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=200,000 Ncrit=9


Details Polar file
Airfoil: GOE 180 (MVA H.26) AIRFOIL (goe180-il)
Reynolds number: 200,000
Max Cl/Cd: 78.19 at α=6.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe180-il-200000.txt
Download as CSV file: xf-goe180-il-200000.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 =     0.200 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.3467   0.09601   0.09249  -0.0269   1.0000   0.0419
  -8.000  -0.3581   0.09426   0.09083  -0.0289   1.0000   0.0427
  -7.750  -0.3659   0.09196   0.08861  -0.0322   1.0000   0.0431
  -7.500  -0.3663   0.08873   0.08543  -0.0366   1.0000   0.0433
  -7.250  -0.3668   0.08418   0.08094  -0.0324   1.0000   0.0440
  -7.000  -0.3621   0.08185   0.07864  -0.0287   1.0000   0.0449
  -6.750  -0.3600   0.07979   0.07663  -0.0269   1.0000   0.0462
  -6.500  -0.3607   0.07763   0.07451  -0.0262   1.0000   0.0477
  -6.250  -0.3615   0.07519   0.07210  -0.0265   1.0000   0.0496
  -6.000  -0.3478   0.07101   0.06775  -0.0386   1.0000   0.0534
  -5.750  -0.3445   0.06588   0.06261  -0.0399   1.0000   0.0542
  -5.500  -0.3297   0.06288   0.05967  -0.0388   0.9978   0.0554
  -5.250  -0.2984   0.05965   0.05642  -0.0424   0.9941   0.0583
  -5.000  -0.2413   0.05182   0.04812  -0.0577   0.9879   0.0665
  -4.750  -0.2141   0.04935   0.04576  -0.0591   0.9845   0.0692
  -4.500  -0.1624   0.04439   0.04018  -0.0686   0.9777   0.0798
  -4.250  -0.1315   0.04074   0.03667  -0.0713   0.9738   0.0821
  -4.000  -0.0821   0.04010   0.03534  -0.0762   0.9665   0.0936
  -3.750  -0.0531   0.03437   0.02989  -0.0795   0.9620   0.0974
  -3.500  -0.0095   0.03181   0.02685  -0.0837   0.9563   0.1102
  -3.250   0.0254   0.02909   0.02420  -0.0861   0.9498   0.1139
  -3.000   0.0773   0.02304   0.01711  -0.0900   0.9466   0.0823
  -2.750   0.1187   0.01943   0.01271  -0.0913   0.9397   0.0641
  -2.500   0.1600   0.01738   0.01036  -0.0936   0.9335   0.0628
  -2.250   0.1971   0.01599   0.00876  -0.0949   0.9237   0.0627
  -2.000   0.2360   0.01510   0.00769  -0.0965   0.9138   0.0646
  -1.750   0.2729   0.01422   0.00668  -0.0978   0.9016   0.0655
  -1.500   0.3040   0.01315   0.00560  -0.0979   0.8845   0.0671
  -1.250   0.3327   0.01255   0.00501  -0.0976   0.8632   0.0698
  -1.000   0.3610   0.01214   0.00453  -0.0971   0.8390   0.0737
  -0.750   0.3878   0.01183   0.00412  -0.0963   0.8100   0.0794
  -0.500   0.4140   0.01150   0.00372  -0.0954   0.7773   0.0939
  -0.250   0.4392   0.01065   0.00341  -0.0949   0.7452   0.2693
   0.000   0.4648   0.00892   0.00332  -0.0936   0.7169   1.0000
   0.250   0.4908   0.00914   0.00322  -0.0930   0.6916   1.0000
   0.500   0.5170   0.00936   0.00319  -0.0924   0.6691   1.0000
   0.750   0.5433   0.00960   0.00320  -0.0919   0.6502   1.0000
   1.000   0.5699   0.00983   0.00324  -0.0915   0.6348   1.0000
   1.250   0.5967   0.01005   0.00331  -0.0912   0.6214   1.0000
   1.500   0.6235   0.01028   0.00340  -0.0908   0.6092   1.0000
   1.750   0.6503   0.01053   0.00349  -0.0905   0.5982   1.0000
   2.000   0.6771   0.01074   0.00361  -0.0902   0.5869   1.0000
   2.250   0.7038   0.01095   0.00374  -0.0899   0.5756   1.0000
   2.500   0.7306   0.01118   0.00388  -0.0897   0.5655   1.0000
   2.750   0.7573   0.01141   0.00403  -0.0894   0.5557   1.0000
   3.000   0.7841   0.01161   0.00421  -0.0891   0.5461   1.0000
   3.250   0.8109   0.01187   0.00439  -0.0889   0.5380   1.0000
   3.500   0.8375   0.01206   0.00458  -0.0886   0.5283   1.0000
   3.750   0.8639   0.01228   0.00478  -0.0882   0.5188   1.0000
   4.000   0.8903   0.01252   0.00497  -0.0879   0.5091   1.0000
   4.250   0.9163   0.01269   0.00518  -0.0875   0.4984   1.0000
   4.500   0.9423   0.01290   0.00539  -0.0871   0.4881   1.0000
   4.750   0.9683   0.01314   0.00557  -0.0867   0.4778   1.0000
   5.000   0.9939   0.01330   0.00582  -0.0862   0.4660   1.0000
   5.250   1.0194   0.01350   0.00606  -0.0857   0.4543   1.0000
   5.500   1.0447   0.01370   0.00629  -0.0852   0.4419   1.0000
   5.750   1.0696   0.01390   0.00653  -0.0846   0.4279   1.0000
   6.000   1.0941   0.01411   0.00678  -0.0839   0.4119   1.0000
   6.250   1.1182   0.01433   0.00703  -0.0832   0.3934   1.0000
   6.500   1.1416   0.01461   0.00728  -0.0824   0.3732   1.0000
   6.750   1.1650   0.01490   0.00762  -0.0816   0.3498   1.0000
   7.000   1.1874   0.01530   0.00797  -0.0806   0.3269   1.0000
   7.250   1.2086   0.01579   0.00835  -0.0796   0.3009   1.0000
   7.500   1.2287   0.01640   0.00883  -0.0784   0.2745   1.0000
   7.750   1.2486   0.01708   0.00944  -0.0773   0.2534   1.0000
   8.000   1.2690   0.01774   0.01008  -0.0763   0.2360   1.0000
   8.250   1.2893   0.01839   0.01073  -0.0752   0.2205   1.0000
   8.500   1.3096   0.01899   0.01136  -0.0742   0.2046   1.0000
   8.750   1.3300   0.01954   0.01196  -0.0732   0.1871   1.0000
   9.000   1.3501   0.02007   0.01248  -0.0721   0.1584   1.0000
   9.250   1.3595   0.02161   0.01361  -0.0698   0.0993   1.0000
   9.500   1.3601   0.02395   0.01547  -0.0662   0.0505   1.0000
   9.750   1.3676   0.02550   0.01709  -0.0633   0.0432   1.0000
  10.000   1.3705   0.02710   0.01877  -0.0599   0.0394   1.0000
  10.250   1.3708   0.02877   0.02056  -0.0563   0.0373   1.0000
  10.500   1.3748   0.03028   0.02224  -0.0534   0.0357   1.0000
  10.750   1.3768   0.03202   0.02409  -0.0506   0.0342   1.0000
  11.000   1.3776   0.03396   0.02612  -0.0482   0.0328   1.0000
  11.250   1.3761   0.03623   0.02846  -0.0459   0.0315   1.0000
  11.500   1.3712   0.03906   0.03132  -0.0436   0.0303   1.0000
  11.750   1.3738   0.04135   0.03371  -0.0419   0.0294   1.0000
  12.000   1.3788   0.04346   0.03596  -0.0405   0.0287   1.0000
  12.250   1.3839   0.04567   0.03829  -0.0391   0.0280   1.0000
  12.500   1.3898   0.04791   0.04065  -0.0377   0.0273   1.0000
  12.750   1.3964   0.05018   0.04302  -0.0364   0.0267   1.0000
  13.000   1.4022   0.05257   0.04555  -0.0352   0.0261   1.0000
  13.250   1.4065   0.05513   0.04823  -0.0342   0.0255   1.0000
  13.500   1.4100   0.05779   0.05098  -0.0333   0.0248   1.0000
  13.750   1.4166   0.06052   0.05376  -0.0322   0.0240   1.0000
  14.000   1.4234   0.06474   0.05814  -0.0305   0.0233   1.0000
  14.250   1.4168   0.06833   0.06196  -0.0302   0.0232   1.0000
  14.500   1.4086   0.07228   0.06617  -0.0303   0.0231   1.0000
  14.750   1.3990   0.07662   0.07074  -0.0307   0.0231   1.0000
  15.000   1.3879   0.08130   0.07565  -0.0315   0.0231   1.0000
  15.250   1.3748   0.08637   0.08094  -0.0327   0.0231   1.0000
  15.500   1.3606   0.09182   0.08660  -0.0344   0.0231   1.0000
  15.750   1.3457   0.09760   0.09259  -0.0364   0.0232   1.0000
  16.000   1.3301   0.10380   0.09898  -0.0387   0.0232   1.0000
  16.250   1.3158   0.10986   0.10524  -0.0416   0.0234   1.0000
  16.500   1.3006   0.11615   0.11171  -0.0454   0.0234   1.0000
  16.750   1.2853   0.12282   0.11858  -0.0498   0.0236   1.0000
  17.000   1.2676   0.13054   0.12650  -0.0552   0.0238   1.0000
  17.250   1.2441   0.14045   0.13666  -0.0627   0.0241   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)