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EPPLER 1200 AIRFOIL (e1200-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: EPPLER 1200 AIRFOIL (e1200-il)
Reynolds number: 50,000
Max Cl/Cd: 9.29 at α=8.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e1200-il-50000.txt
Download as CSV file: xf-e1200-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 1200 AIRFOIL                             
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.250  -0.3614   0.11633   0.11122  -0.0032   1.0000   0.3615
  -7.000  -0.6856   0.08002   0.07414  -0.0299   1.0000   0.1515
  -6.750  -0.7087   0.07172   0.06511  -0.0320   1.0000   0.1376
  -6.500  -0.7025   0.06773   0.06099  -0.0314   1.0000   0.1360
  -6.250  -0.6958   0.06362   0.05661  -0.0312   1.0000   0.1338
  -6.000  -0.6867   0.05933   0.05189  -0.0314   1.0000   0.1308
  -5.750  -0.6735   0.05503   0.04697  -0.0317   1.0000   0.1276
  -5.500  -0.6563   0.05152   0.04282  -0.0319   1.0000   0.1259
  -5.250  -0.6373   0.04887   0.03976  -0.0316   1.0000   0.1262
  -5.000  -0.6181   0.04691   0.03757  -0.0312   1.0000   0.1287
  -4.750  -0.5979   0.04516   0.03553  -0.0308   1.0000   0.1330
  -4.500  -0.5760   0.04350   0.03342  -0.0303   1.0000   0.1368
  -4.250  -0.5546   0.04182   0.03156  -0.0296   1.0000   0.1408
  -4.000  -0.5341   0.04063   0.03035  -0.0287   1.0000   0.1474
  -3.750  -0.5132   0.03956   0.02917  -0.0278   1.0000   0.1575
  -3.500  -0.4927   0.03867   0.02827  -0.0265   1.0000   0.1716
  -3.250  -0.4736   0.03773   0.02758  -0.0250   1.0000   0.1928
  -3.000  -0.4541   0.03656   0.02675  -0.0237   1.0000   0.2378
  -2.750  -0.4359   0.03371   0.02605  -0.0228   1.0000   0.4345
  -2.500  -0.4473   0.03539   0.02884  -0.0105   1.0000   0.6742
  -2.250  -0.4538   0.03690   0.03031  -0.0004   1.0000   0.7399
  -2.000  -0.4584   0.03780   0.03112   0.0085   1.0000   0.7863
  -1.750  -0.4624   0.03829   0.03152   0.0170   1.0000   0.8249
  -1.500  -0.4627   0.03857   0.03167   0.0242   1.0000   0.8640
  -1.250  -0.4485   0.03915   0.03207   0.0290   1.0000   0.9059
  -1.000  -0.2075   0.04658   0.03845  -0.0076   1.0000   0.9854
  -0.750  -0.1521   0.04767   0.03924  -0.0158   1.0000   1.0000
  -0.500  -0.1509   0.04733   0.03880  -0.0138   1.0000   1.0000
  -0.250  -0.1498   0.04699   0.03837  -0.0118   1.0000   1.0000
   0.000  -0.1488   0.04667   0.03796  -0.0097   1.0000   1.0000
   0.250  -0.1479   0.04635   0.03756  -0.0076   1.0000   1.0000
   0.500  -0.1473   0.04602   0.03716  -0.0054   1.0000   1.0000
   0.750  -0.1468   0.04568   0.03675  -0.0032   1.0000   1.0000
   1.000  -0.1463   0.04534   0.03634  -0.0011   1.0000   1.0000
   1.250  -0.1457   0.04499   0.03594   0.0010   1.0000   1.0000
   1.500  -0.1445   0.04468   0.03558   0.0030   1.0000   1.0000
   1.750  -0.1415   0.04449   0.03532   0.0046   1.0000   1.0000
   2.000  -0.1342   0.04458   0.03534   0.0054   1.0000   1.0000
   2.250  -0.1226   0.04499   0.03568   0.0054   1.0000   1.0000
   2.500  -0.0905   0.04706   0.03764   0.0015   0.9936   1.0000
   2.750  -0.0516   0.04970   0.04015  -0.0037   0.9824   1.0000
   3.000  -0.0142   0.05220   0.04255  -0.0084   0.9692   1.0000
   3.250   0.0168   0.05401   0.04430  -0.0120   0.9543   1.0000
   3.500   0.0464   0.05569   0.04593  -0.0152   0.9367   1.0000
   3.750   0.0773   0.05762   0.04780  -0.0184   0.9187   1.0000
   4.000   0.1094   0.05968   0.04982  -0.0218   0.8996   1.0000
   4.250   0.1438   0.06196   0.05205  -0.0254   0.8800   1.0000
   4.500   0.1854   0.06492   0.05496  -0.0299   0.8599   1.0000
   4.750   0.2063   0.06629   0.05632  -0.0312   0.8393   1.0000
   5.000   0.2533   0.06868   0.05867  -0.0353   0.8081   1.0000
   5.250   0.3403   0.06647   0.05633  -0.0381   0.7161   1.0000
   5.500   0.3879   0.06731   0.05715  -0.0408   0.6950   1.0000
   5.750   0.4165   0.06815   0.05798  -0.0418   0.6753   1.0000
   6.000   0.4389   0.06911   0.05896  -0.0423   0.6566   1.0000
   6.250   0.4658   0.07005   0.05991  -0.0431   0.6389   1.0000
   6.500   0.4944   0.07088   0.06076  -0.0440   0.6217   1.0000
   6.750   0.5215   0.07181   0.06172  -0.0447   0.6057   1.0000
   7.000   0.5487   0.07264   0.06259  -0.0454   0.5897   1.0000
   7.250   0.5741   0.07356   0.06354  -0.0458   0.5742   1.0000
   7.500   0.5981   0.07454   0.06456  -0.0462   0.5591   1.0000
   7.750   0.6214   0.07555   0.06561  -0.0464   0.5444   1.0000
   8.000   0.6452   0.07655   0.06667  -0.0467   0.5306   1.0000
   8.250   0.6764   0.07690   0.06706  -0.0469   0.5177   1.0000
   8.500   0.7125   0.07670   0.06694  -0.0472   0.5056   1.0000
   8.750   0.7182   0.07919   0.06948  -0.0469   0.4912   1.0000
   9.000   0.7225   0.08198   0.07232  -0.0468   0.4775   1.0000
   9.250   0.7300   0.08468   0.07506  -0.0469   0.4652   1.0000
   9.500   0.7704   0.08393   0.07441  -0.0467   0.4551   1.0000
   9.750   0.7779   0.08659   0.07713  -0.0467   0.4428   1.0000
  10.000   0.7592   0.09251   0.08308  -0.0473   0.4313   1.0000
  10.250   0.7827   0.09370   0.08434  -0.0472   0.4215   1.0000
  10.500   0.7871   0.09708   0.08778  -0.0476   0.4112   1.0000
  10.750   0.7728   0.10297   0.09373  -0.0487   0.4029   1.0000
  11.000   0.8104   0.10252   0.09338  -0.0481   0.3932   1.0000
  11.250   0.7739   0.11131   0.10218  -0.0504   0.3869   1.0000
  11.500   0.7809   0.11484   0.10578  -0.0511   0.3794   1.0000
  11.750   0.7857   0.11894   0.10994  -0.0521   0.3736   1.0000
  12.000   0.7666   0.12577   0.11681  -0.0544   0.3712   1.0000
  12.250   0.7575   0.13149   0.12259  -0.0564   0.3699   1.0000
  12.500   0.7575   0.13703   0.12819  -0.0585   0.3731   1.0000
  12.750   0.7706   0.14207   0.13333  -0.0604   0.3763   1.0000
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