[DeTomaso] Rake as drag / Just a guess.

Wed Feb 22 13:44:55 EST 2006

So......isn't the whole notion of a wing being "half a Bernoulli tube",
generating lift while leaving the air undisturbed behind it, in conflict
with Newton's laws of motion?

IOW, if an aircraft is accelerating enough air downward, won't it will
overcome gravity, regardless of any "longer path over the top of the wing"
type hooey?  After all, aircraft fly upside down, on their sides, etc.

Hey, if you want to see some old cars and bikes at Bonneville, go see "The
World's Fastest Indian".  Saccharin story, little technical detail beyond
the guy casting his own pistons at home, some talk of getting the center of
gravity ahead of the center of pressure, but lots of old cars, some
apparently hauled out of museums to make the movie.  Question:  Is "Jim
Moffat from San Jose" a real historical person?

Guy D.

----- Original Message ----- 
From: "Daniel C Jones" <daniel.c.jones2 at gmail.com>
To: <detomaso at realbig.com>
Sent: Wednesday, February 22, 2006 9:20 AM
Subject: Re: [DeTomaso] Rake as drag / Just a guess.

> > Why do you say a wing will load the engine?
>
> Wings produce drag as a direct consequence of generating lift/downforce.
> The drag component is termed induced drag and is proportional to the
square
> of the lift/downforce produced:
>
>    Cdi = Cl**2/(pi*e*AR)
>
>  where:
>
>    Cdi = induced drag coefficient
>    Cl  = coefficient of lift
>    AR  = the aspect ratio (wing span squared/wing area) of the wing
>    pi  = mathematical constant (approximately 3.14159)
>    e   = wing efficiency factor (1 for an elliptical wing planform like
>          that used on the WWII Spitfire fighter planes, less than 1 for
>          other planforms)
>
> This drag is in *addition* to the wing's basic profile drag (the drag at
zero
> lift) which varies with the square of velocity:
>
>     Cdo = Do / (q*S)
>
>   where:
>
>     Cdo = coefficient of draq at zero lift
>     D   = airfoil aerodynamic drag force in lbs at zero lift
>     S   = wing planform area in square feet (ft**2)
>     q   = dynamic pressure in lbs/ft**2
>
>   and:
>
>     q  = (rho * V**2)/2
>
>   where:
>
>     V   = velocity
>     rho = air density (a function of temperature and altitude)
>
> Also remember that while drag varies with the square of speed, the power
> required to overcome drag varies with the cube of speed:
>
>        D = 1/2 Rho V*V * Cd * Area
>
> Power is a rate so
>
>        Preqd = 1/2 * Rho * V ** 3 * Cd * Area
>
> To estimate a car's top speed, use the following formula:
>
>                           /------------
>                15        /  1100 P
>       Vmax =  ---- \ 3  / -------------
>                22   \  /   Cd S rho
>                      \/
>
> where
>
>         P      = power in horsepower
>         Cd     = total drag coefficient
>         S      = frontal Area in square feet
>         rho    = density of air in slug/cu. ft.
>         Vmax   = speed in miles/hour
>         rho    = 0.002378 slug/cu ft. (standard sea level density)
>         1 HP   = 550 LBF * FT / SEC
>         1 mile = 5280 feet
>
> Note: The above formula ignores the effect of tire rolling resistance.
>
> > Does anyone know the ratio of lift (or down force) to drag for an
efficent
> > wing design for 150 to 200 MPH?
>
> You'll find race car wings are usually designed not to minimized drag for
a
> given lift but rather to maximize lift (downforce) within the rule
limitations
> (usually some sort of a physical constraint on the allowable dimensions
like
> width).  Bob Liebeck is a fellow Boeing engineer and designed a series of
> race car airfoils ("Liebeck Airfoils") that go back 20 or 30 years and
stem
> from research on the maximum pressure gradient a boundary layer can
sustain
> without separation.  As I recall, they were inverse airfoil designs with
> Stratford pressure-recovery on the aft end and laminar flow forward
sections.
> Liebeck's airfoils were designed to provide maximum lift (downforce) but
do
> so at the cost of drag.  Depending upon the application, there may be
lower
> drag airfoils that can produce similar downforce in a larger physical
space.
> I met a guy who crewed on a Trans Am race team and the rules say their car
> must use Liebeck airfoil #1LD104E.  I don't have that particular airfoil
in
> my database but I do have an AIAA book with the following Liebeck
airfoils:
>
>  L1003
>  L1004
>  LA203A
>  LA5104E
>  FX63-137
>  LS2573A
>  LC111A
>  LW101B
>  LW108A
>  LAB121A15
>  LSB119A
>  L175
>  LPT102B
>
> I found this airfoil database on the web you might be interested in.
> It has a couple Lieback airfoils in it:
>
>  http://www.nasg.com/afdb/list-airfoil-e.phtml
>
> Remember that the body in front of the wing influences the effective angle
> of attack and the angle of the relative wind.  This is especially
important
> for aircraft as a wing in front of another (the horizontal tail, for
> instance) has a big influence on the effective angle of attack.  For
> three-dimensional wings, downwash, induced by trailing wingtip vortices,
> alters the local flowfield, resulting in an effective angle of attack
> which is less than the geometric angle of attack.  The difference between
> the geometric and effective angles of attack is termed the induced angle
> of attack:
>
>   AOAi = AOA - AOAeff
>
>   where:
>
>    AOA    = geometric angle of attack
>    AOAi   = induced angle of attack
>    AOAeff = effective angle of attack
>
> For an elliptical lift distribution, the induced angle of attack can be
> calculated as:
>
>     AOAi = Cl/(pi*AR)
>
> This occurs in both symmeric and cambered airfoils, and is a three-
> dimensional effect.  Air foil sections are two-dimensional and therefore
> the geometric and effective angle of attacks are the same (i.e. the
> induced angle of attack is zero).  Even for 2-D airfoil sections, in
> general, there is a non-zero lift component at zero angle of attack
> for cambered airfoils.  Note that air foil data can be applied to
> finite wings since flow over an inifinite wing at geometric angle
> of attack is similar to a finite wing at effective angle of attack.
> Since it can be measured directly, wind tunnel data for finite wings
> is generally taken at the geometric angle of attack.
>
>
> > The Liebeck profile is non-symetrical, so it generates lift/downforce
> > at 0 degrees AofA.
>
> Usually referred to as a cambered airfoil.  A little wing theory and
> several definitions are in order here.  This would be easier to explain
> with illustrations, but I'll give it a shot with words.  An airfoil is
> the 2-dimensional cross-sectional shape obtained by the intersection of a
> wing and a perpendicular plane.  The mean camber line of an airfoil is the
> locus of points halfway between the upper and lower surfaces (measured
> perpendicular to the mean camber line itself).  The chord of an airfoil is
> the straight line connecting its leading edge to its trailing edge. Camber
> is the maximum distance between the mean camber line and the chord line,
> measured perpendicular to the chord line.
>
> An airfoil's angle of attack is the angle between the relative wind (the
> local airflow direction) and the airfoil's chord line.  Drag is the
> component of aerodynamic force parallel to the relative wind and lift is
> the perpendicular component.
>
> If an airfoil is symmetric (top-to-bottom), it has no camber.  A sheet of
> plywood has no camber.  A Venetian blind slat is a shape that has camber
but
> (practically) no thickness.  The camber, the shape of the mean camber
line,
> and the thickness distribution of an airfoil determine its lift and moment
> characteristics.  Surface roughness also plays a roll but is usually
treated
> as a separate design issue.
>
> Because of camber, wings can have lift at zero degrees angle of attack and
> because of angle of attack, wings (and sheets of plywood) with no camber
can
> still produce lift.  To separate these effects, aerodynamicists break an
> airfoil's lift into two components:
>
>     Cl = Clo + (Cla * alpha)
>
>   where:
>
>     Cl = coefficient of lift
>     Clo = coefficient of lift at zero angle of attack
>     Cla = lift curve slope (the slope of Cl versus alpha)
>     alpha = angle of attack
>
> Dan Jones
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