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

GOE 683 AIRFOIL (goe683-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: GOE 683 AIRFOIL (goe683-il)
Reynolds number: 100,000
Max Cl/Cd: 31.25 at α=3°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe683-il-100000.txt
Download as CSV file: xf-goe683-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 683 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -12.500  -0.5488   0.07618   0.07053  -0.0736   1.0000   0.1234
 -12.250  -0.5941   0.06879   0.06308  -0.0750   1.0000   0.1226
 -12.000  -0.6448   0.06318   0.05739  -0.0734   1.0000   0.1218
 -11.750  -0.6920   0.05955   0.05369  -0.0688   1.0000   0.1212
 -11.500  -0.7393   0.05797   0.05211  -0.0613   1.0000   0.1206
 -11.250  -0.7898   0.05778   0.05194  -0.0519   1.0000   0.1200
 -11.000  -0.8370   0.05722   0.05133  -0.0433   0.9992   0.1196
 -10.750  -0.8518   0.05303   0.04669  -0.0419   0.9855   0.1202
 -10.500  -0.8575   0.04926   0.04241  -0.0403   0.9723   0.1216
 -10.250  -0.8583   0.04607   0.03864  -0.0384   0.9596   0.1231
 -10.000  -0.8247   0.04385   0.03639  -0.0409   0.9513   0.1257
  -9.750  -0.7936   0.04231   0.03478  -0.0425   0.9421   0.1283
  -9.500  -0.7654   0.04041   0.03262  -0.0437   0.9336   0.1317
  -9.250  -0.7527   0.03869   0.03049  -0.0419   0.9221   0.1345
  -9.000  -0.7122   0.03674   0.02843  -0.0450   0.9161   0.1383
  -8.750  -0.6726   0.03543   0.02706  -0.0476   0.9092   0.1426
  -8.500  -0.6450   0.03412   0.02545  -0.0479   0.8997   0.1475
  -8.250  -0.5919   0.03234   0.02366  -0.0529   0.8954   0.1534
  -8.000  -0.5693   0.03161   0.02284  -0.0521   0.8837   0.1584
  -7.750  -0.5271   0.03015   0.02120  -0.0549   0.8774   0.1659
  -7.500  -0.4991   0.02947   0.02058  -0.0550   0.8668   0.1727
  -7.250  -0.4633   0.02834   0.01933  -0.0565   0.8586   0.1821
  -7.000  -0.4402   0.02783   0.01883  -0.0556   0.8483   0.1917
  -6.750  -0.4103   0.02701   0.01806  -0.0559   0.8390   0.2044
  -6.500  -0.3917   0.02653   0.01758  -0.0541   0.8290   0.2188
  -6.250  -0.3727   0.02608   0.01713  -0.0522   0.8191   0.2357
  -6.000  -0.3584   0.02585   0.01687  -0.0495   0.8094   0.2520
  -5.750  -0.3436   0.02570   0.01665  -0.0467   0.7993   0.2672
  -5.500  -0.3201   0.02536   0.01638  -0.0457   0.7910   0.2820
  -5.250  -0.3037   0.02515   0.01617  -0.0435   0.7804   0.2944
  -5.000  -0.2737   0.02469   0.01558  -0.0434   0.7740   0.3097
  -4.750  -0.2621   0.02467   0.01567  -0.0405   0.7619   0.3204
  -4.500  -0.2304   0.02415   0.01512  -0.0409   0.7551   0.3355
  -4.250  -0.2199   0.02423   0.01515  -0.0376   0.7445   0.3481
  -4.000  -0.1905   0.02375   0.01475  -0.0376   0.7367   0.3654
  -3.750  -0.1675   0.02349   0.01458  -0.0366   0.7281   0.3823
  -3.500  -0.1449   0.02321   0.01439  -0.0354   0.7189   0.4009
  -3.250  -0.1128   0.02268   0.01389  -0.0358   0.7131   0.4281
  -3.000  -0.1028   0.02275   0.01414  -0.0325   0.7023   0.4522
  -2.750  -0.0733   0.02219   0.01375  -0.0324   0.6955   0.4919
  -2.500  -0.0488   0.02186   0.01376  -0.0316   0.6870   0.5403
  -2.250  -0.0055   0.02127   0.01364  -0.0342   0.6784   0.6235
  -2.000   0.1395   0.02114   0.01390  -0.0548   0.6706   0.7555
  -1.750   0.1985   0.02223   0.01499  -0.0591   0.6599   0.7964
  -1.500   0.2487   0.02293   0.01548  -0.0616   0.6533   0.8242
  -1.250   0.2868   0.02407   0.01658  -0.0625   0.6437   0.8449
  -1.000   0.3353   0.02478   0.01713  -0.0651   0.6361   0.8640
  -0.750   0.3823   0.02551   0.01771  -0.0679   0.6287   0.8824
  -0.500   0.4265   0.02619   0.01833  -0.0705   0.6196   0.9003
  -0.250   0.4844   0.02655   0.01847  -0.0754   0.6133   0.9175
   0.000   0.5290   0.02714   0.01906  -0.0786   0.6044   0.9358
   0.250   0.5800   0.02724   0.01905  -0.0830   0.5969   0.9528
   0.500   0.6357   0.02711   0.01874  -0.0885   0.5905   0.9677
   0.750   0.6863   0.02693   0.01860  -0.0937   0.5815   0.9828
   1.000   0.7528   0.02596   0.01746  -0.1015   0.5748   0.9985
   1.250   0.7700   0.02609   0.01756  -0.1006   0.5691   1.0000
   1.500   0.7781   0.02645   0.01797  -0.0980   0.5625   1.0000
   1.750   0.7938   0.02655   0.01800  -0.0964   0.5574   1.0000
   2.000   0.8147   0.02656   0.01786  -0.0954   0.5533   1.0000
   2.250   0.8189   0.02723   0.01866  -0.0922   0.5466   1.0000
   2.500   0.8313   0.02760   0.01903  -0.0900   0.5409   1.0000
   2.750   0.8514   0.02767   0.01901  -0.0888   0.5364   1.0000
   3.000   0.8715   0.02789   0.01914  -0.0877   0.5325   1.0000
   3.250   0.8719   0.02891   0.02032  -0.0838   0.5258   1.0000
   3.500   0.8869   0.02927   0.02068  -0.0819   0.5206   1.0000
   3.750   0.9105   0.02929   0.02059  -0.0811   0.5166   1.0000
   4.000   0.9245   0.02988   0.02119  -0.0791   0.5121   1.0000
   4.250   0.9224   0.03118   0.02264  -0.0749   0.5059   1.0000
   4.500   0.9374   0.03165   0.02312  -0.0730   0.5013   1.0000
   4.750   0.9627   0.03167   0.02305  -0.0725   0.4977   1.0000
   5.000   0.9765   0.03234   0.02372  -0.0704   0.4933   1.0000
   5.250   0.9611   0.03431   0.02589  -0.0645   0.4867   1.0000
   5.500   0.9753   0.03489   0.02648  -0.0625   0.4824   1.0000
   5.750   1.0004   0.03499   0.02653  -0.0619   0.4792   1.0000
   6.000   1.0348   0.03488   0.02631  -0.0627   0.4765   1.0000
   6.250   0.9387   0.04073   0.03257  -0.0468   0.4674   1.0000
   6.500   0.9494   0.04144   0.03328  -0.0443   0.4633   1.0000
   6.750   1.0038   0.04006   0.03181  -0.0472   0.4606   1.0000
   7.000   1.0616   0.03871   0.03032  -0.0507   0.4579   1.0000
   7.250   0.5036   0.08154   0.07363  -0.0126   0.4285   1.0000
   7.500   0.4881   0.08595   0.07807  -0.0117   0.4281   1.0000
   7.750   0.4692   0.09129   0.08345  -0.0114   0.4335   1.0000
   8.000   0.4871   0.09410   0.08628  -0.0117   0.4407   1.0000
   8.250   0.5025   0.09634   0.08853  -0.0113   0.4398   1.0000
   8.500   0.5162   0.09891   0.09110  -0.0109   0.4393   1.0000
   8.750   0.5391   0.09688   0.08903  -0.0084   0.4200   1.0000
   9.000   0.5437   0.10012   0.09229  -0.0081   0.4203   1.0000
   9.250   0.4302   0.11206   0.10435  -0.0088   0.4472   1.0000
   9.500   0.4767   0.11300   0.10530  -0.0088   0.4406   1.0000
<< Back to GOE 683 AIRFOIL (goe683-il)

Polar data table (+)

Polar graphs


<< Back to GOE 683 AIRFOIL (goe683-il)