Lift-to-Drag Ratio?

Alan Baker wrote in
:

OTOH, look at an F-16 fighter. Virtually no "lift" from the airfoil
design at all. The lift comes from Angle of Attack (AOA), which
causes induced drag.

How can people write something as incorrect as AOA generated lift
causes induced drag that "airfoil design" wouldn't????

Lack of sleep? Rushed? G

Still, it's not that incorrect. I didn't actually say what you
paraphrased... that lift due to airfoil design didn't cause induced drag.

Increasing the angle of attack (until stall) increases lift, with a
corresponding increase in induced drag. The alternative is a higher lift
airfoil that produces the same amount of lift at a much lower airspeed,
where lower parasitic drag will be encountered.

Please feel free to give him a much more detailed (and hopefully more
specifically accurate) explanation. But remember that he is looking for a
a basic understanding... not Reynolds numbers.

-----------------------------------------------
James M. Knox
TriSoft ph 512-385-0316
1109-A Shady Lane fax 512-366-4331
Austin, Tx 78721
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"James M. Knox" wrote in message ...
Alan Baker wrote in
:

OTOH, look at an F-16 fighter. Virtually no "lift" from the airfoil
design at all. The lift comes from Angle of Attack (AOA), which
causes induced drag.

....

Please feel free to give him a much more detailed (and hopefully more
specifically accurate) explanation. But remember that he is looking for a
a basic understanding... not Reynolds numbers.


In supersonic flight the passage of the leading edge of the airfoil
through the air generates a shock wave. Ahead of the shock wave the
air pressure is ambient, of course because the shock wave hasn't
arrived there yet. At the shock wave the pressure is higher than
ambient, which why it is a wave. Behind the shock wave the pressure
is less than ambient because some of that air has been pushed forward
into the shock wave.

Now, there actually are two shock waves generated, one over the top of
the wing and and one under the bottom. Both trail back at some angle
relative to the wing. For various angles of attack the angle between
the wing and the upper shock wave will be different from the angle
between the wing and the lower shock wave. There is also a difference
in the pressure behind the wave over the wing and the wave under the
wing. The lift is approximately the product of this pressure
differential and the area of the wing, e.g. (Po - Pu) * Ah = Lift
where Po is the pressure over the wing, Pu under the wing and Ah
is the cross sectional area of the wing in a plane parallel to the
direction of motion.

Now, there is a region of reduced pressure behind the wing which is
somewhere between Po and Pu in value, call that Pb.
Remembering that ahead of the wing is the shock wave (Pw) and ahead of
that ambient air (Pa) then to a first approximation:
[(Pa - Pw) + (Pw - Pb)] * Av = (Pa -Pb) * Av = Drag, where Av is
the cross-sectional area of the wing perpendicular to the direction
of motion.

Now of course the pressures in these regions are not constant so to
get a more exact answer you would integrate all the pressures over
the entire surface of the aircraft (not just the wings) to arrive
at one force which is the net superposition of lift and drag.

At least that is the way I remember it from aeronautics.

--

FF