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

MARSKE XM-1D AIRFOIL (F14-3.0) (marske1-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: MARSKE XM-1D AIRFOIL (F14-3.0) (marske1-il)
Reynolds number: 50,000
Max Cl/Cd: 17.82 at α=4.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-marske1-il-50000-n5.txt
Download as CSV file: xf-marske1-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: MARSKE XM-1D AIRFOIL (F14-3.0)                  
  
 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
  ------ -------- --------- --------- -------- -------- --------
 -10.000  -0.5005   0.10282   0.09660   0.0017   1.0000   0.0921
  -9.750  -0.4834   0.09961   0.09344   0.0015   1.0000   0.0898
  -9.500  -0.5064   0.09116   0.08492  -0.0067   0.8528   0.0840
  -9.250  -0.5019   0.08709   0.08066  -0.0088   0.8133   0.0825
  -9.000  -0.5125   0.08197   0.07538  -0.0116   0.7946   0.0806
  -8.750  -0.5561   0.07470   0.06781  -0.0141   0.7863   0.0774
  -8.500  -0.5568   0.07141   0.06436  -0.0137   0.7722   0.0766
  -8.250  -0.5596   0.06781   0.06055  -0.0133   0.7604   0.0757
  -7.750  -0.5629   0.06056   0.05272  -0.0116   0.7404   0.0744
  -7.500  -0.5599   0.05736   0.04916  -0.0103   0.7313   0.0743
  -7.250  -0.5529   0.05433   0.04579  -0.0091   0.7214   0.0744
  -7.000  -0.5437   0.05158   0.04265  -0.0076   0.7126   0.0744
  -6.750  -0.5311   0.04892   0.03962  -0.0064   0.7032   0.0743
  -6.500  -0.5169   0.04644   0.03671  -0.0049   0.6944   0.0739
  -6.250  -0.5000   0.04412   0.03399  -0.0037   0.6853   0.0735
  -6.000  -0.4811   0.04208   0.03155  -0.0024   0.6762   0.0733
  -5.750  -0.4597   0.04027   0.02939  -0.0014   0.6676   0.0732
  -5.500  -0.4368   0.03869   0.02752  -0.0006   0.6583   0.0736
  -5.250  -0.4131   0.03735   0.02593   0.0002   0.6497   0.0746
  -5.000  -0.3878   0.03610   0.02443   0.0008   0.6402   0.0759
  -4.750  -0.3616   0.03494   0.02298   0.0015   0.6322   0.0769
  -4.500  -0.3333   0.03384   0.02168   0.0017   0.6227   0.0776
  -4.250  -0.3051   0.03290   0.02064   0.0021   0.6149   0.0782
  -4.000  -0.2750   0.03205   0.01979   0.0019   0.6052   0.0791
  -3.750  -0.2458   0.03135   0.01896   0.0023   0.5982   0.0803
  -3.500  -0.2151   0.03077   0.01836   0.0020   0.5882   0.0826
  -3.250  -0.1862   0.03026   0.01772   0.0024   0.5814   0.0851
  -3.000  -0.1574   0.02983   0.01731   0.0022   0.5720   0.0877
  -2.750  -0.1309   0.02942   0.01678   0.0029   0.5653   0.0901
  -2.500  -0.1045   0.02911   0.01644   0.0031   0.5572   0.0926
  -2.250  -0.0794   0.02877   0.01606   0.0038   0.5499   0.0963
  -2.000  -0.0549   0.02849   0.01571   0.0046   0.5437   0.1021
  -1.750  -0.0305   0.02828   0.01554   0.0050   0.5356   0.1104
  -1.500  -0.0079   0.02789   0.01518   0.0062   0.5302   0.1228
  -1.250   0.0122   0.02743   0.01515   0.0071   0.5235   0.1610
  -0.750   0.1019   0.02603   0.01638   0.0058   0.5105   0.8772
  -0.500   0.1498   0.02683   0.01706   0.0028   0.5010   0.9133
  -0.250   0.2125   0.02718   0.01712  -0.0023   0.4945   0.9497
   0.000   0.2803   0.02754   0.01730  -0.0097   0.4857   0.9796
   0.250   0.3420   0.02743   0.01698  -0.0163   0.4785   1.0000
   0.500   0.3623   0.02746   0.01683  -0.0150   0.4748   1.0000
   0.750   0.3839   0.02804   0.01742  -0.0150   0.4679   1.0000
   1.000   0.4044   0.02838   0.01769  -0.0142   0.4623   1.0000
   1.250   0.4245   0.02852   0.01770  -0.0130   0.4583   1.0000
   1.500   0.4446   0.02878   0.01785  -0.0119   0.4543   1.0000
   1.750   0.4645   0.02970   0.01884  -0.0119   0.4474   1.0000
   2.000   0.4840   0.03012   0.01920  -0.0109   0.4426   1.0000
   2.250   0.5037   0.03031   0.01927  -0.0095   0.4390   1.0000
   2.500   0.5227   0.03088   0.01980  -0.0085   0.4345   1.0000
   2.750   0.5398   0.03209   0.02111  -0.0083   0.4280   1.0000
   3.000   0.5580   0.03271   0.02170  -0.0073   0.4236   1.0000
   3.250   0.5771   0.03299   0.02190  -0.0058   0.4203   1.0000
   3.500   0.5956   0.03347   0.02231  -0.0044   0.4168   1.0000
   3.750   0.6066   0.03549   0.02451  -0.0044   0.4091   1.0000
   4.000   0.6223   0.03637   0.02540  -0.0033   0.4048   1.0000
   4.250   0.6402   0.03686   0.02583  -0.0018   0.4017   1.0000
   4.500   0.6602   0.03705   0.02593  -0.0001   0.3992   1.0000
   4.750   0.6565   0.04064   0.02976  -0.0002   0.3901   1.0000
   5.000   0.6690   0.04173   0.03086   0.0011   0.3857   1.0000
   5.250   0.6859   0.04228   0.03138   0.0027   0.3826   1.0000
   5.500   0.7060   0.04254   0.03158   0.0044   0.3804   1.0000
   5.750   0.6713   0.04895   0.03821   0.0040   0.3690   1.0000
   6.000   0.6843   0.04985   0.03909   0.0055   0.3653   1.0000
   6.250   0.7046   0.05002   0.03923   0.0073   0.3629   1.0000
   6.750   0.6527   0.05995   0.04919   0.0075   0.3468   1.0000
   7.000   0.6746   0.06015   0.04938   0.0091   0.3443   1.0000
   7.500   0.6305   0.06852   0.05769   0.0100   0.3299   1.0000
   7.750   0.6510   0.06898   0.05815   0.0113   0.3266   1.0000
   8.250   0.6338   0.07567   0.06483   0.0112   0.3127   1.0000
   8.500   0.6564   0.07593   0.06510   0.0125   0.3090   1.0000
   8.750   0.6410   0.08016   0.06933   0.0117   0.3008   1.0000
   9.000   0.6490   0.08194   0.07113   0.0121   0.2947   1.0000
   9.250   0.6735   0.08197   0.07117   0.0134   0.2912   1.0000
   9.500   0.6528   0.08706   0.07628   0.0119   0.2818   1.0000
   9.750   0.6676   0.08819   0.07743   0.0126   0.2767   1.0000
  10.000   0.6941   0.08793   0.07721   0.0140   0.2736   1.0000
  10.250   0.6675   0.09407   0.08338   0.0118   0.2632   1.0000
  10.500   0.6875   0.09460   0.08393   0.0127   0.2590   1.0000
  10.750   0.6755   0.09919   0.08855   0.0112   0.2508   1.0000
  11.000   0.6844   0.10121   0.09060   0.0113   0.2452   1.0000
  11.250   0.7071   0.10140   0.09084   0.0124   0.2419   1.0000
  11.500   0.6855   0.10754   0.09701   0.0099   0.2323   1.0000
  11.750   0.7017   0.10860   0.09811   0.0104   0.2280   1.0000
<< Back to MARSKE XM-1D AIRFOIL (F14-3.0) (marske1-il)

Polar data table (+)

Polar graphs


<< Back to MARSKE XM-1D AIRFOIL (F14-3.0) (marske1-il)