Pitching Rolling Yawing
Lateral axis Longitudinal axis Vertical axis
Figure 5-18. Axes of an airplane.
During the takeoff phase of flight, ground effect produces
some important relationships. An aircraft leaving ground
effect after takeoff encounters just the reverse of an aircraft
entering ground effect during landing. The aircraft leaving
ground effect will:
• Require an increase in AOA to maintain the same CL
• Experience an increase in induced drag and thrust
required
• Experience a decrease in stability and a nose-up
change in moment
• Experience a reduction in static source pressure and
increase in indicated airspeed
Ground effect must be considered during takeoffs and landings.
For example, if a pilot fails to understand the relationship
between the aircraft and ground effect during takeoff, a
hazardous situation is possible because the recommended
takeoff speed may not be achieved. Due to the reduced drag
in ground effect, the aircraft may seem capable of takeoff well
below the recommended speed. As the aircraft rises out of
ground effect with a deficiency of speed, the greater induced
drag may result in marginal initial climb performance. In
extreme conditions, such as high gross weight, high density
altitude, and high temperature, a deficiency of airspeed during
takeoff may permit the aircraft to become airborne but be
incapable of sustaining flight out of ground effect. In this case,
the aircraft may become airborne initially with a deficiency
of speed and then settle back to the runway.
A pilot should not attempt to force an aircraft to become
airborne with a deficiency of speed. The manufacturer’s
recommended takeoff speed is necessary to provide adequate
initial climb performance. It is also important that a definite
climb be established before a pilot retracts the landing gear
or flaps. Never retract the landing gear or flaps prior to
establishing a positive rate of climb and only after achieving
a safe altitude.
If, during the landing phase of flight, the aircraft is brought
into ground effect with a constant AOA, the aircraft
experiences an increase in C L and a reduction in the thrust
required, and a “floating” effect may occur. Because of the
reduced drag and lack of power-off deceleration in ground
effect, any excess speed at the point of flare may incur a
considerable “float” distance. As the aircraft nears the point
of touchdown, ground effect is most realized at altitudes less
than the wingspan. During the final phases of the approach
as the aircraft nears the ground, a reduction of power is
necessary to offset the increase in lift caused from ground
effect otherwise the aircraft will have a tendency to climb
above the desired glidepath (GP).
Axes of an Aircraft
The axes of an aircraft are three imaginary lines that pass
through an aircraft’s CG. The axes can be considered as
imaginary axles around which the aircraft turns. The three
axes pass through the CG at 90° angles to each other. The
axis passes through the CG and parallel to a line from nose
to tail is the longitudinal axis, the axis that passes parallel
to a line from wingtip to wingtip is the lateral axis, and the
axis that passes through the CG at right angles to the other
two axes is the vertical axis. Whenever an aircraft changes
its flight attitude or position in flight, it rotates about one or
more of the three axes. [Figure 5-18]
The aircraft’s motion about its longitudinal axis resembles
the roll of a ship from side to side. In fact, the names
used to describe the motion about an aircraft’s three axes
were originally nautical terms. They have been adapted to
aeronautical terminology due to the similarity of motion of
aircraft and seagoing ships. The motion about the aircraft’s
longitudinal axis is “roll,” the motion about its lateral axis is
Figure 5-20. A powered parachute.
Figure 5-19. A weight-shift control aircraft.
“pitch,” and the motion about its vertical axis is “yaw.” Yaw
is the left and right movement of the aircraft’s nose.
The three motions of the conventional airplane (roll, pitch,
and yaw) are controlled by three control surfaces. Roll is
controlled by the ailerons; pitch is controlled by the elevators;
yaw is controlled by the rudder. The use of these controls
is explained in Chapter 6, Flight Controls. Other types of
aircraft may utilize different methods of controlling the
movements about the various axes.
For example, weight-shift control aircraft control two axes
(roll and pitch) using an “A” frame suspended from the
flexible wing attached to a three-wheeled carriage. These
aircraft are controlled by moving a horizontal bar (called a
control bar) in roughly the same way hang glider pilots fly.
[Figure 5-19] They are termed weight-shift control aircraft
because the pilot controls the aircraft by shifting the CG.
For more information on weight-shift control aircraft, see
the Federal Aviation Administration (FAA) Weight-Shift
Control Flying Handbook, FAA-H-8083-5. In the case of
powered parachutes, aircraft control is accomplished by
altering the airfoil via steering lines.
A powered parachute wing is a parachute that has a cambered
upper surface and a flatter under surface. The two surfaces are
separated by ribs that act as cells, which open to the airflow
at the leading edge and have internal ports to allow lateral
airflow. The principle at work holds that the cell pressure is
greater than the outside pressure, thereby forming a wing that
maintains its airfoil shape in flight. The pilot and passenger
sit in tandem in front of the engine, which is located at the
rear of a vehicle. The airframe is attached to the parachute
via two attachment points and lines. Control is accomplished
by both power and the changing of the airfoil via the control
lines. [Figure 5-20]
Moment and Moment Arm
A study of physics shows that a body that is free to rotate
will always turn about its CG. In aerodynamic terms, the
mathematical measure of an aircraft’s tendency to rotate
about its CG is called a “moment.” A moment is said to be
equal to the product of the force applied and the distance at
which the force is applied. (A moment arm is the distance
from a datum [reference point or line] to the applied force.)
For aircraft weight and balance computations, “moments”
are expressed in terms of the distance of the arm times the
aircraft’s weight, or simply, inch-pounds.
Aircraft designers locate the fore and aft position of the
aircraft’s CG as nearly as possible to the 20 percent point
of the mean aerodynamic chord (MAC). If the thrust line
is designed to pass horizontally through the CG, it will not
cause the aircraft to pitch when power is changed, and there
will be no difference in moment due to thrust for a power-on
or power-off condition of flight. Although designers have
some control over the location of the drag forces, they are not
always able to make the resultant drag forces pass through the
CG of the aircraft. However, the one item over which they
have the greatest control is the size and location of the tail.
The objective is to make the moments (due to thrust, drag, and
lift) as small as possible and, by proper location of the tail,
to provide the means of balancing an aircraft longitudinally
for any condition of flight.
The pilot has no direct control over the location of forces
acting on the aircraft in flight, except for controlling the
center of lift by changing the AOA. The pilot can control
the magnitude of the forces. Such a change, however,
immediately involves changes in other forces. Therefore,
the pilot cannot independently change the location of one
force without changing the effect of others. For example,
a change in airspeed involves a change in lift, as well as a
change in drag and a change in the up or down force on the
Positive Static Stability Neutral Static Stability Negative Static Stability
CG
CG
CG
CG
Applied
force
Applied
force
Applied
force
Figure 5-21. Types of static stability.
tail. As forces such as turbulence and gusts act to displace
the aircraft, the pilot reacts by providing opposing control
forces to counteract this displacement.
Some aircraft are subject to changes in the location of the CG
with variations of load. Trimming devices, such as elevator
trim tabs and adjustable horizontal stabilizers, are used to
counteract the moments set up by fuel burnoff and loading
or off-loading of passengers or cargo.
Aircraft Design Characteristics
Each aircraft handles somewhat differently because each
resists or responds to control pressures in its own way. For
example, a training aircraft is quick to respond to control
applications, while a transport aircraft feels heavy on the
controls and responds to control pressures more slowly.
These features can be designed into an aircraft to facilitate
the particular purpose of the aircraft by considering certain
stability and maneuvering requirements. The following
discussion summarizes the more important aspects of an
aircraft’s stability, maneuverability, and controllability
qualities; how they are analyzed; and their relationship to
various flight conditions.
Stability
Stability is the inherent quality of an aircraft to correct for
conditions that may disturb its equilibrium and to return to
or to continue on the original flight path. It is primarily an
aircraft design characteristic. The flight paths and attitudes an
aircraft flies are limited by the aerodynamic characteristics of
the aircraft, its propulsion system, and its structural strength.
These limitations indicate the maximum performance and
maneuverability of the aircraft. If the aircraft is to provide
maximum utility, it must be safely controllable to the full
extent of these limits without exceeding the pilot’s strength
or requiring exceptional flying ability. If an aircraft is to fly
straight and steady along any arbitrary flight path, the forces
acting on it must be in static equilibrium. The reaction of
any body when its equilibrium is disturbed is referred to as
stability. The two types of stability are static and dynamic.
Static Stability
Static stability refers to the initial tendency, or direction of
movement, back to equilibrium. In aviation, it refers to the
aircraft’s initial response when disturbed from a given pitch,
yaw, or bank.
• Positive static stability—the initial tendency of the
aircraft to return to the original state of equilibrium
after being disturbed. [Figure 5-21]
• Neutral static stability—the initial tendency of
the aircraft to remain in a new condition after its
equilibrium has been disturbed. [Figure 5-21]
• Negative static stability—the initial tendency of the
aircraft to continue away from the original state of
equilibrium after being disturbed. [Figure 5-21]
Dynamic Stability
Static stability has been defined as the initial tendency to
return to equilibrium that the aircraft displays after being
disturbed from its trimmed condition. Occasionally, the
initial tendency is different or opposite from the overall
tendency, so a distinction must be made between the two.
Dynamic stability refers to the aircraft response over time
Displacement
Positive static
(neutral dynamic)
Positive static
(positive dynamic)
Positive Static
(negative dynamic)
Time
Damped oscillation
Undamped oscillation
Divergent oscillation
Figure 5-22. Damped versus undamped stability.
when disturbed from a given pitch, yaw, or bank. This type
of stability also has three subtypes: [Figure 5-22]
• Positive dynamic stability—over time, the motion
of the displaced object decreases in amplitude and,
because it is positive, the object displaced returns
toward the equilibrium state.
• Neutral dynamic stability—once displaced, the
displaced object neither decreases nor increases in
amplitude. A worn automobile shock absorber exhibits
this tendency.
• Negative dynamic stability—over time, the motion
of the displaced object increases and becomes more
divergent.
Stability in an aircraft affects two areas significantly:
• Maneuverability—the quality of an aircraft that
permits it to be maneuvered easily and to withstand
the stresses imposed by maneuvers. It is governed by
the aircraft’s weight, inertia, size and location of flight
controls, structural strength, and powerplant. It too is
an aircraft design characteristic.
• Controllability—the capability of an aircraft to
respond to the pilot’s control, especially with regard
to flight path and attitude. It is the quality of the
aircraft’s response to the pilot’s control application
when maneuvering the aircraft, regardless of its
stability characteristics.
Longitudinal Stability (Pitching)
In designing an aircraft, a great deal of effort is spent in
developing the desired degree of stability around all three
axes. But longitudinal stability about the lateral axis is
considered to be the most affected by certain variables in
various flight conditions.
Longitudinal stability is the quality that makes an aircraft
stable about its lateral axis. It involves the pitching motion
as the aircraft’s nose moves up and down in flight. A
longitudinally unstable aircraft has a tendency to dive or
climb progressively into a very steep dive or climb, or even
a stall. Thus, an aircraft with longitudinal instability becomes
difficult and sometimes dangerous to fly.
Static longitudinal stability, or instability in an aircraft, is
dependent upon three factors:
1. Location of the wing with respect to the CG
2. Location of the horizontal tail surfaces with respect
to the CG
3. Area or size of the tail surfaces
In analyzing stability, it should be recalled that a body free
to rotate always turns about its CG.
To obtain static longitudinal stability, the relation of the
wing and tail moments must be such that, if the moments
are initially balanced and the aircraft is suddenly nose up,
the wing moments and tail moments change so that the sum
of their forces provides an unbalanced but restoring moment
which, in turn, brings the nose down again. Similarly, if the
aircraft is nose down, the resulting change in moments brings
the nose back up.
The Center of Lift (CL) in most asymmetrical airfoils has a
tendency to change its fore and aft positions with a change in
the AOA. The CL tends to move forward with an increase in
AOA and to move aft with a decrease in AOA. This means
CG
CL
CG
CL
Figure 5-23. Longitudinal stability.
CG
CG
CG
Balanced tail load
Lesser downward tail load
Greater downward tail load
Cruise Speed
High Speed
Low Speed
Figure 5-24. Effect of speed on downwash.
Normal downwash
Reduced downwash
Thrust
Weight
CG
Lift
Thrust
Weight
CG
Lift
Figure 5-25. Reduced power allows pitch down.
that when the AOA of an airfoil is increased, the CL, by
moving forward, tends to lift the leading edge of the wing
still more. This tendency gives the wing an inherent quality
of instability. (NOTE: CL is also known as the center of
pressure (CP).)
Figure 5-23 shows an aircraft in straight-and-level flight. The
line CG-CL-T represents the aircraft’s longitudinal axis from
the CG to a point T on the horizontal stabilizer.
Most aircraft are designed so that the wing’s CL is to the rear
of the CG. This makes the aircraft “nose heavy” and requires
that there be a slight downward force on the horizontal
stabilizer in order to balance the aircraft and keep the nose
from continually pitching downward. Compensation for this
nose heaviness is provided by setting the horizontal stabilizer
at a slight negative AOA. The downward force thus produced
holds the tail down, counterbalancing the “heavy” nose. It
is as if the line CG-CL-T were a lever with an upward force
at CL and two downward forces balancing each other, one
a strong force at the CG point and the other, a much lesser
force, at point T (downward air pressure on the stabilizer).
To better visualize this physics principle: If an iron bar were
suspended at point CL, with a heavy weight hanging on it at
the CG, it would take downward pressure at point T to keep
the “lever” in balance.
Even though the horizontal stabilizer may be level when the
aircraft is in level flight, there is a downwash of air from the
wings. This downwash strikes the top of the stabilizer and
produces a downward pressure, which at a certain speed is
just enough to balance the “lever.” The faster the aircraft
is flying, the greater this downwash and the greater the
downward force on the horizontal stabilizer (except T-tails).
[Figure 5-24] In aircraft with fixed-position horizontal
stabilizers, the aircraft manufacturer sets the stabilizer at an
angle that provides the best stability (or balance) during flight
at the design cruising speed and power setting.
If the aircraft’s speed decreases, the speed of the airflow
over the wing is decreased. As a result of this decreased
flow of air over the wing, the downwash is reduced, causing
a lesser downward force on the horizontal stabilizer. In turn,
the characteristic nose heaviness is accentuated, causing the
aircraft’s nose to pitch down more. [Figure 5-25] This places
the aircraft in a nose-low attitude, lessening the wing’s AOA
and drag and allowing the airspeed to increase. As the aircraft
continues in the nose-low attitude and its speed increases,
the downward force on the horizontal stabilizer is once again
increased. Consequently, the tail is again pushed downward
and the nose rises into a climbing attitude.
Below center of gravity
Through center of gravity
Above center of gravity
Thrust CG
Thrust CG
Thrust
CG
Figure 5-26. Thrust line affects longitudinal stability.
Cruise power
Idle power
Full power
Lift
Thrust CG
Lift
Thrust CG
Lift
Thrust CG
Figure 5-27. Power changes affect longitudinal stability.
As this climb continues, the airspeed again decreases, causing
the downward force on the tail to decrease until the nose
lowers once more. Because the aircraft is dynamically stable,
the nose does not lower as far this time as it did before. The
aircraft acquires enough speed in this more gradual dive to
start it into another climb, but the climb is not as steep as
the preceding one.
After several of these diminishing oscillations, in which
the nose alternately rises and lowers, the aircraft finally
settles down to a speed at which the downward force on the
tail exactly counteracts the tendency of the aircraft to dive.
When this condition is attained, the aircraft is once again in
balanced flight and continues in stabilized flight as long as
this attitude and airspeed are not changed.
A similar effect is noted upon closing the throttle. The
downwash of the wings is reduced and the force at T in
Figure 5-23 is not enough to hold the horizontal stabilizer
down. It seems as if the force at T on the lever were allowing
the force of gravity to pull the nose down. This is a desirable
characteristic because the aircraft is inherently trying to
regain airspeed and reestablish the proper balance.
Power or thrust can also have a destabilizing effect in that
an increase of power may tend to make the nose rise. The
aircraft designer can offset this by establishing a “high
thrust line” wherein the line of thrust passes above the CG.
[Figures 5-26 and 5-27] In this case, as power or thrust is
increased a moment is produced to counteract the down load
on the tail. On the other hand, a very “low thrust line” would
tend to add to the nose-up effect of the horizontal tail surface.
Conclusion: with CG forward of the CL and with an
aerodynamic tail-down force, the aircraft usually tries to
return to a safe flying attitude.
The following is a simple demonstration of longitudinal
stability. Trim the aircraft for “hands off” control in level
flight. Then, momentarily give the controls a slight push to
nose the aircraft down. If, within a brief period, the nose rises
towards the original position, the aircraft is statically stable.
Ordinarily, the nose passes the original position (that of level
flight) and a series of slow pitching oscillations follows. If the
oscillations gradually cease, the aircraft has positive stability;
if they continue unevenly, the aircraft has neutral stability;
if they increase, the aircraft is unstable.
Lateral Stability (Rolling)
Stability about the aircraft’s longitudinal axis, which extends
from the nose of the aircraft to its tail, is called lateral
stability. Positive lateral stability helps to stabilize the lateral
or “rolling effect” when one wing gets lower than the wing
on the opposite side of the aircraft. There are four main
design factors that make an aircraft laterally stable: dihedral,
sweepback, keel effect, and weight distribution.
DihedralDihedral
Figure 5-28. Dihedral is the upward angle of the wings from a horizontal (front/rear view) axis of the plane as shown in the graphic
depiction and the rear view of a Ryanair Boeing 737.
Wing has decreased
AOA, hence reduced lift
due to sideslip.
Wing has increased AOA, hence
increased lift due to sideslip.
Sideslip
Restoring lift
Figure 5-29. Sideslip causing different AOA on each blade.
Figure 5-30. Sweepback wings.
Dihedral
Some aircraft are designed so that the outer tips of the wings
are higher than the wing roots. The upward angle thus formed
by the wings is called dihedral. [Figure 5-28] When a gust
causes a roll, a sideslip will result. This sideslip causes the
relative wind affecting the entire airplane to be from the
direction of the slip. When the relative wind comes from the
side, the wing slipping into the wind is subject to an increase
in AOA and develops an increase in lift. The wing away
from the wind is subject to a decrease in angle of attack, and
develops a decrease in lift. The changes in lift effect a rolling
moment tending to raise the windward wing, hence dihedral
contributes to a stable roll due to sideslip. [Figure 5-29]
Sweepback and Wing Location
Many aspects of an aircraft's configuration can affect its
effective dihedral, but two major components are wing
sweepback and the wing location with respect to the fuselage
(such as a low wing or high wing). As a rough estimation,
10° of sweepback on a wing provides about 1° of effective
dihedral, while a high wing configuration can provide about
5° of effective dihedral over a low wing configuration.
A sweptback wing is one in which the leading edge slopes
backward. [Figure 5-30] When a disturbance causes an
aircraft with sweepback to slip or drop a wing, the low
wing presents its leading edge at an angle that is more
perpendicular to the relative airflow. As a result, the low
wing acquires more lift, rises, and the aircraft is restored to
its original flight attitude.
Keel Effect and Weight Distribution
A high wing aircraft always has the tendency to turn the
longitudinal axis of the aircraft into the relative wind, which
is often referred to as the keel effect. These aircraft are
laterally stable simply because the wings are attached in a
high position on the fuselage, making the fuselage behave like
a keel exerting a steadying influence on the aircraft laterally
about the longitudinal axis. When a high-winged aircraft is
disturbed and one wing dips, the fuselage weight acts like a
pendulum returning the aircraft to the horizontal level.
Laterally stable aircraft are constructed so that the greater
portion of the keel area is above the CG. [Figure 5-31] Thus,
when the aircraft slips to one side, the combination of the
Area
forward
of CG
Area aft of CG
CG
CG
Relative wind
Yaw
yaw
Figure 5-32. Fuselage and fin for directional stability.
CG
CG
CG centerline
Figure 5-31. Keel area for lateral stability.
aircraft’s weight and the pressure of the airflow against the
upper portion of the keel area (both acting about the CG)
tends to roll the aircraft back to wings-level flight.
Directional Stability (Yawing)
Stability about the aircraft’s vertical axis (the sideways
moment) is called yawing or directional stability. Yawing
or directional stability is the most easily achieved stability
in aircraft design. The area of the vertical fin and the sides
of the fuselage aft of the CG are the prime contributors that
make the aircraft act like the well known weather vane or
arrow, pointing its nose into the relative wind.
In examining a weather vane, it can be seen that if exactly the
same amount of surface were exposed to the wind in front
of the pivot point as behind it, the forces fore and aft would
be in balance and little or no directional movement would
result. Consequently, it is necessary to have a greater surface
aft of the pivot point than forward of it.
Similarly, the aircraft designer must ensure positive
directional stability by making the side surface greater aft
than ahead of the CG. [Figure 5-32] To provide additional
positive stability to that provided by the fuselage, a vertical
fin is added. The fin acts similar to the feather on an arrow
in maintaining straight flight. Like the weather vane and the
arrow, the farther aft this fin is placed and the larger its size,
the greater the aircraft’s directional stability.
If an aircraft is flying in a straight line, and a sideward gust
of air gives the aircraft a slight rotation about its vertical
axis (i.e., the right), the motion is retarded and stopped by
the fin because while the aircraft is rotating to the right, the
air is striking the left side of the fin at an angle. This causes
pressure on the left side of the fin, which resists the turning
motion and slows down the aircraft’s yaw. In doing so, it
acts somewhat like the weather vane by turning the aircraft
into the relative wind. The initial change in direction of the
aircraft’s flight path is generally slightly behind its change
of heading. Therefore, after a slight yawing of the aircraft
to the right, there is a brief moment when the aircraft is still
moving along its original path, but its longitudinal axis is
pointed slightly to the right.
The aircraft is then momentarily skidding sideways and,
during that moment (since it is assumed that although the
yawing motion has stopped, the excess pressure on the left
side of the fin still persists), there is necessarily a tendency
for the aircraft to be turned partially back to the left. That is,
there is a momentary restoring tendency caused by the fin.
This restoring tendency is relatively slow in developing and
ceases when the aircraft stops skidding. When it ceases, the
aircraft is flying in a direction slightly different from the
original direction. In other words, it will not return of its own
accord to the original heading; the pilot must reestablish the
initial heading.
A minor improvement of directional stability may be obtained
through sweepback. Sweepback is incorporated in the design
of the wing primarily to delay the onset of compressibility
during high-speed flight. In lighter and slower aircraft,
sweepback aids in locating the center of pressure in the
correct relationship with the CG. A longitudinally stable
aircraft is built with the center of pressure aft of the CG.
Because of structural reasons, aircraft designers sometimes
cannot attach the wings to the fuselage at the exact desired
point. If they had to mount the wings too far forward, and at
right angles to the fuselage, the center of pressure would not
be far enough to the rear to result in the desired amount of
longitudinal stability. By building sweepback into the wings,
however, the designers can move the center of pressure toward
the rear. The amount of sweepback and the position of the
wings then place the center of pressure in the correct location.
When turbulence or rudder application causes the aircraft to
yaw to one side, the opposite wing presents a longer leading
edge perpendicular to the relative airflow. The airspeed of
the forward wing increases and it acquires more drag than
the back wing. The additional drag on the forward wing pulls
the wing back, turning the aircraft back to its original path.
The contribution of the wing to static directional stability is
usually small. The swept wing provides a stable contribution
depending on the amount of sweepback, but the contribution
is relatively small when compared with other components.
Free Directional Oscillations (Dutch Roll)
Dutch roll is a coupled lateral/directional oscillation that is
usually dynamically stable but is unsafe in an aircraft because
of the oscillatory nature. The damping of the oscillatory mode
may be weak or strong depending on the properties of the
particular aircraft.
If the aircraft has a right wing pushed down, the positive
sideslip angle corrects the wing laterally before the nose is
realigned with the relative wind. As the wing corrects the
position, a lateral directional oscillation can occur resulting in
the nose of the aircraft making a figure eight on the horizon as
a result of two oscillations (roll and yaw), which, although of
about the same magnitude, are out of phase with each other.
In most modern aircraft, except high-speed swept wing
designs, these free directional oscillations usually die out
automatically in very few cycles unless the air continues to
be gusty or turbulent. Those aircraft with continuing Dutch
roll tendencies are usually equipped with gyro-stabilized yaw
dampers. Manufacturers try to reach a midpoint between too
much and too little directional stability. Because it is more
desirable for the aircraft to have “spiral instability” than
Dutch roll tendencies, most aircraft are designed with that
characteristic.
Spiral Instability
Spiral instability exists when the static directional stability
of the aircraft is very strong as compared to the effect of its
dihedral in maintaining lateral equilibrium. When the lateral
equilibrium of the aircraft is disturbed by a gust of air and a
sideslip is introduced, the strong directional stability tends
to yaw the nose into the resultant relative wind while the
comparatively weak dihedral lags in restoring the lateral
balance. Due to this yaw, the wing on the outside of the
turning moment travels forward faster than the inside wing
and, as a consequence, its lift becomes greater. This produces
an overbanking tendency which, if not corrected by the pilot,
results in the bank angle becoming steeper and steeper. At
the same time, the strong directional stability that yaws the
aircraft into the relative wind is actually forcing the nose
to a lower pitch attitude. A slow downward spiral begins
which, if not counteracted by the pilot, gradually increases
into a steep spiral dive. Usually the rate of divergence in the
spiral motion is so gradual the pilot can control the tendency
without any difficulty.
Many aircraft are affected to some degree by this characteristic,
although they may be inherently stable in all other normal
parameters. This tendency explains why an aircraft cannot
be flown “hands off” indefinitely.
Much research has gone into the development of control
devices (wing leveler) to correct or eliminate this instability.
The pilot must be careful in application of recovery controls
during advanced stages of this spiral condition or excessive
loads may be imposed on the structure. Improper recovery
from spiral instability leading to inflight structural failures
has probably contributed to more fatalities in general aviation
aircraft than any other factor. Since the airspeed in the spiral
condition builds up rapidly, the application of back elevator
force to reduce this speed and to pull the nose up only
“tightens the turn,” increasing the load factor. The results
of the prolonged uncontrolled spiral are inflight structural
failure, crashing into the ground, or both. Common recorded
causes for pilots who get into this situation are loss of horizon
reference, inability to control the aircraft by reference to
instruments, or a combination of both.
Effect of Wing Planform
Understanding the effects of different wing planforms
is important when learning about wing performance and
airplane flight characteristics. A planform is the shape of the
wing as viewed from directly above and deals with airflow
in three dimensions. Aspect ratio, taper ratio, and sweepback
are factors in planform design that are very important to the
overall aerodynamic characteristic of a wing. [Figure 5-33]
Aspect ratio is the ratio of wing span to wing chord. Taper
ratio can be either in planform or thickness, or both. In its
simplest terms, it is a decrease from wing root to wingtip in
wing chord or wing thickness. Sweepback is the rearward
slant of a wing, horizontal tail, or other airfoil surface.
There are two general means by which the designer can
change the planform of a wing and both will affect the
Elliptical wing Regular wing
Moderate taper wing High taper wing
Pointed tip wing Sweepback wing
Figure 5-33. Different types of wing planforms.
aerodynamic characteristics of the wing. The first is to effect
a change in the aspect ratio. Aspect ratio is the primary factor
in determining the three dimensional characteristics of the
ordinary wing and its lift/drag ratio. An increase in aspect
ratio with constant velocity will decrease the drag, especially
at high angles of attack, improving the performance of the
wing when in a climbing attitude.
A decrease in aspect ratio will give a corresponding increase
in drag. It should be noted, however, that with an increase in
aspect ratio there is an increase in the length of span, with a
corresponding increase in the weight of the wing structure,
which means the wing must be heavier to carry the same load.
For this reason, part of the gain (due to a decrease in drag) is
lost because of the increased weight, and a compromise in