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

GOE 55 AIRFOIL (goe55-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: GOE 55 AIRFOIL (goe55-il)
Reynolds number: 200,000
Max Cl/Cd: 51.9 at α=5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe55-il-200000-n5.txt
Download as CSV file: xf-goe55-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 55 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
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.6284   0.10872   0.10508   0.0376   1.0000   0.0352
  -8.500  -0.6315   0.10429   0.10069   0.0316   1.0000   0.0378
  -8.250  -0.6313   0.09985   0.09628   0.0276   1.0000   0.0379
  -7.250  -0.5982   0.08303   0.07948   0.0181   1.0000   0.0359
  -7.000  -0.5862   0.07701   0.07342   0.0119   1.0000   0.0340
  -6.750  -0.5695   0.06930   0.06561   0.0028   1.0000   0.0329
  -6.500  -0.5501   0.06546   0.06172  -0.0009   1.0000   0.0341
  -6.250  -0.5277   0.06014   0.05628  -0.0064   1.0000   0.0349
  -6.000  -0.5029   0.05382   0.04976  -0.0122   1.0000   0.0347
  -5.750  -0.4759   0.04756   0.04324  -0.0173   1.0000   0.0346
  -5.500  -0.4474   0.04195   0.03732  -0.0213   1.0000   0.0355
  -5.250  -0.4170   0.03564   0.03055  -0.0250   1.0000   0.0363
  -5.000  -0.3869   0.03026   0.02464  -0.0274   1.0000   0.0366
  -4.750  -0.3565   0.02575   0.01953  -0.0291   1.0000   0.0372
  -4.500  -0.3262   0.02304   0.01620  -0.0299   1.0000   0.0383
  -4.250  -0.2975   0.02087   0.01377  -0.0306   1.0000   0.0388
  -4.000  -0.2688   0.01940   0.01211  -0.0309   1.0000   0.0393
  -3.750  -0.2401   0.01824   0.01080  -0.0310   1.0000   0.0398
  -3.500  -0.2114   0.01726   0.00969  -0.0311   1.0000   0.0405
  -3.250  -0.1827   0.01637   0.00870  -0.0312   1.0000   0.0412
  -3.000  -0.1540   0.01559   0.00784  -0.0311   1.0000   0.0421
  -2.750  -0.1255   0.01497   0.00715  -0.0311   1.0000   0.0436
  -2.500  -0.0968   0.01433   0.00647  -0.0310   1.0000   0.0450
  -2.250  -0.0659   0.01413   0.00588  -0.0307   0.7804   0.0459
  -2.000  -0.0441   0.01434   0.00539  -0.0287   0.6127   0.0466
  -1.750  -0.0173   0.01422   0.00481  -0.0284   0.5030   0.0477
  -1.500   0.0109   0.01403   0.00442  -0.0284   0.4555   0.0493
  -1.250   0.0393   0.01386   0.00412  -0.0283   0.4290   0.0512
  -1.000   0.0679   0.01372   0.00388  -0.0283   0.4113   0.0542
  -0.750   0.0965   0.01358   0.00366  -0.0283   0.3987   0.0579
  -0.500   0.1253   0.01343   0.00350  -0.0283   0.3891   0.0628
  -0.250   0.1541   0.01333   0.00337  -0.0283   0.3815   0.0698
   0.000   0.1828   0.01325   0.00333  -0.0282   0.3747   0.0832
   0.250   0.2115   0.01323   0.00329  -0.0282   0.3689   0.1033
   0.500   0.2402   0.01311   0.00325  -0.0282   0.3623   0.1261
   0.750   0.2689   0.01291   0.00328  -0.0284   0.3546   0.2052
   1.250   0.3165   0.01087   0.00320  -0.0261   0.3379   1.0000
   1.500   0.3452   0.01099   0.00325  -0.0260   0.3299   1.0000
   1.750   0.3737   0.01115   0.00331  -0.0259   0.3236   1.0000
   2.000   0.4023   0.01131   0.00340  -0.0258   0.3173   1.0000
   2.250   0.4307   0.01144   0.00347  -0.0257   0.3072   1.0000
   2.500   0.4592   0.01155   0.00353  -0.0256   0.2949   1.0000
   2.750   0.4877   0.01167   0.00360  -0.0255   0.2833   1.0000
   3.000   0.5160   0.01181   0.00370  -0.0254   0.2731   1.0000
   3.250   0.5445   0.01193   0.00381  -0.0253   0.2607   1.0000
   3.500   0.5729   0.01206   0.00391  -0.0252   0.2448   1.0000
   3.750   0.6012   0.01222   0.00405  -0.0251   0.2262   1.0000
   4.000   0.6293   0.01245   0.00417  -0.0250   0.1977   1.0000
   4.250   0.6571   0.01285   0.00437  -0.0250   0.1653   1.0000
   4.500   0.6848   0.01331   0.00471  -0.0250   0.1419   1.0000
   4.750   0.7125   0.01375   0.00507  -0.0250   0.1223   1.0000
   5.000   0.7401   0.01426   0.00548  -0.0250   0.0986   1.0000
   5.250   0.7675   0.01481   0.00591  -0.0250   0.0686   1.0000
   5.500   0.7948   0.01547   0.00643  -0.0250   0.0536   1.0000
   5.750   0.8222   0.01599   0.00697  -0.0249   0.0496   1.0000
   6.000   0.8494   0.01655   0.00757  -0.0248   0.0470   1.0000
   6.250   0.8764   0.01718   0.00826  -0.0247   0.0451   1.0000
   6.500   0.9034   0.01775   0.00890  -0.0245   0.0440   1.0000
   6.750   0.9301   0.01836   0.00960  -0.0244   0.0428   1.0000
   7.000   0.9564   0.01902   0.01036  -0.0242   0.0416   1.0000
   7.250   0.9824   0.01974   0.01117  -0.0240   0.0404   1.0000
   7.500   1.0080   0.02053   0.01205  -0.0238   0.0396   1.0000
   7.750   1.0331   0.02140   0.01301  -0.0236   0.0389   1.0000
   8.000   1.0576   0.02238   0.01405  -0.0233   0.0382   1.0000
   8.250   1.0814   0.02350   0.01523  -0.0230   0.0375   1.0000
   8.500   1.1041   0.02489   0.01668  -0.0226   0.0367   1.0000
   8.750   1.1275   0.02602   0.01794  -0.0222   0.0361   1.0000
   9.000   1.1510   0.02707   0.01915  -0.0217   0.0355   1.0000
   9.250   1.1736   0.02829   0.02052  -0.0213   0.0349   1.0000
   9.500   1.1955   0.02963   0.02204  -0.0208   0.0343   1.0000
   9.750   1.2167   0.03106   0.02366  -0.0203   0.0336   1.0000
  10.000   1.2370   0.03253   0.02531  -0.0198   0.0328   1.0000
  10.250   1.2564   0.03406   0.02701  -0.0193   0.0320   1.0000
  10.500   1.2746   0.03571   0.02883  -0.0187   0.0313   1.0000
  10.750   1.2913   0.03755   0.03082  -0.0181   0.0307   1.0000
  11.000   1.3061   0.03964   0.03304  -0.0175   0.0300   1.0000
  11.250   1.3157   0.04282   0.03639  -0.0168   0.0292   1.0000
  11.500   1.3243   0.04512   0.03903  -0.0162   0.0288   1.0000
  11.750   1.3287   0.04773   0.04199  -0.0155   0.0283   1.0000
  12.000   1.3271   0.05085   0.04543  -0.0150   0.0279   1.0000
  12.250   1.3162   0.05433   0.04918  -0.0143   0.0276   1.0000
  12.500   1.2999   0.05920   0.05431  -0.0167   0.0275   1.0000
  12.750   1.2816   0.06653   0.06190  -0.0234   0.0274   1.0000
  13.000   1.2591   0.07582   0.07144  -0.0317   0.0274   1.0000
  13.250   1.2300   0.08639   0.08221  -0.0399   0.0276   1.0000
  13.500   1.1936   0.09833   0.09433  -0.0483   0.0278   1.0000
  13.750   1.1448   0.11352   0.10967  -0.0583   0.0282   1.0000
<< Back to GOE 55 AIRFOIL (goe55-il)

Polar data table (+)

Polar graphs


<< Back to GOE 55 AIRFOIL (goe55-il)