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

GOE 143 (MVA H.20) AIRFOIL (goe143-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: GOE 143 (MVA H.20) AIRFOIL (goe143-il)
Reynolds number: 200,000
Max Cl/Cd: 72.29 at α=5.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe143-il-200000.txt
Download as CSV file: xf-goe143-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 143 (MVA H.20) 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.250  -0.4318   0.11272   0.10913  -0.0041   1.0000   0.0400
  -9.000  -0.4327   0.11053   0.10699  -0.0090   1.0000   0.0404
  -8.750  -0.4310   0.10767   0.10417  -0.0129   1.0000   0.0406
  -8.500  -0.4286   0.10446   0.10101  -0.0167   1.0000   0.0407
  -8.250  -0.4185   0.09790   0.09446  -0.0132   1.0000   0.0415
  -8.000  -0.4053   0.09441   0.09096  -0.0110   1.0000   0.0427
  -7.750  -0.3979   0.09142   0.08799  -0.0116   1.0000   0.0439
  -7.500  -0.3925   0.08842   0.08502  -0.0130   1.0000   0.0450
  -7.250  -0.3855   0.08518   0.08181  -0.0155   1.0000   0.0464
  -7.000  -0.3756   0.08163   0.07828  -0.0200   1.0000   0.0484
  -6.750  -0.3520   0.07748   0.07405  -0.0362   1.0000   0.0507
  -6.500  -0.3344   0.07336   0.06982  -0.0423   1.0000   0.0509
  -6.250  -0.3353   0.06763   0.06421  -0.0397   1.0000   0.0520
  -6.000  -0.3293   0.06529   0.06194  -0.0365   1.0000   0.0531
  -5.750  -0.3199   0.06278   0.05945  -0.0359   1.0000   0.0545
  -5.500  -0.3090   0.05998   0.05664  -0.0366   1.0000   0.0563
  -5.250  -0.2966   0.05699   0.05359  -0.0381   1.0000   0.0589
  -4.750  -0.2609   0.04877   0.04497  -0.0448   1.0000   0.0648
  -4.500  -0.2502   0.04681   0.04308  -0.0435   0.9994   0.0663
  -4.250  -0.2107   0.04358   0.03974  -0.0484   0.9952   0.0706
  -4.000  -0.1589   0.03858   0.03422  -0.0564   0.9905   0.0785
  -3.750  -0.1236   0.03613   0.03181  -0.0596   0.9858   0.0825
  -3.500  -0.0753   0.03283   0.02798  -0.0651   0.9811   0.0924
  -3.250  -0.0391   0.03037   0.02555  -0.0681   0.9751   0.0962
  -3.000   0.0062   0.02791   0.02274  -0.0725   0.9711   0.1075
  -2.750   0.0471   0.02621   0.02073  -0.0756   0.9645   0.1203
  -2.500   0.0861   0.02409   0.01864  -0.0785   0.9587   0.1270
  -2.250   0.1237   0.02240   0.01679  -0.0809   0.9508   0.1406
  -2.000   0.1702   0.01910   0.01250  -0.0819   0.9432   0.0931
  -1.750   0.2042   0.01736   0.01054  -0.0824   0.9319   0.0841
  -1.500   0.2367   0.01602   0.00892  -0.0825   0.9196   0.0814
  -1.250   0.2664   0.01510   0.00784  -0.0821   0.9054   0.0815
  -1.000   0.2939   0.01449   0.00716  -0.0814   0.8894   0.0844
  -0.750   0.3207   0.01395   0.00650  -0.0804   0.8723   0.0870
  -0.500   0.3469   0.01342   0.00585  -0.0793   0.8544   0.0886
  -0.250   0.3728   0.01309   0.00538  -0.0781   0.8359   0.0905
   0.000   0.3983   0.01242   0.00475  -0.0772   0.8149   0.0941
   0.250   0.4240   0.01217   0.00447  -0.0763   0.7932   0.1006
   0.500   0.4499   0.01188   0.00413  -0.0754   0.7696   0.1063
   0.750   0.4759   0.01159   0.00382  -0.0746   0.7449   0.1141
   1.000   0.5023   0.01134   0.00353  -0.0740   0.7165   0.1266
   1.250   0.5287   0.01095   0.00331  -0.0735   0.6866   0.1923
   1.500   0.5504   0.00919   0.00318  -0.0716   0.6562   1.0000
   1.750   0.5764   0.00943   0.00314  -0.0709   0.6231   1.0000
   2.000   0.6026   0.00970   0.00316  -0.0703   0.5914   1.0000
   2.250   0.6288   0.00999   0.00323  -0.0698   0.5640   1.0000
   2.500   0.6554   0.01026   0.00333  -0.0694   0.5393   1.0000
   2.750   0.6818   0.01056   0.00345  -0.0691   0.5175   1.0000
   3.000   0.7084   0.01082   0.00360  -0.0687   0.4958   1.0000
   3.250   0.7349   0.01110   0.00375  -0.0684   0.4768   1.0000
   3.500   0.7616   0.01138   0.00393  -0.0681   0.4607   1.0000
   3.750   0.7884   0.01165   0.00413  -0.0678   0.4458   1.0000
   4.000   0.8151   0.01191   0.00437  -0.0676   0.4315   1.0000
   4.250   0.8417   0.01218   0.00459  -0.0673   0.4174   1.0000
   4.500   0.8683   0.01244   0.00483  -0.0670   0.4031   1.0000
   4.750   0.8947   0.01270   0.00509  -0.0667   0.3884   1.0000
   5.000   0.9210   0.01296   0.00534  -0.0664   0.3729   1.0000
   5.250   0.9470   0.01323   0.00560  -0.0660   0.3559   1.0000
   5.500   0.9729   0.01350   0.00587  -0.0657   0.3361   1.0000
   5.750   0.9983   0.01381   0.00617  -0.0653   0.3106   1.0000
   6.000   1.0227   0.01424   0.00648  -0.0648   0.2711   1.0000
   6.250   1.0453   0.01496   0.00693  -0.0641   0.2232   1.0000
   6.500   1.0679   0.01573   0.00749  -0.0635   0.1929   1.0000
   6.750   1.0913   0.01638   0.00808  -0.0629   0.1753   1.0000
   7.000   1.1147   0.01701   0.00865  -0.0624   0.1624   1.0000
   7.250   1.1379   0.01764   0.00924  -0.0619   0.1524   1.0000
   7.500   1.1622   0.01810   0.00978  -0.0614   0.1437   1.0000
   7.750   1.1857   0.01863   0.01035  -0.0609   0.1339   1.0000
   8.000   1.2094   0.01910   0.01085  -0.0604   0.1212   1.0000
   8.250   1.2337   0.01951   0.01129  -0.0600   0.1050   1.0000
   8.500   1.2569   0.02012   0.01181  -0.0593   0.0768   1.0000
   8.750   1.2728   0.02167   0.01300  -0.0579   0.0413   1.0000
   9.000   1.2897   0.02303   0.01444  -0.0564   0.0369   1.0000
   9.250   1.3059   0.02436   0.01592  -0.0549   0.0344   1.0000
   9.500   1.3215   0.02565   0.01739  -0.0533   0.0325   1.0000
   9.750   1.3341   0.02714   0.01903  -0.0515   0.0307   1.0000
  10.000   1.3436   0.02881   0.02084  -0.0494   0.0296   1.0000
  10.250   1.3497   0.03069   0.02280  -0.0470   0.0287   1.0000
  10.500   1.3529   0.03276   0.02493  -0.0442   0.0280   1.0000
  10.750   1.3555   0.03506   0.02728  -0.0413   0.0275   1.0000
  11.000   1.3635   0.03712   0.02942  -0.0392   0.0271   1.0000
  11.250   1.3720   0.03879   0.03125  -0.0373   0.0266   1.0000
  11.500   1.3793   0.04060   0.03322  -0.0354   0.0260   1.0000
  11.750   1.3857   0.04262   0.03540  -0.0338   0.0253   1.0000
  12.000   1.3917   0.04486   0.03780  -0.0322   0.0248   1.0000
  12.250   1.3970   0.04732   0.04045  -0.0307   0.0245   1.0000
  12.500   1.3996   0.05004   0.04336  -0.0294   0.0243   1.0000
  12.750   1.3998   0.05302   0.04656  -0.0282   0.0242   1.0000
  13.000   1.3973   0.05632   0.05008  -0.0274   0.0242   1.0000
  13.250   1.3918   0.05999   0.05399  -0.0269   0.0242   1.0000
  13.500   1.3836   0.06407   0.05831  -0.0269   0.0243   1.0000
  13.750   1.3726   0.06862   0.06312  -0.0275   0.0244   1.0000
  14.000   1.3593   0.07369   0.06843  -0.0289   0.0246   1.0000
  14.250   1.3440   0.07929   0.07426  -0.0309   0.0248   1.0000
  14.500   1.3269   0.08546   0.08066  -0.0336   0.0250   1.0000
  14.750   1.3082   0.09214   0.08756  -0.0369   0.0252   1.0000
  15.000   1.2884   0.09939   0.09502  -0.0409   0.0254   1.0000
  15.250   1.2675   0.10729   0.10311  -0.0456   0.0256   1.0000
  15.500   1.2461   0.11578   0.11178  -0.0509   0.0259   1.0000
  15.750   1.2237   0.12503   0.12120  -0.0568   0.0261   1.0000
<< Back to GOE 143 (MVA H.20) AIRFOIL (goe143-il)

Polar data table (+)

Polar graphs


<< Back to GOE 143 (MVA H.20) AIRFOIL (goe143-il)