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=10% T=5% R=1.3 (cp-100-050-gn) Xfoil prediction polar at RE=1,000,000 Ncrit=9


Details Polar file
Airfoil: Cambered plate C=10% T=5% R=1.3 (cp-100-050-gn)
Reynolds number: 1,000,000
Max Cl/Cd: 48.24 at α=6.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-cp-100-050-gn-1000000.txt
Download as CSV file: xf-cp-100-050-gn-1000000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: Cambered plate C=10% T=5% R=1.3                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     1.000 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.500  -0.0086   0.11347   0.11129  -0.1329   0.9720   0.0283
 -10.250  -0.0060   0.11122   0.10904  -0.1329   0.9677   0.0284
 -10.000   0.0056   0.10810   0.10592  -0.1345   0.9660   0.0285
  -9.750   0.0207   0.10563   0.10344  -0.1364   0.9647   0.0286
  -9.500   0.0363   0.10314   0.10095  -0.1388   0.9636   0.0288
  -9.250   0.0529   0.10074   0.09855  -0.1413   0.9625   0.0291
  -9.000   0.0699   0.09826   0.09606  -0.1441   0.9614   0.0295
  -8.750   0.0879   0.09564   0.09344  -0.1473   0.9603   0.0303
  -8.500   0.1060   0.09213   0.08990  -0.1521   0.9592   0.0314
  -8.250   0.1125   0.08954   0.08730  -0.1543   0.9545   0.0315
  -8.000   0.1268   0.08658   0.08434  -0.1565   0.9516   0.0316
  -7.750   0.1456   0.08416   0.08191  -0.1586   0.9495   0.0318
  -7.500   0.1680   0.08176   0.07950  -0.1621   0.9480   0.0320
  -7.250   0.1922   0.07925   0.07697  -0.1664   0.9465   0.0323
  -7.000   0.2180   0.07676   0.07446  -0.1711   0.9447   0.0329
  -6.750   0.2256   0.07509   0.07280  -0.1709   0.9387   0.0335
  -6.500   0.2421   0.07240   0.07008  -0.1745   0.9339   0.0347
  -6.250   0.2594   0.06924   0.06686  -0.1817   0.9280   0.0349
  -6.000   0.2671   0.06726   0.06489  -0.1801   0.9227   0.0350
  -5.750   0.2821   0.06539   0.06302  -0.1805   0.9185   0.0351
  -5.500   0.3019   0.06348   0.06108  -0.1828   0.9146   0.0353
  -5.250   0.3084   0.06226   0.05987  -0.1817   0.9087   0.0356
  -5.000   0.3160   0.06107   0.05867  -0.1810   0.9026   0.0361
  -4.750   0.3326   0.05927   0.05682  -0.1830   0.8957   0.0366
  -4.500   0.3401   0.05775   0.05532  -0.1828   0.8912   0.0375
  -4.250   0.3516   0.05497   0.05246  -0.1893   0.8832   0.0383
  -4.000   0.3664   0.05272   0.05019  -0.1893   0.8784   0.0385
  -3.750   0.3761   0.05128   0.04877  -0.1878   0.8720   0.0386
  -3.500   0.3932   0.04973   0.04717  -0.1883   0.8645   0.0387
  -3.250   0.4075   0.04837   0.04580  -0.1883   0.8588   0.0389
  -3.000   0.4231   0.04699   0.04441  -0.1887   0.8529   0.0393
  -2.750   0.4439   0.04538   0.04275  -0.1903   0.8460   0.0398
  -2.500   0.4596   0.04388   0.04125  -0.1907   0.8396   0.0405
  -2.250   0.4981   0.04029   0.03750  -0.1999   0.8327   0.0420
  -2.000   0.5138   0.03842   0.03560  -0.1994   0.8268   0.0421
  -1.750   0.5297   0.03711   0.03430  -0.1985   0.8195   0.0422
  -1.500   0.5495   0.03585   0.03298  -0.1985   0.8115   0.0424
  -1.250   0.5686   0.03467   0.03179  -0.1983   0.8034   0.0427
  -1.000   0.5893   0.03348   0.03053  -0.1984   0.7932   0.0431
  -0.750   0.6092   0.03226   0.02926  -0.1983   0.7805   0.0436
  -0.500   0.6315   0.03089   0.02781  -0.1986   0.7677   0.0448
   0.000   0.6800   0.02659   0.02316  -0.2003   0.7276   0.0461
   0.250   0.6778   0.02631   0.02263  -0.1944   0.6765   0.0463
   0.500   0.6603   0.02640   0.02238  -0.1850   0.6073   0.0463
   0.750   0.6393   0.02698   0.02245  -0.1752   0.4991   0.0464
   1.000   0.6078   0.02851   0.02310  -0.1638   0.2988   0.0464
   1.250   0.5894   0.02978   0.02357  -0.1554   0.0519   0.0464
   1.500   0.6100   0.02907   0.02280  -0.1547   0.0438   0.0467
   1.750   0.6325   0.02830   0.02199  -0.1542   0.0416   0.0472
   2.000   0.6563   0.02747   0.02109  -0.1540   0.0401   0.0481
   2.250   0.6961   0.02457   0.01791  -0.1565   0.0389   0.0505
   2.500   0.7168   0.02417   0.01749  -0.1555   0.0384   0.0507
   2.750   0.7383   0.02377   0.01705  -0.1546   0.0376   0.0510
   3.000   0.7603   0.02337   0.01661  -0.1537   0.0370   0.0515
   3.250   0.7830   0.02290   0.01608  -0.1528   0.0367   0.0525
   3.500   0.8140   0.02100   0.01387  -0.1529   0.0364   0.0555
   3.750   0.8334   0.02096   0.01381  -0.1514   0.0359   0.0559
   4.000   0.8535   0.02080   0.01362  -0.1500   0.0356   0.0564
   4.250   0.8737   0.02062   0.01337  -0.1485   0.0353   0.0575
   4.500   0.8970   0.01966   0.01211  -0.1470   0.0351   0.0607
   4.750   0.9145   0.01974   0.01219  -0.1451   0.0349   0.0612
   5.000   0.9317   0.01987   0.01230  -0.1431   0.0345   0.0617
   5.250   0.9487   0.02003   0.01244  -0.1410   0.0344   0.0628
   5.500   0.9654   0.02005   0.01227  -0.1386   0.0341   0.0667
   5.750   0.9775   0.02047   0.01272  -0.1358   0.0338   0.0673
   6.000   0.9898   0.02098   0.01323  -0.1330   0.0336   0.0679
   6.250   1.0070   0.02123   0.01348  -0.1310   0.0335   0.0695
   6.500   1.0263   0.02133   0.01345  -0.1292   0.0332   0.0734
   6.750   1.0424   0.02161   0.01375  -0.1272   0.0329   0.0743
   7.000   1.0578   0.02203   0.01419  -0.1249   0.0328   0.0755
   7.250   1.0733   0.02250   0.01464  -0.1227   0.0326   0.0774
   7.500   1.0887   0.02293   0.01505  -0.1206   0.0323   0.0820
   7.750   1.1039   0.02345   0.01559  -0.1184   0.0321   0.0841
   8.000   1.1197   0.02402   0.01615  -0.1163   0.0319   0.0864
   8.250   1.1372   0.02386   0.01576  -0.1137   0.0318   0.0603
   8.500   1.1529   0.02444   0.01636  -0.1117   0.0316   0.0606
   8.750   1.1693   0.02504   0.01699  -0.1098   0.0314   0.0609
   9.000   1.1866   0.02567   0.01765  -0.1080   0.0313   0.0617
   9.250   1.2045   0.02625   0.01825  -0.1064   0.0310   0.0622
   9.500   1.2235   0.02687   0.01889  -0.1049   0.0309   0.0627
   9.750   1.2429   0.02746   0.01950  -0.1035   0.0306   0.0630
  10.000   1.2631   0.02807   0.02011  -0.1022   0.0304   0.0642
  10.250   1.2812   0.02865   0.02070  -0.1007   0.0300   0.0648
  10.500   1.3025   0.02929   0.02134  -0.0997   0.0298   0.0657
  10.750   1.3279   0.02996   0.02200  -0.0992   0.0294   0.0677
  11.000   1.4070   0.03130   0.02331  -0.1066   0.0289   0.0741
  11.250   1.4144   0.03182   0.02399  -0.1033   0.0288   0.1164
  11.500   1.4561   0.03206   0.02553  -0.1071   0.0285   1.0000
  11.750   1.4767   0.03290   0.02642  -0.1058   0.0284   1.0000
  12.000   1.4998   0.03387   0.02746  -0.1050   0.0282   1.0000
  12.250   1.5250   0.03502   0.02868  -0.1045   0.0281   1.0000
  12.500   1.5503   0.03635   0.03010  -0.1040   0.0280   1.0000
  12.750   1.5690   0.03760   0.03146  -0.1026   0.0276   1.0000
  13.000   1.5865   0.03892   0.03288  -0.1011   0.0273   1.0000
  13.250   1.6055   0.04063   0.03472  -0.0998   0.0271   1.0000
  13.500   1.6215   0.04257   0.03682  -0.0982   0.0270   1.0000
  13.750   1.6347   0.04475   0.03917  -0.0963   0.0268   1.0000
  14.000   1.6447   0.04830   0.04300  -0.0940   0.0270   1.0000
  14.250   1.6447   0.05257   0.04758  -0.0908   0.0277   1.0000
  14.500   1.6422   0.05592   0.05111  -0.0874   0.0281   1.0000
  14.750   1.6036   0.05442   0.04973  -0.0814   0.0293   1.0000
  15.750   1.5598   0.07352   0.06978  -0.0671   0.0277   1.0000
  16.000   1.5389   0.07756   0.07403  -0.0641   0.0272   1.0000
  16.250   1.5158   0.08211   0.07876  -0.0617   0.0271   1.0000
  16.500   1.4944   0.08651   0.08332  -0.0600   0.0269   1.0000
  16.750   1.4710   0.09140   0.08838  -0.0589   0.0267   1.0000
  17.000   1.4458   0.09690   0.09405  -0.0585   0.0266   1.0000
  17.250   1.4189   0.10305   0.10036  -0.0590   0.0265   1.0000
  17.500   1.3902   0.10990   0.10737  -0.0605   0.0264   1.0000
  17.750   1.3582   0.11791   0.11554  -0.0634   0.0263   1.0000
  18.000   1.3143   0.12920   0.12703  -0.0691   0.0264   1.0000
<< Back to Cambered plate C=10% T=5% R=1.3 (cp-100-050-gn)

Polar data table (+)

Polar graphs


<< Back to Cambered plate C=10% T=5% R=1.3 (cp-100-050-gn)