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

Cambered plate C=12% T=5% R=1.1 (cp-120-050-gn) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: Cambered plate C=12% T=5% R=1.1 (cp-120-050-gn)
Reynolds number: 500,000
Max Cl/Cd: 42.84 at α=2°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-cp-120-050-gn-500000.txt
Download as CSV file: xf-cp-120-050-gn-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: Cambered plate C=12% T=5% R=1.1                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.750   0.1490   0.10479   0.10175  -0.1642   0.9405   0.0288
  -9.500   0.1647   0.10230   0.09925  -0.1667   0.9390   0.0291
  -9.250   0.1821   0.09975   0.09669  -0.1696   0.9378   0.0298
  -9.000   0.2008   0.09696   0.09389  -0.1731   0.9368   0.0305
  -8.750   0.2217   0.09396   0.09087  -0.1776   0.9359   0.0313
  -8.500   0.2444   0.09115   0.08803  -0.1833   0.9350   0.0315
  -8.250   0.2662   0.08850   0.08535  -0.1889   0.9339   0.0316
  -8.000   0.2880   0.08493   0.08177  -0.1915   0.9332   0.0318
  -7.750   0.2825   0.08416   0.08103  -0.1873   0.9265   0.0319
  -7.500   0.2982   0.08206   0.07892  -0.1891   0.9238   0.0322
  -7.250   0.3184   0.07972   0.07657  -0.1922   0.9217   0.0328
  -7.000   0.3401   0.07725   0.07408  -0.1960   0.9200   0.0334
  -6.750   0.3627   0.07470   0.07149  -0.2001   0.9181   0.0344
  -6.500   0.3563   0.07408   0.07090  -0.1961   0.9111   0.0346
  -6.250   0.3662   0.07259   0.06939  -0.1979   0.9063   0.0348
  -6.000   0.3813   0.07057   0.06734  -0.2013   0.9029   0.0349
  -5.750   0.4043   0.06754   0.06429  -0.2034   0.9015   0.0351
  -5.500   0.3896   0.06730   0.06411  -0.1966   0.8939   0.0352
  -5.250   0.4040   0.06556   0.06236  -0.1976   0.8905   0.0355
  -5.000   0.4245   0.06357   0.06034  -0.2004   0.8879   0.0360
  -4.750   0.4479   0.06151   0.05825  -0.2040   0.8854   0.0366
  -4.500   0.4369   0.06113   0.05792  -0.1985   0.8782   0.0369
  -4.250   0.4536   0.05935   0.05612  -0.2004   0.8745   0.0377
  -4.000   0.4823   0.05746   0.05418  -0.2067   0.8713   0.0382
  -3.750   0.4985   0.05632   0.05301  -0.2104   0.8661   0.0384
  -3.500   0.5023   0.05445   0.05116  -0.2071   0.8615   0.0385
  -3.250   0.5184   0.05239   0.04910  -0.2071   0.8585   0.0386
  -3.000   0.5423   0.05044   0.04713  -0.2094   0.8559   0.0389
  -2.750   0.5706   0.04854   0.04520  -0.2130   0.8537   0.0394
  -2.500   0.5756   0.04761   0.04429  -0.2108   0.8484   0.0398
  -2.250   0.5927   0.04618   0.04286  -0.2115   0.8441   0.0404
  -2.000   0.6254   0.04430   0.04094  -0.2162   0.8411   0.0417
  -1.750   0.6813   0.04183   0.03834  -0.2278   0.8387   0.0422
  -1.500   0.7047   0.03974   0.03623  -0.2287   0.8364   0.0424
  -1.250   0.7050   0.03888   0.03543  -0.2244   0.8304   0.0426
  -1.000   0.7299   0.03727   0.03379  -0.2258   0.8252   0.0430
  -0.750   0.7689   0.03542   0.03187  -0.2305   0.8209   0.0438
  -0.500   0.7818   0.03432   0.03076  -0.2289   0.8134   0.0445
  -0.250   0.8437   0.03204   0.02828  -0.2392   0.8072   0.0465
   0.000   0.8608   0.03051   0.02673  -0.2381   0.8002   0.0467
   0.250   0.8705   0.02945   0.02569  -0.2353   0.7926   0.0469
   0.500   0.9011   0.02814   0.02430  -0.2372   0.7855   0.0475
   0.750   0.9049   0.02737   0.02356  -0.2328   0.7769   0.0480
   1.000   0.9322   0.02620   0.02229  -0.2336   0.7676   0.0492
   1.250   0.9687   0.02455   0.02047  -0.2361   0.7576   0.0514
   1.500   0.9764   0.02371   0.01961  -0.2325   0.7444   0.0516
   1.750   0.9801   0.02314   0.01895  -0.2278   0.7231   0.0519
   2.000   0.9793   0.02286   0.01851  -0.2222   0.6902   0.0522
   2.250   0.9702   0.02306   0.01843  -0.2148   0.6394   0.0525
   2.500   0.9531   0.02386   0.01886  -0.2061   0.5724   0.0528
   2.750   0.9256   0.02568   0.02011  -0.1960   0.4689   0.0528
   3.000   0.8977   0.02802   0.02172  -0.1865   0.3223   0.0529
   3.250   0.8706   0.03070   0.02357  -0.1778   0.1071   0.0530
   3.500   0.8836   0.03101   0.02364  -0.1759   0.0469   0.0541
   3.750   0.9229   0.02990   0.02231  -0.1782   0.0443   0.0570
   4.000   0.9398   0.02976   0.02219  -0.1769   0.0422   0.0575
   4.250   0.9579   0.02972   0.02215  -0.1757   0.0408   0.0582
   4.500   0.9763   0.02975   0.02215  -0.1745   0.0398   0.0596
   4.750   1.0065   0.02906   0.02122  -0.1748   0.0394   0.0634
   5.000   1.0231   0.02917   0.02135  -0.1733   0.0386   0.0639
   5.250   1.0398   0.02937   0.02155  -0.1718   0.0381   0.0649
   5.500   1.0572   0.02959   0.02174  -0.1703   0.0375   0.0668
   5.750   1.0784   0.02953   0.02151  -0.1691   0.0372   0.0712
   6.000   1.0926   0.03000   0.02200  -0.1672   0.0369   0.0723
   6.250   1.1072   0.03051   0.02249  -0.1653   0.0365   0.0746
   6.500   1.1244   0.03086   0.02269  -0.1636   0.0363   0.0799
   6.750   1.1369   0.03153   0.02339  -0.1615   0.0360   0.0813
   7.000   1.1497   0.03229   0.02415  -0.1593   0.0357   0.0841
   8.500   1.2303   0.03696   0.02826  -0.1467   0.0344   0.0711
   8.750   1.2423   0.03823   0.02949  -0.1445   0.0340   0.0700
   9.000   1.2589   0.03909   0.03030  -0.1429   0.0336   0.0692
   9.250   1.2806   0.03972   0.03089  -0.1418   0.0335   0.0690
   9.500   1.3073   0.04017   0.03131  -0.1412   0.0335   0.0695
   9.750   1.3402   0.04040   0.03150  -0.1414   0.0334   0.0695
  10.000   1.3752   0.04064   0.03172  -0.1419   0.0335   0.0698
  10.250   1.4052   0.04105   0.03215  -0.1418   0.0336   0.0705
  10.500   1.4355   0.04152   0.03267  -0.1419   0.0337   0.0725
  10.750   1.4519   0.04227   0.03352  -0.1401   0.0341   0.0748
  11.000   1.6809   0.04563   0.03826  -0.1673   0.0453   1.0000
  11.250   1.6806   0.04622   0.03900  -0.1625   0.0441   1.0000
  11.500   1.6961   0.04754   0.04041  -0.1605   0.0428   1.0000
  11.750   1.7730   0.05265   0.04571  -0.1715   0.0395   1.0000
  12.000   1.7636   0.05377   0.04704  -0.1652   0.0394   1.0000
  12.250   1.7432   0.05408   0.04754  -0.1570   0.0392   1.0000
  12.500   1.7143   0.05421   0.04789  -0.1478   0.0388   1.0000
  12.750   1.6846   0.05527   0.04922  -0.1393   0.0381   1.0000
  13.000   1.6720   0.05737   0.05155  -0.1339   0.0372   1.0000
  13.250   1.6668   0.06003   0.05438  -0.1299   0.0364   1.0000
  13.500   1.6624   0.06286   0.05738  -0.1264   0.0359   1.0000
  13.750   1.6578   0.06582   0.06047  -0.1230   0.0355   1.0000
  14.000   1.6532   0.06883   0.06362  -0.1199   0.0351   1.0000
  14.250   1.6481   0.07193   0.06684  -0.1170   0.0349   1.0000
  14.500   1.6410   0.07515   0.07019  -0.1141   0.0346   1.0000
  14.750   1.6310   0.07849   0.07367  -0.1112   0.0344   1.0000
  15.000   1.6278   0.08176   0.07703  -0.1091   0.0342   1.0000
  15.250   1.7749   0.08906   0.08389  -0.1226   0.0332   1.0000
  15.500   1.7477   0.09150   0.08655  -0.1173   0.0331   1.0000
  15.750   1.7194   0.09414   0.08941  -0.1125   0.0330   1.0000
  16.000   1.6818   0.09730   0.09285  -0.1078   0.0329   1.0000
  16.250   1.6374   0.10076   0.09666  -0.1035   0.0325   1.0000
<< Back to Cambered plate C=12% T=5% R=1.1 (cp-120-050-gn)

Polar data table (+)

Polar graphs


<< Back to Cambered plate C=12% T=5% R=1.1 (cp-120-050-gn)