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

GOE 79 (PFALZ 11) AIRFOIL (goe79-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: GOE 79 (PFALZ 11) AIRFOIL (goe79-il)
Reynolds number: 200,000
Max Cl/Cd: 75.6 at α=6.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe79-il-200000-n5.txt
Download as CSV file: xf-goe79-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 79 (PFALZ 11) 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.2890   0.09897   0.09542  -0.0144   1.0000   0.0233
  -7.750  -0.2833   0.09651   0.09300  -0.0152   1.0000   0.0237
  -7.500  -0.2783   0.09418   0.09072  -0.0160   1.0000   0.0240
  -7.250  -0.2705   0.09174   0.08832  -0.0179   1.0000   0.0243
  -7.000  -0.2619   0.08945   0.08608  -0.0204   1.0000   0.0245
  -6.750  -0.2468   0.08718   0.08383  -0.0260   0.9973   0.0246
  -6.500  -0.2237   0.08288   0.07954  -0.0298   0.9869   0.0247
  -6.000  -0.1755   0.07532   0.07195  -0.0379   0.9542   0.0253
  -5.750  -0.1448   0.07167   0.06825  -0.0439   0.9371   0.0258
  -5.500  -0.1121   0.06806   0.06458  -0.0502   0.9214   0.0264
  -5.250  -0.0791   0.06466   0.06109  -0.0562   0.9056   0.0273
  -5.000  -0.0351   0.06158   0.05784  -0.0655   0.8889   0.0281
  -4.750   0.0055   0.05863   0.05470  -0.0729   0.8726   0.0282
  -4.500   0.0257   0.05544   0.05146  -0.0740   0.8571   0.0284
  -4.250   0.0451   0.05259   0.04857  -0.0748   0.8422   0.0286
  -4.000   0.0698   0.05004   0.04592  -0.0768   0.8266   0.0291
  -3.750   0.0975   0.04761   0.04337  -0.0793   0.8094   0.0296
  -3.500   0.1271   0.04527   0.04089  -0.0820   0.7908   0.0306
  -3.250   0.1663   0.04316   0.03853  -0.0860   0.7718   0.0322
  -3.000   0.2093   0.04123   0.03626  -0.0901   0.7540   0.0325
  -2.750   0.2340   0.03861   0.03354  -0.0913   0.7366   0.0327
  -2.500   0.2578   0.03639   0.03122  -0.0922   0.7191   0.0332
  -2.250   0.2858   0.03452   0.02920  -0.0934   0.7015   0.0338
  -2.000   0.3161   0.03280   0.02728  -0.0948   0.6848   0.0348
  -1.750   0.3486   0.03120   0.02546  -0.0963   0.6698   0.0363
  -1.500   0.3867   0.02973   0.02363  -0.0979   0.6574   0.0379
  -1.250   0.4131   0.02800   0.02184  -0.0988   0.6446   0.0387
  -1.000   0.4427   0.02666   0.02033  -0.0996   0.6329   0.0398
  -0.750   0.4738   0.02546   0.01892  -0.1003   0.6218   0.0416
  -0.500   0.5099   0.02464   0.01770  -0.1008   0.6107   0.0436
  -0.250   0.5370   0.02306   0.01609  -0.1016   0.5999   0.0445
   0.250   0.5983   0.02135   0.01392  -0.1022   0.5772   0.0503
   0.500   0.6260   0.02031   0.01283  -0.1027   0.5660   0.0527
   0.750   0.6553   0.01960   0.01195  -0.1029   0.5541   0.0594
   1.000   0.6842   0.01896   0.01117  -0.1031   0.5410   0.0677
   1.250   0.7125   0.01840   0.01048  -0.1032   0.5261   0.0774
   1.500   0.7403   0.01792   0.00988  -0.1033   0.5088   0.0902
   1.750   0.7728   0.01713   0.00861  -0.1021   0.4921   0.0491
   2.000   0.8011   0.01643   0.00774  -0.1018   0.4711   0.0412
   2.250   0.8290   0.01602   0.00708  -0.1014   0.4467   0.0387
   2.500   0.8559   0.01579   0.00668  -0.1011   0.4197   0.0391
   2.750   0.8825   0.01568   0.00641  -0.1008   0.3939   0.0408
   3.000   0.9089   0.01559   0.00614  -0.1004   0.3721   0.0404
   3.250   0.9352   0.01555   0.00598  -0.1000   0.3544   0.0400
   3.500   0.9615   0.01557   0.00590  -0.0997   0.3389   0.0399
   3.750   0.9877   0.01561   0.00588  -0.0993   0.3250   0.0402
   4.000   1.0139   0.01570   0.00592  -0.0990   0.3133   0.0408
   4.250   1.0400   0.01586   0.00604  -0.0987   0.3037   0.0431
   4.500   1.0665   0.01592   0.00612  -0.0985   0.2955   0.0449
   4.750   1.0926   0.01611   0.00626  -0.0982   0.2875   0.0461
   5.000   1.1187   0.01631   0.00644  -0.0979   0.2807   0.0477
   5.250   1.1446   0.01652   0.00664  -0.0975   0.2740   0.0502
   5.750   1.1949   0.01607   0.00729  -0.0968   0.2619   0.7867
   6.250   1.2433   0.01648   0.00781  -0.0953   0.2490   1.0000
   6.500   1.2686   0.01678   0.00812  -0.0949   0.2420   1.0000
   6.750   1.2931   0.01718   0.00848  -0.0944   0.2347   1.0000
   7.000   1.3183   0.01748   0.00884  -0.0939   0.2276   1.0000
   7.250   1.3424   0.01788   0.00922  -0.0934   0.2198   1.0000
   7.500   1.3671   0.01822   0.00962  -0.0929   0.2113   1.0000
   7.750   1.3906   0.01865   0.01006  -0.0924   0.2006   1.0000
   8.000   1.4138   0.01911   0.01051  -0.0917   0.1874   1.0000
   8.250   1.4365   0.01961   0.01100  -0.0910   0.1727   1.0000
   8.500   1.4580   0.02022   0.01158  -0.0902   0.1507   1.0000
   8.750   1.4764   0.02115   0.01233  -0.0891   0.1211   1.0000
   9.000   1.4940   0.02215   0.01321  -0.0879   0.1060   1.0000
   9.250   1.5123   0.02301   0.01407  -0.0867   0.1001   1.0000
   9.500   1.5293   0.02394   0.01504  -0.0854   0.0954   1.0000
   9.750   1.5464   0.02480   0.01597  -0.0840   0.0923   1.0000
  10.000   1.5626   0.02568   0.01695  -0.0826   0.0898   1.0000
  10.250   1.5767   0.02667   0.01802  -0.0809   0.0874   1.0000
  10.500   1.5879   0.02781   0.01924  -0.0788   0.0851   1.0000
  10.750   1.5942   0.02917   0.02064  -0.0762   0.0826   1.0000
  11.000   1.6033   0.03017   0.02176  -0.0738   0.0806   1.0000
  11.250   1.6119   0.03127   0.02300  -0.0716   0.0781   1.0000
  11.500   1.6184   0.03259   0.02445  -0.0695   0.0754   1.0000
  11.750   1.6213   0.03430   0.02622  -0.0674   0.0729   1.0000
  12.000   1.6202   0.03649   0.02846  -0.0653   0.0707   1.0000
  12.250   1.6293   0.03792   0.03006  -0.0642   0.0689   1.0000
  12.500   1.6366   0.03959   0.03190  -0.0632   0.0667   1.0000
  12.750   1.6418   0.04152   0.03398  -0.0623   0.0645   1.0000
  13.000   1.6452   0.04373   0.03631  -0.0616   0.0625   1.0000
  13.250   1.6449   0.04643   0.03909  -0.0610   0.0607   1.0000
  13.500   1.6481   0.04887   0.04169  -0.0607   0.0588   1.0000
  13.750   1.6522   0.05129   0.04430  -0.0605   0.0564   1.0000
  14.000   1.6538   0.05407   0.04723  -0.0605   0.0540   1.0000
  14.250   1.6523   0.05733   0.05061  -0.0609   0.0521   1.0000
  14.500   1.6499   0.06079   0.05419  -0.0612   0.0502   1.0000
  14.750   1.6484   0.06416   0.05774  -0.0616   0.0477   1.0000
  15.000   1.6440   0.06803   0.06174  -0.0624   0.0455   1.0000
  15.250   1.6360   0.07251   0.06632  -0.0634   0.0437   1.0000
  15.500   1.6285   0.07695   0.07090  -0.0644   0.0418   1.0000
  15.750   1.6202   0.08167   0.07577  -0.0657   0.0398   1.0000
  16.000   1.6095   0.08693   0.08115  -0.0673   0.0381   1.0000
  16.250   1.5965   0.09270   0.08702  -0.0693   0.0368   1.0000
  16.500   1.5849   0.09829   0.09274  -0.0712   0.0353   1.0000
  16.750   1.5739   0.10391   0.09851  -0.0732   0.0338   1.0000
  17.000   1.5617   0.10989   0.10460  -0.0756   0.0325   1.0000
  17.250   1.5484   0.11614   0.11095  -0.0782   0.0314   1.0000
<< Back to GOE 79 (PFALZ 11) AIRFOIL (goe79-il)

Polar data table (+)

Polar graphs


<< Back to GOE 79 (PFALZ 11) AIRFOIL (goe79-il)