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

GOE 627 AIRFOIL (goe627-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: GOE 627 AIRFOIL (goe627-il)
Reynolds number: 50,000
Max Cl/Cd: 25.4 at α=3.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe627-il-50000.txt
Download as CSV file: xf-goe627-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 627 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
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.3236   0.11169   0.10532  -0.0146   1.0000   0.3311
  -8.500  -0.3121   0.10816   0.10184  -0.0128   1.0000   0.3401
  -8.250  -0.3574   0.10905   0.10293  -0.0098   1.0000   0.3513
  -8.000  -0.3184   0.10390   0.09773  -0.0086   1.0000   0.3620
  -7.750  -0.3362   0.10231   0.09627  -0.0059   1.0000   0.3732
  -7.500  -0.3572   0.10166   0.09575  -0.0025   1.0000   0.3874
  -7.250  -0.3381   0.09774   0.09186  -0.0013   1.0000   0.3946
  -7.000  -0.3630   0.09673   0.09099   0.0025   1.0000   0.4088
  -6.750  -0.3511   0.09389   0.08819   0.0043   1.0000   0.4186
  -6.500  -0.3597   0.09195   0.08635   0.0074   1.0000   0.4315
  -5.750  -0.5944   0.06218   0.05545  -0.0095   1.0000   0.1720
  -5.500  -0.5890   0.05868   0.05181  -0.0079   1.0000   0.1666
  -5.250  -0.5886   0.05400   0.04632  -0.0063   1.0000   0.1555
  -5.000  -0.5793   0.05124   0.04334  -0.0045   1.0000   0.1527
  -4.750  -0.5697   0.04857   0.04032  -0.0027   1.0000   0.1500
  -4.500  -0.5585   0.04639   0.03777  -0.0009   1.0000   0.1495
  -4.250  -0.5460   0.04457   0.03561   0.0008   1.0000   0.1500
  -4.000  -0.5322   0.04293   0.03363   0.0025   1.0000   0.1507
  -3.750  -0.5173   0.04145   0.03183   0.0041   1.0000   0.1511
  -3.500  -0.5014   0.04017   0.03026   0.0056   1.0000   0.1519
  -3.250  -0.4849   0.03912   0.02892   0.0069   1.0000   0.1535
  -3.000  -0.4684   0.03837   0.02787   0.0082   1.0000   0.1563
  -2.750  -0.4325   0.03792   0.02741   0.0060   0.9932   0.1629
  -2.500  -0.3721   0.03796   0.02718   0.0001   0.9739   0.1735
  -2.250  -0.3116   0.03785   0.02721  -0.0058   0.9561   0.1928
  -2.000  -0.0503   0.03701   0.02999  -0.0401   0.9469   1.0000
  -1.750  -0.0069   0.03730   0.02984  -0.0445   0.9271   1.0000
  -1.500   0.0386   0.03758   0.02976  -0.0488   0.9083   1.0000
  -1.250   0.0755   0.03771   0.02960  -0.0514   0.8887   1.0000
  -1.000   0.1077   0.03777   0.02941  -0.0528   0.8682   1.0000
  -0.750   0.1451   0.03779   0.02921  -0.0548   0.8488   1.0000
  -0.500   0.1858   0.03769   0.02889  -0.0570   0.8299   1.0000
  -0.250   0.2278   0.03748   0.02847  -0.0592   0.8116   1.0000
   0.000   0.2716   0.03710   0.02791  -0.0613   0.7937   1.0000
   0.250   0.3154   0.03659   0.02724  -0.0631   0.7761   1.0000
   0.500   0.3605   0.03590   0.02640  -0.0649   0.7589   1.0000
   0.750   0.4068   0.03505   0.02542  -0.0666   0.7423   1.0000
   1.000   0.4588   0.03406   0.02429  -0.0692   0.7258   1.0000
   1.250   0.5135   0.03301   0.02312  -0.0722   0.7089   1.0000
   1.500   0.5658   0.03195   0.02192  -0.0747   0.6915   1.0000
   1.750   0.6142   0.03109   0.02093  -0.0768   0.6735   1.0000
   2.000   0.6581   0.03048   0.02017  -0.0783   0.6550   1.0000
   2.250   0.6960   0.03017   0.01970  -0.0790   0.6363   1.0000
   2.500   0.7234   0.03030   0.01970  -0.0783   0.6186   1.0000
   2.750   0.7450   0.03065   0.01993  -0.0769   0.6016   1.0000
   3.000   0.7610   0.03124   0.02044  -0.0746   0.5857   1.0000
   3.250   0.7756   0.03194   0.02106  -0.0723   0.5712   1.0000
   3.500   0.7972   0.03251   0.02154  -0.0710   0.5583   1.0000
   3.750   0.8316   0.03274   0.02160  -0.0716   0.5460   1.0000
   4.000   0.8231   0.03425   0.02319  -0.0661   0.5343   1.0000
   4.250   0.8410   0.03505   0.02392  -0.0644   0.5240   1.0000
   4.500   0.8547   0.03599   0.02483  -0.0622   0.5140   1.0000
   4.750   0.8550   0.03752   0.02637  -0.0582   0.5055   1.0000
   5.000   0.8646   0.03876   0.02762  -0.0556   0.4973   1.0000
   5.250   0.8722   0.04020   0.02905  -0.0528   0.4901   1.0000
   5.500   0.8370   0.04326   0.03221  -0.0448   0.4832   1.0000
   5.750   0.9066   0.04238   0.03116  -0.0498   0.4754   1.0000
   6.000   0.7844   0.04941   0.03841  -0.0324   0.4709   1.0000
   6.250   0.5739   0.06879   0.05783  -0.0218   0.4661   1.0000
   6.500   0.5490   0.07441   0.06344  -0.0212   0.4651   1.0000
   6.750   0.5323   0.07925   0.06828  -0.0209   0.4649   1.0000
   7.000   0.5155   0.08411   0.07314  -0.0208   0.4668   1.0000
   7.250   0.5092   0.08848   0.07751  -0.0212   0.4713   1.0000
   7.500   0.5264   0.09178   0.08082  -0.0219   0.4751   1.0000
<< Back to GOE 627 AIRFOIL (goe627-il)

Polar data table (+)

Polar graphs


<< Back to GOE 627 AIRFOIL (goe627-il)