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NASA RC(5)-10 AIRFOIL (rc510-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: NASA RC(5)-10 AIRFOIL (rc510-il)
Reynolds number: 50,000
Max Cl/Cd: 31.62 at α=5.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-rc510-il-50000.txt
Download as CSV file: xf-rc510-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NASA RC(5)-10 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.5382   0.12215   0.11570   0.0077   1.0000   0.1830
  -9.250  -0.5080   0.11590   0.10942   0.0109   1.0000   0.1944
  -9.000  -0.5324   0.11473   0.10839   0.0052   1.0000   0.1986
  -8.750  -0.5065   0.10926   0.10291   0.0080   1.0000   0.2120
  -8.500  -0.4960   0.10504   0.09872   0.0082   1.0000   0.2211
  -8.250  -0.5199   0.10337   0.09722   0.0029   1.0000   0.2297
  -8.000  -0.5016   0.09889   0.09274   0.0047   1.0000   0.2441
  -7.750  -0.4900   0.09493   0.08883   0.0053   1.0000   0.2579
  -7.500  -0.4868   0.09152   0.08550   0.0050   1.0000   0.2736
  -7.250  -0.4936   0.08862   0.08271   0.0036   1.0000   0.2913
  -7.000  -0.4727   0.08472   0.07884   0.0064   1.0000   0.3148
  -6.250  -0.4256   0.07665   0.07090   0.0179   1.0000   0.4533
  -6.000  -0.3957   0.07454   0.06879   0.0246   1.0000   0.5373
  -5.750  -0.3388   0.07057   0.06473   0.0295   1.0000   0.6412
  -5.500  -0.2618   0.06499   0.05903   0.0299   1.0000   0.7823
  -4.750  -0.3118   0.05865   0.05312   0.0272   1.0000   0.6774
  -4.500  -0.3852   0.05774   0.05256   0.0288   1.0000   0.6052
  -4.250  -0.4386   0.05762   0.05263   0.0355   1.0000   0.6017
  -4.000  -0.4273   0.04398   0.03629  -0.0106   1.0000   0.1956
  -3.750  -0.4035   0.04098   0.03260  -0.0101   1.0000   0.1643
  -3.500  -0.3815   0.03876   0.02970  -0.0091   1.0000   0.1512
  -3.250  -0.3613   0.03663   0.02724  -0.0082   1.0000   0.1504
  -3.000  -0.3395   0.03478   0.02499  -0.0074   1.0000   0.1506
  -2.750  -0.3159   0.03299   0.02281  -0.0067   1.0000   0.1504
  -2.500  -0.2912   0.03140   0.02085  -0.0062   1.0000   0.1537
  -2.250  -0.2671   0.02993   0.01928  -0.0059   1.0000   0.1628
  -2.000  -0.2413   0.02861   0.01772  -0.0055   1.0000   0.1717
  -1.750  -0.2152   0.02742   0.01644  -0.0052   1.0000   0.1875
  -1.500  -0.1874   0.02635   0.01534  -0.0051   1.0000   0.2105
  -1.250  -0.1600   0.02529   0.01435  -0.0050   1.0000   0.2428
  -1.000  -0.0974   0.02118   0.01356  -0.0096   1.0000   1.0000
  -0.750  -0.0784   0.02157   0.01319  -0.0079   1.0000   1.0000
  -0.500  -0.0617   0.02193   0.01312  -0.0067   1.0000   1.0000
  -0.250  -0.0449   0.02231   0.01318  -0.0057   1.0000   1.0000
   0.000  -0.0277   0.02273   0.01333  -0.0049   1.0000   1.0000
   0.250  -0.0101   0.02318   0.01356  -0.0042   1.0000   1.0000
   0.500   0.0077   0.02367   0.01385  -0.0036   1.0000   1.0000
   0.750   0.0344   0.02432   0.01432  -0.0048   0.9967   1.0000
   1.000   0.0734   0.02521   0.01503  -0.0084   0.9882   1.0000
   1.250   0.1160   0.02620   0.01587  -0.0125   0.9772   1.0000
   1.500   0.1756   0.02737   0.01692  -0.0196   0.9580   1.0000
   1.750   0.2339   0.02833   0.01782  -0.0258   0.9333   1.0000
   2.000   0.2978   0.02917   0.01865  -0.0325   0.9089   1.0000
   2.250   0.3403   0.02982   0.01931  -0.0353   0.8873   1.0000
   2.500   0.4071   0.03028   0.01986  -0.0415   0.8648   1.0000
   2.750   0.4428   0.03073   0.02038  -0.0425   0.8417   1.0000
   3.000   0.4982   0.03073   0.02051  -0.0454   0.8181   1.0000
   3.250   0.5322   0.03077   0.02065  -0.0448   0.7926   1.0000
   3.500   0.5648   0.03061   0.02059  -0.0433   0.7660   1.0000
   3.750   0.5971   0.03005   0.02016  -0.0410   0.7380   1.0000
   4.000   0.6253   0.02926   0.01946  -0.0375   0.7078   1.0000
   4.250   0.6506   0.02818   0.01845  -0.0333   0.6733   1.0000
   4.500   0.6771   0.02665   0.01695  -0.0287   0.6372   1.0000
   4.750   0.7000   0.02537   0.01560  -0.0242   0.5928   1.0000
   5.000   0.7231   0.02443   0.01446  -0.0204   0.5468   1.0000
   5.250   0.7452   0.02419   0.01398  -0.0176   0.5031   1.0000
   5.500   0.7675   0.02441   0.01398  -0.0154   0.4665   1.0000
   5.750   0.7903   0.02499   0.01432  -0.0138   0.4370   1.0000
   6.000   0.8137   0.02577   0.01487  -0.0125   0.4122   1.0000
   6.250   0.8366   0.02687   0.01590  -0.0115   0.3908   1.0000
   6.500   0.8600   0.02807   0.01708  -0.0107   0.3726   1.0000
   6.750   0.8833   0.02937   0.01834  -0.0099   0.3560   1.0000
   7.000   0.9061   0.03085   0.01987  -0.0092   0.3418   1.0000
   7.250   0.9284   0.03245   0.02156  -0.0085   0.3287   1.0000
   7.500   0.9502   0.03413   0.02335  -0.0078   0.3161   1.0000
   7.750   0.9741   0.03586   0.02504  -0.0072   0.3051   1.0000
   8.000   0.9904   0.03799   0.02755  -0.0064   0.2944   1.0000
   8.250   1.0071   0.04024   0.03005  -0.0055   0.2842   1.0000
   8.500   1.0306   0.04223   0.03203  -0.0049   0.2748   1.0000
   8.750   1.0358   0.04539   0.03577  -0.0038   0.2668   1.0000
   9.000   1.0543   0.04768   0.03812  -0.0030   0.2575   1.0000
   9.250   1.0531   0.05158   0.04254  -0.0019   0.2515   1.0000
   9.500   1.0731   0.05399   0.04497  -0.0011   0.2432   1.0000
   9.750   1.0538   0.05934   0.05084   0.0000   0.2399   1.0000
  10.000   1.0266   0.06551   0.05733   0.0004   0.2382   1.0000
  10.250   0.9814   0.07364   0.06560  -0.0007   0.2406   1.0000
  10.500   0.9402   0.08275   0.07469  -0.0039   0.2431   1.0000
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