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

GOE 735 AIRFOIL (goe735-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: GOE 735 AIRFOIL (goe735-il)
Reynolds number: 50,000
Max Cl/Cd: 16.36 at α=0.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe735-il-50000-n5.txt
Download as CSV file: xf-goe735-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 735 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -12.750  -0.2437   0.11922   0.11235  -0.0453   0.9351   0.1131
 -12.500  -0.2463   0.11324   0.10629  -0.0508   0.9176   0.1152
 -12.250  -0.2434   0.10835   0.10134  -0.0547   0.9007   0.1169
 -12.000  -0.2261   0.10617   0.09906  -0.0557   0.8822   0.1192
 -11.750  -0.2219   0.10285   0.09567  -0.0572   0.8670   0.1216
 -11.500  -0.2269   0.09847   0.09123  -0.0595   0.8548   0.1239
 -11.250  -0.2448   0.09234   0.08504  -0.0630   0.8450   0.1261
 -11.000  -0.3207   0.07683   0.06943  -0.0746   0.8418   0.1282
 -10.750  -0.2794   0.08024   0.07285  -0.0696   0.8254   0.1304
 -10.500  -0.2947   0.07473   0.06729  -0.0726   0.8150   0.1325
 -10.250  -0.3492   0.06638   0.05878  -0.0768   0.8071   0.1337
 -10.000  -0.3926   0.06133   0.05353  -0.0762   0.7970   0.1349
  -9.750  -0.4269   0.05812   0.05008  -0.0727   0.7872   0.1361
  -9.500  -0.4526   0.05530   0.04698  -0.0693   0.7773   0.1379
  -9.250  -0.4734   0.05256   0.04384  -0.0654   0.7680   0.1398
  -9.000  -0.4833   0.05034   0.04126  -0.0621   0.7584   0.1422
  -8.750  -0.4662   0.04945   0.04037  -0.0607   0.7484   0.1449
  -8.500  -0.4551   0.04830   0.03909  -0.0589   0.7389   0.1477
  -8.250  -0.4490   0.04684   0.03739  -0.0565   0.7294   0.1507
  -8.000  -0.4457   0.04520   0.03535  -0.0537   0.7216   0.1541
  -7.750  -0.4375   0.04383   0.03370  -0.0515   0.7114   0.1571
  -7.500  -0.4176   0.04286   0.03267  -0.0501   0.7041   0.1600
  -7.250  -0.4013   0.04204   0.03177  -0.0489   0.6936   0.1634
  -7.000  -0.3854   0.04096   0.03043  -0.0471   0.6860   0.1674
  -6.750  -0.3711   0.03991   0.02901  -0.0453   0.6777   0.1711
  -6.500  -0.3503   0.03905   0.02810  -0.0444   0.6687   0.1744
  -6.250  -0.3277   0.03823   0.02719  -0.0432   0.6626   0.1784
  -6.000  -0.3086   0.03769   0.02656  -0.0423   0.6525   0.1827
  -5.750  -0.2875   0.03687   0.02543  -0.0411   0.6456   0.1873
  -5.500  -0.2637   0.03618   0.02469  -0.0404   0.6384   0.1916
  -5.250  -0.2411   0.03575   0.02427  -0.0398   0.6296   0.1964
  -5.000  -0.2158   0.03507   0.02339  -0.0390   0.6239   0.2024
  -4.750  -0.1925   0.03471   0.02292  -0.0385   0.6156   0.2079
  -4.500  -0.1670   0.03428   0.02254  -0.0381   0.6083   0.2140
  -4.250  -0.1388   0.03367   0.02177  -0.0376   0.6033   0.2220
  -4.000  -0.1150   0.03356   0.02169  -0.0374   0.5946   0.2292
  -3.750  -0.0883   0.03322   0.02132  -0.0372   0.5882   0.2384
  -3.500  -0.0581   0.03266   0.02068  -0.0371   0.5836   0.2497
  -3.250  -0.0365   0.03277   0.02084  -0.0367   0.5753   0.2613
  -3.000  -0.0119   0.03250   0.02062  -0.0362   0.5690   0.2755
  -2.750   0.0151   0.03199   0.02013  -0.0356   0.5647   0.2946
  -2.500   0.0333   0.03213   0.02040  -0.0347   0.5576   0.3153
  -2.250   0.0522   0.03206   0.02047  -0.0335   0.5512   0.3439
  -2.000   0.0741   0.03156   0.02015  -0.0322   0.5468   0.3847
  -1.750   0.0938   0.03100   0.01989  -0.0304   0.5429   0.4427
  -1.500   0.1007   0.03150   0.02093  -0.0279   0.5347   0.5133
  -1.250   0.1489   0.03123   0.02138  -0.0299   0.5294   0.6687
  -1.000   0.2323   0.03140   0.02158  -0.0366   0.5252   0.7912
  -0.750   0.2637   0.03270   0.02285  -0.0375   0.5176   0.8393
  -0.500   0.2996   0.03357   0.02359  -0.0384   0.5117   0.8780
  -0.250   0.3502   0.03396   0.02374  -0.0412   0.5075   0.9132
   0.000   0.4180   0.03435   0.02387  -0.0473   0.5031   0.9456
   0.250   0.4723   0.03564   0.02515  -0.0538   0.4948   0.9717
   0.500   0.5311   0.03561   0.02491  -0.0598   0.4897   0.9902
   0.750   0.5773   0.03528   0.02435  -0.0632   0.4863   1.0000
   1.000   0.5816   0.03647   0.02554  -0.0613   0.4812   1.0000
   1.250   0.5797   0.03806   0.02718  -0.0588   0.4751   1.0000
   1.500   0.5886   0.03874   0.02778  -0.0567   0.4710   1.0000
   1.750   0.6045   0.03893   0.02783  -0.0550   0.4680   1.0000
   2.000   0.6252   0.03887   0.02761  -0.0537   0.4657   1.0000
   2.250   0.5735   0.04348   0.03247  -0.0473   0.4558   1.0000
   2.500   0.5727   0.04469   0.03362  -0.0443   0.4515   1.0000
   2.750   0.5854   0.04514   0.03396  -0.0423   0.4488   1.0000
   3.000   0.6072   0.04512   0.03380  -0.0409   0.4468   1.0000
   3.500   0.4851   0.05458   0.04339  -0.0273   0.4294   1.0000
   3.750   0.4126   0.06175   0.05064  -0.0228   0.4183   1.0000
   4.000   0.4137   0.06349   0.05230  -0.0210   0.4142   1.0000
   4.250   0.4287   0.06418   0.05288  -0.0195   0.4115   1.0000
   4.500   0.4503   0.06438   0.05297  -0.0182   0.4096   1.0000
   5.000   0.4047   0.07237   0.06094  -0.0148   0.3963   1.0000
   5.250   0.4183   0.07347   0.06196  -0.0137   0.3937   1.0000
   5.500   0.4376   0.07409   0.06249  -0.0126   0.3915   1.0000
   5.750   0.4119   0.07860   0.06701  -0.0115   0.3846   1.0000
   6.000   0.4107   0.08109   0.06947  -0.0106   0.3800   1.0000
   6.250   0.4207   0.08269   0.07102  -0.0098   0.3770   1.0000
   6.500   0.4379   0.08367   0.07192  -0.0089   0.3746   1.0000
   6.750   0.4603   0.08422   0.07241  -0.0081   0.3727   1.0000
   7.000   0.4278   0.08960   0.07783  -0.0078   0.3648   1.0000
   7.250   0.4334   0.09172   0.07992  -0.0072   0.3611   1.0000
   7.500   0.4472   0.09312   0.08129  -0.0066   0.3583   1.0000
   7.750   0.4671   0.09400   0.08210  -0.0059   0.3560   1.0000
   8.000   0.4562   0.09763   0.08575  -0.0057   0.3505   1.0000
   8.250   0.4538   0.10054   0.08867  -0.0055   0.3458   1.0000
   8.500   0.4635   0.10242   0.09053  -0.0051   0.3425   1.0000
   8.750   0.4810   0.10357   0.09163  -0.0045   0.3398   1.0000
   9.000   0.5026   0.10442   0.09244  -0.0039   0.3379   1.0000
   9.250   0.4794   0.10923   0.09731  -0.0043   0.3309   1.0000
   9.500   0.4842   0.11163   0.09971  -0.0042   0.3273   1.0000
<< Back to GOE 735 AIRFOIL (goe735-il)

Polar data table (+)

Polar graphs


<< Back to GOE 735 AIRFOIL (goe735-il)