Slope of lift curve
Ed Haering
>Can someone advise me on how to locate or calculate the slope of
>the lift curve for airfoils other than NACA 0012. Of paticular
>interest is NACA 8-H-12.
>
>Thanks;
>Dave
>
Check out "Theory of Wing Sections" by Ira H. Abbott and Albert E. Von
Doenhoff, Appendix IV. Among other things, it has CL-alpha0 plots, but
I didn't find NACA 8-H-12. When in doubt, lift curve slope equals
2*pi.
Ed Haering
NASA Dryden Flight Research Center
Edwards, CA
Stone Engineering
<75123...@CompuServe.COM> wrote:
> Can someone advise me on how to locate or calculate the slope of
> the lift curve for airfoils other than NACA 0012. Of paticular
> interest is NACA 8-H-12.
>
Depending upon how detailed you want to get, the theoretical lift curve
slope is two pi. Many of the NACA airfoils (and others, for that matter)
approach that slope (around 5.8-6.0). I don't have the lift curve, but
have you tried searching NASA's "library?" There is an option that allows
you to search only the NACA database. The URL is:
http://www.sti.nasa.gov/RECONselectv3.html
Joe
dave lawson
Stone Engineering wrote:
>>
> > Can someone advise me on how to locate or calculate the slope of
> > the lift curve for airfoils other than NACA 0012. Of paticular
> > interest is NACA 8-H-12.
> >
> Depending upon how detailed you want to get, the theoretical lift curve
> slope is two pi. Many of the NACA airfoils (and others, for that matter)
> approach that slope (around 5.8-6.0). I don't have the lift curve, but
> have you tried searching NASA's "library?" There is an option that allows
> you to search only the NACA database. The URL is:
> http://www.sti.nasa.gov/RECONselectv3.html
>
> Joe
Try looking at NACA report 460, "The Characteristics of 78 Related
Airfoil Sections from Tests in the Variable Density Wind Tunnel." by
Jacobs, Ward and Pinkerton. My reference, "Introduction to Aerodynamics"
by C.F. Toms shows data from this report and it may help you in you
quest. It has representative data points from airfoils between 5 and 25%
thickness (whick seems to be a main determinator of the slope).
The data at 20% reads to be about 0.096 and at 5% it reads 0.104 (per
degree). If you took these numbers and used a straight line
interpolation, I don't think you would be too far off the mark.
Hope this helps.
Dave
Ben Grocholsky
In article <rddDtp...@netcom.com>, Hae...@cs1.dfrc.nasa.gov says...
>
>dave jackson <75123...@CompuServe.COM> wrote:
>>Can someone advise me on how to locate or calculate the slope of
>>the lift curve for airfoils other than NACA 0012. Of paticular
>>interest is NACA 8-H-12.
>>
>>Dave
>>
>
>Check out "Theory of Wing Sections" by Ira H. Abbott and Albert E. Von
>Doenhoff, Appendix IV. Among other things, it has CL-alpha0 plots, but
>I didn't find NACA 8-H-12. When in doubt, lift curve slope equals
>2*pi.
>
>Ed Haering
>NASA Dryden Flight Research Center
>Edwards, CA
>
If you know the geometry, the methods in "Theoretical Aerodynamics" by
Milne-Thomson, L,M will give the inviscid result. These methods will give
results close to published high Reynolds number tests, as in "Theory of Wing
Sections". (for lift curve slope)
It may be obvious, but for low Reynolds number applications (such as model
aircraft) the inviscid solution does not apply. Boundary layer effects will
lower the lift curve slope and probably make it a function of angle of attack.
Search the public domain CFD sites for 2D aerofoil programs that include
viscous approximations.
Mark Drela
In article <rddDtp...@netcom.com>, st...@stoneeng.com (Stone Engineering) writes:
|> In article <rddDtC...@netcom.com>, dave jackson
|> <75123...@CompuServe.COM> wrote:
|>
|> > Can someone advise me on how to locate or calculate the slope of
|> > the lift curve for airfoils other than NACA 0012. Of paticular
|> > interest is NACA 8-H-12.
|> >
|> slope is two pi.
The theoretical lift-curve slope is 2*pi only for zero-thickness
airfoils. It is somewhat higher for finite-thickness airfoils.
Batchelor gives a very good estimate based on the thickness/chord
ratio t/c:
dCL/da = (1 + 0.77 t/c) * 2 * pi
This gives 2.185*pi for a 12% thick inviscid airfoil.
In reality, this will be decreased somewhat by viscous displacement
effects near the trailing edge. The real viscous lift-slope can be
less than or greater than 2*pi, depending on the airfoil thickness
and the Reynolds number.
Mark Drela First Law of Aviation:
MIT Aero & Astro "Takeoff is optional, landing is compulsory"