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

NASA/LANGLEY RC08-64C AIRFOIL (rc0864c-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: NASA/LANGLEY RC08-64C AIRFOIL (rc0864c-il)
Reynolds number: 200,000
Max Cl/Cd: 33.78 at α=1.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-rc0864c-il-200000-n5.txt
Download as CSV file: xf-rc0864c-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NASA/LANGLEY RC08-64C AIRFOIL                   
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.5629   0.10414   0.10035  -0.0073   1.0000   0.0203
  -9.250  -0.5643   0.09869   0.09494  -0.0125   1.0000   0.0206
  -9.000  -0.5652   0.09337   0.08965  -0.0173   1.0000   0.0207
  -8.750  -0.5652   0.09003   0.08635  -0.0161   1.0000   0.0213
  -8.500  -0.5618   0.08725   0.08358  -0.0164   1.0000   0.0219
  -8.250  -0.5620   0.08336   0.07972  -0.0190   1.0000   0.0225
  -8.000  -0.5664   0.07890   0.07527  -0.0226   1.0000   0.0229
  -7.750  -0.5672   0.07462   0.07096  -0.0251   1.0000   0.0237
  -7.500  -0.5664   0.07002   0.06629  -0.0275   1.0000   0.0248
  -7.250  -0.5655   0.06410   0.05997  -0.0307   1.0000   0.0266
  -7.000  -0.5659   0.05929   0.05517  -0.0301   1.0000   0.0273
  -6.750  -0.5591   0.05699   0.05291  -0.0290   1.0000   0.0281
  -6.250  -0.5439   0.05151   0.04724  -0.0265   1.0000   0.0310
  -4.250  -0.4302   0.02522   0.01818  -0.0103   0.9967   0.0160
  -4.000  -0.3991   0.02291   0.01564  -0.0112   0.9939   0.0148
  -3.750  -0.3651   0.02148   0.01356  -0.0114   0.9903   0.0130
  -3.500  -0.3328   0.01964   0.01164  -0.0124   0.9876   0.0126
  -3.250  -0.2997   0.01830   0.01013  -0.0134   0.9848   0.0123
  -3.000  -0.2680   0.01717   0.00890  -0.0141   0.9807   0.0121
  -2.750  -0.2347   0.01618   0.00783  -0.0151   0.9774   0.0121
  -2.500  -0.2010   0.01531   0.00692  -0.0162   0.9743   0.0122
  -2.250  -0.1707   0.01460   0.00618  -0.0168   0.9688   0.0133
  -2.000  -0.1378   0.01381   0.00545  -0.0180   0.9649   0.0148
  -1.750  -0.1041   0.01326   0.00488  -0.0193   0.9609   0.0152
  -1.500  -0.0740   0.01285   0.00442  -0.0197   0.9541   0.0156
  -1.250  -0.0400   0.01249   0.00401  -0.0209   0.9494   0.0163
  -1.000  -0.0101   0.01217   0.00366  -0.0212   0.9419   0.0174
  -0.750   0.0222   0.01189   0.00337  -0.0220   0.9354   0.0215
  -0.500   0.0427   0.01013   0.00311  -0.0214   0.9268   0.4073
  -0.250   0.0639   0.00874   0.00351  -0.0185   0.9212   0.8889
   0.000   0.1150   0.00892   0.00374  -0.0223   0.9188   0.9535
   0.250   0.1665   0.00900   0.00378  -0.0268   0.9094   0.9766
   0.500   0.2273   0.00883   0.00348  -0.0326   0.8643   0.9934
   0.750   0.2651   0.00875   0.00328  -0.0342   0.8240   0.9989
   1.000   0.2882   0.00888   0.00298  -0.0324   0.7237   1.0000
   1.250   0.3101   0.00918   0.00287  -0.0309   0.6330   1.0000
   1.500   0.3278   0.01025   0.00287  -0.0290   0.4071   1.0000
   1.750   0.3425   0.01252   0.00335  -0.0278   0.0225   1.0000
   2.000   0.3655   0.01279   0.00358  -0.0269   0.0181   1.0000
   2.250   0.3885   0.01303   0.00385  -0.0259   0.0169   1.0000
   2.500   0.4112   0.01331   0.00418  -0.0249   0.0161   1.0000
   2.750   0.4336   0.01366   0.00458  -0.0239   0.0155   1.0000
   3.000   0.4555   0.01408   0.00506  -0.0227   0.0150   1.0000
   3.250   0.4766   0.01463   0.00566  -0.0214   0.0145   1.0000
   3.500   0.4972   0.01522   0.00628  -0.0201   0.0137   1.0000
   3.750   0.5180   0.01577   0.00687  -0.0187   0.0126   1.0000
   4.000   0.5381   0.01648   0.00759  -0.0173   0.0120   1.0000
   4.250   0.5582   0.01726   0.00838  -0.0158   0.0118   1.0000
   4.500   0.5786   0.01813   0.00926  -0.0144   0.0116   1.0000
   4.750   0.5995   0.01908   0.01025  -0.0130   0.0115   1.0000
   5.000   0.6211   0.02011   0.01133  -0.0118   0.0114   1.0000
   5.250   0.6431   0.02120   0.01252  -0.0105   0.0115   1.0000
   5.500   0.6652   0.02240   0.01389  -0.0093   0.0116   1.0000
   5.750   0.6872   0.02378   0.01551  -0.0079   0.0119   1.0000
   6.000   0.7086   0.02540   0.01740  -0.0065   0.0122   1.0000
   6.250   0.7290   0.02726   0.01954  -0.0050   0.0127   1.0000
   6.500   0.7480   0.02932   0.02190  -0.0034   0.0131   1.0000
   6.750   0.7653   0.03161   0.02445  -0.0017   0.0136   1.0000
   7.000   0.7811   0.03458   0.02757  -0.0004   0.0141   1.0000
  11.500   0.5592   0.11926   0.11562  -0.0025   0.0353   1.0000
  11.750   0.5572   0.12294   0.11929  -0.0037   0.0341   1.0000
<< Back to NASA/LANGLEY RC08-64C AIRFOIL (rc0864c-il)

Polar data table (+)

Polar graphs


<< Back to NASA/LANGLEY RC08-64C AIRFOIL (rc0864c-il)