Effect of Weight on Flight Performance
The takeoff/climb and landing performance of an aircraft are
determined on the basis of its maximum allowable takeoff and
landing weights. A heavier gross weight results in a longer
takeoff run and shallower climb, and a faster touchdown
speed and longer landing roll. Even a minor overload may
make it impossible for the aircraft to clear an obstacle that
normally would not be a problem during takeoff under more
favorable conditions.
The detrimental effects of overloading on performance are
not limited to the immediate hazards involved with takeoffs
and landings. Overloading has an adverse effect on all
climb and cruise performance, which leads to overheating
during climbs, added wear on engine parts, increased fuel
consumption, slower cruising speeds, and reduced range.
The manufacturers of modern aircraft furnish weight and
balance data with each aircraft produced. Generally, this
information may be found in the FAA-approved AFM/POH
and easy-to-read charts for determining weight and balance
data are now provided. Increased performance and load-
carrying capability of these aircraft require strict adherence
to the operating limitations prescribed by the manufacturer.
Deviations from the recommendations can result in structural
damage or complete failure of the aircraft’s structure. Even
if an aircraft is loaded well within the maximum weight
limitations, it is imperative that weight distribution be
within the limits of CG location. The preceding brief study
of aerodynamics and load factors points out the reasons for
this precaution. The following discussion is background
information into some of the reasons why weight and balance
conditions are important to the safe flight of an aircraft.
In some aircraft, it is not possible to fill all seats, baggage
compartments, and fuel tanks, and still remain within
approved weight or balance limits. For example, in several
popular four-place aircraft, the fuel tanks may not be filled to
capacity when four occupants and their baggage are carried.
In a certain two-place aircraft, no baggage may be carried
in the compartment aft of the seats when spins are to be
practiced. It is important for a pilot to be aware of the weight
and balance limitations of the aircraft being flown and the
reasons for these limitations.
Effect of Weight on Aircraft Structure
The effect of additional weight on the wing structure of an
aircraft is not readily apparent. Airworthiness requirements
prescribe that the structure of an aircraft certificated in the
normal category (in which acrobatics are prohibited) must
be strong enough to withstand a load factor of 3.8 Gs to take
care of dynamic loads caused by maneuvering and gusts. This
means that the primary structure of the aircraft can withstand
a load of 3.8 times the approved gross weight of the aircraft
without structural failure occurring. If this is accepted as
indicative of the load factors that may be imposed during
operations for which the aircraft is intended, a 100-pound
overload imposes a potential structural overload of 380
pounds. The same consideration is even more impressive
in the case of utility and acrobatic category aircraft, which
have load factor requirements of 4.4 and 6.0, respectively.
Structural failures that result from overloading may be
dramatic and catastrophic, but more often they affect
structural components progressively in a manner that
is difficult to detect and expensive to repair. Habitual
overloading tends to cause cumulative stress and damage
that may not be detected during preflight inspections and
result in structural failure later during completely normal
operations. The additional stress placed on structural parts
by overloading is believed to accelerate the occurrence of
metallic fatigue failures.
A knowledge of load factors imposed by flight maneuvers
and gusts emphasizes the consequences of an increase in the
gross weight of an aircraft. The structure of an aircraft about to
undergo a load factor of 3 Gs, as in recovery from a steep dive,
must be prepared to withstand an added load of 300 pounds
for each 100-pound increase in weight. It should be noted that
this would be imposed by the addition of about 16 gallons
of unneeded fuel in a particular aircraft. FAA-certificated
civil aircraft have been analyzed structurally and tested for
flight at the maximum gross weight authorized and within the
speeds posted for the type of flights to be performed. Flights at
weights in excess of this amount are quite possible and often
are well within the performance capabilities of an aircraft.
This fact should not mislead the pilot, as the pilot may not
realize that loads for which the aircraft was not designed are
being imposed on all or some part of the structure.
In loading an aircraft with either passengers or cargo, the
structure must be considered. Seats, baggage compartments,
and cabin floors are designed for a certain load or
concentration of load and no more. For example, a light
plane baggage compartment may be placarded for 20 pounds
because of the limited strength of its supporting structure
even though the aircraft may not be overloaded or out of CG
limits with more weight at that location.
Effect of Weight on Stability and Controllability
Overloading also affects stability. An aircraft that is stable
and controllable when loaded normally may have very
different flight characteristics when overloaded. Although
the distribution of weight has the most direct effect on this,
an increase in the aircraft’s gross weight may be expected
to have an adverse effect on stability, regardless of location
Down load on tail
Stronger
CG
Down load on tail
Lighter
CG
Center of liftForward CG
AFT CG
Load imposed by tail
Gross weight
Load imposed by tail
Gross weight
Figure 5-63. Effect of load distribution on balance.
of the CG. The stability of many certificated aircraft is
completely unsatisfactory if the gross weight is exceeded.
Effect of Load Distribution
The effect of the position of the CG on the load imposed
on an aircraft’s wing in flight is significant to climb and
cruising performance. An aircraft with forward loading is
“heavier” and consequently, slower than the same aircraft
with the CG further aft.
Figure 5-63 illustrates why this is true. With forward loading,
“nose-up” trim is required in most aircraft to maintain level
cruising flight. Nose-up trim involves setting the tail surfaces
to produce a greater down load on the aft portion of the
fuselage, which adds to the wing loading and the total lift
required from the wing if altitude is to be maintained. This
requires a higher AOA of the wing, which results in more
drag and, in turn, produces a higher stalling speed.
With aft loading and “nose-down” trim, the tail surfaces
exert less down load, relieving the wing of that much wing
loading and lift required to maintain altitude. The required
AOA of the wing is less, so the drag is less, allowing for a
faster cruise speed. Theoretically, a neutral load on the tail
surfaces in cruising flight would produce the most efficient
overall performance and fastest cruising speed, but it would
also result in instability. Modern aircraft are designed to
require a down load on the tail for stability and controllability.
A zero indication on the trim tab control is not necessarily
the same as “neutral trim” because of the force exerted by
downwash from the wings and the fuselage on the tail surfaces.
The effects of the distribution of the aircraft’s useful load
have a significant influence on its flight characteristics, even
when the load is within the CG limits and the maximum
permissible gross weight. Important among these effects
are changes in controllability, stability, and the actual load
imposed on the wing.
Generally, an aircraft becomes less controllable, especially
at slow flight speeds, as the CG is moved further aft. An
aircraft that cleanly recovers from a prolonged spin with
the CG at one position may fail completely to respond to
normal recovery attempts when the CG is moved aft by one
or two inches.
It is common practice for aircraft designers to establish
an aft CG limit that is within one inch of the maximum,
which allows normal recovery from a one-turn spin. When
certificating an aircraft in the utility category to permit
intentional spins, the aft CG limit is usually established
at a point several inches forward of that permissible for
certification in the normal category.
Another factor affecting controllability, which has become
more important in current designs of large aircraft, is the
effect of long moment arms to the positions of heavy
equipment and cargo. The same aircraft may be loaded to
maximum gross weight within its CG limits by concentrating
fuel, passengers, and cargo near the design CG, or by
dispersing fuel and cargo loads in wingtip tanks and cargo
bins forward and aft of the cabin.
With the same total weight and CG, maneuvering the
aircraft or maintaining level flight in turbulent air requires
the application of greater control forces when the load is
dispersed. The longer moment arms to the positions of the
heavy fuel and cargo loads must be overcome by the action
of the control surfaces. An aircraft with full outboard wing
tanks or tip tanks tends to be sluggish in roll when control
situations are marginal, while one with full nose and aft cargo
bins tends to be less responsive to the elevator controls.
The rearward CG limit of an aircraft is determined largely
by considerations of stability. The original airworthiness
requirements for a type certificate specify that an aircraft in
flight at a certain speed dampens out vertical displacement of
the nose within a certain number of oscillations. An aircraft
loaded too far rearward may not do this. Instead, when the
nose is momentarily pulled up, it may alternately climb and
dive becoming steeper with each oscillation. This instability
is not only uncomfortable to occupants, but it could even
become dangerous by making the aircraft unmanageable
under certain conditions.
The recovery from a stall in any aircraft becomes progressively
more difficult as its CG moves aft. This is particularly
important in spin recovery, as there is a point in rearward
loading of any aircraft at which a “flat” spin develops. A
flat spin is one in which centrifugal force, acting through a
CG located well to the rear, pulls the tail of the aircraft out
away from the axis of the spin, making it impossible to get
the nose down and recover.
An aircraft loaded to the rear limit of its permissible CG
range handles differently in turns and stall maneuvers and
has different landing characteristics than when it is loaded
near the forward limit.
The forward CG limit is determined by a number of
considerations. As a safety measure, it is required that the
trimming device, whether tab or adjustable stabilizer, be
capable of holding the aircraft in a normal glide with the power
off. A conventional aircraft must be capable of a full stall,
power-off landing in order to ensure minimum landing speed
in emergencies. A tailwheel-type aircraft loaded excessively
nose-heavy is difficult to taxi, particularly in high winds. It
can be nosed over easily by use of the brakes, and it is difficult
to land without bouncing since it tends to pitch down on the
wheels as it is slowed down and flared for landing. Steering
difficulties on the ground may occur in nosewheel-type
aircraft, particularly during the landing roll and takeoff. The
effects of load distribution are summarized as follows:
• The CG position influences the lift and AOA of the
wing, the amount and direction of force on the tail,
and the degree of deflection of the stabilizer needed
to supply the proper tail force for equilibrium. The
latter is very important because of its relationship to
elevator control force.
• The aircraft stalls at a higher speed with a forward CG
location. This is because the stalling AOA is reached
at a higher speed due to increased wing loading.
• Higher elevator control forces normally exist with a
forward CG location due to the increased stabilizer
deflection required to balance the aircraft.
• The aircraft cruises faster with an aft CG location
because of reduced drag. The drag is reduced because
a smaller AOA and less downward deflection of the
stabilizer are required to support the aircraft and
overcome the nose-down pitching tendency.
• The aircraft becomes less stable as the CG is moved
rearward. This is because when the CG is moved
rearward, it causes a decrease in the AOA. Therefore,
the wing contribution to the aircraft’s stability is
now decreased, while the tail contribution is still
stabilizing. When the point is reached that the wing
and tail contributions balance, then neutral stability
exists. Any CG movement further aft results in an
unstable aircraft.
• A forward CG location increases the need for greater
back elevator pressure. The elevator may no longer
be able to oppose any increase in nose-down pitching.
Adequate elevator control is needed to control the
aircraft throughout the airspeed range down to the stall.
A detailed discussion and additional information relating
to weight and balance can be found in Chapter 10, Weight
and Balance.
High Speed Flight
Subsonic Versus Supersonic Flow
In subsonic aerodynamics, the theory of lift is based upon the
forces generated on a body and a moving gas (air) in which
it is immersed. At speeds of approximately 260 knots or
less, air can be considered incompressible in that, at a fixed
altitude, its density remains nearly constant while its pressure
varies. Under this assumption, air acts the same as water and
is classified as a fluid. Subsonic aerodynamic theory also
assumes the effects of viscosity (the property of a fluid that
tends to prevent motion of one part of the fluid with respect
to another) are negligible and classifies air as an ideal fluid
conforming to the principles of ideal-fluid aerodynamics such
as continuity, Bernoulli’s principle, and circulation.
In reality, air is compressible and viscous. While the effects of
these properties are negligible at low speeds, compressibility
effects in particular become increasingly important as speed
increases. Compressibility (and to a lesser extent viscosity) is
of paramount importance at speeds approaching the speed of
sound. In these speed ranges, compressibility causes a change
in the density of the air around an aircraft.
During flight, a wing produces lift by accelerating the airflow
over the upper surface. This accelerated air can, and does,
reach sonic speeds even though the aircraft itself may be flying
subsonic. At some extreme AOAs, in some aircraft, the speed
of the air over the top surface of the wing may be double the
aircraft’s speed. It is therefore entirely possible to have both
supersonic and subsonic airflow on an aircraft at the same time.
When flow velocities reach sonic speeds at some location on
an aircraft (such as the area of maximum camber on the wing),
further acceleration results in the onset of compressibility
effects, such as shock wave formation, drag increase, buffeting,
stability, and control difficulties. Subsonic flow principles are
invalid at all speeds above this point. [Figure 5-64]
Speed Ranges
The speed of sound varies with temperature. Under standard
temperature conditions of 15 °C, the speed of sound at sea
level is 661 knots. At 40,000 feet, where the temperature is
–55 °C, the speed of sound decreases to 574 knots. In high-
Possible separation
M = 0.50
M = 0.72
M = 0.77
Maximum local velocity is less than sonic
Critical mach number
Maximum local velocity equal to sonic
Normal shock wave
Supersonic
flow
Subsonic
Figure 5-64. Wing airflow.
CD (Drag coefficient)
M (Mach number)
0.5 1.0
CL= 0.3
Force divergence Mach number
Critical Mach number
Figure 5-65. Critical Mach.
speed flight and/or high-altitude flight, the measurement of
speed is expressed in terms of a “Mach number”—the ratio
of the true airspeed of the aircraft to the speed of sound in
the same atmospheric conditions. An aircraft traveling at
the speed of sound is traveling at Mach 1.0. Aircraft speed
regimes are defined approximately as follows:
Subsonic—Mach numbers below 0.75
Transonic—Mach numbers from 0.75 to 1.20
Supersonic—Mach numbers from 1.20 to 5.00
Hypersonic—Mach numbers above 5.00
While flights in the transonic and supersonic ranges are
common occurrences for military aircraft, civilian jet aircraft
normally operate in a cruise speed range of Mach 0.7 to
Mach 0.90.
The speed of an aircraft in which airflow over any part of
the aircraft or structure under consideration first reaches
(but does not exceed) Mach 1.0 is termed “critical Mach
number” or “Mach Crit.” Thus, critical Mach number is
the boundary between subsonic and transonic flight and is
largely dependent on the wing and airfoil design. Critical
Mach number is an important point in transonic flight. When
shock waves form on the aircraft, airflow separation followed
by buffet and aircraft control difficulties can occur. Shock
waves, buffet, and airflow separation take place above critical
Mach number. A jet aircraft typically is most efficient when
cruising at or near its critical Mach number. At speeds 5–10
percent above the critical Mach number, compressibility
effects begin. Drag begins to rise sharply. Associated with
the “drag rise” are buffet, trim, and stability changes and a
decrease in control surface effectiveness. This is the point
of “drag divergence.” [Figure 5-65]
VMO/MMO is defined as the maximum operating limit speed.
VMO is expressed in knots calibrated airspeed (KCAS), while
MMO is expressed in Mach number. The VMO limit is usually
associated with operations at lower altitudes and deals with
structural loads and flutter. The MMO limit is associated with
operations at higher altitudes and is usually more concerned
with compressibility effects and flutter. At lower altitudes,
structural loads and flutter are of concern; at higher altitudes,
compressibility effects and flutter are of concern.
Adherence to these speeds prevents structural problems due
to dynamic pressure or flutter, degradation in aircraft control
response due to compressibility effects (e.g., Mach Tuck,
aileron reversal, or buzz), and separated airflow due to shock
waves resulting in loss of lift or vibration and buffet. Any of
these phenomena could prevent the pilot from being able to
adequately control the aircraft.
For example, an early civilian jet aircraft had a VMO limit of
306 KCAS up to approximately FL 310 (on a standard day).
At this altitude (FL 310), an MMO of 0.82 was approximately
equal to 306 KCAS. Above this altitude, an M MO of 0.82
always equaled a KCAS less than 306 KCAS and, thus,
became the operating limit as you could not reach the V MO
limit without first reaching the M MO limit. For example, at
FL 380, an MMO of 0.82 is equal to 261 KCAS.
Mach Number Versus Airspeed
It is important to understand how airspeed varies with Mach
number. As an example, consider how the stall speed of a
jet transport aircraft varies with an increase in altitude. The
increase in altitude results in a corresponding drop in air
density and outside temperature. Suppose this jet transport
is in the clean configuration (gear and flaps up) and weighs
550,000 pounds. The aircraft might stall at approximately 152
KCAS at sea level. This is equal to (on a standard day) a true
velocity of 152 KTAS and a Mach number of 0.23. At FL 380,
the aircraft will still stall at approximately 152 KCAS, but the
true velocity is about 287 KTAS with a Mach number of 0.50.
Although the stalling speed has remained the same for our
purposes, both the Mach number and TAS have increased.
With increasing altitude, the air density has decreased; this
requires a faster true airspeed in order to have the same
pressure sensed by the pitot tube for the same KCAS, or KIAS
(for our purposes, KCAS and KIAS are relatively close to
each other). The dynamic pressure the wing experiences at
FL 380 at 287 KTAS is the same as at sea level at 152 KTAS.
However, it is flying at higher Mach number.
Another factor to consider is the speed of sound. A decrease
in temperature in a gas results in a decrease in the speed of
sound. Thus, as the aircraft climbs in altitude with outside
temperature dropping, the speed of sound is dropping. At
sea level, the speed of sound is approximately 661 KCAS,
while at FL 380 it is 574 KCAS. Thus, for our jet transport
aircraft, the stall speed (in KTAS) has gone from 152 at sea
level to 287 at FL 380. Simultaneously, the speed of sound
(in KCAS) has decreased from 661 to 574 and the Mach
number has increased from 0.23 (152 KTAS divided by 661
KTAS) to 0.50 (287 KTAS divided by 574 KTAS). All the
while, the KCAS for stall has remained constant at 152. This
describes what happens when the aircraft is at a constant
KCAS with increasing altitude, but what happens when the
pilot keeps Mach constant during the climb? In normal jet
flight operations, the climb is at 250 KIAS (or higher (e.g.
heavy)) to 10,000 feet and then at a specified en route climb
airspeed (about 330 if a DC10) until reaching an altitude in
the “mid-twenties” where the pilot then climbs at a constant
Mach number to cruise altitude.
Assuming for illustration purposes that the pilot climbs at a
MMO of 0.82 from sea level up to FL 380. KCAS goes from
543 to 261. The KIAS at each altitude would follow the
same behavior and just differ by a few knots. Recall from
the earlier discussion that the speed of sound is decreasing
with the drop in temperature as the aircraft climbs. The Mach
number is simply the ratio of the true airspeed to the speed
of sound at flight conditions. The significance of this is that
at a constant Mach number climb, the KCAS (and KTAS or
KIAS as well) is falling off.
If the aircraft climbed high enough at this constant M MO
with decreasing KIAS, KCAS, and KTAS, it would begin to
approach its stall speed. At some point, the stall speed of the
aircraft in Mach number could equal the MMO of the aircraft,
and the pilot could neither slow down (without stalling) nor
speed up (without exceeding the max operating speed of the
aircraft). This has been dubbed the “coffin corner.”
Boundary Layer
The viscous nature of airflow reduces the local velocities on
a surface and is responsible for skin friction. As discussed
earlier in the chapter, the layer of air over the wing’s surface
that is slowed down or stopped by viscosity is the boundary
layer. There are two different types of boundary layer flow:
laminar and turbulent.
Laminar Boundary Layer Flow
The laminar boundary layer is a very smooth flow, while
the turbulent boundary layer contains swirls or eddies.
The laminar flow creates less skin friction drag than the
turbulent flow but is less stable. Boundary layer flow over a
wing surface begins as a smooth laminar flow. As the flow
continues back from the leading edge, the laminar boundary
layer increases in thickness.
Turbulent Boundary Layer Flow
At some distance back from the leading edge, the smooth
laminar flow breaks down and transitions to a turbulent flow.
From a drag standpoint, it is advisable to have the transition
from laminar to turbulent flow as far aft on the wing as
possible or have a large amount of the wing surface within
the laminar portion of the boundary layer. The low energy
laminar flow, however, tends to break down more suddenly
than the turbulent layer.
Boundary Layer Separation
Another phenomenon associated with viscous flow is
separation. Separation occurs when the airflow breaks away
from an airfoil. The natural progression is from laminar
boundary layer to turbulent boundary layer and then to
airflow separation. Airflow separation produces high drag
and ultimately destroys lift. The boundary layer separation
point moves forward on the wing as the AOA is increased.
[Figure 5-66]
Vortex generators are used to delay or prevent shock wave
induced boundary layer separation encountered in transonic
flight. They are small low aspect ratio airfoils placed at a 12°
to 15° AOA to the airstream. Usually spaced a few inches
apart along the wing ahead of the ailerons or other control
surfaces, vortex generators create a vortex that mixes the
boundary airflow with the high energy airflow just above the
surface. This produces higher surface velocities and increases
the energy of the boundary layer. Thus, a stronger shock wave
is necessary to produce airflow separation.
Shock Waves
When an airplane flies at subsonic speeds, the air ahead is
“warned” of the airplane’s coming by a pressure change
transmitted ahead of the airplane at the speed of sound.
Because of this warning, the air begins to move aside before
the airplane arrives and is prepared to let it pass easily. When
the airplane’s speed reaches the speed of sound, the pressure
Transition region
Laminar sublayer
Laminar boundary layer
Turbulent boundary layer
Figure 5-66. Boundary layer.
M = 0.82
M = 0.95
M = 1.05
Normal shock wave
Supersonic flow
Normal shock wave
Supersonic flow
Subsonic airflow
Bow wave
Figure 5-67. Shock waves.
change can no longer warn the air ahead because the airplane
is keeping up with its own pressure waves. Rather, the air
particles pile up in front of the airplane causing a sharp
decrease in the flow velocity directly in front of the airplane
with a corresponding increase in air pressure and density.
As the airplane’s speed increases beyond the speed of sound,
the pressure and density of the compressed air ahead of it
increase, the area of compression extending some distance
ahead of the airplane. At some point in the airstream, the air
particles are completely undisturbed, having had no advanced
warning of the airplane’s approach, and in the next instant the
same air particles are forced to undergo sudden and drastic
changes in temperature, pressure, density, and velocity.
The boundary between the undisturbed air and the region
of compressed air is called a shock or “compression” wave.
This same type of wave is formed whenever a supersonic
airstream is slowed to subsonic without a change in direction,
such as when the airstream is accelerated to sonic speed
over the cambered portion of a wing, and then decelerated
to subsonic speed as the area of maximum camber is passed.
A shock wave forms as a boundary between the supersonic
and subsonic ranges.
Whenever a shock wave forms perpendicular to the airflow, it
is termed a “normal” shock wave, and the flow immediately
behind the wave is subsonic. A supersonic airstream passing
through a normal shock wave experiences these changes:
• The airstream is slowed to subsonic.
• The airflow immediately behind the shock wave does
not change direction.
• The static pressure and density of the airstream behind
the wave is greatly increased.
• The energy of the airstream (indicated by total
pressure—dynamic plus static) is greatly reduced.
Shock wave formation causes an increase in drag. One of
the principal effects of a shock wave is the formation of a
dense high pressure region immediately behind the wave.
The instability of the high pressure region, and the fact that
part of the velocity energy of the airstream is converted to
heat as it flows through the wave, is a contributing factor
in the drag increase, but the drag resulting from airflow
separation is much greater. If the shock wave is strong,
the boundary layer may not have sufficient kinetic energy
to withstand airflow separation. The drag incurred in the
transonic region due to shock wave formation and airflow
separation is known as “wave drag.” When speed exceeds
the critical Mach number by about 10 percent, wave drag
increases sharply. A considerable increase in thrust (power)
is required to increase flight speed beyond this point into the
supersonic range where, depending on the airfoil shape and
the AOA, the boundary layer may reattach.
Normal shock waves form on the wing’s upper surface and
form an additional area of supersonic flow and a normal shock
wave on the lower surface. As flight speed approaches the
speed of sound, the areas of supersonic flow enlarge and the
shock waves move nearer the trailing edge. [Figure 5-67]
Spanwise flow
True airspeed
Mach 0.85
Airspeed sensed
by wing Mach 0.70
Figure 5-68. Sweepback effect.
Associated with “drag rise” are buffet (known as Mach
buffet), trim, and stability changes and a decrease in control
force effectiveness. The loss of lift due to airflow separation
results in a loss of downwash and a change in the position of
the center pressure on the wing. Airflow separation produces
a turbulent wake behind the wing, which causes the tail
surfaces to buffet (vibrate). The nose-up and nose-down pitch
control provided by the horizontal tail is dependent on the
downwash behind the wing. Thus, an increase in downwash
decreases the horizontal tail’s pitch control effectiveness
since it effectively increases the AOA that the tail surface is
seeing. Movement of the wing CP affects the wing pitching
moment. If the CP moves aft, a diving moment referred to
as “Mach tuck” or “tuck under” is produced, and if it moves
forward, a nose-up moment is produced. This is the primary
reason for the development of the T-tail configuration on
many turbine-powered aircraft, which places the horizontal
stabilizer as far as practical from the turbulence of the wings.
Sweepback
Most of the difficulties of transonic flight are associated with
shock wave induced flow separation. Therefore, any means of
delaying or alleviating the shock induced separation improves
aerodynamic performance. One method is wing sweepback.
Sweepback theory is based upon the concept that it is only the
component of the airflow perpendicular to the leading edge
of the wing that affects pressure distribution and formation
of shock waves. [Figure 5-68]
On a straight wing aircraft, the airflow strikes the wing
leading edge at 90°, and its full impact produces pressure and
lift. A wing with sweepback is struck by the same airflow at
an angle smaller than 90°. This airflow on the swept wing has
the effect of persuading the wing into believing that it is flying
slower than it really is; thus the formation of shock waves is
delayed. Advantages of wing sweep include an increase in
critical Mach number, force divergence Mach number, and
the Mach number at which drag rise peaks. In other words,
sweep delays the onset of compressibility effects.
The Mach number that produces a sharp change in coefficient
of drag is termed the “force divergence” Mach number and,
for most airfoils, usually exceeds the critical Mach number by
5 to 10 percent. At this speed, the airflow separation induced
by shock wave formation can create significant variations in
the drag, lift, or pitching moment coefficients. In addition to
the delay of the onset of compressibility effects, sweepback
reduces the magnitude in the changes of drag, lift, or moment
coefficients. In other words, the use of sweepback “softens”
the force divergence.
A disadvantage of swept wings is that they tend to stall at the
wingtips rather than at the wing roots. [Figure 5-69] This is
because the boundary layer tends to flow spanwise toward
the tips and to separate near the leading edges. Because the
tips of a swept wing are on the aft part of the wing (behind
the CL), a wingtip stall causes the CL to move forward on
the wing, forcing the nose to rise further. The tendency for
tip stall is greatest when wing sweep and taper are combined.
The stall situation can be aggravated by a T-tail configuration,
which affords little or no pre-stall warning in the form of tail
control surface buffet. [Figure 5-70] The T-tail, being above
the wing wake remains effective even after the wing has begun
to stall, allowing the pilot to inadvertently drive the wing
into a deeper stall at a much greater AOA. If the horizontal
tail surfaces then become buried in the wing’s wake, the
elevator may lose all effectiveness, making it impossible to
reduce pitch attitude and break the stall. In the pre-stall and
immediate post-stall regimes, the lift/drag qualities of a swept
wing aircraft (specifically the enormous increase in drag
at low speeds) can cause an increasingly descending flight
path with no change in pitch attitude, further increasing the
Stalled
Pre-stall
Figure 5-69. Wingtip pre-stall.
Figure 5-70. T-tail stall.
Stalled
Prestall
AOA. In this situation, without reliable AOA information,
a nose-down pitch attitude with an increasing airspeed is no
guarantee that recovery has been affected, and up-elevator
movement at this stage may merely keep the aircraft stalled.
It is a characteristic of T-tail aircraft to pitch up viciously
when stalled in extreme nose-high attitudes, making
recovery difficult or violent. The stick pusher inhibits this
type of stall. At approximately one knot above stall speed,
pre-programmed stick forces automatically move the stick
forward, preventing the stall from developing. A G-limiter
may also be incorporated into the system to prevent the pitch
down generated by the stick pusher from imposing excessive
loads on the aircraft. A “stick shaker,” on the other hand,
provides stall warning when the airspeed is five to seven
percent above stall speed.
Mach Buffet Boundaries
Mach buffet is a function of the speed of the airflow over the
wing—not necessarily the speed of the aircraft. Any time that
too great a lift demand is made on the wing, whether from too
fast an airspeed or from too high an AOA near the MMO, the
“high-speed” buffet occurs. There are also occasions when
the buffet can be experienced at much lower speeds known
as the “low-speed Mach buffet.”
An aircraft flown at a speed too slow for its weight and
altitude necessitating a high AOA is the most likely situation
to cause a low-speed Mach buffet. This very high AOA has
the effect of increasing airflow velocity over the upper surface
of the wing until the same effects of the shock waves and
buffet occur as in the high-speed buffet situation. The AOA
of the wing has the greatest effect on inducing the Mach
buffet at either the high-speed or low-speed boundaries for
the aircraft. The conditions that increase the AOA, the speed
of the airflow over the wing, and chances of Mach buffet are:
• High altitudes—the higher an aircraft flies, the thinner
the air and the greater the AOA required to produce
the lift needed to maintain level flight.
• Heavy weights—the heavier the aircraft, the greater
the lift required of the wing, and all other factors being
equal, the greater the AOA.
• G loading—an increase in the G loading on the aircraft
has the same effect as increasing the weight of the
aircraft. Whether the increase in G forces is caused
by turns, rough control usage, or turbulence, the effect
of increasing the wing’s AOA is the same.
High Speed Flight Controls
On high-speed aircraft, flight controls are divided into
primary flight controls and secondary or auxiliary flight
controls. The primary flight controls maneuver the aircraft
about the pitch, roll, and yaw axes. They include the ailerons,
elevator, and rudder. Secondary or auxiliary flight controls
include tabs, leading edge flaps, trailing edge flaps, spoilers,
and slats.
Spoilers are used on the upper surface of the wing to spoil or
reduce lift. High speed aircraft, due to their clean low drag
design, use spoilers as speed brakes to slow them down.
Spoilers are extended immediately after touchdown to dump
lift and thus transfer the weight of the aircraft from the wings
onto the wheels for better braking performance. [Figure 5-71]
Jet transport aircraft have small ailerons. The space for
ailerons is limited because as much of the wing trailing
edge as possible is needed for flaps. Also, a conventional
size aileron would cause wing twist at high speed. For that
reason, spoilers are used in unison with ailerons to provide
additional roll control.
Some jet transports have two sets of ailerons, a pair of outboard
low-speed ailerons and a pair of high-speed inboard ailerons.
When the flaps are fully retracted after takeoff, the outboard
ailerons are automatically locked out in the faired position.
When used for roll control, the spoiler on the side of the
up-going aileron extends and reduces the lift on that side,
causing the wing to drop. If the spoilers are extended as speed
brakes, they can still be used for roll control. If they are the
Figure 5-71. Control surfaces.
737 Control Surfaces
727 Control Surfaces
Takeoff Flaps
Flaps Retracted
Landing Flaps
Aileron
Leading edge slat
Leading edge flap
Foreflap Midflap Aftflap
Inboard wing
Outboard wing
Inboard wing
Outboard wing
Foreflap
Midflap
Aftflap
Aileron
Inboard wing
Outboard wing
Aileron
Rudder
Elevator
Tab
Stabilizer
Inboard flap
Flight spoilers
Ground spoiler
Outboard flap
Tab
Aileron
Leading edge slats
Leading edge flaps
Stabilizer
Elevator
Control tab
Upper rudder
Anti-balance tabs
Control tab
Inboard aileron
Leading edge flaps
Lower rudder
Vortex generators
Pitot tubes
Ground spoilers
Inboard flaps
Flight spoilers
Outboard flap
Balance tab
Outboard aileron
Slats
Fence
differential type, they extend further on one side and retract
on the other side. If they are the non-differential type, they
extend further on one side but do not retract on the other side.
When fully extended as speed brakes, the non-differential
spoilers remain extended and do not supplement the ailerons.
To obtain a smooth stall and a higher AOA without airflow
separation, the wing’s leading edge should have a well-
rounded almost blunt shape that the airflow can adhere to
at the higher AOA. With this shape, the airflow separation
starts at the trailing edge and progresses forward gradually
as AOA is increased.
The pointed leading edge necessary for high-speed flight
results in an abrupt stall and restricts the use of trailing edge
flaps because the airflow cannot follow the sharp curve
around the wing leading edge. The airflow tends to tear loose
rather suddenly from the upper surface at a moderate AOA.
To utilize trailing edge flaps, and thus increase the CL-MAX,
the wing must go to a higher AOA without airflow separation.
Therefore, leading edge slots, slats, and flaps are used to
improve the low-speed characteristics during takeoff, climb,
and landing. Although these devices are not as powerful as
trailing edge flaps, they are effective when used full span in
combination with high-lift trailing edge flaps. With the aid
of these sophisticated high-lift devices, airflow separation is
delayed and the CL-MAX is increased considerably. In fact, a
50 knot reduction in stall speed is not uncommon.
The operational requirements of a large jet transport aircraft
necessitate large pitch trim changes. Some requirements are:
• A large CG range
• A large speed range
• The ability to perform large trim changes due to
wing leading edge and trailing edge high-lift devices
without limiting the amount of elevator remaining
• Maintaining trim drag to a minimum
These requirements are met by the use of a variable incidence
horizontal stabilizer. Large trim changes on a fixed-tail
aircraft require large elevator deflections. At these large
deflections, little further elevator movement remains in the
same direction. A variable incidence horizontal stabilizer
is designed to take out the trim changes. The stabilizer is
larger than the elevator, and consequently does not need to
be moved through as large an angle. This leaves the elevator
streamlining the tail plane with a full range of movement up
and down. The variable incidence horizontal stabilizer can
be set to handle the bulk of the pitch control demand, with
the elevator handling the rest. On aircraft equipped with a
variable incidence horizontal stabilizer, the elevator is smaller
and less effective in isolation than it is on a fixed-tail aircraft.
In comparison to other flight controls, the variable incidence
horizontal stabilizer is enormously powerful in its effect.
Because of the size and high speeds of jet transport aircraft,
the forces required to move the control surfaces can be beyond
the strength of the pilot. Consequently, the control surfaces are
actuated by hydraulic or electrical power units. Moving the
controls in the flight deck signals the control angle required,
and the power unit positions the actual control surface. In the
event of complete power unit failure, movement of the control
surface can be affected by manually controlling the control
tabs. Moving the control tab upsets the aerodynamic balance,
which causes the control surface to move.
Chapter Summary
In order to sustain an aircraft in flight, a pilot must understand
how thrust, drag, lift, and weight act on the aircraft. By
understanding the aerodynamics of flight, how design,
weight, load factors, and gravity affect an aircraft during
flight maneuvers from stalls to high speed flight, the pilot
learns how to control the balance between these forces. For
information on stall speeds, load factors, and other important
aircraft data, always consult the AFM/POH for specific
information pertaining to the aircraft being flown.