NACA 0020 Data

Discussion in 'Software' started by PI Design, May 14, 2010.

  1. PI Design
    Joined: Oct 2006
    Posts: 673
    Likes: 21, Points: 0, Legacy Rep: 328
    Location: England

    PI Design Senior Member

    Hi all,

    Not really a software discussion, but I'm after some help. Despite being a fairly common section, I am struggling to find lift, drag and pitching momet data for a NACA 0020 section. Can anyone point me to a reference?

    Whilst I'm here, does pitching moment coefficient change with aspect ratio and Reynolds number?

    Many thanks!
     
  2. daiquiri
    Joined: May 2004
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    Location: Italy (Garda Lake) and Croatia (Istria)

    daiquiri Engineering and Design

    Hello,
    It is true - there are virtually no available data in internet about NACA0020 and the Abbott&Doenhoff book doesn't treat that airfoil either.

    Depending on the purpose of your research and the precision required, I could suggest you to perform an analysis with XFLR5 ( http://xflr5.sourceforge.net/xflr5.htm ), which is a GUI version of XFOIL. When a correct transition parameter is used (n_crit = 6 appears to be a good choice), it will probably give fairly accurate polar curves up to angles of attack 3-4° short of stall. But be prudent if you need to analyse the stall region - it is generally off-limits for 2D methods (and sometimes for experimental methods too).

    You can use this little program to generate NACA airfoil coordinates to feed into XFOIL/XFLR5:
    http://tracfoil.free.fr/airfoils/downloads/nacalte.exe
    I have tested it on my PC, no viruses found.

    The Cm will not change significantly with Re number if you are operating in turbulent Re zone. If it is a laminar flow then, whilst I don't have data here to support my claim, I believe it very probably will change, because laminar flows are generally very instable.

    As about variations of Cm with AR, I suggest you to look at the page 17 of Abbott&Doenhoff's "Theory of Wing Sections", where you will find a formula for the calculation of Cm for finite wings as a function of AR, taper ratio, sweep angle and twist. I can anticipate that, if your rudder/keel/wing has no spanwise twist, you can safely assume that it's Cm will be about 1.03-1.04 times the Cm of the relative 2D airfoil. The multiplier will vary by a very little amount for different AR's, if the AR is bigger than 3.

    finally, you could download the following document from the NACA Technical Reports site:
    http://naca.central.cranfield.ac.uk/reports/1949/naca-tn-1945.pdf
    It will show you the variation of various coefficients for 15 NACA foils at 7 different Reynolds numbers.

    I hope that will help.

    Cheers!
     
  3. PI Design
    Joined: Oct 2006
    Posts: 673
    Likes: 21, Points: 0, Legacy Rep: 328
    Location: England

    PI Design Senior Member

    Thanks Daiquiri - that's really helpful. I feared I'd have to use x-foil to get some results, but was nervous as I have not used it before. The GUI interface should make it easier though. I'm not getting near stall angles, so at least that's something in my favour! Thanks for the heads up on the effect of AR in Abbot and Doenhoff, I was in too much of a hurry to see that. More haste, less speed!
    I have run tank tests at three speeds and the lowest of these speeds is in laminar flow, so that may be a problem.
    Thanks also for the other links - rep points coming your way...
    Off to learn/use x-foil now.
    Edit: Darn, have to spread the love before giving you more points!
     
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