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RAF 25 AIRFOIL (raf25-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: RAF 25 AIRFOIL (raf25-il)
Reynolds number: 50,000
Max Cl/Cd: 30.55 at α=3.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-raf25-il-50000-n5.txt
Download as CSV file: xf-raf25-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: RAF 25 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
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.5729   0.09123   0.08438  -0.0124   1.0000   0.0443
  -8.250  -0.5757   0.08659   0.07982  -0.0152   1.0000   0.0432
  -8.000  -0.5789   0.08160   0.07491  -0.0189   1.0000   0.0421
  -7.750  -0.5809   0.07571   0.06903  -0.0235   1.0000   0.0406
  -7.250  -0.5830   0.06374   0.05660  -0.0309   1.0000   0.0378
  -7.000  -0.5770   0.05920   0.05188  -0.0316   1.0000   0.0375
  -6.750  -0.5692   0.05469   0.04709  -0.0322   1.0000   0.0374
  -6.500  -0.5594   0.05023   0.04218  -0.0324   1.0000   0.0378
  -6.250  -0.5475   0.04664   0.03852  -0.0321   1.0000   0.0399
  -6.000  -0.5327   0.04346   0.03506  -0.0317   1.0000   0.0425
  -5.750  -0.5163   0.03978   0.03079  -0.0311   1.0000   0.0439
  -5.500  -0.4977   0.03627   0.02672  -0.0302   1.0000   0.0451
  -5.250  -0.4769   0.03315   0.02294  -0.0290   1.0000   0.0475
  -5.000  -0.4563   0.03058   0.02004  -0.0281   1.0000   0.0529
  -4.750  -0.4334   0.02829   0.01739  -0.0270   1.0000   0.0578
  -4.500  -0.4100   0.02622   0.01498  -0.0258   1.0000   0.0656
  -4.250  -0.3864   0.02462   0.01308  -0.0246   1.0000   0.0761
  -4.000  -0.3636   0.02297   0.01134  -0.0234   1.0000   0.0857
  -3.750  -0.3403   0.02181   0.01002  -0.0224   1.0000   0.1036
  -3.500  -0.3161   0.02059   0.00869  -0.0216   1.0000   0.1184
  -3.250  -0.2936   0.01933   0.00752  -0.0207   1.0000   0.1498
  -3.000  -0.2730   0.01765   0.00663  -0.0200   1.0000   0.2869
  -2.750  -0.2585   0.01622   0.00632  -0.0173   1.0000   0.5345
  -2.500  -0.2336   0.01511   0.00630  -0.0137   1.0000   0.8264
  -2.250  -0.1449   0.01487   0.00549  -0.0251   1.0000   1.0000
  -2.000  -0.1279   0.01476   0.00504  -0.0236   1.0000   1.0000
  -1.750  -0.1099   0.01469   0.00471  -0.0221   1.0000   1.0000
  -1.500  -0.0912   0.01466   0.00445  -0.0208   1.0000   1.0000
  -1.250  -0.0720   0.01465   0.00425  -0.0195   1.0000   1.0000
  -1.000  -0.0526   0.01467   0.00407  -0.0182   1.0000   1.0000
  -0.750  -0.0330   0.01472   0.00397  -0.0170   1.0000   1.0000
  -0.500  -0.0133   0.01479   0.00392  -0.0158   1.0000   1.0000
  -0.250   0.0064   0.01489   0.00392  -0.0147   1.0000   1.0000
   0.000   0.0261   0.01501   0.00395  -0.0135   1.0000   1.0000
   0.250   0.0456   0.01517   0.00404  -0.0124   1.0000   1.0000
   0.500   0.0651   0.01535   0.00418  -0.0114   1.0000   1.0000
   0.750   0.0844   0.01556   0.00437  -0.0104   1.0000   1.0000
   1.000   0.1036   0.01580   0.00462  -0.0094   1.0000   1.0000
   1.250   0.1226   0.01608   0.00492  -0.0085   1.0000   1.0000
   1.500   0.1522   0.01643   0.00533  -0.0098   0.9940   1.0000
   1.750   0.1964   0.01682   0.00588  -0.0138   0.9793   1.0000
   2.000   0.2407   0.01715   0.00638  -0.0177   0.9631   1.0000
   2.250   0.2852   0.01741   0.00686  -0.0214   0.9448   1.0000
   2.500   0.3326   0.01760   0.00738  -0.0254   0.9272   1.0000
   2.750   0.3774   0.01776   0.00787  -0.0287   0.9083   1.0000
   3.000   0.4204   0.01788   0.00841  -0.0313   0.8880   1.0000
   3.250   0.4707   0.01739   0.00833  -0.0330   0.8340   1.0000
   3.500   0.5039   0.01694   0.00796  -0.0305   0.7357   1.0000
   3.750   0.5240   0.01715   0.00769  -0.0260   0.5667   1.0000
   4.000   0.5363   0.01850   0.00800  -0.0220   0.3392   1.0000
   4.250   0.5479   0.02128   0.00933  -0.0201   0.1369   1.0000
   4.500   0.5688   0.02296   0.01092  -0.0189   0.1020   1.0000
   4.750   0.5902   0.02449   0.01244  -0.0179   0.0787   1.0000
   5.000   0.6123   0.02627   0.01429  -0.0166   0.0669   1.0000
   5.250   0.6357   0.02825   0.01639  -0.0155   0.0555   1.0000
   5.500   0.6631   0.03047   0.01897  -0.0144   0.0498   1.0000
   5.750   0.6881   0.03284   0.02158  -0.0135   0.0450   1.0000
   6.000   0.7106   0.03566   0.02472  -0.0125   0.0406   1.0000
   6.250   0.7328   0.03864   0.02828  -0.0112   0.0388   1.0000
   6.500   0.7520   0.04197   0.03216  -0.0097   0.0378   1.0000
   6.750   0.7681   0.04549   0.03620  -0.0083   0.0370   1.0000
   7.000   0.7814   0.04897   0.04013  -0.0070   0.0354   1.0000
   7.250   0.7939   0.05215   0.04354  -0.0062   0.0333   1.0000
   7.500   0.8017   0.05625   0.04793  -0.0051   0.0323   1.0000
   7.750   0.8064   0.06038   0.05251  -0.0040   0.0321   1.0000
   8.000   0.8087   0.06468   0.05712  -0.0032   0.0322   1.0000
   8.250   0.8095   0.06896   0.06163  -0.0027   0.0324   1.0000
   8.500   0.8086   0.07331   0.06615  -0.0023   0.0326   1.0000
   9.000   0.7729   0.08307   0.07643  -0.0046   0.0346   1.0000
   9.250   0.7578   0.08904   0.08242  -0.0085   0.0354   1.0000
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