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=9


Details Polar file
Airfoil: GOE 79 (PFALZ 11) AIRFOIL (goe79-il)
Reynolds number: 200,000
Max Cl/Cd: 75.69 at α=5.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe79-il-200000.txt
Download as CSV file: xf-goe79-il-200000.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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.3043   0.10096   0.09742  -0.0143   1.0000   0.0261
  -7.750  -0.2964   0.09751   0.09401  -0.0141   1.0000   0.0263
  -7.500  -0.2897   0.09474   0.09127  -0.0146   1.0000   0.0266
  -7.250  -0.2814   0.09203   0.08861  -0.0158   1.0000   0.0269
  -7.000  -0.2727   0.08941   0.08603  -0.0172   1.0000   0.0273
  -6.750  -0.2639   0.08687   0.08353  -0.0186   1.0000   0.0278
  -6.500  -0.2553   0.08443   0.08114  -0.0200   1.0000   0.0284
  -6.250  -0.2472   0.08215   0.07892  -0.0215   1.0000   0.0289
  -6.000  -0.2430   0.08042   0.07724  -0.0223   1.0000   0.0293
  -5.750  -0.2380   0.07919   0.07606  -0.0242   0.9992   0.0296
  -5.500  -0.1850   0.07467   0.07144  -0.0379   0.9931   0.0299
  -5.250  -0.1616   0.07003   0.06682  -0.0398   0.9872   0.0302
  -5.000  -0.1265   0.06597   0.06273  -0.0458   0.9806   0.0308
  -4.750  -0.0856   0.06212   0.05883  -0.0531   0.9733   0.0318
  -4.500  -0.0392   0.05844   0.05507  -0.0616   0.9648   0.0334
  -4.250   0.0320   0.05514   0.05150  -0.0759   0.9566   0.0342
  -4.000   0.0557   0.05102   0.04743  -0.0776   0.9448   0.0346
  -3.750   0.0896   0.04777   0.04413  -0.0814   0.9309   0.0354
  -3.500   0.1266   0.04495   0.04120  -0.0856   0.9146   0.0366
  -3.250   0.1635   0.04257   0.03863  -0.0891   0.8959   0.0383
  -3.000   0.2054   0.04051   0.03627  -0.0930   0.8763   0.0396
  -2.750   0.2255   0.03812   0.03383  -0.0931   0.8563   0.0404
  -2.500   0.2518   0.03620   0.03178  -0.0939   0.8372   0.0415
  -2.250   0.2815   0.03444   0.02984  -0.0950   0.8187   0.0432
  -2.000   0.3256   0.03365   0.02857  -0.0972   0.8019   0.0456
  -1.750   0.3469   0.03098   0.02591  -0.0976   0.7861   0.0464
  -1.500   0.3734   0.02932   0.02414  -0.0982   0.7712   0.0480
  -1.250   0.4135   0.02952   0.02380  -0.0988   0.7571   0.0526
  -1.000   0.4372   0.02658   0.02090  -0.0997   0.7433   0.0539
  -0.750   0.4644   0.02526   0.01949  -0.1002   0.7291   0.0574
  -0.500   0.4972   0.02422   0.01816  -0.1007   0.7158   0.0628
  -0.250   0.5244   0.02299   0.01684  -0.1011   0.7029   0.0664
   0.000   0.5556   0.02217   0.01572  -0.1013   0.6906   0.0736
   0.250   0.5835   0.02126   0.01472  -0.1016   0.6774   0.0802
   0.500   0.6126   0.02029   0.01361  -0.1019   0.6644   0.0884
   0.750   0.6415   0.01955   0.01270  -0.1020   0.6518   0.1003
   1.000   0.6698   0.01885   0.01185  -0.1021   0.6389   0.1153
   1.250   0.6976   0.01815   0.01101  -0.1023   0.6254   0.1414
   1.500   0.7244   0.01766   0.01048  -0.1023   0.6106   0.1825
   2.500   0.8393   0.01589   0.00801  -0.1006   0.5409   0.1851
   3.000   0.8976   0.01529   0.00688  -0.0983   0.4982   0.0823
   3.250   0.9242   0.01491   0.00639  -0.0977   0.4735   0.0765
   3.500   0.9507   0.01480   0.00619  -0.0972   0.4455   0.0767
   3.750   0.9768   0.01479   0.00603  -0.0966   0.4199   0.0738
   4.000   1.0028   0.01490   0.00602  -0.0961   0.3978   0.0725
   4.250   1.0286   0.01503   0.00604  -0.0957   0.3798   0.0727
   4.500   1.0546   0.01524   0.00614  -0.0954   0.3649   0.0745
   4.750   1.0805   0.01552   0.00636  -0.0950   0.3530   0.0812
   5.000   1.1061   0.01588   0.00660  -0.0945   0.3432   0.0917
   5.250   1.1286   0.01491   0.00691  -0.0934   0.3343   1.0000
   5.500   1.1541   0.01539   0.00721  -0.0929   0.3264   1.0000
   5.750   1.1801   0.01574   0.00755  -0.0925   0.3184   1.0000
   6.000   1.2052   0.01628   0.00793  -0.0921   0.3113   1.0000
   6.250   1.2310   0.01657   0.00829  -0.0917   0.3032   1.0000
   6.750   1.2812   0.01734   0.00908  -0.0909   0.2877   1.0000
   7.000   1.3060   0.01775   0.00947  -0.0904   0.2810   1.0000
   7.250   1.3309   0.01813   0.00991  -0.0899   0.2743   1.0000
   7.500   1.3553   0.01844   0.01028  -0.0894   0.2657   1.0000
   7.750   1.3794   0.01877   0.01068  -0.0889   0.2567   1.0000
   8.000   1.4028   0.01914   0.01105  -0.0882   0.2474   1.0000
   8.250   1.4268   0.01939   0.01145  -0.0876   0.2361   1.0000
   8.500   1.4497   0.01972   0.01186  -0.0869   0.2218   1.0000
   8.750   1.4715   0.02014   0.01229  -0.0861   0.2032   1.0000
   9.000   1.4923   0.02071   0.01284  -0.0852   0.1782   1.0000
   9.250   1.5105   0.02158   0.01359  -0.0839   0.1532   1.0000
   9.500   1.5278   0.02256   0.01450  -0.0826   0.1365   1.0000
   9.750   1.5440   0.02359   0.01549  -0.0811   0.1264   1.0000
  10.000   1.5584   0.02473   0.01657  -0.0794   0.1185   1.0000
  10.250   1.5738   0.02573   0.01764  -0.0779   0.1119   1.0000
  10.500   1.5840   0.02708   0.01895  -0.0758   0.1061   1.0000
  10.750   1.5962   0.02823   0.02020  -0.0738   0.1014   1.0000
  11.000   1.6052   0.02951   0.02154  -0.0715   0.0972   1.0000
  11.250   1.6069   0.03122   0.02319  -0.0683   0.0934   1.0000
  11.500   1.6139   0.03257   0.02469  -0.0659   0.0902   1.0000
  11.750   1.6195   0.03407   0.02632  -0.0636   0.0868   1.0000
  12.000   1.6228   0.03590   0.02819  -0.0614   0.0833   1.0000
  12.250   1.6274   0.03811   0.03038  -0.0594   0.0796   1.0000
  12.500   1.6305   0.03989   0.03239  -0.0577   0.0766   1.0000
  12.750   1.6329   0.04192   0.03456  -0.0563   0.0731   1.0000
  13.000   1.6353   0.04424   0.03685  -0.0548   0.0696   1.0000
  13.250   1.6354   0.04678   0.03956  -0.0535   0.0664   1.0000
  13.500   1.6336   0.04940   0.04241  -0.0529   0.0633   1.0000
  13.750   1.6325   0.05209   0.04519  -0.0524   0.0604   1.0000
  14.000   1.6317   0.05496   0.04802  -0.0515   0.0574   1.0000
  14.250   1.6257   0.05848   0.05185  -0.0519   0.0551   1.0000
  14.500   1.6213   0.06197   0.05550  -0.0523   0.0526   1.0000
  14.750   1.6174   0.06545   0.05907  -0.0530   0.0506   1.0000
  15.000   1.6135   0.06886   0.06246  -0.0529   0.0485   1.0000
  15.250   1.6046   0.07333   0.06720  -0.0539   0.0469   1.0000
  15.500   1.5973   0.07763   0.07169  -0.0549   0.0452   1.0000
  15.750   1.5912   0.08182   0.07599  -0.0560   0.0437   1.0000
  16.000   1.5858   0.08596   0.08021  -0.0572   0.0425   1.0000
  16.250   1.5822   0.08960   0.08383  -0.0575   0.0411   1.0000
  16.500   1.5718   0.09483   0.08928  -0.0592   0.0401   1.0000
  16.750   1.5609   0.10031   0.09498  -0.0612   0.0392   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)