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

GOE 101 AIRFOIL (goe101-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: GOE 101 AIRFOIL (goe101-il)
Reynolds number: 50,000
Max Cl/Cd: 39.67 at α=7.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe101-il-50000.txt
Download as CSV file: xf-goe101-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 101 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.3874   0.10211   0.09546  -0.0132   1.0000   0.1418
  -7.750  -0.3900   0.10033   0.09379  -0.0145   1.0000   0.1481
  -7.500  -0.4016   0.10026   0.09388  -0.0195   1.0000   0.1501
  -7.250  -0.3823   0.09409   0.08771  -0.0150   1.0000   0.1608
  -7.000  -0.3906   0.09391   0.08765  -0.0223   1.0000   0.1645
  -6.750  -0.3751   0.08842   0.08220  -0.0175   1.0000   0.1761
  -6.250  -0.3676   0.08274   0.07665  -0.0212   1.0000   0.1944
  -6.000  -0.3627   0.08042   0.07438  -0.0236   1.0000   0.2078
  -5.750  -0.3552   0.07670   0.07074  -0.0216   1.0000   0.2231
  -5.250  -0.3440   0.07104   0.06523  -0.0198   1.0000   0.2660
  -5.000  -0.3393   0.06856   0.06283  -0.0185   1.0000   0.2942
  -4.750  -0.3351   0.06565   0.06002  -0.0150   1.0000   0.3233
  -4.500   0.0078   0.03875   0.03200  -0.0224   1.0000   1.0000
  -4.250   0.0190   0.03688   0.03020  -0.0233   1.0000   1.0000
  -4.000   0.0301   0.03510   0.02848  -0.0242   1.0000   1.0000
  -3.750   0.0320   0.03399   0.02747  -0.0227   1.0000   0.9960
  -3.500  -0.0189   0.03596   0.02971  -0.0092   1.0000   0.9604
  -3.250  -0.0673   0.03726   0.03130   0.0023   1.0000   0.9214
  -3.000  -0.1139   0.03802   0.03236   0.0124   1.0000   0.8859
  -2.750  -0.1603   0.03843   0.03309   0.0217   1.0000   0.8570
  -2.500  -0.2072   0.03867   0.03363   0.0313   1.0000   0.8414
  -1.750  -0.0612   0.03189   0.02357  -0.0442   1.0000   0.1991
  -1.500  -0.0332   0.03083   0.02180  -0.0448   1.0000   0.1781
  -1.250  -0.0118   0.02960   0.02033  -0.0448   1.0000   0.1684
  -1.000   0.0248   0.02868   0.01896  -0.0473   0.9941   0.1598
  -0.750   0.0838   0.02754   0.01733  -0.0533   0.9770   0.1558
  -0.500   0.1394   0.02679   0.01620  -0.0585   0.9598   0.1621
  -0.250   0.1966   0.02612   0.01519  -0.0636   0.9436   0.1648
   0.000   0.2499   0.02541   0.01444  -0.0682   0.9279   0.1741
   0.250   0.3005   0.02486   0.01389  -0.0724   0.9123   0.1950
   0.500   0.3552   0.02222   0.01342  -0.0770   0.8993   1.0000
   0.750   0.4032   0.02259   0.01309  -0.0801   0.8842   1.0000
   1.000   0.4395   0.02306   0.01323  -0.0818   0.8680   1.0000
   1.250   0.4754   0.02351   0.01347  -0.0834   0.8532   1.0000
   1.500   0.5103   0.02397   0.01379  -0.0847   0.8396   1.0000
   1.750   0.5458   0.02439   0.01409  -0.0860   0.8273   1.0000
   2.000   0.5733   0.02503   0.01467  -0.0862   0.8141   1.0000
   2.250   0.5986   0.02576   0.01536  -0.0862   0.8014   1.0000
   2.500   0.6231   0.02655   0.01611  -0.0860   0.7892   1.0000
   2.750   0.6479   0.02736   0.01693  -0.0859   0.7779   1.0000
   3.000   0.6752   0.02807   0.01768  -0.0859   0.7672   1.0000
   3.250   0.7046   0.02866   0.01831  -0.0860   0.7569   1.0000
   3.500   0.7246   0.02970   0.01940  -0.0854   0.7447   1.0000
   3.750   0.7464   0.03065   0.02041  -0.0847   0.7323   1.0000
   4.000   0.7712   0.03129   0.02113  -0.0838   0.7184   1.0000
   4.250   0.7977   0.03170   0.02168  -0.0827   0.7034   1.0000
   4.500   0.8241   0.03206   0.02214  -0.0814   0.6881   1.0000
   4.750   0.8492   0.03255   0.02275  -0.0801   0.6732   1.0000
   5.000   0.8735   0.03314   0.02348  -0.0789   0.6585   1.0000
   5.250   0.8984   0.03365   0.02421  -0.0775   0.6435   1.0000
   5.500   0.9246   0.03403   0.02476  -0.0760   0.6280   1.0000
   5.750   0.9411   0.03519   0.02613  -0.0744   0.6090   1.0000
   6.000   0.9667   0.03534   0.02648  -0.0723   0.5892   1.0000
   6.250   0.9922   0.03521   0.02658  -0.0697   0.5662   1.0000
   6.500   1.0222   0.03431   0.02582  -0.0667   0.5410   1.0000
   6.750   1.0503   0.03360   0.02525  -0.0638   0.5143   1.0000
   7.000   1.0735   0.03367   0.02553  -0.0613   0.4865   1.0000
   7.250   1.0956   0.03261   0.02465  -0.0577   0.4485   1.0000
   7.500   1.1160   0.02979   0.02173  -0.0529   0.3966   1.0000
   7.750   1.1246   0.02835   0.02027  -0.0483   0.3303   1.0000
   8.000   1.1182   0.02991   0.02085  -0.0431   0.2203   1.0000
   8.250   1.1112   0.03407   0.02409  -0.0391   0.1487   1.0000
   8.500   1.1194   0.03682   0.02664  -0.0366   0.1194   1.0000
   8.750   1.1388   0.03941   0.02914  -0.0349   0.1044   1.0000
   9.000   1.1673   0.04251   0.03233  -0.0339   0.0948   1.0000
   9.250   1.1937   0.04595   0.03566  -0.0335   0.0867   1.0000
   9.500   1.2086   0.04924   0.03956  -0.0319   0.0827   1.0000
   9.750   1.2226   0.05321   0.04400  -0.0304   0.0807   1.0000
  10.000   1.2303   0.05745   0.04872  -0.0288   0.0801   1.0000
  10.250   1.2304   0.06191   0.05366  -0.0270   0.0802   1.0000
  10.500   1.2233   0.06642   0.05861  -0.0253   0.0807   1.0000
  10.750   1.2098   0.07096   0.06351  -0.0236   0.0815   1.0000
  11.000   1.1906   0.07522   0.06803  -0.0220   0.0822   1.0000
  11.250   1.1682   0.07988   0.07290  -0.0216   0.0830   1.0000
  11.500   1.1448   0.08517   0.07837  -0.0228   0.0838   1.0000
  11.750   1.1221   0.09125   0.08457  -0.0253   0.0848   1.0000
  12.000   1.1009   0.09804   0.09145  -0.0287   0.0857   1.0000
  12.250   1.0860   0.10506   0.09852  -0.0319   0.0867   1.0000
<< Back to GOE 101 AIRFOIL (goe101-il)

Polar data table (+)

Polar graphs


<< Back to GOE 101 AIRFOIL (goe101-il)