.6165   1.0000
   3.750   0.4730   0.05170   0.04061  -0.0557   0.6056   1.0000
   4.000   0.5060   0.05196   0.04080  -0.0563   0.6011   1.0000
   4.250   0.4754   0.05688   0.04571  -0.0529   0.5933   1.0000
   4.500   0.4852   0.05894   0.04772  -0.0522   0.5883   1.0000
   4.750   0.5247   0.05890   0.04762  -0.0531   0.5845   1.0000
   5.000   0.4811   0.06493   0.05365  -0.0498   0.5801   1.0000
   5.250   0.4715   0.06864   0.05735  -0.0486   0.5789   1.0000
   5.500   0.4659   0.07214   0.06085  -0.0477   0.5791   1.0000
   5.750   0.4630   0.07528   0.06398  -0.0469   0.5782   1.0000
   6.000   0.4631   0.07815   0.06684  -0.0462   0.5767   1.0000
   6.250   0.4973   0.07872   0.06738  -0.0466   0.5674   1.0000
   6.500   0.4906   0.08204   0.07071  -0.0457   0.5663   1.0000
   6.750   0.4875   0.08469   0.07336  -0.0446   0.5610   1.0000
   7.000   0.5104   0.08547   0.07412  -0.0440   0.5472   1.0000
   7.250   0.5395   0.08599   0.07463  -0.0437   0.5344   1.0000
   7.500   0.5376   0.08876   0.07743  -0.0428   0.5280   1.0000
   7.750   0.5649   0.08990   0.07858  -0.0428   0.5193   1.0000
   8.000   0.5525   0.09378   0.08250  -0.0421   0.5163   1.0000
   8.250   0.5543   0.09675   0.08551  -0.0418   0.5111   1.0000
   8.500   0.5678   0.09875   0.08754  -0.0413   0.5000   1.0000
   8.750   0.5965   0.09969   0.08851  -0.0410   0.4866   1.0000
   9.000   0.5798   0.10393   0.09280  -0.0405   0.4796   1.0000
   9.250   0.5969   0.10599   0.09488  -0.0402   0.4683   1.0000
   9.500   0.5871   0.11087   0.09982  -0.0407   0.4683   1.0000
   9.750   0.5248   0.12469   0.11380  -0.0463   0.5536   1.0000
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