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RAF 33 AIRFOIL (raf33-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: RAF 33 AIRFOIL (raf33-il)
Reynolds number: 50,000
Max Cl/Cd: 19.38 at α=3.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-raf33-il-50000-n5.txt
Download as CSV file: xf-raf33-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: RAF 33 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
  ------ -------- --------- --------- -------- -------- --------
 -10.000  -0.2562   0.11116   0.10486  -0.0374   0.9012   0.1280
  -9.750  -0.2633   0.10971   0.10337  -0.0423   0.8854   0.1323
  -9.500  -0.2774   0.10845   0.10211  -0.0473   0.8706   0.1332
  -9.250  -0.2419   0.10220   0.09576  -0.0443   0.8611   0.1366
  -9.000  -0.2312   0.09909   0.09259  -0.0448   0.8505   0.1395
  -8.750  -0.2263   0.09635   0.08983  -0.0459   0.8403   0.1425
  -8.500  -0.2286   0.09395   0.08743  -0.0479   0.8301   0.1466
  -8.250  -0.2600   0.09321   0.08674  -0.0522   0.8196   0.1498
  -8.000  -0.2628   0.09035   0.08389  -0.0531   0.8110   0.1506
  -7.750  -0.2377   0.08575   0.07925  -0.0519   0.8045   0.1529
  -7.500  -0.2266   0.08288   0.07637  -0.0515   0.7967   0.1567
  -7.250  -0.2489   0.07658   0.06972  -0.0580   0.7904   0.0989
  -7.000  -0.2384   0.07340   0.06654  -0.0578   0.7827   0.0968
  -6.500  -0.2393   0.06580   0.05840  -0.0590   0.7703   0.0847
  -6.250  -0.2282   0.06308   0.05562  -0.0584   0.7636   0.0836
  -6.000  -0.2188   0.06039   0.05277  -0.0574   0.7583   0.0825
  -5.750  -0.2089   0.05778   0.04998  -0.0568   0.7516   0.0814
  -5.500  -0.1987   0.05519   0.04717  -0.0557   0.7457   0.0805
  -5.250  -0.1873   0.05274   0.04443  -0.0543   0.7410   0.0808
  -5.000  -0.1753   0.05058   0.04199  -0.0533   0.7343   0.0815
  -4.750  -0.1616   0.04842   0.03950  -0.0518   0.7286   0.0822
  -4.500  -0.1461   0.04627   0.03698  -0.0502   0.7243   0.0824
  -4.250  -0.1304   0.04447   0.03487  -0.0490   0.7178   0.0822
  -4.000  -0.1133   0.04268   0.03272  -0.0476   0.7123   0.0822
  -3.750  -0.0938   0.04095   0.03060  -0.0460   0.7081   0.0824
  -3.500  -0.0749   0.03961   0.02892  -0.0449   0.7018   0.0831
  -3.250  -0.0540   0.03839   0.02725  -0.0436   0.6963   0.0849
  -3.000  -0.0312   0.03717   0.02587  -0.0428   0.6922   0.0874
  -2.750  -0.0088   0.03638   0.02494  -0.0422   0.6864   0.0896
  -2.500   0.0155   0.03555   0.02389  -0.0417   0.6806   0.0914
  -2.250   0.0436   0.03458   0.02265  -0.0415   0.6763   0.0936
  -2.000   0.0728   0.03382   0.02164  -0.0416   0.6717   0.0968
  -1.750   0.1041   0.03337   0.02103  -0.0426   0.6653   0.1018
  -1.500   0.1426   0.03271   0.02029  -0.0445   0.6607   0.1091
  -1.250   0.1832   0.03196   0.01928  -0.0461   0.6573   0.1168
  -1.000   0.2083   0.03201   0.01943  -0.0467   0.6500   0.1271
  -0.750   0.2349   0.03169   0.01910  -0.0465   0.6451   0.1409
  -0.500   0.2612   0.03122   0.01864  -0.0459   0.6415   0.1637
  -0.250   0.2779   0.03140   0.01905  -0.0449   0.6347   0.1956
   0.000   0.2956   0.03090   0.01905  -0.0435   0.6294   0.2912
   0.250   0.4782   0.02893   0.01824  -0.0708   0.6260   1.0000
   0.500   0.4969   0.02958   0.01874  -0.0698   0.6202   1.0000
   0.750   0.5151   0.03022   0.01926  -0.0688   0.6139   1.0000
   1.000   0.5366   0.03045   0.01932  -0.0675   0.6097   1.0000
   1.250   0.5536   0.03120   0.01999  -0.0664   0.6036   1.0000
   1.500   0.5702   0.03194   0.02065  -0.0651   0.5972   1.0000
   1.750   0.5919   0.03215   0.02073  -0.0638   0.5930   1.0000
   2.000   0.6053   0.03317   0.02172  -0.0625   0.5861   1.0000
   2.250   0.6215   0.03386   0.02236  -0.0611   0.5800   1.0000
   2.500   0.6442   0.03397   0.02236  -0.0597   0.5761   1.0000
   2.750   0.6511   0.03545   0.02385  -0.0580   0.5676   1.0000
   3.000   0.6697   0.03589   0.02424  -0.0565   0.5623   1.0000
   3.250   0.6945   0.03583   0.02410  -0.0553   0.5589   1.0000
   3.500   0.6932   0.03782   0.02614  -0.0530   0.5487   1.0000
   3.750   0.7161   0.03790   0.02616  -0.0517   0.5445   1.0000
   4.250   0.7353   0.04001   0.02828  -0.0477   0.5300   1.0000
   4.500   0.7633   0.03969   0.02791  -0.0467   0.5268   1.0000
   4.750   0.7523   0.04216   0.03042  -0.0435   0.5152   1.0000
   5.000   0.7797   0.04188   0.03012  -0.0425   0.5117   1.0000
   5.250   0.7665   0.04446   0.03274  -0.0392   0.5001   1.0000
   5.500   0.7937   0.04415   0.03241  -0.0381   0.4964   1.0000
   6.000   0.8049   0.04664   0.03494  -0.0337   0.4809   1.0000
   6.500   0.8089   0.04952   0.03783  -0.0290   0.4650   1.0000
   7.000   0.8159   0.05248   0.04082  -0.0250   0.4493   1.0000
   7.500   0.8250   0.05559   0.04400  -0.0216   0.4338   1.0000
   8.000   0.8338   0.05906   0.04753  -0.0188   0.4184   1.0000
   8.500   0.8420   0.06286   0.05140  -0.0164   0.4032   1.0000
   9.000   0.8469   0.06735   0.05600  -0.0145   0.3883   1.0000
   9.500   0.8501   0.07228   0.06103  -0.0130   0.3737   1.0000
   9.750   0.8686   0.07291   0.06173  -0.0120   0.3702   1.0000
  10.000   0.8473   0.07837   0.06723  -0.0121   0.3598   1.0000
  10.250   0.8670   0.07898   0.06792  -0.0112   0.3570   1.0000
  10.500   0.8453   0.08466   0.07363  -0.0116   0.3472   1.0000
  10.750   0.8590   0.08601   0.07509  -0.0110   0.3432   1.0000
  11.000   0.8803   0.08647   0.07566  -0.0101   0.3408   1.0000
  11.250   0.8495   0.09374   0.08295  -0.0113   0.3315   1.0000
  11.500   0.8599   0.09575   0.08505  -0.0110   0.3281   1.0000
  11.750   0.8780   0.09673   0.08615  -0.0103   0.3257   1.0000
  12.000   0.8518   0.10348   0.09294  -0.0117   0.3175   1.0000
  12.250   0.8603   0.10575   0.09531  -0.0116   0.3136   1.0000
  12.500   0.8800   0.10644   0.09615  -0.0108   0.3108   1.0000
  12.750   0.8593   0.11266   0.10242  -0.0123   0.3039   1.0000
  13.000   0.8712   0.11404   0.10392  -0.0118   0.2981   1.0000
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