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RAE 103 AIRFOIL (rae103-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: RAE 103 AIRFOIL (rae103-il)
Reynolds number: 500,000
Max Cl/Cd: 54.84 at α=3.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-rae103-il-500000.txt
Download as CSV file: xf-rae103-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: RAE 103 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.750  -0.5764   0.07874   0.07663  -0.0214   1.0000   0.0238
 -10.500  -0.6726   0.05434   0.05198  -0.0371   1.0000   0.0219
  -9.750  -0.8439   0.04431   0.04093  -0.0283   1.0000   0.0139
  -9.500  -0.8559   0.04018   0.03650  -0.0249   1.0000   0.0137
  -9.250  -0.8626   0.03611   0.03209  -0.0215   1.0000   0.0135
  -9.000  -0.8623   0.03252   0.02813  -0.0184   1.0000   0.0135
  -8.750  -0.8546   0.02986   0.02514  -0.0158   1.0000   0.0137
  -8.500  -0.8411   0.02831   0.02333  -0.0137   1.0000   0.0141
  -8.250  -0.8258   0.02729   0.02208  -0.0118   1.0000   0.0144
  -8.000  -0.8145   0.02417   0.01861  -0.0094   1.0000   0.0145
  -7.750  -0.8021   0.02063   0.01471  -0.0070   1.0000   0.0150
  -7.500  -0.7852   0.01893   0.01289  -0.0053   1.0000   0.0156
  -7.250  -0.7666   0.01799   0.01190  -0.0038   1.0000   0.0163
  -7.000  -0.7476   0.01721   0.01105  -0.0023   1.0000   0.0172
  -6.750  -0.7291   0.01639   0.01016  -0.0006   1.0000   0.0181
  -6.500  -0.7110   0.01564   0.00934   0.0012   1.0000   0.0191
  -6.250  -0.6910   0.01533   0.00897   0.0026   1.0000   0.0202
  -6.000  -0.6794   0.01400   0.00755   0.0055   1.0000   0.0220
  -5.750  -0.6642   0.01338   0.00691   0.0078   1.0000   0.0237
  -5.500  -0.6449   0.01293   0.00643   0.0092   0.9995   0.0258
  -5.250  -0.6080   0.01249   0.00593   0.0070   0.9961   0.0285
  -5.000  -0.5745   0.01162   0.00504   0.0054   0.9917   0.0350
  -4.750  -0.5395   0.01107   0.00444   0.0035   0.9870   0.0424
  -4.500  -0.5029   0.01052   0.00397   0.0013   0.9834   0.0592
  -4.250  -0.4711   0.00988   0.00352   0.0001   0.9765   0.1003
  -4.000  -0.4373   0.00902   0.00309  -0.0019   0.9716   0.1936
  -3.750  -0.4089   0.00799   0.00267  -0.0028   0.9633   0.3382
  -3.500  -0.3807   0.00693   0.00235  -0.0036   0.9557   0.5163
  -3.250  -0.3521   0.00664   0.00226  -0.0038   0.9443   0.5885
  -3.000  -0.3220   0.00651   0.00216  -0.0041   0.9328   0.6251
  -2.750  -0.2930   0.00641   0.00208  -0.0042   0.9204   0.6497
  -2.500  -0.2652   0.00636   0.00200  -0.0041   0.9070   0.6696
  -2.250  -0.2383   0.00632   0.00195  -0.0037   0.8935   0.6864
  -2.000  -0.2119   0.00630   0.00189  -0.0032   0.8806   0.7020
  -1.750  -0.1855   0.00630   0.00185  -0.0027   0.8685   0.7164
  -1.500  -0.1593   0.00629   0.00182  -0.0023   0.8565   0.7298
  -1.250  -0.1330   0.00629   0.00179  -0.0018   0.8450   0.7423
  -1.000  -0.1066   0.00631   0.00178  -0.0014   0.8343   0.7546
  -0.750  -0.0800   0.00630   0.00177  -0.0010   0.8246   0.7653
  -0.500  -0.0534   0.00629   0.00176  -0.0006   0.8145   0.7755
  -0.250  -0.0267   0.00630   0.00175  -0.0003   0.8048   0.7858
   0.000   0.0000   0.00631   0.00175   0.0000   0.7957   0.7957
   0.250   0.0268   0.00630   0.00175   0.0003   0.7858   0.8048
   0.500   0.0534   0.00629   0.00176   0.0006   0.7755   0.8145
   0.750   0.0800   0.00630   0.00177   0.0010   0.7653   0.8247
   1.000   0.1066   0.00631   0.00178   0.0014   0.7546   0.8343
   1.250   0.1330   0.00629   0.00180   0.0018   0.7424   0.8450
   1.500   0.1593   0.00629   0.00182   0.0023   0.7298   0.8566
   1.750   0.1855   0.00630   0.00185   0.0027   0.7163   0.8685
   2.000   0.2119   0.00630   0.00189   0.0032   0.7020   0.8806
   2.250   0.2384   0.00632   0.00194   0.0037   0.6866   0.8935
   2.500   0.2652   0.00636   0.00200   0.0041   0.6696   0.9070
   2.750   0.2930   0.00641   0.00208   0.0042   0.6500   0.9204
   3.000   0.3220   0.00651   0.00216   0.0041   0.6255   0.9328
   3.250   0.3520   0.00664   0.00226   0.0038   0.5883   0.9443
   3.500   0.3806   0.00694   0.00235   0.0036   0.5147   0.9557
   3.750   0.4089   0.00799   0.00267   0.0028   0.3381   0.9633
   4.000   0.4374   0.00901   0.00309   0.0019   0.1965   0.9716
   4.250   0.4711   0.00987   0.00352  -0.0001   0.1005   0.9765
   4.500   0.5029   0.01051   0.00396  -0.0013   0.0590   0.9834
   4.750   0.5395   0.01106   0.00444  -0.0035   0.0426   0.9870
   5.000   0.5745   0.01162   0.00503  -0.0053   0.0349   0.9917
   5.250   0.6079   0.01250   0.00594  -0.0070   0.0285   0.9961
   5.500   0.6449   0.01292   0.00642  -0.0092   0.0258   0.9995
   5.750   0.6640   0.01340   0.00692  -0.0077   0.0237   1.0000
   6.000   0.6793   0.01401   0.00755  -0.0055   0.0220   1.0000
   6.250   0.6910   0.01532   0.00896  -0.0026   0.0202   1.0000
   6.500   0.7104   0.01574   0.00944  -0.0010   0.0192   1.0000
   6.750   0.7290   0.01641   0.01018   0.0007   0.0181   1.0000
   7.000   0.7476   0.01719   0.01103   0.0023   0.0172   1.0000
   7.250   0.7665   0.01802   0.01193   0.0038   0.0164   1.0000
   7.500   0.7851   0.01899   0.01296   0.0054   0.0157   1.0000
   7.750   0.8015   0.02084   0.01493   0.0071   0.0149   1.0000
   8.000   0.8162   0.02358   0.01797   0.0091   0.0146   1.0000
   8.250   0.8257   0.02723   0.02201   0.0118   0.0144   1.0000
   8.500   0.8416   0.02817   0.02318   0.0136   0.0141   1.0000
   8.750   0.8550   0.02978   0.02505   0.0157   0.0137   1.0000
   9.000   0.8638   0.03224   0.02782   0.0182   0.0133   1.0000
   9.250   0.8632   0.03602   0.03199   0.0214   0.0134   1.0000
   9.500   0.8573   0.04001   0.03634   0.0248   0.0136   1.0000
   9.750   0.8428   0.04449   0.04113   0.0284   0.0139   1.0000
  10.000   0.8494   0.04675   0.04342   0.0300   0.0148   1.0000
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