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NACA 0010-34 a=0.8 c(li)=0.2 (naca001034a08cli02-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: NACA 0010-34 a=0.8 c(li)=0.2 (naca001034a08cli02-il)
Reynolds number: 200,000
Max Cl/Cd: 67.14 at α=4°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-naca001034a08cli02-il-200000.txt
Download as CSV file: xf-naca001034a08cli02-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 0010-34 a=0.8 c(li)=0.2                    
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.750  -0.3928   0.08858   0.08534  -0.0393   1.0000   0.0381
  -9.500  -0.3995   0.08435   0.08114  -0.0402   1.0000   0.0389
  -9.250  -0.4077   0.08001   0.07684  -0.0411   1.0000   0.0396
  -9.000  -0.4180   0.07551   0.07239  -0.0422   1.0000   0.0401
  -8.750  -0.4309   0.07093   0.06787  -0.0433   1.0000   0.0403
  -8.500  -0.4476   0.06632   0.06330  -0.0445   1.0000   0.0400
  -8.250  -0.4707   0.06218   0.05919  -0.0451   1.0000   0.0391
  -8.000  -0.4982   0.05949   0.05652  -0.0426   1.0000   0.0383
  -7.750  -0.5251   0.05735   0.05438  -0.0385   1.0000   0.0377
  -7.500  -0.5450   0.05452   0.05151  -0.0355   1.0000   0.0377
  -7.250  -0.5610   0.05148   0.04839  -0.0326   1.0000   0.0380
  -7.000  -0.5732   0.04829   0.04510  -0.0299   1.0000   0.0388
  -6.750  -0.5817   0.04512   0.04176  -0.0273   1.0000   0.0401
  -6.500  -0.5875   0.04462   0.04078  -0.0235   1.0000   0.0430
  -6.250  -0.5885   0.04332   0.03912  -0.0201   1.0000   0.0435
  -6.000  -0.6115   0.04640   0.04164  -0.0203   1.0000   0.0450
  -5.750  -0.6014   0.04313   0.03840  -0.0190   1.0000   0.0470
  -5.500  -0.5900   0.04071   0.03585  -0.0174   1.0000   0.0503
  -5.250  -0.5797   0.03875   0.03323  -0.0147   1.0000   0.0583
  -5.000  -0.5642   0.03556   0.03008  -0.0140   1.0000   0.0620
  -4.750  -0.5493   0.03348   0.02764  -0.0125   1.0000   0.0731
  -4.500  -0.5293   0.03138   0.02534  -0.0121   0.9994   0.0866
  -4.250  -0.4989   0.02905   0.02290  -0.0137   0.9969   0.1039
  -4.000  -0.4492   0.02356   0.01593  -0.0098   0.9969   0.0316
  -3.750  -0.4121   0.02252   0.01465  -0.0110   0.9945   0.0303
  -3.500  -0.3785   0.02040   0.01235  -0.0118   0.9925   0.0297
  -3.250  -0.3449   0.01863   0.01048  -0.0127   0.9904   0.0305
  -3.000  -0.3116   0.01707   0.00894  -0.0141   0.9883   0.0372
  -2.750  -0.2767   0.01631   0.00801  -0.0157   0.9853   0.0433
  -2.500  -0.2478   0.01358   0.00682  -0.0172   0.9835   0.3836
  -2.250  -0.2276   0.01217   0.00714  -0.0148   0.9810   0.7945
  -2.000  -0.2008   0.01243   0.00768  -0.0124   0.9782   0.9207
  -1.750  -0.1423   0.01303   0.00808  -0.0177   0.9791   0.9763
  -1.500  -0.0838   0.01326   0.00810  -0.0245   0.9788   0.9870
  -1.250  -0.0254   0.01342   0.00809  -0.0315   0.9783   0.9954
  -1.000   0.0233   0.01345   0.00798  -0.0366   0.9753   1.0000
  -0.750   0.0592   0.01339   0.00783  -0.0392   0.9692   1.0000
  -0.500   0.0950   0.01335   0.00773  -0.0416   0.9633   1.0000
  -0.250   0.1321   0.01331   0.00761  -0.0442   0.9573   1.0000
   0.000   0.1780   0.01327   0.00753  -0.0485   0.9543   1.0000
   0.250   0.2056   0.01321   0.00745  -0.0490   0.9453   1.0000
   0.500   0.2518   0.01312   0.00736  -0.0532   0.9421   1.0000
   0.750   0.2825   0.01307   0.00731  -0.0542   0.9341   1.0000
   1.000   0.3282   0.01292   0.00720  -0.0582   0.9304   1.0000
   1.250   0.3754   0.01271   0.00706  -0.0624   0.9276   1.0000
   1.500   0.4010   0.01262   0.00701  -0.0621   0.9177   1.0000
   1.750   0.4341   0.01245   0.00691  -0.0632   0.9103   1.0000
   2.000   0.4694   0.01219   0.00675  -0.0646   0.9033   1.0000
   2.250   0.4969   0.01199   0.00664  -0.0644   0.8932   1.0000
   2.500   0.5289   0.01167   0.00642  -0.0648   0.8845   1.0000
   2.750   0.5590   0.01135   0.00624  -0.0648   0.8748   1.0000
   3.000   0.5849   0.01103   0.00602  -0.0638   0.8615   1.0000
   3.250   0.6083   0.01038   0.00541  -0.0615   0.8347   1.0000
   3.500   0.6291   0.00991   0.00493  -0.0587   0.7936   1.0000
   3.750   0.6502   0.00973   0.00465  -0.0563   0.7354   1.0000
   4.000   0.6687   0.00996   0.00459  -0.0536   0.6432   1.0000
   4.250   0.6687   0.01127   0.00480  -0.0476   0.4367   1.0000
   4.500   0.6526   0.01436   0.00590  -0.0402   0.0839   1.0000
   4.750   0.6642   0.01562   0.00698  -0.0371   0.0459   1.0000
   5.000   0.6761   0.01684   0.00826  -0.0339   0.0378   1.0000
   5.250   0.6917   0.01784   0.00935  -0.0314   0.0333   1.0000
   5.500   0.7066   0.01916   0.01065  -0.0291   0.0280   1.0000
   5.750   0.7269   0.02182   0.01334  -0.0275   0.0262   1.0000
   6.000   0.7532   0.02373   0.01541  -0.0267   0.0257   1.0000
   6.250   0.7805   0.02645   0.01836  -0.0260   0.0258   1.0000
   6.500   0.8027   0.02773   0.01993  -0.0245   0.0241   1.0000
   6.750   0.8244   0.03065   0.02327  -0.0228   0.0243   1.0000
   7.000   0.8484   0.03300   0.02585  -0.0211   0.0318   1.0000
  12.000   0.5776   0.11679   0.11358  -0.0003   0.0437   1.0000
  12.250   0.5740   0.12037   0.11715  -0.0012   0.0424   1.0000
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