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

ONERA OA213 AIRFOIL (oa213-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: ONERA OA213 AIRFOIL (oa213-il)
Reynolds number: 50,000
Max Cl/Cd: 29.79 at α=6.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-oa213-il-50000-n5.txt
Download as CSV file: xf-oa213-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: ONERA OA213 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
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.3656   0.10170   0.09599  -0.0204   1.0000   0.1187
  -8.500  -0.3636   0.09775   0.09209  -0.0237   1.0000   0.1180
  -8.250  -0.3955   0.08637   0.08056  -0.0376   1.0000   0.0571
  -8.000  -0.3965   0.08393   0.07817  -0.0361   1.0000   0.0560
  -7.750  -0.3992   0.08074   0.07499  -0.0367   0.9952   0.0548
  -7.500  -0.3909   0.07573   0.06987  -0.0413   0.9857   0.0529
  -7.250  -0.3841   0.07029   0.06423  -0.0458   0.9764   0.0508
  -7.000  -0.3810   0.06401   0.05742  -0.0498   0.9664   0.0478
  -6.750  -0.3667   0.06020   0.05339  -0.0514   0.9586   0.0475
  -6.500  -0.3517   0.05628   0.04909  -0.0529   0.9512   0.0477
  -6.250  -0.3395   0.05296   0.04539  -0.0530   0.9423   0.0483
  -6.000  -0.3210   0.05072   0.04308  -0.0533   0.9349   0.0498
  -5.750  -0.3024   0.04807   0.04011  -0.0534   0.9274   0.0507
  -5.500  -0.2860   0.04547   0.03713  -0.0527   0.9190   0.0510
  -5.250  -0.2637   0.04287   0.03409  -0.0525   0.9124   0.0514
  -5.000  -0.2453   0.04075   0.03159  -0.0514   0.9038   0.0520
  -4.750  -0.2208   0.03881   0.02921  -0.0510   0.8973   0.0542
  -4.500  -0.1987   0.03711   0.02698  -0.0499   0.8889   0.0569
  -4.250  -0.1721   0.03547   0.02512  -0.0496   0.8826   0.0593
  -4.000  -0.1484   0.03420   0.02369  -0.0489   0.8744   0.0618
  -3.750  -0.1190   0.03306   0.02225  -0.0486   0.8684   0.0671
  -3.500  -0.0946   0.03200   0.02110  -0.0479   0.8599   0.0718
  -3.250  -0.0640   0.03112   0.02007  -0.0476   0.8539   0.0793
  -3.000  -0.0386   0.03034   0.01927  -0.0471   0.8454   0.0876
  -2.750  -0.0104   0.02960   0.01844  -0.0466   0.8389   0.1005
  -2.500   0.0120   0.02900   0.01784  -0.0457   0.8299   0.1175
  -2.250   0.0353   0.02825   0.01719  -0.0446   0.8230   0.1472
  -2.000   0.0525   0.02728   0.01670  -0.0433   0.8135   0.2166
  -1.750   0.2117   0.02526   0.01670  -0.0598   0.8163   1.0000
  -1.500   0.2310   0.02529   0.01648  -0.0583   0.8067   1.0000
  -1.250   0.2497   0.02538   0.01636  -0.0571   0.7948   1.0000
  -1.000   0.2683   0.02538   0.01617  -0.0554   0.7837   1.0000
  -0.750   0.2870   0.02526   0.01584  -0.0532   0.7733   1.0000
  -0.500   0.3052   0.02526   0.01568  -0.0516   0.7599   1.0000
  -0.250   0.3236   0.02522   0.01549  -0.0498   0.7470   1.0000
   0.000   0.3422   0.02512   0.01524  -0.0479   0.7349   1.0000
   0.250   0.3613   0.02485   0.01482  -0.0455   0.7247   1.0000
   0.500   0.3801   0.02483   0.01468  -0.0438   0.7110   1.0000
   1.000   0.4181   0.02466   0.01430  -0.0402   0.6845   1.0000
   1.250   0.4377   0.02449   0.01403  -0.0382   0.6717   1.0000
   1.500   0.4578   0.02423   0.01365  -0.0361   0.6595   1.0000
   1.750   0.4779   0.02407   0.01339  -0.0342   0.6457   1.0000
   2.000   0.4980   0.02396   0.01319  -0.0324   0.6304   1.0000
   2.250   0.5183   0.02384   0.01299  -0.0307   0.6147   1.0000
   2.500   0.5390   0.02371   0.01277  -0.0289   0.5983   1.0000
   2.750   0.5601   0.02359   0.01253  -0.0272   0.5812   1.0000
   3.000   0.5807   0.02364   0.01251  -0.0257   0.5609   1.0000
   3.250   0.6019   0.02366   0.01243  -0.0242   0.5405   1.0000
   3.500   0.6237   0.02366   0.01229  -0.0227   0.5200   1.0000
   3.750   0.6453   0.02379   0.01227  -0.0214   0.4977   1.0000
   4.000   0.6667   0.02401   0.01235  -0.0202   0.4747   1.0000
   4.250   0.6886   0.02424   0.01241  -0.0190   0.4537   1.0000
   4.500   0.7097   0.02467   0.01270  -0.0180   0.4313   1.0000
   4.750   0.7309   0.02512   0.01299  -0.0171   0.4113   1.0000
   5.000   0.7522   0.02564   0.01334  -0.0162   0.3935   1.0000
   5.250   0.7735   0.02622   0.01380  -0.0154   0.3772   1.0000
   5.500   0.7948   0.02685   0.01431  -0.0147   0.3622   1.0000
   5.750   0.8162   0.02752   0.01489  -0.0141   0.3486   1.0000
   6.000   0.8380   0.02821   0.01550  -0.0135   0.3368   1.0000
   6.250   0.8604   0.02888   0.01605  -0.0129   0.3264   1.0000
   6.500   0.8819   0.02968   0.01689  -0.0124   0.3155   1.0000
   6.750   0.9044   0.03044   0.01761  -0.0120   0.3065   1.0000
   7.000   0.9270   0.03123   0.01842  -0.0115   0.2981   1.0000
   7.250   0.9493   0.03208   0.01929  -0.0112   0.2903   1.0000
   7.500   0.9713   0.03294   0.02018  -0.0108   0.2827   1.0000
   7.750   0.9939   0.03382   0.02106  -0.0104   0.2762   1.0000
   8.000   1.0145   0.03485   0.02225  -0.0101   0.2695   1.0000
   8.250   1.0386   0.03570   0.02306  -0.0098   0.2644   1.0000
   8.500   1.0581   0.03692   0.02445  -0.0094   0.2589   1.0000
   8.750   1.0774   0.03811   0.02580  -0.0090   0.2535   1.0000
   9.000   1.1006   0.03904   0.02673  -0.0088   0.2490   1.0000
   9.250   1.1177   0.04045   0.02833  -0.0083   0.2444   1.0000
   9.500   1.1324   0.04203   0.03015  -0.0078   0.2399   1.0000
   9.750   1.1507   0.04342   0.03168  -0.0074   0.2363   1.0000
  10.000   1.1730   0.04460   0.03292  -0.0072   0.2334   1.0000
  10.250   1.1864   0.04641   0.03493  -0.0067   0.2303   1.0000
  10.500   1.1879   0.04892   0.03780  -0.0057   0.2270   1.0000
  10.750   1.1911   0.05129   0.04040  -0.0049   0.2238   1.0000
  11.000   1.1982   0.05331   0.04259  -0.0042   0.2209   1.0000
  11.250   1.2136   0.05485   0.04423  -0.0037   0.2184   1.0000
  11.500   1.2239   0.05685   0.04634  -0.0031   0.2163   1.0000
  11.750   1.1774   0.06300   0.05285  -0.0019   0.2143   1.0000
  12.000   0.9468   0.09838   0.08849  -0.0166   0.2068   1.0000
<< Back to ONERA OA213 AIRFOIL (oa213-il)

Polar data table (+)

Polar graphs


<< Back to ONERA OA213 AIRFOIL (oa213-il)