by: dodi suroto
Minggu, 23 Juni 2013
LIFT AUGMENTATION DEVICES
by: dodi suroto
Lift Augmentation Devices
The wing of an aircraft is designed for high or cruise speed
where lift is mainly created
by forward speed only. For slow flight we need to increase lift somehow, but there are limits to
what can be done by increasing the angle of attack. Aerodynamicists have
devised other ways of increasing the amount of lift generated. These are called lift augmentation devices.
Pilots use these devices on almost every flight.
There are several devices that can be used on an aircraft wing to
increase the lift. We all know
them as leading and trailing edge flaps. Other techniques are also used to
increase the angle of attack (coefficient lift):
vortex generators, wing fences and discontinuous leading edges.
We will look at these devices on this page, exploring them and
seeing how they operate.
Increasing Lift
An aircraft wing, designed for high or
cruise speed, can be characterized by having low camber, low thickness/chord
(T/C) ratio and you will see the maximum thickness somewhere in the middle of
the chord. CL and drag will also be low compared to
high lift wings.
To be able to fly and maneuver safely at
low airspeeds the wing must have a high lift aerofoil
(high camber, T/C ratio and maximum thickness well forward).
Flaps
All these seemingly conflicting wing
properties can be incorporated into one design. To accomplish this, the wing
designer uses leading and trailing edge flaps. When deployed, these devices
change certain properties of the wing: camber, wing area, etc.
If we revisit the lift formula: L = 1/2 ρ V2 x
S x CL, we can see that we have several ways to
increase the lift for a given
wing. Speed (V), wing area (S) and the coefficient lift (CL) can be
varied for given angle of attack (AOA) to change lift. The use of flaps will change the wing area and
coefficient lift.
Flaps
General Aviation
aircraft normally use the plain flap, although some manufacturers (Dyn Aero)
use the fowler flap on their aircraft. The image shows the types of flaps in
use today. From top to bottom: the simple or plain flap, slotted flap, split
and fowler flap.
Fowler
flaps
Its easy to see that the fowler flap not
only increases camber of the wing but also the wing area by moving toward the
back, when they are deployed. They have the greatest increase in CL with
the lowest increase in drag, hence the use on high performance aircraft,
airliners and so on.
Aerodynamic
effects of flaps
The use of trailing edge flaps (lowering
them) will have a number of effects aerodynamically. CL increases
for all angle of attacks (AOA), reducing stall speed, results in a higher CD (with
more drag airspeed stabilizes easily), rearward movement of the center of
pressure (CP), a lower stalling AOA, aircraft are more
maneuverable and can land at lower speeds resulting in shorter landing distances.
Lift/Drag
Lift/drag ratio (L/D) is
also reduced when full flaps are extended. Normally, flap settings between 0 -
25° will noticeably increase lift more
than drag, ideal for take off. Flap settings beyond 25° will increase drag much
more than lift, ideal for landing
and steep approaches.
Flaps also increase the camber and wing
area of the wing where they are installed, usually the inner part toward the
wing root.
AIRCRAFT STABILITY
by: dodi suroto
Aircraft
Stability
When an airplane is
in straight-and-level flight at a constant velocity, all the forces acting on
the airplane are in equilibrium. If that straight-and-level flight is disrupted
by a disturbance in the air, such as wake turbulence, the airplane might pitch
up or down, yaw left or right, or go into a roll. If the airplane has what is
characterized as stability, once the disturbance goes away, the airplane will
return to a state of equilibrium.
Static
Stability
The initial response
that an airplane displays after its equilibrium is disrupted is referred to as
its static stability. If the static stability is positive, the airplane will
tend to return to its original position after the disruptive force is removed.
If the static stability is negative, the airplane will continue to move away
from its original position after the disruptive force is removed. If an
airplane with negative static stability has the nose pitch up because of wake
turbulence, the tendency will be for the nose to continue to pitch up even
after the turbulence goes away. If an airplane tends to remain in a displaced
position after the force is removed, but does not continue to move toward even
greater displacement, its static stability is described as being neutral.
Dynamic
Stability
The dynamic stability
of an airplane involves the amount of time it takes for it to react to its
static stability after it has been displaced from a condition of equilibrium.
Dynamic stability involves the oscillations that typically occur as the
airplane tries to return to its original position or attitude. Even though an
airplane may have positive static stability, it may have dynamic stability
which is positive, neutral, or negative.
Imagine that an
airplane in straight-and-level flight is disturbed and pitches nose up. If the
airplane has positive static stability, the nose will pitch back down after the
disturbance is removed. If it immediately returns to straight-and-level flight,
it is also said to have positive dynamic stability. The airplane, however, may
pass through level flight and remain pitched down, and then continue the
recovery process by pitching back up. This pitching up and then down is known
as an oscillation. If the oscillations lessen over time, the airplane is still
classified as having positive dynamic stability. If the oscillations increase
over time, the airplane is classified as having negative dynamic stability. If
the oscillations remain the same over time, the airplane is classified as
having neutral dynamic stability. Figure 3-61 shows the concept of dynamic
stability. In view A, the displacement from equilibrium goes through three
oscillations and then returns to equilibrium. In view B, the displacement from
equilibrium is increasing after two oscillations, and will not return to
equilibrium. In view C, the displacement from equilibrium is staying the same
with each oscillation.
Longitudinal
Stability
Longitudinal
stability for an airplane involves the tendency for the nose to pitch up or
pitch down, rotating around the lateral axis (wingtip to wingtip). If an
airplane is longitudinally stable, it will return to a properly trimmed angle
of attack after the force that upset its flightpath is removed. The weight and
balance of an airplane, which is based on both the design characteristics of
the airplane and the way it is loaded, is a major factor in determining
longitudinal stability. There is a point on the wing of an airplane, called the
center of pressure or center of lift, where all the lifting forces concentrate.
In flight, the airplane acts like it is being lifted from or supported by this
point. This center of lift runs from wingtip to wingtip. There is also a point
on the airplane, called the center of gravity, where the mass or weight of the
airplane is concentrated. For an airplane to have good longitudinal stability,
the center of gravity is typically located forward of the center of lift. This
gives the airplane a nose-down pitching tendency, which is balanced out by the
force generated at the horizontal stabilizer and elevator. The center of
gravity has limits within which it must fall. If it is too far forward, the
forces at the tail might not be able to compensate and it may not be possible
to keep the nose of the airplane from pitching down. In Figure 3-62, the center
of lift, center of gravity, and center of gravity limits are shown. It can be
seen that the center of gravity is not only forward of the center of lift, it
is also forward of the center of gravity limit. At the back of the airplane,
the elevator trailing edge is deflected upward to create a downward force on
the tail, to try and keep the nose of the airplane up. This airplane would be
highly unstable longitudinally, especially at low speed when trying to land. It
is especially dangerous if the center of gravity is behind the aft limit. The
airplane will now have a tendency to pitch nose up, which can lead to the wing
stalling and possible loss of control of the airplane.
Lateral
Stability
Lateral stability of
an airplane takes place around the longitudinal axis, which is from the
airplane’s nose to its tail. If one wing is lower than the other, good lateral
stability will tend to bring the wings back to a level flight attitude. One
design characteristic that tends to give an airplane good lateral stability is
called dihedral. Dihedral is an upward angle for the wings with respect to the
horizontal, and it is usually just a few degrees. Imagine a low wing airplane
with a few degrees of dihedral experiencing a disruption of its flightpath such
that the left wing drops. When the left wing drops, this will cause the
airplane to experience a sideslip toward the low wing. The sideslip causes the
low wing to experience a higher angle of attack, which increases its lift and
raises it back to a level flight attitude. The dihedral on a wing is shown in
Figure 3-63.
Directional
Stability
Movement of the
airplane around its vertical axis, and the airplane’s ability to not be
adversely affected by a force creating a yaw type of motion, is called
directional stability. The vertical fin gives the airplane this stability,
causing the airplane to align with the relative wind. In flight, the airplane
acts like the weather vane we use around our home to show the direction the
wind is blowing. The distance from the pivot point on a weather vane to its
tail is greater than the distance from its pivot point to the nose. So when the
wind blows, it creates a greater torque force on the tail and forces it to
align with the wind. On an airplane, the same is true. With the CG being the
pivot point, it is a greater distance from the CG to the vertical stabilizer
than it is from the CG to the nose. [Figure 3-64]
Dutch
Roll
The dihedral of the
wing tries to roll the airplane in the opposite direction of how it is
slipping, and the vertical fin will try to yaw the airplane in the direction of
the slip. These two events combine in a way that affects lateral and
directional stability. If the wing dihedral has the greatest effect, the
airplane will have a tendency to experience a Dutch roll. A Dutch roll is a
small amount of oscillation around both the longitudinal and vertical axes.
Although this condition is not considered dangerous, it can produce an
uncomfortable feeling for passengers. Commercial airliners typically have yaw dampers
that sense a Dutch roll condition and cancel it out.
Rabu, 19 Juni 2013
NACA Airfoil Series
NACA Airfoil Series
Please send me information on the NACA 4, 5 and 6 digit airfoils. I
would like to know some general information, their applications, advantages,
disadvantages, and the formulas used to calculate the coordinates.
I'm currently trying
to design a 3D model of the B-58 bomber, but I lack the mathematical
definitions of NACA profiles such as 0003.46-64.069 (root section) and
0004.08-63 (tip section). I've found 4- and 5-digit NACA airfoil generators,
but they don't seem to do the job. Can you provide any help?
As you suggest in your questions, the early NACA airfoil series, the
4-digit, 5-digit, and modified 4-/5-digit, were generated using analytical
equations that describe the camber (curvature) of the mean-line (geometric
centerline) of the airfoil section as well as the section's thickness
distribution along the length of the airfoil. Later families, including the
6-Series, are more complicated shapes derived using theoretical rather than
geometrical methods. Before the National Advisory Committee for Aeronautics
(NACA) developed these series, airfoil design was rather arbitrary with nothing
to guide the designer except past experience with known shapes and
experimentation with modifications to those shapes.
This
methodology began to change in the early 1930s with the publishing of a NACA
report entitled The Characteristics of 78 Related Airfoil Sections from
Tests in the Variable Density Wind Tunnel. In this landmark report, the
authors noted that there were many similarities between the airfoils that were
most successful, and the two primary variables that affect those shapes are the
slope of the airfoil mean camber line and the thickness distribution above and
below this line. They then presented a series of equations incorporating these
two variables that could be used to generate an entire family of related
airfoil shapes. As airfoil design became more sophisticated, this basic
approach was modified to include additional variables, but these two basic
geometrical values remained at the heart of all NACA airfoil series, as
illustrated below.
NACA airfoil geometrical construction
NACA Four-Digit
Series:
The
first family of airfoils designed using this approach became known as the NACA
Four-Digit Series. The first digit specifies the maximum camber (m) in
percentage of the chord (airfoil length), the second indicates the position of
the maximum camber (p) in tenths of chord, and the last two numbers provide the
maximum thickness (t) of the airfoil in percentage of chord. For example, the
NACA 2415 airfoil has a maximum thickness of 15% with a camber of 2% located
40% back from the airfoil leading edge (or 0.4c). Utilizing these m, p, and t
values, we can compute the coordinates for an entire airfoil using the
following relationships:
1.
Pick values of x from 0 to the
maximum chord c.
2.
Compute the mean camber line
coordinates by plugging the values of m and p into the following equations for
each of the x coordinates.
3.
Calculate the thickness
distribution above (+) and below (-) the mean line by plugging the value of t
into the following equation for each of the x coordinates.
4.
Determine the final coordinates
for the airfoil upper surface (xU, yU) and lower surface
(xL, yL) using the following relationships.
NACA Five-Digit Series:
The
NACA Five-Digit Series uses the same thickness forms as the Four-Digit Series
but the mean camber line is defined differently and the naming convention is a
bit more complex. The first digit, when multiplied by 3/2, yields the design
lift coefficient (cl) in tenths. The next two digits, when divided
by 2, give the position of the maximum camber (p) in tenths of chord. The final
two digits again indicate the maximum thickness (t) in percentage of chord. For
example, the NACA 23012 has a maximum thickness of 12%, a design lift
coefficient of 0.3, and a maximum camber located 15% back from the leading
edge. The steps needed to calculate the coordinates of such an airfoil are:
1.
Pick values of x from 0 to the
maximum chord c.
2.
Compute the mean camber line
coordinates for each x location using the following equations, and since we
know p, determine the values of m and k1 using the table shown
below.
3.
Calculate the thickness
distribution using the same equation as the Four-Digit Series.
4.
Determine the final coordinates
using the same equations as the Four-Digit Series.
Modified NACA Four- and Five-Digit Series:
The
airfoil sections you mention for the B-58 bomber are members of the Four-Digit
Series, but the names are slightly different as these shapes have been
modified. Let us consider the root section, the NACA 0003.46-64.069, as an
example. The basic shape is the 0003, a 3% thick airfoil with 0% camber. This
shape is a symmetrical airfoil that is identical above and below the mean
camber line. The first modification we will consider is the 0003-64. The first
digit following the dash refers to the roundedness of the nose. A value of 6
indicates that the nose radius is the same as the original airfoil while a
value of 0 indicates a sharp leading edge. Increasing this value specifies an
increasingly more rounded nose. The second digit determines the location of
maximum thickness in tenths of chord. The default location for all four- and
five-digit airfoils is 30% back from the leading edge. In this example, the
location of maximum thickness has been moved back to 40% chord. Finally, notice
that the 0003.46-64.069 features two sets of digits preceeded by decimals.
These merely indicate slight adjustments to the maximum thickness and location
thereof. Instead of being 3% thick, this airfoil is 3.46% thick. Instead of the
maximum thickness being located at 40% chord, the position on this airfoil is
at 40.69% chord. To compute the coordinates for a modified airfoil shape:
1.
Pick values of x from 0 to the
maximum chord c.
2.
Compute the mean camber line
coordinates using the same equations provided for the Four- or Five-Digit
Series as appropriate.
3.
Calculate the thickness
distribution above (+) and below (-) the mean line using these equations. The
values of the ax and dx coefficients are determined from
the following table (these are derived for a 20% thick airfoil).
4.
Determine the "final"
coordinates using the same equations as the Four-Digit Series.
5.
As noted above, this procedure
yields a 20% thick airfoil. To obtain the desired thickness, simply scale the
airfoil by multiplying the "final" y coordinates by [t / 0.2].
NACA 1-Series or 16-Series:
Unlike
those airfoil families discussed so far, the 1-Series was developed based on
airfoil theory rather than on geometrical relationships. By the time these
airfoils were designed during the late 1930s, many advances had been made in
inverse airfoil design methods. The basic concept behind this design approach
is to specify the desired pressure distribution over the airfoil (this
distribution dictates the lift characteristics of the shape) and then derive
the geometrical shape that produces this pressure distribution. As a result,
these airfoils were not generated using some set of analytical expressions like
the Four- or Five-Digit Series. The 1-Series airfoils are identified by five
digits, as exemplified by the NACA 16-212. The first digit, 1, indicates the
series (this series was designed for airfoils with regions of barely supersonic
flow). The 6 specifies the location of minimum pressure in tenths of chord,
i.e. 60% back from the leading edge in this case. Following a dash, the first
digit indicates the design lift coefficient in tenths (0.2) and the final two
digits specify the maximum thickness in tenths of chord (12%). Since the 16-XXX
airfoils are the only ones that have ever seen much use, this family is often
referred to as the 16-Series rather than as a subset of the 1-Series.
NACA 6-Series:
Although
NACA experimented with approximate theoretical methods that produced the
2-Series through the 5-Series, none of these approaches was found to accurately
produce the desired airfoil behavior. The 6-Series was derived using an
improved theoretical method that, like the 1-Series, relied on specifying the
desired pressure distribution and employed advanced mathematics to derive the
required geometrical shape. The goal of this approach was to design airfoils
that maximized the region over which the airflow remains laminar. In so doing,
the drag over a small range of lift coefficients can be substantially reduced.
The naming convention of the 6-Series is by far the most confusing of any of
the families discussed thus far, especially since many different variations
exist. One of the more common examples is the NACA 641-212, a=0.6.
In
this example, 6 denotes the series and indicates that this family is designed
for greater laminar flow than the Four- or Five-Digit Series. The second digit,
4, is the location of the minimum pressure in tenths of chord (0.4c). The subscript
1 indicates that low drag is maintained at lift coefficients 0.1 above and
below the design lift coefficient (0.2) specified by the first digit after the
dash in tenths. The final two digits specify the thickness in percentage of
chord, 12%. The fraction specified by a=___ indicates the percentage of the
airfoil chord over which the pressure distribution on the airfoil is uniform,
60% chord in this case. If not specified, the quantity is assumed to be 1, or
the distribution is constant over the entire airfoil.
NACA 7-Series:
The
7-Series was a further attempt to maximize the regions of laminar flow over an
airfoil differentiating the locations of the minimum pressure on the upper and
lower surfaces. An example is the NACA 747A315. The 7 denotes the series, the 4
provides the location of the minimum pressure on the upper surface in tenths of
chord (40%), and the 7 provides the location of the minimum pressure on the
lower surface in tenths of chord (70%). The fourth character, a letter,
indicates the thickness distribution and mean line forms used. A series of
standaradized forms derived from earlier families are designated by different
letters. Again, the fifth digit incidates the design lift coefficient in tenths
(0.3) and the final two integers are the airfoil thickness in perecentage of
chord (15%).
NACA 8-Series:
A
final variation on the 6- and 7-Series methodology was the NACA 8-Series
designed for flight at supercritical speeds. Like the earlier airfoils, the
goal was to maximize the extent of laminar flow on the upper and lower surfaces
independently. The naming convention is very similar to the 7-Series, an
example being the NACA 835A216. The 8 designates the series, 3 is the location
of minimum pressure on the upper surface in tenths of chord (0.3c), 5 is the
location of minimum pressure on the lower surface in tenths of chord (50%), the
letter A distinguishes airfoils having different camber or thickness forms, 2
denotes the design lift coefficient in tenths (0.2), and 16 provides the airfoil
thickness in percentage of chord (16%).
Further Sources:
This
is probably the most theoretical and mathematically-intense answer we have yet
given on this site, but let me point out that coordinates for many of these
airfoils already exist in print or on the web. In addition, many programs and
web sites now exist that can automatically compute the coordinates once the
user enters the desired airfoil name or characteristics. Some excellent tools
include:
- NACA 4-Digit Series Airfoil Generator
- NACA 5-Digit Series Airfoil Generator
- SNACK -- download this program that includes NACA airfoil coordinate generators for all of the families we have discussed
- XFOIL -- download this airfoil analysis code that includes a 4-Digit and 5-Digit airfoil generation tool, but this program is difficult for a novice to use (program is 100% free)
- UIUC Airfoil Coordinates Database -- vast library of coordinates for many airfoils, including those of the NACA families discussed above
- The Incomplete Guide to Airfoil Usage -- check out the airfoils used on a huge assortment of aircraft
Summary:
Though
we have introduced the primary airfoil families developed in the United States
before the advent of supersonic flight, we haven't said anything about their
uses. So let's briefly explore the advantages, disadvantages, and applications
of each of these families.
|
Family
|
Advantages
|
Disadvantages
|
Applications
|
|
4-Digit
|
1. Good stall
characteristics
2. Small center of pressure movement across large
speed range
3. Roughness has little effect
|
1. Low maximum lift
coefficient
2. Relatively high drag
3. High pitching moment
|
1. General aviation
2. Horizontal tails
Symmetrical:
3. Supersonic jets
4. Helicopter blades 5. Shrouds 6. Missile/rocket fins |
|
5-Digit
|
1. Higher maximum lift
coefficient
2. Low pitching moment
3. Roughness has little effect
|
1. Poor stall behavior
2. Relatively high drag
|
1. General aviation
2. Piston-powered bombers, transports 3. Commuters 4. Business jets |
|
16-Series
|
1. Avoids low pressure
peaks
2. Low drag at high speed
|
1. Relatively low lift
|
1. Aircraft propellers
2. Ship propellers |
|
6-Series
|
1. High maximum lift
coefficient
2. Very low drag over a small range of operating
conditions
3. Optimized for high speed
|
1. High drag outside of
the optimum range of operating conditions
2. High pitching moment
3. Poor stall behavior
4. Very susceptible to roughness
|
1. Piston-powered fighters
2. Business jets 3. Jet trainers 4. Supersonic jets |
|
7-Series
|
1. Very low drag over a
small range of operating conditions
2. Low pitching moment
|
1. Reduced maximum lift
coefficient
2. High drag outside of the optimum range of
operating conditions
3. Poor stall behavior
4. Very susceptible to roughness
|
Seldom used
|
|
8-Series
|
Unknown
|
Unknown
|
Very seldom used
|
Today,
airfoil design has in many ways returned to an earlier time before the NACA
families were created. The computational resources available now allow the
designer to quickly design and optimize an airfoil specifically tailored to a
particular application rather than making a selection from an existing family.
To learn about some more recent developments, check out other questions on
supercritical airfoils and airfoil design.
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