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

GOE 180 (MVA H.26) AIRFOIL (goe180-il) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: GOE 180 (MVA H.26) AIRFOIL (goe180-il)
Reynolds number: 500,000
Max Cl/Cd: 99.2 at α=4.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe180-il-500000-n5.txt
Download as CSV file: xf-goe180-il-500000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 180 (MVA H.26) AIRFOIL                      
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.3742   0.08941   0.08716  -0.0287   1.0000   0.0105
  -8.500  -0.3838   0.08507   0.08287  -0.0293   1.0000   0.0108
  -8.250  -0.3881   0.08286   0.08071  -0.0285   1.0000   0.0109
  -8.000  -0.3772   0.07968   0.07754  -0.0319   0.9965   0.0111
  -7.750  -0.3564   0.07606   0.07391  -0.0377   0.9912   0.0115
  -7.500  -0.3361   0.07173   0.06957  -0.0444   0.9846   0.0121
  -7.250  -0.3176   0.06472   0.06253  -0.0543   0.9730   0.0129
  -6.750  -0.2674   0.02263   0.01880  -0.0946   0.9432   0.0182
  -6.500  -0.2334   0.02217   0.01828  -0.0964   0.9362   0.0186
  -6.250  -0.2000   0.02129   0.01725  -0.0982   0.9272   0.0192
  -6.000  -0.1682   0.02024   0.01600  -0.0997   0.9169   0.0201
  -5.750  -0.1387   0.01842   0.01379  -0.1006   0.9065   0.0215
  -5.500  -0.1106   0.01702   0.01200  -0.1010   0.8959   0.0224
  -5.250  -0.0826   0.01653   0.01126  -0.1010   0.8851   0.0230
  -5.000  -0.0567   0.01492   0.00943  -0.1012   0.8739   0.0240
  -4.750  -0.0298   0.01429   0.00865  -0.1011   0.8617   0.0245
  -4.500  -0.0029   0.01370   0.00791  -0.1009   0.8483   0.0251
  -4.250   0.0241   0.01319   0.00726  -0.1006   0.8334   0.0258
  -4.000   0.0510   0.01275   0.00667  -0.1004   0.8167   0.0266
  -3.750   0.0777   0.01228   0.00603  -0.1000   0.7962   0.0272
  -3.500   0.1041   0.01186   0.00543  -0.0996   0.7718   0.0276
  -3.250   0.1303   0.01152   0.00489  -0.0991   0.7431   0.0280
  -3.000   0.1561   0.01127   0.00442  -0.0985   0.7101   0.0284
  -2.750   0.1818   0.01109   0.00402  -0.0979   0.6733   0.0287
  -2.500   0.2075   0.01100   0.00372  -0.0974   0.6386   0.0291
  -2.250   0.2339   0.01093   0.00349  -0.0970   0.6108   0.0295
  -2.000   0.2605   0.01073   0.00317  -0.0967   0.5891   0.0298
  -1.750   0.2873   0.01045   0.00277  -0.0965   0.5704   0.0304
  -1.500   0.3143   0.01027   0.00250  -0.0963   0.5537   0.0311
  -1.250   0.3415   0.01015   0.00230  -0.0961   0.5399   0.0318
  -1.000   0.3690   0.01007   0.00215  -0.0960   0.5289   0.0326
  -0.750   0.3966   0.01001   0.00203  -0.0958   0.5189   0.0333
  -0.500   0.4242   0.00996   0.00194  -0.0957   0.5101   0.0341
  -0.250   0.4517   0.00995   0.00186  -0.0956   0.5014   0.0349
   0.000   0.4795   0.00993   0.00181  -0.0955   0.4934   0.0362
   0.500   0.5346   0.00996   0.00175  -0.0953   0.4753   0.0405
   0.750   0.5620   0.00992   0.00177  -0.0952   0.4668   0.0637
   1.000   0.5896   0.00989   0.00179  -0.0951   0.4586   0.0951
   1.250   0.6170   0.00981   0.00185  -0.0951   0.4517   0.1559
   1.500   0.6446   0.00966   0.00193  -0.0951   0.4454   0.2503
   1.750   0.6716   0.00941   0.00203  -0.0952   0.4387   0.4059
   2.250   0.7221   0.00830   0.00222  -0.0940   0.4229   1.0000
   2.500   0.7497   0.00840   0.00229  -0.0939   0.4154   1.0000
   2.750   0.7768   0.00854   0.00237  -0.0937   0.4076   1.0000
   3.000   0.8042   0.00866   0.00246  -0.0936   0.3992   1.0000
   3.250   0.8313   0.00880   0.00256  -0.0935   0.3913   1.0000
   3.500   0.8584   0.00893   0.00267  -0.0933   0.3813   1.0000
   3.750   0.8854   0.00908   0.00278  -0.0931   0.3707   1.0000
   4.000   0.9121   0.00925   0.00291  -0.0929   0.3586   1.0000
   4.250   0.9384   0.00946   0.00305  -0.0927   0.3395   1.0000
   4.500   0.9639   0.00975   0.00322  -0.0923   0.3114   1.0000
   4.750   0.9884   0.01016   0.00346  -0.0918   0.2785   1.0000
   5.000   1.0132   0.01054   0.00374  -0.0914   0.2570   1.0000
   5.250   1.0385   0.01086   0.00400  -0.0910   0.2440   1.0000
   5.500   1.0640   0.01115   0.00426  -0.0907   0.2343   1.0000
   5.750   1.0896   0.01142   0.00453  -0.0903   0.2269   1.0000
   6.000   1.1152   0.01169   0.00480  -0.0900   0.2199   1.0000
   6.250   1.1404   0.01198   0.00509  -0.0896   0.2125   1.0000
   6.500   1.1652   0.01231   0.00541  -0.0892   0.2023   1.0000
   6.750   1.1901   0.01261   0.00571  -0.0888   0.1908   1.0000
   7.000   1.2139   0.01304   0.00605  -0.0882   0.1711   1.0000
   7.250   1.2366   0.01356   0.00643  -0.0875   0.1441   1.0000
   7.500   1.2587   0.01412   0.00690  -0.0868   0.1256   1.0000
   7.750   1.2813   0.01461   0.00736  -0.0860   0.1145   1.0000
   8.000   1.3031   0.01516   0.00787  -0.0852   0.1028   1.0000
   8.500   1.3374   0.01710   0.00942  -0.0823   0.0371   1.0000
   8.750   1.3551   0.01796   0.01025  -0.0809   0.0232   1.0000
   9.000   1.3741   0.01866   0.01095  -0.0796   0.0179   1.0000
   9.250   1.3930   0.01934   0.01167  -0.0784   0.0152   1.0000
   9.500   1.4121   0.01996   0.01236  -0.0771   0.0137   1.0000
   9.750   1.4295   0.02068   0.01313  -0.0757   0.0124   1.0000
  10.000   1.4447   0.02153   0.01406  -0.0739   0.0111   1.0000
  10.250   1.4611   0.02221   0.01483  -0.0723   0.0105   1.0000
  10.500   1.4745   0.02297   0.01567  -0.0703   0.0099   1.0000
  10.750   1.4859   0.02380   0.01658  -0.0679   0.0094   1.0000
  11.000   1.4958   0.02474   0.01760  -0.0655   0.0089   1.0000
  11.250   1.5034   0.02586   0.01880  -0.0630   0.0085   1.0000
  11.500   1.5071   0.02730   0.02036  -0.0602   0.0081   1.0000
  11.750   1.5152   0.02849   0.02165  -0.0582   0.0078   1.0000
  12.000   1.5220   0.02984   0.02311  -0.0562   0.0075   1.0000
  12.250   1.5274   0.03138   0.02476  -0.0544   0.0072   1.0000
  12.500   1.5320   0.03309   0.02658  -0.0528   0.0069   1.0000
  12.750   1.5353   0.03501   0.02860  -0.0514   0.0067   1.0000
  13.000   1.5374   0.03719   0.03089  -0.0503   0.0065   1.0000
  13.250   1.5378   0.03968   0.03349  -0.0494   0.0063   1.0000
  13.500   1.5364   0.04253   0.03645  -0.0488   0.0062   1.0000
  13.750   1.5326   0.04587   0.03992  -0.0487   0.0060   1.0000
  14.000   1.5263   0.04971   0.04390  -0.0489   0.0059   1.0000
  14.250   1.5170   0.05414   0.04846  -0.0495   0.0058   1.0000
  14.500   1.5078   0.05872   0.05317  -0.0504   0.0057   1.0000
  14.750   1.5002   0.06314   0.05772  -0.0513   0.0057   1.0000
  15.000   1.4919   0.06775   0.06247  -0.0523   0.0056   1.0000
  15.250   1.4825   0.07260   0.06745  -0.0535   0.0056   1.0000
  15.500   1.4721   0.07762   0.07261  -0.0548   0.0055   1.0000
  15.750   1.4612   0.08284   0.07795  -0.0563   0.0055   1.0000
  16.000   1.4504   0.08815   0.08338  -0.0579   0.0054   1.0000
  16.250   1.4390   0.09358   0.08894  -0.0596   0.0054   1.0000
<< Back to GOE 180 (MVA H.26) AIRFOIL (goe180-il)

Polar data table (+)

Polar graphs


<< Back to GOE 180 (MVA H.26) AIRFOIL (goe180-il)