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RAE(NPL) 5212 AIRFOIL (rae5212-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: RAE(NPL) 5212 AIRFOIL (rae5212-il)
Reynolds number: 100,000
Max Cl/Cd: 47.68 at α=4.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-rae5212-il-100000.txt
Download as CSV file: xf-rae5212-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: RAE(NPL) 5212 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.250  -0.5366   0.09149   0.08655  -0.0441   1.0000   0.1538
  -9.000  -0.4659   0.09059   0.08553  -0.0348   1.0000   0.1624
  -8.750  -0.5064   0.08623   0.08134  -0.0385   1.0000   0.1686
  -8.500  -0.4847   0.08392   0.07903  -0.0348   1.0000   0.1756
  -8.250  -0.5097   0.08059   0.07583  -0.0352   1.0000   0.1828
  -8.000  -0.5620   0.07613   0.07152  -0.0365   1.0000   0.1839
  -7.750  -0.5343   0.07513   0.07056  -0.0313   1.0000   0.1959
  -7.500  -0.5869   0.07036   0.06584  -0.0330   1.0000   0.1993
  -7.000  -0.6615   0.04950   0.04274  -0.0361   1.0000   0.0942
  -6.750  -0.6465   0.04470   0.03813  -0.0353   1.0000   0.0901
  -6.500  -0.6369   0.04031   0.03281  -0.0331   1.0000   0.0795
  -6.250  -0.6221   0.03701   0.02930  -0.0319   1.0000   0.0776
  -6.000  -0.6055   0.03423   0.02617  -0.0306   1.0000   0.0760
  -5.750  -0.5869   0.03189   0.02346  -0.0293   1.0000   0.0752
  -5.500  -0.5671   0.03013   0.02139  -0.0281   1.0000   0.0766
  -5.250  -0.5464   0.02878   0.01969  -0.0268   1.0000   0.0790
  -5.000  -0.5250   0.02753   0.01812  -0.0256   1.0000   0.0803
  -4.750  -0.5033   0.02566   0.01623  -0.0246   1.0000   0.0825
  -4.500  -0.4830   0.02462   0.01524  -0.0237   1.0000   0.0874
  -4.250  -0.4619   0.02377   0.01428  -0.0225   1.0000   0.0920
  -4.000  -0.4414   0.02264   0.01320  -0.0212   1.0000   0.0968
  -3.750  -0.4215   0.02192   0.01256  -0.0201   1.0000   0.1050
  -3.500  -0.4020   0.02107   0.01185  -0.0189   1.0000   0.1144
  -3.250  -0.3823   0.02034   0.01123  -0.0178   1.0000   0.1301
  -3.000  -0.3621   0.01937   0.01050  -0.0171   1.0000   0.1672
  -2.750  -0.3528   0.01772   0.01113  -0.0137   1.0000   0.6035
  -2.500  -0.3397   0.01855   0.01197  -0.0099   1.0000   0.6569
  -2.250  -0.3258   0.01910   0.01251  -0.0066   1.0000   0.6887
  -2.000  -0.3112   0.01954   0.01291  -0.0036   0.9998   0.7165
  -1.750  -0.2840   0.02033   0.01372  -0.0025   0.9938   0.7469
  -1.500  -0.2633   0.02090   0.01434   0.0002   0.9871   0.7808
  -1.250  -0.2453   0.02132   0.01481   0.0035   0.9805   0.8126
  -1.000  -0.2218   0.02153   0.01501   0.0049   0.9739   0.8327
  -0.750  -0.1916   0.02162   0.01500   0.0038   0.9670   0.8422
  -0.500  -0.1566   0.02180   0.01508   0.0017   0.9607   0.8496
  -0.250  -0.1271   0.02186   0.01508   0.0007   0.9531   0.8577
   0.000  -0.0891   0.02208   0.01523  -0.0019   0.9464   0.8649
   0.250  -0.0602   0.02212   0.01522  -0.0028   0.9378   0.8729
   0.500  -0.0183   0.02237   0.01544  -0.0060   0.9310   0.8805
   0.750   0.0089   0.02239   0.01544  -0.0066   0.9215   0.8892
   1.000   0.0518   0.02261   0.01564  -0.0098   0.9146   0.8972
   1.250   0.0819   0.02264   0.01569  -0.0108   0.9045   0.9069
   1.500   0.1203   0.02274   0.01582  -0.0132   0.8953   0.9162
   1.750   0.1676   0.02278   0.01589  -0.0170   0.8863   0.9261
   2.000   0.2096   0.02275   0.01592  -0.0198   0.8742   0.9369
   2.250   0.2638   0.02259   0.01584  -0.0245   0.8629   0.9465
   2.500   0.3288   0.02212   0.01545  -0.0305   0.8532   0.9542
   2.750   0.4014   0.02068   0.01409  -0.0363   0.8369   0.9576
   3.000   0.4731   0.01868   0.01218  -0.0410   0.8153   0.9595
   3.250   0.5313   0.01748   0.01107  -0.0447   0.7961   0.9652
   3.500   0.5856   0.01647   0.01014  -0.0480   0.7744   0.9718
   3.750   0.6344   0.01571   0.00946  -0.0507   0.7444   0.9805
   4.000   0.6806   0.01510   0.00888  -0.0531   0.7013   0.9912
   4.250   0.7017   0.01476   0.00844  -0.0514   0.6402   1.0000
   4.500   0.7047   0.01478   0.00795  -0.0463   0.5271   1.0000
   4.750   0.7001   0.01635   0.00817  -0.0408   0.3352   1.0000
   5.000   0.7064   0.01811   0.00902  -0.0379   0.2399   1.0000
   5.250   0.7247   0.01933   0.00983  -0.0368   0.2074   1.0000
   5.500   0.7477   0.02041   0.01064  -0.0363   0.1901   1.0000
   5.750   0.7733   0.02132   0.01145  -0.0362   0.1774   1.0000
   6.000   0.8004   0.02231   0.01236  -0.0363   0.1678   1.0000
   6.250   0.8296   0.02347   0.01332  -0.0368   0.1606   1.0000
   6.500   0.8585   0.02454   0.01447  -0.0371   0.1548   1.0000
   6.750   0.8865   0.02558   0.01549  -0.0373   0.1488   1.0000
   7.000   0.9178   0.02717   0.01692  -0.0383   0.1439   1.0000
   7.250   0.9447   0.02834   0.01831  -0.0381   0.1406   1.0000
   7.500   0.9712   0.02966   0.01978  -0.0380   0.1370   1.0000
   7.750   0.9968   0.03098   0.02118  -0.0378   0.1332   1.0000
   8.000   1.0249   0.03309   0.02322  -0.0384   0.1298   1.0000
   8.250   1.0461   0.03479   0.02526  -0.0374   0.1280   1.0000
   8.500   1.0657   0.03663   0.02745  -0.0361   0.1264   1.0000
   8.750   1.0832   0.03863   0.02980  -0.0347   0.1245   1.0000
   9.000   1.0994   0.04069   0.03216  -0.0333   0.1223   1.0000
   9.250   1.1151   0.04284   0.03458  -0.0319   0.1202   1.0000
   9.500   1.1282   0.04543   0.03745  -0.0302   0.1192   1.0000
   9.750   1.1393   0.04818   0.04049  -0.0285   0.1183   1.0000
  10.000   1.1573   0.05081   0.04317  -0.0279   0.1165   1.0000
  10.250   1.1672   0.05519   0.04770  -0.0269   0.1149   1.0000
  10.500   1.1654   0.05883   0.05170  -0.0242   0.1146   1.0000
  10.750   1.1631   0.06287   0.05605  -0.0218   0.1146   1.0000
  11.000   1.1567   0.06706   0.06054  -0.0192   0.1147   1.0000
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