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

GOE 335 (D.F.W.) AIRFOIL (goe335-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: GOE 335 (D.F.W.) AIRFOIL (goe335-il)
Reynolds number: 100,000
Max Cl/Cd: 57.92 at α=7°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe335-il-100000-n5.txt
Download as CSV file: xf-goe335-il-100000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 335 (D.F.W.) 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.750  -0.3048   0.09818   0.09352  -0.0265   1.0000   0.0365
  -7.500  -0.3096   0.09678   0.09223  -0.0265   1.0000   0.0369
  -7.250  -0.3132   0.09530   0.09085  -0.0275   1.0000   0.0372
  -7.000  -0.3120   0.09314   0.08877  -0.0290   1.0000   0.0374
  -6.750  -0.3096   0.09079   0.08648  -0.0306   1.0000   0.0376
  -6.500  -0.3066   0.08829   0.08405  -0.0316   1.0000   0.0376
  -6.250  -0.3041   0.08582   0.08163  -0.0323   1.0000   0.0377
  -6.000  -0.2945   0.08273   0.07856  -0.0346   0.9983   0.0378
  -5.750  -0.2616   0.07764   0.07338  -0.0423   0.9909   0.0379
  -5.500  -0.2294   0.07259   0.06824  -0.0490   0.9835   0.0379
  -5.250  -0.1959   0.06752   0.06305  -0.0554   0.9767   0.0380
  -5.000  -0.1952   0.06305   0.05876  -0.0516   0.9698   0.0423
  -4.750  -0.1400   0.05934   0.05472  -0.0645   0.9602   0.0513
  -4.500  -0.1141   0.05255   0.04786  -0.0675   0.9539   0.0361
  -4.250  -0.0821   0.04801   0.04315  -0.0720   0.9438   0.0347
  -4.000  -0.0390   0.04295   0.03776  -0.0783   0.9350   0.0359
  -3.750   0.0075   0.03784   0.03212  -0.0838   0.9264   0.0377
  -3.500   0.0420   0.03322   0.02706  -0.0866   0.9150   0.0383
  -3.250   0.0703   0.03141   0.02522  -0.0884   0.9032   0.0431
  -3.000   0.1070   0.02815   0.02147  -0.0906   0.8932   0.0456
  -2.750   0.1446   0.02515   0.01784  -0.0924   0.8831   0.0483
  -2.500   0.1799   0.02315   0.01513  -0.0933   0.8712   0.0516
  -2.250   0.2128   0.02105   0.01261  -0.0942   0.8600   0.0527
  -2.000   0.2452   0.01966   0.01092  -0.0949   0.8495   0.0544
  -1.750   0.2772   0.01894   0.01002  -0.0956   0.8391   0.0596
  -1.500   0.3090   0.01800   0.00878  -0.0960   0.8285   0.0611
  -1.250   0.3386   0.01719   0.00777  -0.0959   0.8167   0.0614
  -1.000   0.3674   0.01654   0.00695  -0.0957   0.8048   0.0621
  -0.750   0.3955   0.01601   0.00630  -0.0955   0.7926   0.0630
  -0.500   0.4230   0.01559   0.00577  -0.0951   0.7799   0.0645
  -0.250   0.4501   0.01523   0.00532  -0.0946   0.7667   0.0670
   0.000   0.4767   0.01497   0.00500  -0.0940   0.7528   0.0723
   0.250   0.5034   0.01476   0.00473  -0.0935   0.7383   0.0826
   0.500   0.5300   0.01450   0.00452  -0.0930   0.7228   0.1224
   0.750   0.5509   0.01310   0.00454  -0.0919   0.7068   0.6068
   1.000   0.5901   0.01258   0.00434  -0.0936   0.6876   1.0000
   1.250   0.6155   0.01269   0.00425  -0.0928   0.6684   1.0000
   1.500   0.6404   0.01283   0.00422  -0.0919   0.6470   1.0000
   1.750   0.6654   0.01298   0.00419  -0.0911   0.6248   1.0000
   2.000   0.6901   0.01316   0.00421  -0.0902   0.6001   1.0000
   2.250   0.7145   0.01338   0.00426  -0.0894   0.5743   1.0000
   2.500   0.7388   0.01362   0.00434  -0.0885   0.5492   1.0000
   2.750   0.7631   0.01390   0.00445  -0.0877   0.5265   1.0000
   3.000   0.7874   0.01420   0.00463  -0.0870   0.5070   1.0000
   3.250   0.8118   0.01452   0.00483  -0.0862   0.4903   1.0000
   3.500   0.8363   0.01486   0.00507  -0.0856   0.4760   1.0000
   3.750   0.8610   0.01522   0.00536  -0.0850   0.4640   1.0000
   4.000   0.8858   0.01559   0.00567  -0.0844   0.4539   1.0000
   4.250   0.9109   0.01595   0.00603  -0.0839   0.4446   1.0000
   4.500   0.9358   0.01635   0.00640  -0.0834   0.4358   1.0000
   4.750   0.9606   0.01674   0.00683  -0.0829   0.4266   1.0000
   5.000   0.9853   0.01714   0.00726  -0.0823   0.4173   1.0000
   5.250   1.0099   0.01757   0.00769  -0.0818   0.4087   1.0000
   5.500   1.0345   0.01796   0.00822  -0.0813   0.3999   1.0000
   5.750   1.0593   0.01840   0.00873  -0.0808   0.3927   1.0000
   6.000   1.0835   0.01881   0.00927  -0.0802   0.3841   1.0000
   6.250   1.1076   0.01924   0.00982  -0.0796   0.3758   1.0000
   6.500   1.1291   0.01958   0.01027  -0.0785   0.3600   1.0000
   6.750   1.1468   0.01981   0.01056  -0.0768   0.3315   1.0000
   7.000   1.1670   0.02015   0.01101  -0.0755   0.3101   1.0000
   7.250   1.1867   0.02052   0.01149  -0.0742   0.2828   1.0000
   7.500   1.2045   0.02106   0.01201  -0.0726   0.2327   1.0000
   7.750   1.2055   0.02349   0.01338  -0.0695   0.0937   1.0000
   8.000   1.2086   0.02603   0.01541  -0.0666   0.0341   1.0000
   8.250   1.2202   0.02749   0.01700  -0.0644   0.0274   1.0000
   8.750   1.2386   0.03048   0.02035  -0.0596   0.0219   1.0000
   9.000   1.2447   0.03195   0.02206  -0.0568   0.0205   1.0000
   9.250   1.2467   0.03354   0.02387  -0.0537   0.0196   1.0000
   9.500   1.2464   0.03532   0.02585  -0.0505   0.0190   1.0000
   9.750   1.2446   0.03731   0.02802  -0.0477   0.0185   1.0000
  10.000   1.2425   0.03946   0.03035  -0.0453   0.0181   1.0000
  10.250   1.2397   0.04188   0.03292  -0.0434   0.0175   1.0000
  10.500   1.2377   0.04444   0.03562  -0.0418   0.0169   1.0000
  10.750   1.2359   0.04722   0.03851  -0.0404   0.0161   1.0000
  11.000   1.2357   0.05008   0.04147  -0.0391   0.0153   1.0000
  11.250   1.2405   0.05302   0.04450  -0.0375   0.0145   1.0000
  11.500   1.2521   0.05614   0.04779  -0.0358   0.0141   1.0000
  11.750   1.2585   0.05905   0.05097  -0.0348   0.0139   1.0000
  12.000   1.2623   0.06221   0.05437  -0.0341   0.0138   1.0000
  12.250   1.2630   0.06569   0.05810  -0.0336   0.0137   1.0000
  12.500   1.2600   0.06955   0.06223  -0.0334   0.0136   1.0000
  12.750   1.2541   0.07377   0.06671  -0.0337   0.0136   1.0000
  13.000   1.2460   0.07834   0.07153  -0.0344   0.0136   1.0000
  13.250   1.2360   0.08326   0.07671  -0.0357   0.0136   1.0000
  13.500   1.2241   0.08862   0.08231  -0.0375   0.0136   1.0000
  13.750   1.2110   0.09436   0.08828  -0.0399   0.0136   1.0000
  14.000   1.1974   0.10044   0.09457  -0.0427   0.0136   1.0000
  14.250   1.1831   0.10697   0.10131  -0.0462   0.0137   1.0000
  14.500   1.1686   0.11395   0.10848  -0.0502   0.0137   1.0000
  14.750   1.1539   0.12137   0.11608  -0.0547   0.0138   1.0000
  15.000   1.1395   0.12911   0.12398  -0.0594   0.0139   1.0000
  15.250   1.1254   0.13721   0.13222  -0.0645   0.0140   1.0000
  15.750   0.9216   0.14168   0.13746  -0.0639   0.0149   1.0000
  16.000   0.9040   0.15002   0.14594  -0.0692   0.0153   1.0000
<< Back to GOE 335 (D.F.W.) AIRFOIL (goe335-il)

Polar data table (+)

Polar graphs


<< Back to GOE 335 (D.F.W.) AIRFOIL (goe335-il)