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

FX 63-147 AIRFOIL (fx63147-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: FX 63-147 AIRFOIL (fx63147-il)
Reynolds number: 100,000
Max Cl/Cd: 48.6 at α=8.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-fx63147-il-100000-n5.txt
Download as CSV file: xf-fx63147-il-100000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: FX 63-147 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
  ------ -------- --------- --------- -------- -------- --------
 -10.500  -0.3677   0.10546   0.10057  -0.0485   1.0001   0.0283
 -10.250  -0.3632   0.10313   0.09825  -0.0476   1.0001   0.0268
  -9.750  -0.3925   0.08746   0.08273  -0.0552   1.0001   0.0225
  -9.500  -0.3974   0.08409   0.07942  -0.0556   1.0001   0.0223
  -9.250  -0.4085   0.07994   0.07533  -0.0568   1.0001   0.0220
  -9.000  -0.4255   0.07598   0.07140  -0.0573   1.0001   0.0218
  -8.750  -0.4470   0.07272   0.06817  -0.0565   1.0001   0.0216
  -8.500  -0.4726   0.07029   0.06576  -0.0539   1.0001   0.0215
  -8.000  -0.4819   0.06117   0.05631  -0.0602   0.9844   0.0207
  -7.750  -0.4751   0.05572   0.05053  -0.0636   0.9731   0.0201
  -7.500  -0.4680   0.05050   0.04488  -0.0652   0.9613   0.0193
  -7.000  -0.4429   0.04130   0.03443  -0.0658   0.9384   0.0179
  -6.750  -0.4186   0.03820   0.03096  -0.0670   0.9307   0.0178
  -6.500  -0.3950   0.03546   0.02784  -0.0674   0.9215   0.0177
  -6.250  -0.3692   0.03300   0.02499  -0.0677   0.9128   0.0176
  -6.000  -0.3389   0.03070   0.02234  -0.0685   0.9060   0.0176
  -5.750  -0.3110   0.02882   0.02017  -0.0686   0.8975   0.0177
  -5.500  -0.2790   0.02709   0.01819  -0.0693   0.8909   0.0179
  -5.250  -0.2505   0.02563   0.01660  -0.0693   0.8821   0.0181
  -5.000  -0.2184   0.02422   0.01511  -0.0702   0.8750   0.0186
  -4.750  -0.1904   0.02308   0.01392  -0.0704   0.8649   0.0192
  -4.500  -0.1546   0.02192   0.01264  -0.0722   0.8583   0.0202
  -4.250  -0.1256   0.02104   0.01161  -0.0726   0.8468   0.0215
  -4.000  -0.0921   0.02018   0.01066  -0.0740   0.8370   0.0244
  -3.750  -0.0537   0.01939   0.00979  -0.0762   0.8283   0.0304
  -3.500  -0.0222   0.01857   0.00895  -0.0771   0.8166   0.0472
  -3.250   0.0052   0.01678   0.00820  -0.0783   0.8057   0.2545
  -3.000   0.0252   0.01549   0.00844  -0.0767   0.7956   0.5827
  -2.750   0.0441   0.01619   0.00937  -0.0723   0.7836   0.6861
  -2.500   0.0676   0.01720   0.01028  -0.0693   0.7719   0.7432
  -2.250   0.0930   0.01815   0.01105  -0.0666   0.7612   0.7700
  -2.000   0.1219   0.01870   0.01139  -0.0654   0.7510   0.7872
  -1.750   0.1483   0.01882   0.01133  -0.0649   0.7393   0.7964
  -1.500   0.1765   0.01888   0.01118  -0.0648   0.7286   0.8041
  -1.250   0.2067   0.01891   0.01100  -0.0652   0.7190   0.8110
  -1.000   0.2312   0.01899   0.01094  -0.0643   0.7078   0.8180
  -0.750   0.2569   0.01901   0.01080  -0.0640   0.6976   0.8254
  -0.500   0.2856   0.01904   0.01066  -0.0640   0.6886   0.8303
  -0.250   0.3080   0.01905   0.01057  -0.0634   0.6783   0.8385
   0.000   0.3344   0.01909   0.01049  -0.0628   0.6699   0.8428
   0.250   0.3585   0.01913   0.01042  -0.0621   0.6609   0.8493
   0.500   0.3827   0.01916   0.01035  -0.0617   0.6527   0.8555
   0.750   0.4079   0.01917   0.01028  -0.0611   0.6446   0.8599
   1.000   0.4312   0.01921   0.01024  -0.0605   0.6366   0.8662
   1.250   0.4558   0.01923   0.01019  -0.0600   0.6290   0.8711
   1.500   0.4792   0.01926   0.01017  -0.0592   0.6214   0.8759
   1.750   0.5031   0.01930   0.01014  -0.0588   0.6138   0.8810
   2.000   0.5263   0.01932   0.01013  -0.0581   0.6061   0.8850
   2.250   0.5502   0.01933   0.01010  -0.0575   0.5985   0.8887
   2.500   0.5734   0.01938   0.01012  -0.0569   0.5912   0.8931
   2.750   0.5963   0.01945   0.01018  -0.0564   0.5838   0.8977
   3.000   0.6202   0.01947   0.01018  -0.0557   0.5772   0.9013
   3.250   0.6419   0.01953   0.01026  -0.0548   0.5698   0.9057
   3.500   0.6678   0.01962   0.01030  -0.0548   0.5641   0.9099
   3.750   0.6877   0.01969   0.01046  -0.0536   0.5567   0.9136
   4.000   0.7129   0.01975   0.01050  -0.0533   0.5506   0.9169
   4.250   0.7341   0.01986   0.01066  -0.0524   0.5436   0.9208
   4.500   0.7578   0.01998   0.01080  -0.0520   0.5368   0.9243
   4.750   0.7803   0.02006   0.01094  -0.0513   0.5301   0.9275
   5.000   0.8012   0.02015   0.01109  -0.0503   0.5226   0.9312
   5.250   0.8251   0.02027   0.01122  -0.0499   0.5161   0.9346
   5.500   0.8449   0.02044   0.01151  -0.0489   0.5086   0.9385
   5.750   0.8708   0.02053   0.01161  -0.0488   0.5029   0.9414
   6.000   0.8887   0.02070   0.01194  -0.0474   0.4952   0.9453
   6.250   0.9131   0.02086   0.01214  -0.0471   0.4891   0.9487
   6.500   0.9336   0.02109   0.01249  -0.0462   0.4823   0.9523
   6.750   0.9561   0.02123   0.01272  -0.0456   0.4750   0.9553
   7.000   0.9770   0.02142   0.01303  -0.0448   0.4676   0.9589
   7.250   0.9979   0.02159   0.01329  -0.0440   0.4593   0.9627
   7.500   1.0167   0.02177   0.01360  -0.0428   0.4502   0.9665
   7.750   1.0369   0.02184   0.01374  -0.0418   0.4396   0.9701
   8.000   1.0530   0.02196   0.01394  -0.0402   0.4273   0.9745
   8.250   1.0669   0.02206   0.01414  -0.0383   0.4130   0.9796
   8.750   1.0974   0.02258   0.01483  -0.0354   0.3826   0.9921
   9.000   1.1103   0.02285   0.01520  -0.0337   0.3680   0.9998
   9.250   1.1256   0.02339   0.01583  -0.0326   0.3529   0.9999
   9.500   1.1405   0.02405   0.01655  -0.0317   0.3370   0.9999
   9.750   1.1547   0.02480   0.01740  -0.0307   0.3206   0.9999
  10.000   1.1683   0.02568   0.01839  -0.0299   0.3037   0.9999
  10.250   1.1811   0.02668   0.01948  -0.0291   0.2868   0.9999
  10.500   1.1922   0.02786   0.02075  -0.0284   0.2681   0.9999
  10.750   1.2003   0.02929   0.02222  -0.0275   0.2473   0.9999
  11.000   1.2056   0.03101   0.02396  -0.0267   0.2262   0.9999
  11.500   1.2058   0.03575   0.02865  -0.0255   0.1777   0.9999
  11.750   1.2037   0.03859   0.03149  -0.0253   0.1559   0.9999
  12.000   1.1998   0.04175   0.03465  -0.0254   0.1368   0.9999
  12.250   1.1950   0.04518   0.03809  -0.0257   0.1207   0.9999
  12.750   1.1822   0.05282   0.04575  -0.0270   0.0926   0.9999
  13.000   1.1773   0.05664   0.04964  -0.0278   0.0823   0.9999
  13.250   1.1724   0.06055   0.05364  -0.0287   0.0729   0.9999
  13.500   1.1661   0.06470   0.05785  -0.0298   0.0653   0.9999
  13.750   1.1596   0.06898   0.06220  -0.0309   0.0602   0.9999
  14.000   1.1515   0.07357   0.06682  -0.0323   0.0563   0.9999
  14.250   1.1476   0.07768   0.07108  -0.0335   0.0525   0.9999
  14.500   1.1444   0.08172   0.07527  -0.0348   0.0488   0.9999
  14.750   1.1398   0.08608   0.07972  -0.0363   0.0460   0.9999
  15.000   1.1364   0.09021   0.08394  -0.0376   0.0436   0.9999
  15.250   1.1368   0.09393   0.08781  -0.0389   0.0410   0.9999
<< Back to FX 63-147 AIRFOIL (fx63147-il)

Polar data table (+)

Polar graphs


<< Back to FX 63-147 AIRFOIL (fx63147-il)