On hearing the sound of an aircraft, few people can resist the temptation to look to the sky.
Many will wonder how something so big and heavy as an aircraft can stay in the air.
To the casual observer this question is one of curiosity, but to the pilot it's essential knowledge.
The answer lies partly in the air that the aircraft flies through. To be precise, this air is called the
atmosphere.
An understanding of the earth's atmosphere is fundamental to a pilot, since it's not only the air that we breathe,
but the environment in which an aircraft flies. There are many physical processes at work within the earth's atmosphere,
but for now let's examine both what the atmosphere is and what its main properties are.
An atmosphere is defined as being the gaseous envelope surrounding a celestial body.
In more simple words, our atmosphere is a collection of gases that is held around the earth by gravity.
There are well over 20 of these gases, some more abundant than others. We'll concentrate on the four most abundant
gases.
By far the most common gas is nitrogen, which accounts for about 78% of the atmosphere.
Oxygen makes up 21% of the atmosphere. Oxygen is essential to sustaining life and is required to allow the
combustion of fuel to take place in aircraft engines.
Just less than 1% of the atmosphere is argon, with carbon dioxide making up most of the remainder.
There are many more gases making up the remainder of the atmosphere, but these are found in such small quantities
that we regard them as trace gases.
These include carbon monoxide, helium, methane, ozone, and hydrogen.
There is, however, one other gas that can be found in our atmosphere which is important to the performance of
an aircraft. This is water vapour.
Although it's invisible as a gas, it can be seen in its liquid or solid state.
Clouds and rain are examples of its liquid state, whereas snow, ice, and hail are examples of its solid state.
The amount of water vapour present in the air is expressed as a percentage of the amount of water vapour
in the air compared with the amount of vapour needed to make the air saturated at the air's current temperature.
This is known as relative humidity. 100% humidity would be visible as a cloud.
The relative amounts of the gases in the atmosphere stay fairly constant up to about 60km above the earth. Thereafter,
gravitational separation alters the composition of the atmosphere.
For the purposes of the principles of flight, we only consider the lower parts of the atmosphere, so we may
assume the constituents remain constant.
The atmosphere is structured into four concentric segments or layers. Starting with the nearest to the earth, these are
The troposphere, which is approximately 11km above the earth. The stratosphere then extends to approximately 50km.
From then on, the mesosphere continues up to 80km. And finally, the thermosphere.
Between the troposphere and the stratosphere there is a boundary known as the tropopause. The height of the tropopause varies,
but averages out at 36,090ft.
Aircraft are limited to flight in the troposphere and the stratosphere.
Only rocket research aircraft and spaceships fly in the mesosphere and thermosphere. These layers of the atmosphere are determined by
the change in temperature they experience with increasing altitude.
The earth is heated by the sun. Being closest to the earth, the troposphere is then heated by the earth.
With increasing altitude, the atmosphere is heated less and less, and its temperature drops as it gets further away from
the source of heat.
We'll only concern ourselves with the effect of temperature on the troposphere and stratosphere, because, as we've already stated, it's
within these layers that aircraft fly.
The temperature remains constant from the tropopause through the lower parts of the stratosphere before rising again.
As the holder of a private pilot's licence, you'll generally fly at altitudes of less than 10,000ft.
You'll be restricted in altitude by a number of factors – the weather, aircraft performance, and, more importantly, the need
to breathe.
We've already stated that the relative amounts of the gases in the atmosphere are constant.
Our bodies need to breathe some oxygen in order for us to survive, but to get the oxygen into our
lungs, we must consider another property of the atmosphere, which is that of atmospheric pressure.
One of the problems that people have when they're discussing the atmosphere, or the air, as it's generally referred to,
is that we cannot see it.
If it was visible to us, then these next few statements would be easier to accept.
Air has substance. It's made up of molecules, and they have mass.
Admittedly not a lot when compared to, say, a piece of wood, but nevertheless it does possess these qualities.
It's not the intention of these lessons to make you into a physicist, but a few technical definitions and formula
are required in order to allow us to study this subject.
Pressure is defined as the force acting on an area, or force per unit area.
A force is defined as a push or a pull, and it will be produced when a mass is accelerated.
The units of measurement of mass and acceleration are those stated in the Systeme Internationale, commonly referred to as the
SI system.
The units of measurement then are the kilogram for mass and the metre per second squared for acceleration.
The product of mass times acceleration then is force, and force is measured in newtons.
The mass of the air, the atmosphere, is accelerated towards the Earth's surface by gravity. This of course gives us
a force.
If we considered that force acting on an area of one square metre, then what we have is the force
exerted by a column of air one metre square extending up into the atmosphere.
This gives us force per square metre, which is termed pressure. This pressure acting on the Earth's surface is called
the atmospheric pressure.
The SI pressure unit is newtons per square metre. One newton per square metre is called the Pascal. The Pascal
is a very small unit, so in aviation the hectopascal is used.
Hecto means one hundred, so a hectopascal is one hundred Pascals.
Just to add to the confusion, one hectopascal is also known as a millibar. You're also likely to hear pressure
referred to as being measured in pounds per square inch or inches of mercury.
These are not SI units and consequently will not be used in this study of the principles of flight.
We've only looked at one column of air, but of course each metre of the Earth's surface will support a
similar column.
The air pressure acts in all directions, and if we were to place an aircraft within the columns of air,
it too would have air pressure acting on it.
The atmospheric pressure that acts on anything immersed into the air is also known as static pressure.
This is the term that will be commonly used in these principles of flight lessons.
It was explained previously that the pressure in the atmosphere was due to gravity accelerating the mass of the air
towards the Earth's surface.
This will lead to an increase in the number of air molecules as we get nearer to the surface of
the Earth.
The larger the number of molecules in a substance, the greater its mass.
This means that the pressure is not constant as we move away from the surface of the Earth.
So, static pressure reduces with altitude.
It's for this reason that breathing is difficult at altitude, there being less pressure to force oxygen into our lungs.
It will also reduce the flow of oxygen into an internal combustion engine, which is propelling an aircraft through the
air.
This fact requires us to consider the question of how many molecules are there in certain levels of the atmosphere.
This property is known as air density.
Air density is defined as the amount of a mass of a substance in a known volume of that substance.
This is technically stated as mass per unit volume.
With regard to the air, we specify the volume as being one cubic metre.
This gives the SI unit of density as being the kilogram per cubic metre.
Similarly to pressure, density also reduces with altitude.
The density of the air is also affected by its temperature.
An increase in air temperature tends to reduce the density of the air, and a reduction in temperature will increase
its density.
Earlier we stated that both pressure and temperature reduce with an increase in altitude.
However, it's the effect of pressure that has the greatest effect on density, and as we've just said, density reduces
with altitude.
The amount of water vapour in the atmosphere will also affect density.
It's the mass of air in each cubic metre we consider, not the amount of water.
If there is water in the cubic metre, as a consequence, there must be a reduction in the amount of
air in it.
The way in which the properties of the atmosphere change greatly affects the way in which an aircraft flies.
In 1964, the International Civil Aviation Organisation, IKO, defined the properties of this average standard atmosphere.
This is known as the IKO standard atmosphere, or ISA.
ISA assumes that at mean sea level the temperature is 15°C, the pressure is 1013.25 hPa, and the density
is 1.225 kg per cubic metre.
From mean sea level to 11 kilometres, or 36,090 feet, the temperature decreases by 1.98°C per 1,000
feet, or 0.65°C per 100 metres.
For quick mental calculations, this is generally rounded up to 2°C per 1,000 feet.
From 11 kilometres to 20 kilometres, the temperature remains constant, at minus 56.5°C.
Notice that at 11 kilometres the temperature no longer falls with height, and this therefore marks the top of the
troposphere.
You may recall that we called this point the tropopause.
The layer above this is called the stratosphere, and because it's well above the height at which a private pilot
normally flies, it's not the intention in this subject to look beyond the stratosphere.
The meteorology lessons in this series of CD-ROMs will however cover in detail the changes which occur in the
atmosphere above these layers.
This concludes the lesson on the atmosphere. ISA is only one of several types of standard atmosphere, but it is
the most recognised, and the one we shall refer to during the following lessons.
It must be stressed however that ISA does not refer to the actual atmospheric conditions of the day.
Having examined the properties of the air in lesson 1, the atmosphere, we'll now consider the principles of aerodynamics.
Aerodynamics is the study of the power of the air, both when it's at rest and in motion.
For the issue of the private pilot's licence, we'll be examining the physical properties of air that is either static,
or moving at speeds below the speed of sound.
The speed of sound at sea level in ISA conditions is approximately 660 knots, and air moving at this speed
is said to be sonic.
We'll be considering speeds lower than sonic, which are termed subsonic.
The aircraft in which you will do your initial flying training will probably cruise at speeds of around 100 knots.
This is well below the speed of sound.
The fact that we cannot see or touch the air makes it very difficult for us to imagine that it
does move.
In the lesson on the atmosphere, we already considered the density of air as being that mass contained within a
volume of one cubic metre.
Therefore, in the following graphics, we'll view the movement of air as a one metre cubed box, shaded to stand
out against the surrounding atmosphere.
The air, you'll recall, has mass. The mass of one cubic metre of air at sea level is 1.225
kilograms.
If the air is moving, it will also possess energy, and will exert a pressure on any object in its
path.
Again, this is often a difficult concept to accept when we can't see the air, but take a normal one
kilogram bag of sugar and throw it against a wall.
We can be in no doubt upon hearing the thud of the bag as it hits the wall that a
force has acted against it.
Air behaves in exactly the same way.
The term body refers to any object that gets in the way of the air flow. In our case, that
body is an aircraft.
You will recall that static pressure is already acting on the aircraft, and now if the air is moving, the
aircraft will also be subjected to dynamic pressure.
When a body like an aircraft is subjected to both static and dynamic pressure, it's said to have a total
pressure acting on it.
This total pressure is the sum of both static and dynamic pressure.
You may also hear total pressure referred to as pitot, or stagnation pressure.
The term total pressure will be used throughout these lessons on the principles of flight.
In reality, it's not the air that moves, but the aircraft which moves through the air.
This relative airflow, as it's called, is always parallel to, and opposite to, the path of the aircraft through the
air.
It's important to note that the direction of the flight path of the aircraft is often not the direction in
which its nose is pointing.
It makes no difference if it's the air that moves, or the aircraft. The end result will be the same.
There will be a dynamic pressure.
As a pilot, you need to know the value of the dynamic pressure, as it's fundamental to how an aircraft
flies.
Dynamic pressure cannot be measured on its own because static pressure will always be present.
But, mathematically, it is easy to resolve the value of dynamic pressure by simply rearranging the formula.
Static pressure plus dynamic pressure equals total pressure.
A simple transposition of the formula shows us that total pressure minus static pressure equals dynamic pressure.
In the aircraft, it's the task of the pitot static system to resolve the problem of how to separate static
pressure from dynamic pressure.
This system pneumatically subtracts static pressure from total pressure to deliver dynamic pressure alone to the airspeed indicator,
which is calibrated to show the pilot is indicated airspeed, or IAS. Pressure instruments can be studied elsewhere on this
disk.
How significant dynamic pressure is to the understanding of the principles of flight cannot be over-emphasised.
For a conventional aircraft to fly, it must have air moving over it, and it will of course then be
subjected to dynamic pressure.
Let's assume that to fly, a certain aircraft needs to be moving through the air at 60 nautical miles per
hour.
60 nautical miles per hour is commonly termed a speed of 60 knots, the knot being the unit of measurement
of speed normally used in aircraft.
As the aircraft accelerates along the runway, the dynamic pressure acting on it will increase because of its increasing speed.
Now let's apply some of the knowledge we've learnt so far.
Dynamic pressure equals half rho, the density, multiplied by the velocity squared.
At take-off from an airfield at sea level, with ISA conditions prevailing, the dynamic pressure formula would give us
the figures shown on the screen.
You'll see from the formula that as we're working in SI units, we've converted 60 knots into 31 metres per
second.
The formula gives us a dynamic pressure of 589 Newtons per square metre.
We can see therefore that at the take-off airspeed of 60 knots, a force of 589 Newtons, that is
60 kilograms force or 132 pounds force, is acting on every square metre of the aircraft's frontal surface.
589 Newtons per square metre may also be expressed in SI units as 589 pascals, or 5.89 hectopascals, which
is the normal unit of pressure in the SI system.
Naturally, the mathematics of dynamic pressure need not be considered by the pilot every time he takes off.
The pilot simply reads the dynamic pressure as an airspeed on his airspeed indicator. In this present example, 60 knots.
The indicated airspeed gauge is giving you all the information you require, but it should be appreciated that this instrument
is hardly ever measuring the true speed of the aircraft through the air.
It's measuring the dynamic pressure being exerted by the airflow.
Because dynamic pressure is dependent upon both the air density and the speed of the aircraft, it's necessary to recall
that density changes with pressure, temperature and humidity.
It's for this reason that the only occasion that the true airspeed of the aircraft would be displayed on the
indicated airspeed gauge is when the atmospheric conditions correspond to those of ISA sea level.
If the density were to reduce due to an increase in temperature or an increase in altitude, to maintain the
same dynamic pressure then the true airspeed would have to increase.
Changes in true airspeed are of concern for navigation.
In the study of the principles of flight, it's the dynamic pressure or indicated airspeed that we are interested in.
Just as thick tomato ketchup is slow to flow out of its bottle, so thick dense air is reluctant to
flow.
This affects the way the air will flow around a body.
The measure of a fluid's resistance to flow is known as viscosity.
The thicker or more dense the air, the greater is its resistance to flow.
Air which is more dense is said to be more viscous.
We will concern ourselves with the behaviour of the airflow closest to the aircraft as the aircraft moves through the
air.
This stream of air will flow smoothly around the aircraft as long as it's not decelerated rapidly or required to
make large and abrupt changes in direction.
The airflow, which is in immediate contact with the aircraft and which, at the most, is only a few millimetres
thick, is called the boundary layer.
We'll now examine the airflow and boundary layer around an aircraft's wing.
Friction between the wing surface and the air molecules that are moving over it will reduce the speed of the
air.
The air directly in contact with the wing surface may be slowed to zero.
This slowing effect reduces the further away the airflow is from the surface of the wing.
Eventually there will be no slowing effect on the airflow.
This point is called the free stream airflow.
The boundary layer can be categorised into two types of flow.
These are known as laminar flow and turbulent flow.
Laminar flow is a smooth flow of air around a body, or in our case the aircraft's wing.
Each molecule of air follows the path of the preceding molecules and is subjected to the same changes in pressure
and velocity at different points along its path.
This type of flow is shown graphically as streamlines that do not cross each other.
Changes in velocity are represented by changing the spacing between the lines.
A reduction in spacing represents an increase in velocity and an increase in spacing represents a decrease in velocity.
In turbulent flow the air molecules do not keep to the path of the preceding molecules.
In fact they may travel in any direction with varying speeds.
This has the effect of slightly thickening the boundary layer.
Turbulent flow is not desirable for most phases of flight.
It's not represented graphically by streamlines as its parameters vary so much, but it's normally shown as curls.
The boundary layer can be a combination of both laminar flow and turbulent flow.
The boundary layer can start as laminar flow and then become turbulent flow.
The point at which it changes from one to the other is called the transition point.
One of the fundamental laws of the universe is that energy and mass can neither be created nor destroyed, only
changed from one form to another.
Consider a streamlined flow of air passing through a tube.
The air mass flow will be dependent on the cross-sectional area of the tube, the flow velocity and the
density of the air.
The mass flow will remain constant at all points along the tube.
This is known as the equation of continuity.
We've already stated that in these lessons we'll only consider air flows below the speed of sound.
At these speeds the air density changes are insignificant and can be disregarded.
The equation of continuity can now be simplified to area times velocity equals a constant.
Let's now consider a streamlined flow of air through a tube which had a reduced cross-section in the middle.
This type of tube is known as a Venturi.
For a clearer view of what will happen to the air as it flows through this Venturi tube, we'll view
it from the side.
To explain this principle, let's apply some simple dimensions and values to the diagram.
Let's assume the tube initially has a cross-sectional area of one square metre.
For simplicity, let's assume that the air has a density of one kilogram per cubic metre.
If the velocity of the air is 52 metres per second, this will give a mass flow of 52 kilograms
a second.
At the middle of the Venturi, let's assume the cross-sectional area is reduced to only half a metre squared.
Because mass flow must remain constant, the air flow will increase in velocity.
In this example, the velocity will be doubled.
It can be seen from the continuity equation that the reduction in the cross-sectional area results in an increase
in velocity of the air.
As the cross-sectional area increases once again, the velocity of the air reduces to 52 metres per second.
The principle of continuity was concerned with constant mass flows. Bernoulli's theorem is concerned with the energy of that flow.
Daniel Bernoulli was an 18th century Swiss scientist.
His theorem states that the total sum of the energy in a streamlined flow of an ideal fluid remains constant.
An ideal fluid is both incompressible and has no viscosity.
Although in fact air does not meet these requirements, at the subsonic speeds we are considering we can assume that
it does.
The energies in the airflow that we will consider will be its static pressure energy and kinetic energy.
We've already defined the kinetic energy of the air as dynamic pressure.
Thus the statement we've just made can be expressed as static pressure plus dynamic pressure equals a constant.
Let's now return to the flow through the Venturi tube.
This time though, we'll look at the effects on the static and dynamic pressures.
Just as we did when studying the continuity equation, let's give some simple values to the static and dynamic pressure
of the air flowing through the Venturi.
Assuming the air to be at sea level ISA conditions, we will accept that the static pressure at the opening
of the Venturi is approximately 1013 hectopascals or millibars.
This pressure may also be expressed as 101300 newtons per square metre or 101.3 kilonewtons per square metre.
Dynamic pressure as we know is given by the equation half rho V squared.
At the opening of the Venturi therefore, at a flow velocity of 52 metres per second, we calculate that the
dynamic pressure of the air is 1.656 kilonewtons per square metre.
This is 16.56 hectopascals.
We see then that the total pressure energy at the opening of the Venturi will be equal to static pressure
plus dynamic pressure.
Giving us 103 kilonewtons per square metre rounded up to whole numbers or, if you prefer, 1030 hectopascals or 1030
millibars.
We'll look now at the situation in the throat of the Venturi. That is the area of the smallest cross
section.
We've already seen that the speed of the air flow increases through the throat of the Venturi.
In our Venturi the speed doubles to 104 metres per second.
The dynamic pressure equation half rho V squared allows us to calculate that the dynamic pressure of the air flow
will now increase to about 6.62 kilonewtons per square metre or 66.2 hectopascals or millibars.
Remember though that Bernoulli's theorem states that the total pressure energy must remain constant.
In our case the total pressure remains constant at 103 kilonewtons per square metre.
So a simple arithmetical calculation enables us to see that if the dynamic pressure increases from 1.656 kilonewtons per
square metre to 6.62 kilonewtons per square metre the static pressure must decrease from 101.3 kilonewtons per square
metre to approximately 96.4 kilonewtons per square metre.
Or if you prefer 964 hectopascals or millibars.
We've learnt that as our Venturi widens again the speed of the air flow reduces to its original value of
52 metres per second.
Consequently the dynamic pressure exerted by the air also decreases to its original value.
Because the total pressure remains constant it is thus evident that the static pressure too will increase once more to
its original value.
From this examination of the continuity equation, Bernoulli's theorem and the flow of air through a Venturi tube we've learnt
several fundamental facts about subsonic aerodynamics which are crucial to our understanding of the principles of flight.
Firstly, the mass flow of air remains constant.
Secondly, the velocity of air flow is affected by the shape of the objects around which it flows.
Thirdly, the total pressure energy of air, that is the sum of its dynamic and static pressures, remains constant.
And finally, combining all that we have learnt, when the velocity of air flow increases the dynamic pressure it exerts
will increase and its static pressure decrease.
Conversely, as air flow decreases in velocity its dynamic pressure decreases and its static pressure increases.
To demonstrate the principles of continuity and Bernoulli's theorem we have used a Venturi tube.
In fact, air will behave the same if we remove the tube leaving only part of the curved surface for
the air to flow over.
Air flowing over such a curved surface will increase in velocity and its static pressure will correspondingly decrease.
If we now take a cross section through the wing of a typical light aircraft we can see that we
have a pronounced curved surface on the top of the wing.
We sometimes refer to such a cross section as an aerofoil.
You'll see that the undersurface of the aerofoil has hardly any curve.
When air flows around the aerofoil the air flow over the upper surface will increase in velocity causing the static
pressure there to decrease.
Underneath the wing section the static pressure remains at a higher value.
As the pressure or force per unit area is greater underneath the wing than above the wing a force is
generated across the wing which tends to move the wing upwards.
This force is called lift.
You can easily demonstrate this effect yourself.
Hold a piece of paper between the thumb and forefinger of each hand.
The paper should curve away from you.
Bring the paper up to your mouth and blow down against the curved edge facing you.
Just like the piece of paper the aircraft's wing will generate a force acting upwards as long as the pressure
difference between the top and bottom of the wing is large enough.
That of course will depend on the speed of the air flow.
It's this simple principle that will keep an aircraft flying.
Let's expand our perception of the curved section of the Venturi and make it into the top of an aircraft's
wing.
Of course atmospheric pressure will be acting under the wing.
So now we've found the answer to the question asked in the first lesson which was how does an aircraft
fly?
The basic principle I hope you'll agree is simple.
But there are many other factors to consider.
How is an aircraft controlled?
What is the best shape for a wing?
What forces act on the aircraft in flight?
All these questions and more will be answered in the following lessons.
Aircraft come in different shapes and sizes.
Each is designed for a specific task.
All powered aircraft however will share certain features.
An aircraft will have a fuselage in which to house the pilot, passengers, baggage or cargo.
Wings to generate a lift force for flight.
Stabilising surfaces.
Control surfaces to allow us to change the direction of flight.
An engine or engines to provide for forward movement of the aircraft.
And an undercarriage to allow the aircraft to move on the ground.
This can be fixed or retractable.
Together the fuselage, wings, control surfaces, stabilising surfaces and undercarriage form what is known as the airframe.
The engine or engines will be fitted to the airframe.
Engines can be fitted at the front or the rear of the aircraft or perhaps mounted on top of the
fuselage.
We shall show that the airframe is subjected to four forces.
These forces are very large and the airframe must be strong enough to withstand them.
An aircraft possesses mass by virtue of the amount of matter it contains.
The aircraft's mass is subjected to the acceleration due to gravity, giving rise to a force acting on the aircraft
which is called weight.
Weight acts through a point on the aircraft called its centre of gravity.
The line of action of the weight force is always directed vertically downwards towards the centre of the earth.
The piston engine provides power to drive a propeller.
The propeller generates a force called thrust which moves the aircraft forward.
A jet engine however produces thrust without the need for a propeller.
We showed earlier in this lesson that as the thrust moves the aircraft through the air the aircraft will be
subjected to dynamic pressure.
This results in pressure changes over the wings which produce a force opposing the weight of the aircraft.
This force is called lift.
You'll recall that there will be friction between moving air and any body with which it has contact.
This friction resists the movement of the body.
In our case the moving body is the aircraft and the resistance force is called drag.
To summarise then, the four forces acting on an aircraft in flight are weight, lift, thrust and drag.
For the aircraft to fly in what is known as balanced flight, the force of lift must be equal and
opposite to the weight and thrust must be equal and opposite to drag.
These four forces are closely interrelated.
The greater the weight of the aircraft the greater the lift required.
The greater the lift the greater the drag generated.
The greater the drag the greater the thrust required and so on.
In reality the forces do not all act from the same point as is shown in these diagrams.
The weight could act forwards of the point of lift.
When two forces are acting at a distance from each other they are termed a couple.
The couple will produce a turning moment.
In this example the turning moment is tending to lower the nose of the aircraft.
This movement of the aircraft is around its centre of gravity.
In flight the aircraft will always rotate around the centre of gravity.
Equilibrium in flight is achieved when all the forces are in balance and the sum of all the turning moments
is equal to zero.
Just as lift does not always act in opposition to weight, so thrust and drag also do not always act
directly in opposition to each other.
Normally thrust acts from a point lower than that point from which the force of drag acts.
This would give a turning moment which would tend to raise the nose of the aircraft.
The pilot controls the thrust with the throttle.
On opening the throttle engine power is increased and this will increase the thrust from the propeller.
This change in the moment will cause the nose of the aircraft to rise.
Conversely the nose will drop when the throttle is closed.
It would be convenient for us if the turning moment of lift and weight was cancelled by the turning moment
of thrust and drag, but there is a further problem.
The centre of gravity about which the weight acts can move.
This is caused by the fact that as fuel is used during a flight, not only does the weight reduce,
but the disposition of that weight may also change.
To complicate matter still further, the lift force also moves throughout the flight.
Why this happens will be studied in a later lesson and need not bother us here.
For the time being we only need to concern ourselves with the fact that the lift force does move during
flight.
To keep the aircraft in equilibrium we must compensate for these movements of the centre of gravity and the lift
force.
The horizontal stabiliser or tailplane as it is commonly called, produces the correcting moment.
The tailplane will provide a downforce if the weight acts forward of the lift.
This force will reduce as lift moves forward or the centre of gravity moves rearward.
The aircraft pictured here has what is known as a canard design.
This type of aircraft design has become more popular in recent years.
The canard design claims to be more efficient and able to demonstrate safer flight characteristics than a conventional aircraft design.
As you can see from the picture of the canard, the aircraft does not have a conventional horizontal stabiliser.
The main wing is to the rear and a smaller wing is located at the front of the aircraft.
These smaller wings are referred to as canards.
The word canard is French and the word translated into English means duck.
Looking at a side elevation of the aircraft and adding the forces of lift and weight, you can see that
a nose down moment is produced.
To counter this moment the canard produces an upforce.
Now there are two forces opposing the weight, that produced by the canard and that produced by the main wing.
Thus, if to achieve balanced flight these two forces must be equal to the weight, the lift force developed by
the main wing can be reduced.
We've already stated that all of the forces are closely interrelated and this next point proves it.
The decrease in lift developed by the main wing will result in it having a decrease in drag when comparing
it to the main wing of a conventional aircraft of equal weight.
A propeller driven aircraft can also have a turning moment acting on it produced by what is termed the slipstream
of the propeller.
The operation of the propeller is covered in detail in the propeller section of this CD-ROM, but as a
brief explanation the propeller is a device that converts the engine power to thrust.
It does this by accelerating the air and moving it rearwards.
This movement of air is called the propeller's slipstream, which moves in a helical path behind the propeller.
Mounting the propeller at the front of the aircraft can lead to a turning moment on the aircraft which is
opposite to the direction of rotation of the propeller, as the slipstream strikes the vertical stabiliser.
If the propeller rotates to the right the aircraft will tend to swing to the left, and conversely the aircraft
will tend to swing to the right if the propeller has a left-hand rotation.
The effect of the turning moment can be countered by offsetting the vertical stabiliser, commonly called the fin, of the
aircraft.
This will produce an aerodynamic force which is in opposition to the moment caused by the slipstream.
This offset has been over-emphasised in this diagram to better demonstrate the effect.
When an aircraft flies at a steady speed in equilibrium all the forces are in balance and the sum of
all the moments equals zero.
The pilot, however, will at times want to manoeuvre the aircraft, and also increase or decrease speed.
The balance of forces must be altered during these phases of flight.
The process of producing an aerodynamic force, explained earlier, is fairly straightforward, but as we've just seen aircraft come in
many and various designs.
To obtain the performance required for a particular operation, each type will be configured in a specific way.
High wing, low wing, different wing shapes, every feature has a purpose.
The structural and aerodynamic design of an aircraft is a masterpiece of compromise.
An improvement in one area may lead to a loss of efficiency in another.
The effects of the subtle changes in shape and configuration of an aircraft will be explained in the following lessons.
The wings of an aircraft provide the force to overcome the weight of the aircraft so that it may fly.
We've already defined this force as lift.
The force of lift is an aerodynamic force.
Air moving around the wing will produce a pressure differential between the wing's upper and lower surfaces.
This principle of pressure change was explained in the previous lesson.
If you were to cut through the wing and then look at what is called its cross section, the shape
produced is termed an aerofoil.
An aerofoil is shaped so that it is capable of producing lift with relatively high efficiency.
The following terms are used to describe an aerofoil.
The front of the aerofoil is called the leading edge and the rear part is called the trailing edge.
A straight line drawn from the centres of curvature of the leading and trailing edges is defined as the chord
line.
The distance measured along the chord line between the leading and trailing edges is called the chord.
A line joining the leading and trailing edges of the aerofoil which is equidistant from the upper and lower surfaces
is known as the mean line or camber line.
The chord and camber lines would not actually appear on a real wing. They are only used in the explanation
of theory.
The maximum camber of an aerofoil is defined as the point at which the distance between the mean line and
the chord line is at its maximum.
When the camber line is above the chord line the aerofoil is said to have positive camber.
Some aerofoils have a symmetrical cross section. An aerofoil with a symmetrical cross section would have no camber as the
chord line and the camber line are coincidental.
Aircraft fins have symmetrical aerofoils, as do some tail planes.
The maximum thickness or depth of the aerofoil is given as the maximum distance between the upper and lower surfaces
of the wing, measured along a line at 90 degrees to the chord line.
An aerofoil's maximum thickness is expressed as being at a percentage of the chord measured back from the leading edge.
For example, the maximum thickness of the aerofoil shown on the screen is at about 25% chord.
The thickness-chord ratio is the maximum thickness divided by the chord expressed as a percentage.
Most aerofoils used in light aircraft have a thickness-chord ratio of 12 to 15%, with their maximum thickness being
at around 40% of chord.
The radius of curvature of the leading edge is called the leading edge radius.
The size of the leading edge radius, the amount of camber and the thickness of the aerofoil have a great
influence on the airflow characteristics around the aerofoil and the lift force produced.
The wing doesn't have to be shaped like an aerofoil to produce an aerodynamic force to oppose the weight of
the aircraft.
If, however, the wing is to be efficient, then it does have to be aerofoil shaped.
We'll now look at how lift can be produced by a flat plate.
If the wing of our aircraft was just a flat plate, it could still produce the aerodynamic force to lift
the aircraft.
The flat plate can be likened to the chord line of our aerofoil wing section.
The angular relationship between the chord line and the horizontal datum of the aircraft is called the angle of incidence.
The wing is normally securely attached to the fuselage of the aircraft, and the angle of incidence is fixed.
In this graphic, the angle of incidence is shown as being 2 degrees, which is fairly typical for a light
aircraft.
The horizontal stabiliser of the aircraft also has an angle of incidence.
This can be fixed like the main wing, but can in some designs be variable.
Such a tailplane is called a variable incidence tailplane.
In the previous lesson, we studied the subsonic flow of air around a body, and we stated a convention for
graphically representing the flow.
If we now consider a flat plate parallel to the relative airflow, and study the streamlined flow around it, we
see that there is no change in flow velocity, and hence no change in pressure.
If however the plate is presented to the airflow at an angle, the flow characteristics are altered.
The angle between the relative airflow and the chord line, which in this case is simply the axis of the
flat plate, is called the angle of attack.
It should be noted here that the angle of attack is different from the angle of incidence.
Remember that the angle of incidence is a fixed angle between a wing's chord line and the aircraft's horizontal datum.
The angle of attack on the other hand, as we'll see later, is the angle between the airflow and the
wing's chord line.
That's why when we're referring to angle of attack, we often refer to the airflow as the relative airflow.
You should note that when considering the generation of lift and drag, it's the angle of attack which is of
interest to us.
Previously, when studying the effects of airflow, it was convenient to view the flow through a Venturi tube.
Once again, we'll use an imaginary tube around the aircraft and wing to help explain what's happening to the airflow.
This imaginary tube is called a stream tube.
If we now consider the effect of placing the flat plate at an angle of attack, it can be seen
that the area within the tube has been reduced.
In the previous lesson, the continuity equation stated that the velocity of the air would therefore increase, and Bernoulli's principle
states that the static pressure energy would thus decrease.
This decrease in pressure is happening on the upper surface of the plate.
Beneath the plate, the airflow is being brought to rest and deflected downwards.
This, you'll recall, will have two effects.
Firstly, because the air is slowing down, the static pressure will increase in accordance with Bernoulli's theorem.
Secondly, as we saw in the lesson on the four forces, the force exerted on the air by the flat
plate in order to change the direction of the airflow will generate an equal and opposite force on the flat
plate.
Both these effects will increase the total pressure energy beneath the plate.
Now we have a pressure differential between the upper and lower surfaces, and the plate will have an aerodynamic force
acting on it.
This aerodynamic force can be represented graphically by a vector diagram.
A vector has both magnitude, or size, and acts in a known direction.
On an aerofoil, this vector is called the total reaction.
This is probably the first time you've heard the term total reaction with reference to this aerodynamic effect, but I'm
sure you've experienced its effect in your everyday life.
Whilst out in a car on a sunny day with a window wound down to allow in a cool breeze,
who, as a child, has not placed their hand out of the window and felt the force of the air
blowing against it.
Holding the hand in line with the airflow generates little effect, but tilting the hand slightly in the airflow generates
a force which makes the hand want to move rearwards and up or down.
The hand is thus experiencing the total reaction.
Returning to our flat plate and its total reaction, we see that at this angle of attack the plate will
move rearwards and upwards.
It's the value of the upwards component of the total reaction which is of interest to us, as this will
directly oppose the weight of the aircraft.
We've already mentioned this force, it's called lift.
Now we can define it.
Lift is defined as the component of the total reaction which acts at 90 degrees to the relative airflow.
There's a saying that nothing in life is free, and in this particular case the price we pay for lift
is that there is a component of the total reaction also acting rearwards.
This component of the total reaction acts parallel to the relative airflow and in opposition to the movement of the
plate.
It's defined as drag.
An aircraft designer must always consider the relationship between lift and drag.
We defined an aerofoil as being a shape capable of producing lift efficiently.
That efficiency is the relationship between the lift and drag produced.
Let's now compare the lift and drag characteristics of a flat plate with those of an aerofoil, of the same
chord, representing a typical cross section of the wing of a light aircraft.
This will help us understand why we use this type of aerofoil in the construction of light aircraft.
We'll continue to view the flow around them with the aid of the stream tube.
If there's no movement of air then there's no dynamic pressure. This means no aerodynamic forces.
No total reaction, no lift, no drag.
Thus we can say that both lift and drag are affected by speed.
If both the aerofoil and the flat plate are now subjected to a constant airflow, the flow characteristic around each
can be viewed.
What you will see now is a graphical representation of the airflow which might occur if you carried out this
experiment in a wind tunnel.
Both the plate and the aerofoil are at zero degrees angle of attack.
The aerofoil you'll see, even at zero angle of attack, is modifying the airflow, causing it to accelerate above the
aerofoil and thus reduce the static pressure.
The aerofoil is thus already producing an aerodynamic force.
The flat plate however, as was seen previously, has had little effect on the airflow.
Increasing the angle of attack slightly reduces the area in the stream tubes.
The airflow thus accelerates, reducing the pressure above the aerofoil still more and causing an initial pressure reduction over the
flat plate.
This acceleration is greater over the curved aerofoil than it is over the flat plate.
Notice how, at identical angles of attack, the aerofoil produces more lift and less drag than the flat plate.
Pressure is increased on the lower surfaces as the speed of the airflow is reduced.
Also, over the flat plate, the boundary layer becomes turbulent quite soon after passing the leading edge.
You'll recall that the air will flow smoothly around a body if it's not required to turn through large angles.
The leading edge of the flat plate forms a large angle.
In contrast, the curved leading edge of the aerofoil gently turns the air and the boundary layer remains laminar.
Turbulent flow thickens the boundary layer. This will increase the drag.
Increasing the angle of attack still further again reduces the area within the stream tube.
The airflow accelerates to a higher speed and the pressure reduces still further.
Once again, the effect generated is greater over the aerofoil than it is over the flat plate.
If we continue to increase the angle of attack, even the boundary layer of the aerofoil will start to become
turbulent as the airflow tries to turn through larger angles.
The airflow again accelerates as the area within the stream tube reduces, and there'll be a further drop in pressure
over the upper surfaces.
On the lower surfaces, pressure increases as the velocity of the air is reduced.
This gives a greater pressure differential, increasing the total reaction.
Note also that the turbulent flow starts to separate from the upper surfaces.
This process continues up to an angle of attack of around 16 degrees.
More and more turbulent air separates from the upper surfaces. It's replaced by higher pressure air from the atmosphere.
The pressure difference between the upper and lower surfaces is reduced, with a resultant drop in lift.
The aerodynamic force produced in this condition is predominantly drag.
There is a greater separation from the flat plate than from the aerofoil.
The flat plate produces much less lift and much more drag.
This condition is called the stall.
From the observations we've just seen, we can make some statements regarding the lift-drag relationship of the flat plate
and the aerofoil, and how the two compare.
Firstly, we can state that for the same size chord, airflow or dynamic pressure, and angle of attack, different pressure
changes occurred over the sections.
This would give different values of aerodynamic force.
The two aerodynamic forces we are concerned with are those of lift and drag.
To express these changes in magnitude, a numbering index is used which is called the coefficient of lift and drag.
The coefficients are normally represented by the letters CL for the coefficient of lift, and CD for the coefficient of
drag.
For example, at zero degrees angle of attack, the flat plate did not affect the airflow, whereas the aerofoil did.
For the flat plate, the coefficient of lift would be zero. It produced no force when subjected to dynamic pressure.
The aerofoil would thus be allocated a larger number to reflect this fact. For example, its coefficient of lift would
equal 0.4.
We now have a method to compare aerofoils. The same principle would apply when comparing drag.
Most aircraft do not use flat plates for the main wings. They are however sometimes used for horizontal stabilisers.
Apart from the fact that they are not aerodynamically efficient, it would be difficult to construct a flat plate wing
section of sufficient structural strength to withstand the forces which would act upon it.
Aerofoils come in many different shapes. Generally we can divide them into three categories which would be high lift, general
purpose and high speed.
When discussing the terminology used for aerofoils, it was stated that the camber, thickness and leading edge radius influenced the
airflow and lift characteristics.
The high lift section has a pronounced camber, a high thickness chord ratio and well-rounded leading edge. This would
give good lift at low speeds.
But it would also have relatively high drag. It would be used for aircraft where high speed was not a
consideration.
The general purpose section has lower thickness chord ratio, less camber and a sharper leading edge. This results in the
aerofoil having lower drag and coefficient of lift than the high lift section.
This is probably the type of aerofoil fitted to the aircraft that you'll learn to fly in.
The requirement for the high speed aerofoils are that it has low drag. This is achieved by making a thin
section with little or no camber.
The maximum coefficient of lift for a high speed section is low. The section shown is a symmetrical section which
is typical of a high speed aerofoil.
We'll now look at the change in the coefficient of lift that occurs with a change of angle of attack
for each of the sections starting with the high lift aerofoil.
A cambered aerofoil is capable of producing lift at negative angles of attack down to about minus 4 degrees.
With increasing angle of attack the coefficient of lift increases until about 16 degrees. At this point the coefficient of
lift is at a maximum.
Further increases in the angle of attack result in a reduction in coefficient of lift. This is the point of
stall where the airflow separates from the upper surface.
Note that the coefficient of lift reduces gradually. If we now add the coefficient of lift curves for the general
purpose and high speed aerofoils a comparison can be made between the different sections.
Note the lower C L max values and the shape of the curves at the stall. The coefficient of lift
for the high speed aerofoil reduces very quickly.
This implies that the aircraft would stall without much warning to the pilot. Also note that, like the flat plate,
no lift is produced at zero angle of attack on the high speed section.
Having seen how changes in angle of attack affect the coefficient of lift, we must now go on to look
at how the coefficient of drag varies with angle of attack.
But first we must examine more closely how pressure changes around an aerofoil section.
Pressure changes are represented graphically like these shown here. A negative symbol is used to indicate a pressure lower than
atmospheric and a positive symbol is used to indicate pressures that are higher than atmospheric.
The shape of the pressure pattern would be obtained in reality by experiment. We will now show you a graphical
representation of a typical experiment to obtain the pressure pattern over the upper surface of an aerofoil.
The experiment is very simple to carry out. The aerofoil section under test will have small holes drilled in its
surface across the chord. The holes will then be connected by clear tubes to a container of coloured liquid.
The section is then suspended in an airflow which is produced by a fan.
When air passes over the section we know that it will accelerate, and thus there will be a reduction in
static pressure. Just as you suck a drink up through a straw, the coloured liquid will be sucked up the
tubes.
The pressure patterns for an increase in pressure are similarly represented.
The high pressure at the leading edge of the aerofoil is due to the fact that the moving air has
been brought to rest. This is known as the stagnation point.
Dynamic pressure will act on the leading edge in addition to the static pressure. This now gives a pressure differential
between the leading and trailing edge.
This resists the forward movement of the aerofoil and is another form of drag.
It's already been stated that both lift and drag are affected by speed, so if dynamic pressure is increased the
pressure pattern will change.
The following animation shows the change in pressure distribution as the angle of attack changes at constant airspeed, or otherwise
expressed at constant dynamic pressure.
An increase or decrease in the areas of low and high pressure shown on the diagrams represents an increase or
decrease in the magnitude of the pressure distribution across the aerofoil.
We've seen that the pressure distribution around a wing generates a force known as total reaction, the vertical component of
which is lift.
Lift is represented as acting through an imaginary point on the chord of the wing called the centre of pressure.
At a typical cruising angle of attack of 4 degrees, the centre of pressure is on the chord about a
third of the way back from the leading edge.
This animation shows how the centre of pressure, and thus the line of action of the lift force, moves forward
along the chord as the angle of attack increases.
We assume constant airspeed.
Notice that just before the stall the lift force reaches a maximum value and that the centre of pressure is
at its most forward position.
At the stall the lift force reduces markedly and the centre of pressure moves backwards.
The principles studied so far have been based on the concept of a hypothetical two-dimensional flow.
The air flows across the aerofoil from leading to trailing edge. This gives one dimension.
The second dimension is in the vertical plane.
As the pressure on the upper surface is reduced, the airflow approaching the leading edge, which is at a higher
pressure, flows toward the low pressure.
The approaching airflow flows up towards the upper surface. This is called upwash.
After passing over the aerofoil, the airflow returns to its original position and state. This is termed downwash.
In reality, there is a third dimension, a spanwise flow which gives the true flow over the whole wing.
Upwash was explained as occurring because high pressure air will always flow towards an area of lower pressure.
Wherever flows of different pressures interact, the same thing will occur, no matter how small the pressure difference.
When considering the whole aircraft and not just the aerofoil section, these pressure differences will modify the airflow so far
discussed.
This will change the angle of attack, which will alter the aerodynamic forces on the aircraft.
This modified airflow is called three-dimensional flow and is the true airflow over an aircraft.
To identify where this interaction takes place, we need to introduce some additional terminology relating to the wing.
The wingspan is the distance between the wingtips.
The wing root is at the junction of the wing to the fuselage.
The chord, which has already been defined, is constant on the wing shown.
This type of constant chord wing is often used on training aircraft.
This shape of wing is known as a rectangular wing.
For aerodynamic reasons, which will be discussed later, some aircraft have a reduced chord at their wingtip compared to their
root.
This wing shape is known as a tapered wing.
Since lift depends on the pressure difference between the upper and lower surfaces of the wing and pressure is defined
as force per unit area, the plan area of the wing is a major factor in the magnitude of the
lift force produced.
The wing area is defined as the span times the average chord, the average chord being the geometric average of
the wing.
Although a portion of the area may be covered by the fuselage or, in the case of a multi-engine
aircraft, the engine nacelles, the entire plan area can be considered.
Further consideration must be given to what is called the wing's aspect ratio. This is the ratio of the span
to the chord.
If we just consider half the wingspan as shown here, it can be seen that, although each half of the
wing has the same area, the aspect ratios are very different.
Aspect ratio has a major influence on the aerodynamic efficiency of a wing, high aspect ratio wings being more efficient.
You'll notice that the wings of gliders, for example, are very long and thin. They have a very high aspect
ratio, because with no thrust force available the wings need to be efficient.
Now that we can identify the different sections of the wing, the areas where airflow interaction occurs can be shown.
These are at the trailing edge and the tips of the wings.
The pressure differential modifies the direction of the airflow, inducing a spanwise vector, generally in the direction of the root
of the wing on the upper surface and towards the tip on the lower surface.
Conventionally an aircraft is viewed from the rear. The flow of air from high to low pressure is rotary in
its motion and called a vortex.
An anticlockwise vortex is induced at the right wing and a clockwise vortex at the left wing.
It must be remembered that the changes in pressure are due to the fact that the aircraft is moving. The
vortices are therefore continually rotating and also moving rearwards.
The vortices are normally strongest at the wing tips.
Vortices produce vertical velocity components in the airflow in the vicinity of the wing. This increases both the upwash and
the downwash.
This results in a change of direction of the relative airflow, which in turn will cause a change of the
angle of attack.
The resultant changes to the airflow produce what is called the effective airflow and effective angle of attack.
Both the lift and drag produced by the wing will be affected by these changes.
Let's first consider the component of lift.
It was previously stated in this lesson that lift is directly affected by the angle of attack, up to the
stalling angle.
Lift was also defined as the component of the total reaction acting at 90 degrees to the relative airflow.
Lift can now be redefined as acting at 90 degrees to the effective airflow.
You'll note that this acts rearwards.
As the effective angle of attack is reduced, the magnitude of this lift force will also be reduced. Lift must
of course equal the weight or mass of the aircraft to maintain level flight.
To restore the lift lost due to the downwash, the wing must be selected to a higher angle of attack.
This will increase drag.
This addition in drag is called induced drag or lift dependent drag. In this lesson the term induced drag will
be used.
When we previously introduced the force of drag, it was stated that it was opposed by thrust to achieve balanced
flight.
Every part of the airplane exposed to the airflow produces a resistance to its forward motion. We define this as
drag.
The total drag on an aircraft is made up of two different types of drag. These two types of drag
are called parasite drag and induced drag.
We stated earlier that drag increases with an increase in speed. This statement refers to the parasite drag.
Look at the next graphic which shows a graph showing drag against indicated airspeed. You can see that drag increases
with airspeed. You'll recall that this is because of dynamic pressure.
Because of the squared function of the formula, the graph is a curve.
Induced drag reduces with speed. The reason for this will be explained later, but note that the induced drag graph
is again a curve.
This is because induced drag varies inversely as the square of the speed.
Combined together, the parasite and induced drag curves give the total drag curve.
The actual total drag curve will vary from aircraft to aircraft, but it will be similar to the curve shown.
Now we'll look at the different types of drag in more detail, starting with parasite drag.
Parasite drag is further subdivided into skin friction drag, form pressure drag, and interference drag.
Skin friction and form drag are together known as profile drag.
If an aircraft were to fly at zero lift angle of attack, the only drag present would be parasite drag.
Although skin friction and form drag have already been mentioned when we discussed two-dimensional flow, they were not defined.
It's already been stated that a laminar boundary flow produces less drag than turbulent flow.
The surface condition and area over which the thin boundary layer flows has a marked effect on its flow characteristics.
Any roughness on the skin of a leading portion of an aircraft will cause a transition to turbulent flow and
increases the skin friction drag.
A polished shiny aircraft not only looks good, but will fly more efficiently as the skin friction drag is reduced.
Construction methods can greatly affect skin friction.
Flush riveted or bonded metal skin aircraft will have less skin friction than aircraft with a skin which has rivet
heads that stand proud, or aircraft that are covered with fabric which has pronounced stitching.
Many aircraft today are constructed from composite materials, which produce continuous smooth structures that have low values of skin friction.
The manufacturer will determine the type of construction of the aircraft and its inherent skin friction values.
In service, the aircraft's surface can become contaminated with dirt, water, ice, frost or snow.
These contaminants will affect not only the skin friction drag, but also the amount of lift produced.
Tests have shown that contamination of the leading edge of a wing with the thickness and surface roughness similar to
medium or coarse sandpaper will reduce lift by as much as 30%.
Form drag is caused by the pressure differential between the increase in pressure at the front of an object, due
to the airflow being brought to rest, and the lower pressure at the rear of that object.
The point at which the airflow is brought to rest on an airfoil has previously been defined as the stagnation
point.
The value of the increase in pressure is a function of dynamic pressure. An increase in airspeed will mean an
increase in dynamic pressure.
Form drag will increase with the square of the speed.
Since pressure is defined as the force per unit area, it follows that an increase in the area of the
stagnation point will increase form drag.
A thick aerofoil or a fat aeroplane would have more form drag than a slimmer one for the same dynamic
pressure.
The increased surface area of a thicker wing or fatter aircraft will also affect skin friction drag.
These changes in the aircraft shape are termed its profile.
The profile relates to the cross-sectional area.
The aircraft's profile affects both skin friction and form drag.
Collectively these two types of drag are termed profile drag.
Form drag is also affected by changes in angle of attack.
We've seen that as the angle of attack increases, the transition point moves forward.
Eventually the turbulent airflow will detach from the upper surfaces of the aerofoil.
The point where this happens is called the separation point.
This separation causes air at the trailing edge to flow forwards towards the separation point.
This will cause the pressure at the trailing edge to reduce.
Since form drag is caused by the pressure difference between the leading edge and trailing edge of an aerofoil,
form drag will increase with increasing angle of attack.
Every part of an aircraft will be subjected to form drag.
The example shown in the graphic is of the aircraft's undercarriage.
To reduce form drag, it's necessary to delay separation to a point as close to the trailing edge as possible.
This is achieved by streamlining.
Streamlining increases the ratio between the length and depth of a body.
This delays separation and reduces drag.
The ideal ratio is around 3 to 1.
It should be noted that the shape of the rear cross section of a body has a major effect on
form drag.
When considering a complete aircraft, parasite drag will be greater than the sum of the skin friction and form drag.
Additional drag results from turbulence of the boundary layer caused at the junction of various aircraft structures.
For example, where the wings join the fuselage or where a wing strut is attached to the wing or fuselage.
This additional drag is called interference drag.
Ensuring that there are no sharp angles formed at these junctions can reduce interference drag.
This is achieved by the use of fillets or fairings at these junctions.
This principle is similar to streamlining as the intention is to reduce turbulent flow.
Interference drag will also increase with an increase in speed.
Returning to the parasite drag curve, we can see that parasite drag increases with speed
and we can now identify that it's made up of skin friction and form drag, known collectively as profile drag,
and interference drag.
It's been explained how contamination can increase parasite drag.
Configuration of the aircraft will also affect the drag.
Take for example the lowering of the undercarriage on a retractable undercarriage aircraft.
Note how the drag curve changes.
The nature of induced drag was explained earlier in this lesson.
We'll now look at factors which affect its magnitude.
You may recall that induced drag is one of the undesirable products of lift.
Whenever lift is present, induced drag is present.
That's why it's sometimes referred to as lift-dependent drag.
Induced drag is caused by the rearwards inclination of the lift force under the influence of the vortices produced at
the wingtip.
Since these vortices grow in size and strength as the angle of attack increases, it follows that induced drag will
also increase with increasing angle of attack.
As you've learned, for a given aerofoil, coefficient of lift is directly proportional to angle of attack.
We may also say, therefore, that induced drag increases as the coefficient of lift increases.
In fact, induced drag is directly proportional to the square of the coefficient of lift.
Let's look at the case of two identical aircraft flying at the same speed, one lightly loaded and one heavily
loaded.
At the higher angle of attack, the coefficient of lift will be greater and the induced, or lift-dependent drag,
will increase.
The heavier aircraft therefore creates the greater induced drag.
In flight, an aircraft's weight will reduce as consumables like fuel and oil are used.
Induced drag will therefore reduce throughout the flight.
So far, the relationship between lift and weight has only been considered for straight and level flight.
Certain manoeuvres will require the lift force to be greater than the weight of the aircraft.
The relationship of lift to weight is known as load factor.
This is commonly called g-force.
When lift is equal to weight, the aircraft has a load factor of 1g.
The aircraft is flying at 1g.
If lift is greater than weight, the load factor will increase.
In the example shown, lift is twice the value of weight and the aircraft has a load factor of 2g.
If we consider an aircraft manoeuvring in a steady level banked turn as shown here, lift is no longer balancing
weight.
Weight will always act vertically downwards. The vertical component of lift is now less than weight.
The aircraft would of course descend in this condition.
The coefficient of lift must be increased by increasing the angle of attack until the vertical component of lift is
equal to the weight.
Since lift has been increased, there will be an increase in induced drag and load factor.
The load factor, and hence the change in lift, can be calculated by the formula 1 over the cosine of
the bank angle.
Cosine is a term used in the mathematical study of trigonometry.
It's not the purpose of this course to teach trigonometry.
The cosine value for a given angle can be found from cosine tables, or more often today by electronic calculator.
The value of cosine 60 degrees is 0.5. 1 over 0.5 equals 2.
The aircraft has carried out a 2g manoeuvre, so lift must increase twofold.
The load factor, and therefore the lift and induced drag, will vary depending on the bank angle being flown.
Use the buttons on the graphic to select the value of bank angles that are generally flown in a light
aircraft,
and see how the lift and load factor change for an aircraft of 700kg mass.
This mass must be multiplied by 9.81, the acceleration due to gravity, to give the force of weight equal
to 6,867N.
The induced drag of an aircraft varies depending on the length of the aircraft.
It varies inversely as the square of the speed. With an increase in speed, induced drag reduces.
This reduction in induced drag occurs because the vortices produced on the wing have less effect on the airflow over
the wing when the speed of the aircraft is high.
Think of the air as a football. Imagine that the ball is lightly kicked towards the goal.
At full stretch the goalkeeper just touches the ball and deflects it past the post.
Now imagine that the ball is kicked strongly towards exactly the same area of the goal.
The force the goalkeeper places on the ball as he stretches to touch it is the same, but now the
ball is not deflected as much and it hits the back of the net.
Higher angles of attack are selected at lower speeds. The greater pressure drop on the upper surface increases the strength
of the vortices and this has a greater effect on the airflow, increasing induced drag.
The vortices on the wings are greatest at the wingtip. It follows that if the vortices can be reduced, induced
drag will reduce. This can be achieved in a number of ways.
The use of wings which have a high aspect ratio means that the vortices will affect a smaller proportion of
the span, so the overall change in downwash will be less, and thus induced drag is reduced.
The use of a tapered wing has the same effect.
Another method of reducing induced drag is by using what is called washout. The wing is constructed with a reducing
angle of incidence from root to tip. In effect the wing is twisted along its span.
On a wing that has washout the angle of attack will also alter along the span, being less at the
wingtip. The reduced pressure difference at the tip reduces weaker vortices.
One last method employed to reduce induced drag is to modify the wingtips to reduce the flow of high pressure
air from below the wing to the top.
Wing end plates, wingtip fuel tanks, change of wingtip shape and winglets are all used to reduce the flow of
air that produces the vortices which cause induced drag.
This picture of a high performance glider illustrates the use of winglets.
The explanation that's been given for the formation of vortices and their effect on lift and drag assumes the aircraft
is in flight and well clear of the ground.
However, when an aircraft is taking off or landing, the closeness of the wing to the ground prevents full development
of the vortices, making them much weaker. This is called ground effect.
An aircraft is in ground effect when it's within approximately half the wingspan's distance from the ground.
The reduction in the vortices results in a reduction in drag and an increase in lift.
The change in these forces so close to the ground causes many a student pilot problems.
Consider an aircraft landing. The aircraft is in the final stage of the approach at a steady speed.
On entering ground effect the vortices will reduce and there will be an increase in the effective angle of attack.
Lift will increase, induced drag reduces and the aircraft will accelerate, which will again increase lift.
If the approach has been flown at too high an airspeed, especially with a small flap angle selected, a balloon
landing may result, with the aircraft beginning to climb again.
This demonstrates why pilots need to exercise good speed control on the approach and why the round-out and hold
-off are such critical phases of the landing.
There is one other aspect of the principles of flight that needs mention when discussing downwash, and that is the
effect of downwash on the horizontal stabiliser or tailplane.
The tailplane of the aircraft provides stability and affects controllability of the aircraft.
Downwash will also affect the aerodynamic characteristics of the tailplane as well as the main wing.
The effective angle of attack is modified by the changing downwash from the main wing. This of course will alter
both the aircraft's stability and controllability.
This will be discussed in detail in lessons 4 and 5.
It's been explained how lift and drag both vary with speed and angle of attack.
To maintain straight and level flight at varying speeds, the pilot must fly the aircraft to maintain equilibrium between the
four forces acting on the aircraft.
There are three factors that determine the lift produced by the aircraft's wing.
These are the value of the dynamic pressure, a half rho V squared, which is, as you've learnt, displayed to
the pilot on the airspeed indicator,
the coefficient of lift of a particular aerofoil section, which increases with angle of attack up to the point of
stall, and the wing surface area.
We know that lift force is generated by the pressure differential produced between the upper and lower surfaces of the
wing.
Since pressure is force per unit area, the greater the wing surface area, the greater the lift.
This gives us the basic lift formula, which is, lift equals a half rho V squared times the coefficient of
lift times the surface area of the wing.
Let's take a closer look at the lift considerations for maintaining straight and level flight.
We'll assume that our aircraft has a mass of 700 kg. This makes its weight in SI units 6,867
N.
In straight and level flight, the lift force must, of course, also be 6,867 N to balance the aircraft's
weight.
We'll assume that the aircraft's wing area is 20 square metres, and that the air density is 1.225 kg
per cubic metre.
These two values are constants, so we don't need to consider them again.
In the present screen image, the aircraft is flying at 100 knots, that's 52 metres per second in SI units,
which are the units we have to use in order to make the lift equation balance.
Inserting this value into the lift equation, we see that, at 100 knots, the aircraft must fly at an angle
of attack to achieve a coefficient of lift of 0.2073.
100 knots would be a typical cruise speed for a training aircraft.
But if the pilot flies at a lower speed, obviously the dynamic pressure, half rho V squared, will reduce.
This must be compensated for by an increase in the coefficient of lift.
As we know, the only way the pilot can change the coefficient of lift is by changing the angle of
attack.
Consequently, the pilot must pull back on the control column and select a higher pitch attitude.
Now click on the decrease speed button and note the changes to the coefficient of lift and the angle of
attack as airspeed decreases.
We've learnt that the coefficient of lift of a wing increases only up to approximately 16 degrees angle of attack,
after which the wing stalls.
Therefore, the present airspeed of 37 knots, that is 19 metres per second, corresponds to the aircraft's minimum flight speed.
If the pilot were to attempt to select a higher angle of attack to fly at a slower speed, the
aircraft would stall.
We've just calculated, therefore, what is known as the stall speed.
Now select the increase speed button and you'll see that the angle of attack must be reduced as the aircraft
accelerates, if the pilot is to maintain straight and level flight, with the wing producing a constant lift force to
exactly balance the aircraft's weight.
The simple fact that, as an aircraft slows down, the pilot must select a higher pitch attitude, increasing the angle
of attack, and with an increase in speed, the pilot must select lower pitch attitude, reducing the angle of attack,
is fundamental to the understanding of the principles of flight.
If we consider that lift is good and drag is bad, and that both change with angle of attack, there
must be a point where we obtain the greatest value of the coefficient of lift for a minimum value of
the coefficient of drag.
The relationship between lift and drag will give an appreciation of the efficiency of the lift force produced, and the
comparison of the two is called the lift-drag ratio, a high lift-drag ratio being more efficient.
Let's look at the change in the coefficient of drag with angle of attack, and then add the lift curve
to the graph.
The lift and drag curves are not to scale, and no conclusions should be drawn from the intersection of the
plots.
The proportions of CL and CD can be calculated for each angle of attack, and plotted to give the lift
-drag ratio curve.
The maximum lift-drag ratio occurs at approximately 4 degrees angle of attack, and this angle is known as the
optimum angle of attack.
The maximum lift-drag ratio of a light aircraft is in the region of 10 or 15 to 1. This
means that for every unit of drag, lift will be 10 to 15 times greater.
It was previously shown that for an aircraft in straight and level flight, there was a corresponding speed at which
the aircraft must be flown for every angle of attack.
The speed at which any aircraft must be flown to obtain the optimum angle of attack occurs at the intersection
of the induced drag curve and the parasite drag curve.
This speed is called velocity minimum drag, or VMD, and is situated at the lowest point on the total drag
curve.
This speed is used for the glide and climbing.
The lessons covered so far in this course form the foundations for the following topics.
It's suggested that before continuing further, you ensure you have a good understanding of the principles thus far covered.
The take-off and landing distances of an aircraft depend on the minimum speed required to produce a lift force
equal to the weight of the aircraft.
This has been defined as the stall speed, and is determined by the value of C Lmax of the wing.
For safety, a small factor above the stalling speed is then applied to the take-off and landing speeds.
The highest value of C L would be achieved by having a thick cambered wing section.
This would provide the required value of lift at the lowest speed.
The aircraft at the top of your screen, which is fitted with a higher coefficient of lift wing, takes off
at a lower airspeed, and thus has a shorter take-off run.
For landing, the slower the speed the aircraft can be flown on approach, the shorter the distance required to bring
the aircraft to a full stop.
Take-off and landing only take up a small period of time when considering the whole duration of an aircraft's
flight.
The greater part of flight is spent in the cruise.
At high cruise speeds the cambered section would give high drag, and require a pronounced nose-down pitch attitude of
the aircraft.
It's usual to select a less cambered section aerofoil to optimise cruise, and then modify the section by having movable
leading and trailing edges to increase camber and C L so as to reduce take-off and landing distances.
This is the purpose of C L Max augmentation, or high lift devices as they are known. They reduce take
-off and landing distances.
The most common of the high lift devices used is the trailing edge flap.
A flap is a hinged section of the trailing edge of the wing, which can be deflected downwards and so
produce an increase in camber.
The flap seldom extends the whole length of the trailing edge, as flight controls also occupy this area.
The flap is located inboard of the flight control.
Just as the size of the wing is influential in the magnitude of the lift force produced by the wing,
so the size of the flaps will affect the increase in C L produced.
A compromise must be made between the size of the flap and the flight controls.
There is a combined flap and flight control system used on some aircraft, and this will be discussed in Lesson
5, Controls.
The flaps are operated by the pilot from the cockpit.
The controlling mechanism can be purely mechanical or powered electrically, hydraulically or pneumatically.
Try the electrical flap control on the screen for yourself.
The mechanical system provides the pilot with a lever, which is connected via rods or cables to the flaps.
The pilot provides the force to actuate the system.
To operate a powered system, the pilot simply operates a switch. This will then operate an electric, hydraulic or pneumatic
motor to power the system.
Few aircraft have flap settings of simply up or down. Most have a range of settings that allow the flap
to be selected in stages.
For example 10°, 25° and 40°. These values are actually the flap settings for the Piper PA-28 Warrior. Other
aircraft may have different settings.
The angle of deflection of the flap is measured from the chord line of the flap to the chord line
of the wing section.
Having varying settings requires the pilot to have an indication of the setting selected.
There may be markings at the base of the manual flap lever, detents on the switched powered system or a
gauge indicating flap position.
The various settings of the flap used depends on the flight conditions, and their sequence of operation will be explained
later.
The basic principle of the flap has been developed in many ways. It should be appreciated that, as well as
increasing the coefficient of lift, drag will also increase with use of the flap.
The more commonly used types of trailing edge flap will now be considered.
The plane flap is the type of flap which has been considered so far. It's of simple construction and gives
a good increase in CL max, but also gives a large increase in drag.
It's used mainly on low speed aircraft and where very short take-off and landing performance is not required.
The split flap forms part of the lower surface of the wing trailing edge, the upper surface contour being unaffected
when the flap is lowered.
The split flap gives a slightly higher CL max than the plane flap at higher angles of attack, but drag
is also higher as the depth of turbulent air behind the wing is greater.
The slotted flap is much more complex in construction. However, for the same area of flap it gives a bigger
increase in CL max for less drag than both the plane or split flap.
This is achieved by directing high pressure air from below the wing through the slots formed between the flap and
the wing trailing edge. This has the effect of re-energising the boundary layer, and so delaying separation.
The last type of trailing edge flap considered is the Fowler flap. This type of flap not only increases camber,
but also wing area.
Remember, lift can be calculated by the formula a half row v2 cls, where s is a surface area of
the wing.
The Fowler flap increases both the value of CL max and the surface area. The Fowler flap gives the largest
increase in CL max.
Because the thickness-cord ratio is reduced as the flap increases cord length, there is a reduction in drag produced
by the Fowler flap.
Larger aircraft may have slotted Fowler flaps. These are more efficient, but very complex in their construction.
A comparison of the lift and drag produced by the different types of flap for the same angle of attack
can be seen graphically.
Compared to the basic wing section, it can be seen that a Fowler flap can increase CL by more than
100% with the lowest increase in drag.
It has been stated that the point of CL max occurs just before the stall. Note what is happening to
the stalling angle of attack when flaps are extended.
Consider an aircraft flying close to the stall with its wing at CL max.
If flap is then lowered, the cord line will move and the angle of attack will increase.
It has been stated that a wing will stall if its angle of attack is increased beyond approximately 16 degrees.
To prevent the wing stalling, the angle of attack of the basic wing must be reduced.
Flaps reduce the stalling angle of attack of the wing.
As the aircraft must be pitched nose down to reduce the angle of attack with flaps selected, there will be
an improvement of the pilot's view of the ground in front of the aircraft.
This will be of particular benefit when approaching to land.
Whatever flight condition the aircraft is in, be it climbing, descending or straight and level,
if flying at the same speed the aircraft will always have a lower nose attitude with flaps extended compared to
flaps retracted.
The term clean configuration is used when describing an aircraft that does not have its flaps or undercarriage extended.
We have seen that the pitch attitude of the aircraft needs to be reduced on selection of flap to reduce
the angle of attack.
It is possible that the aircraft may actually make this pitch change by itself.
In lesson 2 we looked at the forces acting on the aircraft.
The relationship between lift and weight produces a nose down pitching moment.
Thrust, drag and the force from the tail plane balance this moment.
The increase in camber due to the use of trailing edge flaps not only increases the value of lift and
drag,
but also alters the pressure distribution around the aerofoil.
A Fowler flap will be used in the explanation of these changes as it clearly illustrates them graphically.
However, it should be understood that the effect applies to all trailing edge flaps.
In the clean configuration at CLmax the pressure distribution will be similar to that shown,
with the centre of pressure acting well forward on the cord.
On selecting flap it is at the rear of the wing that the major change in the pressure pattern takes
place.
The centre of pressure moves towards the trailing edge, altering the lift-weight couple and increasing nose down pitch.
Operation of trailing edge flaps will always produce a nose down pitching moment of the wing.
Since the wing is attached to the aircraft you might assume that the aircraft would pitch nose down,
but we must also consider the forces of drag and the effects of flaps on the tail plane before the
effect on the aircraft can be ascertained.
Selection of flap will also increase drag, and to maintain speed more thrust may be required.
This will alter the thrust drag couple.
Depending on whether the aircraft is of a high, low or mid-wing design will determine if the thrust drag
couple opposes or assists the wing pitching moment.
Selection of flap increases downwash, and this may well change the force of the tail plane.
The position of the tail plane, be it high or low mounted, will determine how much it will be affected
by the downwash.
In conclusion, you cannot say what will happen to the pitch attitude of an aircraft on selection of flap.
Every aircraft will respond differently depending on the location of the main wing, the position of the tail plane and
the position of the thrust line of the engine.
Whichever couple is the greater will determine the eventual pitch attitude change of the aircraft on selection of trailing edge
flaps.
Having studied the purpose, different types, control and effects of flaps on the aircraft, we'll now look at their operational
use.
It's been shown that flap increases both lift and drag, but sadly this is not in the same proportion.
Although the lift is the largest force, the proportional increase in drag is greater.
This reduces the maximum lift-drag ratio.
In Lesson 3, the lift-drag ratio was defined as a measure of aerodynamic efficiency.
In addition to takeoff and landing performance, flaps will also affect the range, climb and glide performance of an aircraft.
Lift-drag maximum occurs in the clean configuration at an angle of attack of approximately 4 degrees.
Use of flap will reduce this ratio and range, climb and glide performance will all reduce.
The greater the flap setting used, the greater will be the effect on these flight manoeuvres.
For takeoff, increasing CL to its maximum value with flap would give the lowest minimum flight speed, but drag would
also be at its maximum, and this could reduce acceleration so that the takeoff distance actually increases.
A lower flap setting than full flap is usually used for takeoff.
The pilot's operating handbook for a particular aircraft type will state the optimum flap setting to be used.
In this example of take-off distance against flap setting, the shortest take-off distance occurs at 25 degrees of
flap.
It can be seen that the take-off distance increases if a lower or greater flap setting is used.
It must be emphasised that this is only an example, and the pilot's operating handbook will specify the flap settings
for a particular aircraft type.
After take-off the flaps will be retracted to improve climb and cruise performance.
Because of the changes in CL, drag and pitching moments with the use of flap, the flaps should be retracted
in stages to prevent large changes to the attitude of the aircraft.
To prevent a sudden loss of lift as the flaps are retracted, the aircraft must first be accelerated with the
flaps extended. This is shown from points A to B on the accompanying graph.
As the flaps are retracted, the angle of attack must be increased to maintain the lift coefficient. This is shown
from points B to C.
As the flap is retracted, drag will reduce and the aircraft will accelerate, shown as points B to C on
the CL-CD curves.
The increase in angle of attack required as the flap retracts may only be required for a short time, depending
on how quickly the aircraft accelerates. However, it is important that the aircraft is not allowed to sink, especially when
close to the ground.
When landing, the high drag associated with the use of higher flap settings is of benefit to the aircraft.
It will assist in slowing the aircraft down on landing and reduce the landing distance. The higher values of CL
available will allow for the approach to be flown at the lowest possible speeds.
As with take-off, the flaps are moved in stages to reduce the effects of the change in lift, drag
and the pitching moments on the aircraft.
There is one additional consideration when landing with flap that is not relevant when selecting flap for take-off. This
is the speed of the aircraft.
Since lift is proportional to speed and CL, if flaps are selected at too high a speed, the large lift
force produced could overstress the airframe or flap operating mechanism.
To ensure the aircraft is not overstressed, flaps should not be selected until the aircraft has slowed down to a
specific speed. This speed is known as VFE, which stands for Velocity Flap Extension Speed.
It's marked on the airspeed indicator by a white arc. The flaps can safely be operated if the aircraft's speed
is within this arc.
The other colour arcs on the airspeed indicator will be discussed in the lesson on limitations.
The speed shown for this value of VFE is an example only and does not relate to any particular aircraft.
Although not a problem when initially selecting flap for take-off, the pilot must ensure that the flaps are retracted
before accelerating through VFE.
The pilot plans the approach of the aircraft to the airfield and gradually slows the aircraft down and progressively selects
more stages of flap in preparation for landing.
Remember that on selecting flap, the pilot must pitch the nose of the aircraft down. This will maintain the value
of CL by reducing the angle of attack. This will also improve the pilot's view of the airfield.
This is shown from point A to B on the graphic. As drag increases, the aircraft can be slowed to
point C on the lift curve.
The pilot should only select the full flap position once he is fully committed to landing.
If the flaps are selected and the pilot does not reduce the angle of attack, lift will for a short
moment be greater than that required and the aircraft would rise. The aircraft is said to balloon. This is not
a pleasant sensation for pilot or passengers.
The same term you recall was used in Lesson 3 to explain the effect of increased lift on entering ground
effect.
This completes the study of trailing edge flaps. We'll now look at leading edge high lift devices.
There are two types of leading edge high lift devices. These are the leading edge flap and slot or slat.
The leading edge flap is a complex system and generally is used on large jet transports. Light general aviation aircraft
are more likely to use slots or slats.
The main effect of leading edge devices is to delay separation of the air from the wing, increasing the stalling
angle of attack and the corresponding value of CLmax.
Note this is the opposite to trailing edge flaps, which reduce the stalling angle.
The principle of how the slot and slat achieve this delay in separation has actually already been explained when discussing
the slotted flap.
The slotted flap directed high pressure air from beneath the wing to energise the boundary layer on the upper surface
at the trailing edge.
The leading edge slot does exactly the same, but is of course positioned at the leading edge of the wing.
A fixed gap from the lower surface to the upper surface of the wing leading edge is formed from the
root to just before the wing tip. This forms the slot.
The horizontal stabiliser is also an aerofoil and as such can stall. Some aircraft may have slots on the tail
plane.
Let's look again at the change in pressure pattern around the basic aerofoil as it increases in angle of attack.
The point of lowest pressure moves forward. This means that the boundary layer flow rearwards from that point is flowing
against a continually rising pressure.
This is called an adverse pressure gradient.
The boundary layer finds it harder and harder to flow against this rising pressure and starts to separate from the
upper surface.
The further forward the point of lowest pressure the greater the adverse pressure gradient.
At the point of stall the large adverse pressure gradient causes most of the air on the upper surface to
separate.
The slot allows the energy of the boundary layer to increase and continue to flow against the adverse pressure gradient.
The angle of attack can then be increased and increase CLmax.
The slot will of course still direct higher pressure air to the upper surface even at low angles of attack.
This thickens the boundary layer and makes it turbulent. This increases drag which is not desirable at cruise speeds where
low angles of attack would be used.
This problem of the fixed slot can be overcome by having a movable slot. This is called a slat.
A slat is a small movable airfoil section attached to the leading edge of the wing. When deployed it moves
forward forming a slot.
Slats can be of the full span leading edge type or partial span.
The slat is either open or closed. There are no intermediate stages of selection like flaps.
Selection of the slat can be manual, by the pilot using a mechanical or powered control system, or as is
often the case on light aircraft, the selection is automatic.
The pressure changes around the leading edge of the wing are used to control the slats.
At low angles of attack the high pressure at the leading edge closes the slat.
As the angle of attack is increased the centre of pressure moves forward. Remember the slat is a small airfoil
in its own right and the pressure difference between its upper and lower surfaces pulls the slat open.
Operation of the slats causes only a small change to the pitch attitude of the aircraft compared to the pitch
change caused by operation of the trailing edge flaps.
This is because the centre of pressure does not move a great deal along the chord. The pressure pattern over
the wing is more evenly spread over the chord.
The disadvantage of the leading edge devices is that although they can increase CLmax to the same magnitude as trailing
edge flaps, it is achieved at higher angles of attack.
This gives the pilot very poor forward vision. For this reason leading edge devices are seldom fitted on their own.
In practice the leading edge and trailing edge devices are used together.
The pilot's view when flying at low speed is then greatly improved.
Very large increases in CLmax can be achieved by using leading and trailing edge devices.
This will then give the shortest take-off and landing distances, which you recall is the purpose of using high
lift devices.
That completes the lesson on lift augmentation devices. Now we'll study aircraft stability.
The dictionary gives the meaning of the word stability as being
firmly fixed, established not to be moved, not changeable.
When these definitions are applied to an aircraft it means that if an aircraft is flying in a certain direction,
at a certain pitch attitude, and a certain bank angle,
the aircraft will tend to stay flying in the selected direction, at the selected pitch attitude, and selected bank angle.
An aircraft which possesses stability will resist any change to its flight path,
and if temporarily moved from a steady state of flight, will tend to return to its original position without any
control input from the pilot.
An aircraft which demonstrates this ability is said to be stable.
Stability is inherent in the aircraft by design.
The level of stability of an aircraft varies depending on the type and role of the aircraft.
An aerobatic aircraft would tend to be less stable than an aircraft used for carrying passengers.
Reference in this lesson will be made to an aircraft being displaced in flight.
This displacement or movement of the aircraft could be caused by a sudden gust of wind or input from the
pilot's controls.
It should be understood that the force producing this displacement is external to the aircraft and of a temporary nature.
Static Stability
An aircraft possesses two types of stability. These are known as static and dynamic.
First we'll define static stability.
Static stability is defined as the initial response of an aircraft if it were to be displaced from its flight
path.
There are three different types of static stability.
The aircraft may have positive static stability, meaning it's stable.
It can be neutral, neutrally stable, or it can possess negative static stability, meaning it's unstable.
Consider the aircraft in the graphic being displaced in pitch.
If the aircraft tends to return to its original position after being displaced, it has demonstrated positive static stability and
is stable.
If the aircraft is again displaced, and this time the aircraft does not tend to return to its original position
but stays at this new position, it has demonstrated neutral static stability.
Starting again from the initial position of the aircraft in the graphic.
If the aircraft is again displaced, but now the aircraft continues to move further away from its original position, the
aircraft has now displayed negative static stability.
It's unstable.
The static stability of the aircraft is determined by the aircraft's initial response to its displacement.
The dynamic stability is determined by the aircraft's response to displacement over a period of time.
Viewing the graphic from a distance, let's now consider the dynamic stability, which, like the static stability, can be positive,
neutral, or negative.
The aircraft is displaced.
It must be remembered that the aircraft is also flying and will move forward.
If the aircraft has positive static stability, it will try to return to its original attitude by pitching down.
Because it's moving, the aircraft possesses kinetic energy and momentum, and it's likely to fly through the line of its
original flight path and oscillate about that line.
If the aircraft has positive dynamic stability, these oscillations will reduce with time and the aircraft will settle in its
original attitude and flight path.
The aircraft has neutral dynamic stability if the oscillations continue and stay at the same magnitude regardless of time once
the aircraft is displaced.
Negative dynamic stability would be shown by an aircraft if, over time, the magnitude of the oscillations increased and the
aircraft moved further and further away from its initial attitude and flight path.
Note that in all these examples, the aircraft demonstrated positive static stability. It tried to return to its original attitude.
An aircraft that has positive dynamic stability must have positive static stability.
The reference axis
The aircraft has three reference axes about which it manoeuvres.
These are known as the longitudinal, lateral, and normal axis.
They all pass through the centre of gravity of the aircraft.
The longitudinal axis runs through the centre of gravity from nose to tail. The aircraft rolls about this axis.
When considering the stability about the axis, to avoid confusion, consider the motion about the axis. The aircraft is stabilised
laterally about the longitudinal axis.
The lateral axis passes through the centre of gravity and runs parallel to a line from wingtip to wingtip. The
aircraft pitches nose up or down about this axis.
Longitudinal stability is about the lateral axis.
The normal axis passes vertically through the centre of gravity at 90 degrees to the longitudinal axis.
Movement of the aircraft about this axis is called yawing.
Directional stability is the term used to express stability about the normal axis.
The magnitude or type of stability about each axis of the aircraft may not be the same.
Generally, an aircraft will be designed with strong longitudinal stability, reduced but still positive stability about the normal axis, and
weak, possibly even neutral, lateral stability.
An aircraft should demonstrate the following qualities during development before it can be certified for general use.
It must possess adequate stability to maintain steady state flight throughout the aircraft's speed range, but also allow proper response
to the flight controls, allowing the required controllability and manoeuvrability of the aircraft.
This leads to a compromise depending on the aircraft's role.
A very stable aircraft would offer a large resistance to the pilot's flight controls and reduce controllability and manoeuvrability.
The pilot would physically have to use a lot of force to move the controls.
The complete opposite will occur if the aircraft has little or no stability, i.e. it's neutral or unstable.
The aircraft would be very manoeuvrable and the flight controls very easy to move.
If we cast our minds back to our childhood, we may recall that we have actually come across the principle
of stability before.
A young child learning to ride a bicycle may benefit from the fitting of stabiliser wheels to help him keep
his balance until he's learnt to balance his bicycle unaided.
With the bicycle upright and in balance, the stabiliser wheels do not touch the ground.
Ignoring any gyroscopic effects from the rotating wheels, the downward forces are balanced by upward forces and the sum of
all moments are equal to zero.
The cycle is in equilibrium.
If balance is lost, the cycle will lean in one direction and the stabiliser wheel will contact the ground.
The displacement of weight on the cycle produces a turning moment, force times distance about the pivot point formed at
the junction of the main wheel and the ground.
The stabiliser wheel produces a moment in the opposite direction preventing the cycle and rider from falling over, preventing the
cycle becoming unstable.
The rider does not require any special skills to achieve this level of stability.
The design of the stabiliser wheel produces the necessary forces when needed automatically.
You will have already worked out for yourself that the type of stability possessed by the bicycle stabilising system is
neutral stability.
You'll be pleased to learn that there are no stabilising wheels fitted to your aircraft for flight.
Aircraft stability is dependent on the use of aerodynamic forces to produce the necessary turning moments, and just like the
cycle stabilising wheels, they do so automatically.
We'll first consider the longitudinal stability of the aircraft, as it can be studied in isolation.
Lateral and directional stability tend to interact, as will be explained later.
The aircraft is stabilised longitudinally about the lateral axis. This maintains pitch attitude.
It's been stated that as angle of attack of an airfoil increases, so the centre of pressure moves forward, and
the magnitude of lift increases.
To simplify the understanding of stability, it's convenient to consider the changes in magnitude of the lift force as angle
of attack increases, acting through a stationary point.
This point is called the aerodynamic centre, or AC.
In reality, we've seen that lift acts aft of the aircraft's centre of gravity, but again, to simplify the subject
of stability, we'll consider the aerodynamic centre being forward of the centre of gravity.
A wing, when considered alone, is statically unstable.
A momentary gust would increase the angle of attack and increase lift.
When multiplied by the distance x, this generates a nose-up pitch, tending to increase the angle of attack further.
Now consider a wing, together with a tailplane.
The tailplane, or horizontal stabiliser, is much smaller than the main wing, and as such produces a much smaller lift
force.
It is, however, positioned at a greater distance from the centre of gravity, and generates a stabilising pitching moment.
If the wing and tailplane are now subjected to a momentary gust, the angle of attack, and hence the lift,
will increase on them both.
As the change in the tail moment is greater than that of the wing, the sum of the moments is
no longer zero.
A nose-down pitching moment will be produced by the tail.
Thus, the aircraft will return to its original attitude, and has demonstrated positive static longitudinal stability.
In terms of longitudinal stability, the moments produced about the centre of gravity of the aircraft depend on the aerodynamic
forces produced at the wing and tailplane, L and LT in the diagram,
and the distance from the centre of gravity at which each force acts.
The centre of gravity of an aircraft is not fixed. Variables like the mass of the pilot, passengers, fuel, oil
and baggage, as well as their distribution within the aircraft, will affect its position.
This subject is covered fully in the Mass and Balance lesson on Disc 6 of this series.
Consider the centre of gravity in the graphic being moved forward.
Note that the moment arm, X, for the wing, has decreased, and the moment arm, Y, for the tailplane, has
increased.
For the same gust and change in angle of attack, the change in lift force will not alter, but the
turning moments will be affected by the change in the length of the moment arms.
Clearly, if the aircraft was stable before we move the centre of gravity forward, the greater turning moment produced by
the tailplane will increase the static stability.
Since increased stability reduces manoeuvrability and increases the force required by the pilot to move the controls, a point must
be reached where the forward position of the centre of gravity will increase stability to a value where the aircraft
cannot satisfy the requirements of certification.
A limit will therefore be set for the most forward position of the centre of gravity which ensures the minimum
requirement of manoeuvrability with proper response to the flight controls.
If the centre of gravity is moved aft, the static stability reduces as the moment arm of the tailplane decreases.
Eventually, a point will be reached where the turning moment of the tailplane is reduced to a value where it
only just balances the turning moment of the wing.
This is called the neutral point.
If this were the case, following any displacement in pitch due to a gust, say, the correcting moment of the
tailplane would not be great enough to return the aircraft to its original pitch attitude.
Because the wing turning moment and the tailplane turning moment would now be equal, the aircraft would stay in the
new pitch attitude to which it had been displaced.
As stability decreases, controllability and manoeuvrability would increase, requiring reduced forces from the pilot to move the flight controls.
If the centre of gravity was moved further aft than the neutral point, the aircraft would become unstable, possessing negative
static stability.
To meet the requirements for certification, the aircraft must have a level of stability longitudinally.
For this reason, an aft limit for the centre of gravity is set forward of the neutral point.
This will provide the minimum positive static stability with controllability and manoeuvrability being at their maximum.
This then gives a range of centre of gravity positions for the aircraft from the forward to aft limit.
The centre of gravity must lie at some point within this range for the take-off, cruise and landing of
the aircraft.
For this reason, the pilot must calculate the centre of gravity position for all stages of the flight to ensure
stability and controllability of the aircraft remain acceptable.
In considering the longitudinal stability of the aircraft, an assumption was made that a gust affecting the angle of attack
of the wing would produce the same change of angle of attack on the tailplane.
In Lesson 3 of this CD, downwash was explained and its effect on the tailplane shown.
Downwash will affect the lift produced by the tailplane.
This in turn affects the turning moment produced opposing the wing moment.
Increased downwash reduces the longitudinal stability of the aircraft.
Downwash increases at low speeds and with the use of flap.
Under these flight conditions, the static longitudinal stability of the aircraft will reduce.
The aircraft will be in this configuration during approach to landing and take-off.
The aircraft must still demonstrate positive static longitudinal stability under these conditions.
You may recall that ground effect reduces downwash.
As the aircraft enters ground effect, the tailplane's angle of attack will be less affected by downwash.
This will cause the longitudinal stability to increase.
An aircraft with positive longitudinal stability is very desirable, as it reduces the pilot's workload.
As previously stated, longitudinal stability is normally the strongest of the stabilised axes.
For the requirements of this course, no further understanding of dynamic stability is required.
The aircraft is stabilised laterally about the longitudinal axis.
The movement about this axis is roll.
The definitions of positive, neutral and negative static stability have not changed.
If the aircraft is displaced in roll and then returns to its original position without any input from the pilot,
it is stable and has positive static lateral stability.
If the aircraft stays in its new position, it has neutral lateral stability.
If the aircraft angle of bank keeps increasing with no tendency to return to its original position,
the aircraft possesses negative static lateral stability and is unstable.
Several factors need to be taken into account when deciding how stable an aircraft should be in roll.
For instance, for certain flight manoeuvres, such as crosswind take-offs and landings,
high positive lateral stability may cause control difficulties.
Generally, therefore, good handling qualities are obtained with a lower value of positive lateral stability.
For certain flight manoeuvres, for example crosswind take-offs and landings,
strong positive lateral stability can cause complications and is undesirable.
Generally, good handling qualities of the aircraft are obtained with relatively weak positive stability.
The principal component of the aircraft contributing to lateral stability is the wing.
Other parts of the airframe do have an effect on lateral stability,
but they are minor compared to that of the wing.
To understand how the aerodynamic forces are produced by the wing to stabilise in roll,
we need to study the changes that occur to airflow as the aircraft is displaced in roll.
If the aircraft is displaced in roll by one wing dropping and the other rising,
the lift will be inclined from the vertical.
Lift is now no longer directly opposing weight,
and the resultant of the lift and weight forces causes a sideways movement of the aircraft
in the direction of the lower wing.
This is called a side-slip, and a relative airflow in opposition to this movement will now affect the aircraft.
The changes in airflow affecting the aircraft produce aerodynamic forces
that will roll the aircraft back to its original wing's level position.
A stabilising turning moment is produced.
The position and the angle formed at the junction of the wing and fuselage
greatly affect the stabilising moment produced.
We'll consider the differences between high and low wing aircraft.
But first, we'll look at the effect of what is called wing dihedral.
Greater stabilising rolling moments can be produced
if the plane of each wing is angled positively above a datum line
parallel to the lateral axis passing through the wing root.
This is called wing dihedral.
The greater the angle of dihedral of the aircraft,
the greater will be the increase in lateral stability.
To explain how wing dihedral increases lateral stability,
we'll initially look at a wing without a fuselage.
In level flight, the lift produced from each side of the wing is equal.
If the wing is displaced, the subsequent airflow due to the side-slip
causes an increase in the angle of attack of the lower wing and increases its lift.
The higher wing is subjected to a reduced angle of attack
and its lift is reduced.
A stabilising moment has been produced
and the wing returns to its original position.
The greater the dihedral angle, the greater the change in angle of attack,
lift and stabilising moment.
So far, the graphics used for illustrating dihedral
have all been on low wing aircraft.
The low wing aircraft has a tendency to be unstable laterally
for the following reasons.
Due to the position of the wing in relation to the fuselage,
the relative airflow due to side-slip
flows down around the fuselage on the lower wing
decreasing its angle of attack
and flows up towards the higher wing increasing its angle of attack.
This does not produce the stabilising moments required.
The other problem with the low wing aircraft
is that if the centre of gravity is above the wing,
the force of the weight which always acts straight down
produces a destabilising turning moment.
This is partly offset by any structure like the fuselage or fin
which is above the centre of gravity
which produces a stabilising moment.
In order that low wing aircraft may have the correct degree of lateral stability
they tend to have wings with pronounced dihedral.
This can be seen in this picture of the Piper PA-28 Warrior.
If we now consider a high wing aircraft,
you'll see that the effects are completely the opposite.
The angle of attack is increased on the lower wing
as the relative airflow due to the side-slip
flows up around the fuselage.
This produces a stabilising moment.
The fuselage and fin above the centre of gravity
are again stabilising,
but the high wing aircraft also benefits from the weight of the aircraft
producing a stabilising moment.
This is known as the pendulum effect.
High wing aircraft have so much inherent lateral stability
that pronounced dihedral angles are not required.
You can see in this picture of the Cessna 172
that dihedral is hardly apparent.
Note the difference in dihedral of the high wing Cessna
compared to the low wing Piper aircraft.
The study of high speed flight is not a requirement for the private pilot.
However, you may have seen the swept back wings used on high speed aircraft.
Swept back increases lateral stability.
For this reason, some aircraft,
and especially those with high mounted swept wings,
have negative dihedral,
so as not to have too high a value of lateral stability.
This negative dihedral is known as anhedral.
You'll recall that flaps reduce to longitudinal stability
due to their influence on downwash.
Lateral stability is also reduced with the use of partial span flaps.
The selection of partial span flaps
causes the lift on the inboard section of the wing to increase.
This reduces their spanwise moment arm.
Therefore, any changes in lift due to sideslip
occur closer inboard and the dihedral effect is reduced.
This will reduce lateral stability.
The propeller slipstream at high power settings and low forward speeds
also reduces the dihedral effect
as the inboard wing sections become more effective.
The reduction in the dihedral effect is most critical when flap
and the power effects of the propeller are combined
as in the powered approach to landing.
The reduction in lateral stability would be conveyed to the pilot
as a reduction in control force to manoeuvre the aircraft in roll.
It has been shown that an aircraft primarily achieves static lateral stability from its wings.
This is due to the increase in lift produced by sideslip.
From our previous lessons, we learnt that as lift increases, so does drag.
This causes the aircraft to behave slightly differently than so far discussed.
Consider an aircraft that has dropped a wing and is about to be returned to its original position
by the increased lift produced from the dihedral effect.
The increased lift on the lower wing will also produce an increase in drag on that wing.
compared to the opposite wing.
If we now look at the aircraft in plan view,
it can be seen that a turning moment has been produced about the centre of gravity of the aircraft.
This movement of the aircraft is about the normal axis and is called yaw.
This interaction between roll and yaw will always occur.
The magnitude of yaw depends on the magnitude of lateral stability.
For this reason, lateral stability should be positive, but light.
The aircraft is stabilised directionally about the normal axis.
Movement about this axis is called yaw.
The primary component of the aircraft providing directional stability is the tail fin,
also called the vertical stabiliser.
Directional stability is built into the aircraft.
The term weathercock is often used to refer to the aircraft tendency to turn nose into wind.
The tail fin is a symmetrical aerofoil.
With the relative airflow straight onto the tail fin, the fin is at zero degrees angle of attack.
No aerodynamic force except drag will therefore be produced by the fin.
As with lateral stability, the term sideslip is used when considering the aircraft being displaced in yaw.
If displaced to the left, the aircraft is sideslipping right,
and sideslipping left when yawed to the right.
The aircraft sideslips opposite to the direction of yaw.
As you can see from the graphics, the sideslip changes the angle of attack on the fin.
This will generate an aerodynamic force at a distance from the centre of gravity.
A stabilising turning moment is produced, and if strong enough,
will return the aircraft to its original position.
The aircraft has demonstrated positive static directional stability.
The terms for neutral static and negative static directional stability
would imply that the moment produced was only equal to the yawing moment when neutral,
and less than the yawing moment when negative or unstable.
An aircraft must demonstrate positive static directional stability.
The turning moment produced by the tail fin depends on three factors.
The sideslip angle or change in angle of attack of the fin,
the size and efficiency of the profile of the airfoil section used,
and the length of the moment arm from the centre of gravity.
Of these factors, as the pilot of the aircraft, you have control of only the centre of gravity position.
This we know must be within certain limits to satisfy the requirements of longitudinal stability.
The further forward the centre of gravity, the greater the directional stability.
The fin is an aerofoil, and as such, its angle of attack can be increased up to the point of
stall.
This imposes a limit to the sideslip of an aircraft,
as the fin could stall and the aircraft become directionally unstable.
The efficiency of the fin depends on its size, plan form, and aspect ratio.
Just like the efficiency of the main wing.
In Lesson 3, we looked at how modifying the wing tips of the main wing reduced induced drag and improved
efficiency.
The same result can be achieved by having a high mounted tail plane on the fin.
You may have seen some aircraft with more than one tail fin.
This would suggest that the aircraft required a greater surface area
than could be reasonably provided by a single fin to provide sufficient directional stability.
Another way of overcoming directional instability is with the use of dorsal and ventral fins.
These are small aerofoil sections positioned above and below the fuselage.
They increase what is called the keel surface area behind the centre of gravity.
The ventral fin is especially of benefit to aircraft when they are flown at high angles of attack.
The design of the fuselage masks the fin from the airflow.
The ventral fin is very evident on this BAE Bulldog aircraft.
At high angles of attack, the drag will be high with a low forward airspeed.
The changes the pilot may be required to select in power and propel a slipstream alter the airflow over the
fin, causing large changes directionally of the aircraft.
We have already seen how rolling movements produce associated yawing motion.
Let's now examine how yawing movements will likewise generate roll.
The stabilising moment produced by the fin causes one wing to advance as the aircraft returns to its original position.
This increases the speed of air over the wing and increases lift.
If we now view the aircraft from the front, it can be seen that the increase in lift from the
advancing wing produces a rolling moment.
The roll can, of course, lead to a side slip, which, as we've learnt, will in turn cause the aircraft
a yaw towards the lower wing.
Although lateral and directional stability have been studied in detail, individually it's clear that they are interconnected.
A corrective yaw will generate a rolling moment.
Likewise, during a side slip the inherent lateral stability of an aircraft will generate both rolling and yawing moments.
Lateral and directional stability have become what is called coupled.
If directional stability is very high compared to lateral stability, the aircraft will experience spiral divergence, slowly turning and descending
in flight.
If the lateral stability is very high compared to the directional stability, the aircraft suffers what is called Dutch roll.
This is a very uncomfortable flight experience for pilot and passengers as the aircraft slowly oscillates in roll and yaw.
Dutch roll is objectionable and spiral divergence tolerable.
The requirement for lateral stability should, as already stated, be sufficient to provide positive stability,
and the directional stability should be adequate to be positive without producing Dutch roll.
At the beginning of this lesson, it was stated that stability is inherent in the aircraft by design.
The study of mass and balance in this series of lessons will ensure that you, the pilot, do not exceed
limits which will reduce stability and controllability of the aircraft.
All aircraft are fitted with a control system to enable the pilot to manoeuvre the aircraft in flight about each
of its three axes.
Control is required about the longitudinal, lateral and normal axis.
Movement of the aircraft is around the centre of gravity.
Control about the axes is achieved by altering the aerodynamic forces at the air force sections of the wingtips, tail
plane and tail fin of the aircraft.
These extreme positions are used because they give the greatest moment arm about the centre of gravity,
and for a given force will generate the largest turning moments.
On general aviation aircraft, the aerodynamic forces are normally changed by altering the camber or angle of incidence of the
air force sections.
The pilot can change the camber of the wingtip, tail plane or fin by deflecting a movable flap with the
use of a control system from the cockpit,
and alter the coefficient of lift of that section for a given angle of attack.
On some aircraft, the angle of incidence of the complete aerofoil is moved to achieve control in flight.
The all moving tail plane is an example.
Changing the angle of incidence is again controlled by the pilot from the cockpit,
and changes to coefficient of lift occur as the angle of attack alters.
The aerofoil sections controlled by the pilot are known as the flying control surfaces.
The magnitude of the force they produce is the product of the angular displacement of the control surface and the
dynamic pressure.
The angular deflection of the control surfaces is limited in range by mechanical stops.
There are usually three separate control systems and three sets of control surfaces.
The aircraft is primarily controlled in yaw by the rudder.
This provides control about the normal axis.
The rudder is a flap type control and is operated from the cockpit by the pilot's feet moving the rudder
bar,
which is connected to the rudder by cables or rods.
Pushing the rudder bar forward with the right foot moves the rudder to the right,
generating an aerodynamic force and the aircraft will yaw to the right.
If the rudder bar is moved forward with the left foot, the reverse happens and the aircraft will yaw to
the left.
It should be appreciated that the yaw of an aircraft is different from a turning manoeuvre.
Longitudinal control is achieved by the use of an elevator,
which is a flap type control, or by the use of a variable incidence tail plane,
also called a stabilator, or all moving tail plane.
These surfaces control the aircraft in pitch about the lateral axis.
The pilot controls the elevator, or stabilator, by the use of a yoke, or control column.
The control column is sometimes simply called the stick.
This control is connected via cables or rods to the flying control surface
and is moved in a fore and aft direction by the pilot to control the aircraft in pitch.
If the pilot pulls the yoke or stick towards him,
then the elevator moves up, as does the trailing edge of the stabilator.
The degree of movement of the cockpit control directly affects the angular displacement of the control surface.
No attempt is made in this graphic to show the operation of the control run.
This movement produces an aerodynamic downforce on the tail which will pitch the aircraft nose up.
If the pilot pushes the yoke or stick away from him,
then the elevator will move down and the trailing edge of the stabilator is lowered.
An upforce will now be produced on the tail which will pitch the nose of the aircraft down.
Lateral control about the longitudinal axis is primarily achieved by the ailerons.
The aircraft rolls about this axis.
There is an aileron positioned outboard on the trailing edge of each wing.
The ailerons are also operated by the control yoke or stick.
The control movement is rotated left and right for the yoke
and moves sideways left and right for the control stick.
The ailerons are flap-type controls but their angular movement is in opposition to each other.
If one aileron is angled up, the other is angled down.
Consider the aircraft in the graphic.
The left wing has the down-going aileron.
This will increase the camber and increase lift.
The up-going aileron on the right wing has a reduced camber and its lift will be reduced.
The lift forces are now out of balance and the aircraft will roll to the right.
It will continue rolling to the right until the ailerons are returned to the neutral position by the pilot.
Movement of the control yoke or stick is instinctive.
If the pilot wishes the aircraft to roll right, the control yoke or stick is moved to the right.
Move the controls to the left and the aircraft will roll left.
The down-going wing determines the direction of roll and will have its aileron deflected up.
In Lesson 4 on Lift Augmentation, reference was made to a combined flap and flight control system.
This is called a flaperon.
As both the aileron and the flap occupy part of the wing trailing edge,
a compromise on the size of each component and therefore its effectiveness must be made.
The solution is to use full span flaperons that can droop down symmetrically as flaps or move differentially as ailerons.
The large flap area gives the aircraft excellent short field performance.
For roll control, the flaperon is moved differentially.
This reduces lift on one side of the wing, producing a powerful rolling moment.
The use of flaperons on an aircraft does not change the pilot's flight controls or the way they are operated.
However, the actual control linkage that operates the system is quite complex.
The controls are designed to give a turning moment around a particular axis,
but may additionally give a moment around a second axis.
This coupling occurs with the yawing and rolling moments and produces a secondary effect to the rudder and aileron controls.
There is no secondary movement associated with the elevator or stabilator,
but there will be a change in speed if the aircraft is pitched up or down.
If the aircraft is yawed to the left with left rudder, the advancing right wing will experience an increase in
airspeed,
increasing its lift compared to the left wing.
This difference in lift will produce a rolling moment to the left.
The secondary effect of the rudder is to produce roll in the direction of yaw.
We now look at the secondary effect of ailerons.
Operation of the ailerons to roll the aircraft to the left increases the lift on the right wing and reduces
lift on the left wing.
As bank increases, the aircraft begins to slip towards the lower wing.
The relative airflow, therefore, now has a considerable sideways component.
Because the keel or side surface of the aircraft behind the centre of gravity is greater than that in front
of the centre of gravity,
the aircraft will yaw to the left as it slips.
The secondary effect of aileron, then, is to produce yaw in the direction of roll.
Let's now take a look at another secondary effect of roll, adverse yaw.
This phenomenon will explain to you why the ailerons are usually operated by the pilot together with rudder.
In rolling the aircraft to the left, we've seen that the lift increases on the right wing and decreases on
the left wing.
Our knowledge of principles of flight tells us that there must be a corresponding increase in drag on the right
wing and a decrease in drag on the left.
This generates a couple, which induces a yawing motion in the opposite direction to the roll.
This is called adverse yaw.
So, as you can see from the screen, when the aircraft rolls to the left, it yaws to the right.
And when it rolls to the right, it yaws to the left.
As a pilot, you can eliminate adverse yaw by applying rudder in the same direction as aileron.
That is, through coordinated movements of the ailerons and the rudder.
The degree of cross-coupling of the controls varies from aircraft to aircraft.
On some aircraft, the effects may be acceptable. On others, they may not.
We'll now consider various ways of reducing these effects.
One method of counteracting the secondary effects of yaw and roll is to interconnect the rudder and aileron controls to
automatically apply the required control correction.
Adverse aileron yaw can also be reduced by the use of differential and freeze type ailerons.
The principle of operation of differential ailerons is that the control linkage causes the up-going aileron to move through
a larger angle of deflection than the down-going aileron.
The greater angle of deflection of the up-going aileron in the differential control system will increase parasite drag more
than on the down-going aileron.
The increase in parasite drag on the up-going aileron helps compensate for the increase in induced drag associated with
the extra lift produced by the down-going aileron.
The difference in total drag between the two wings is thus reduced. Consequently, adverse yaw is also reduced.
Freeze type ailerons achieve the same balance in drag by designing the leading edge of the flying control surface so
as to extend below the lower surface of the wing on the up-going aileron and increasing its drag.
The down-going aileron remains shrouded and produces less drag.
Some aircraft may employ a control system using the two principles together and have differential freeze type ailerons.
Picture the flying control surface as an aerofoil isolated from the main aerofoil.
The aerodynamic force that it will produce would depend on the following parameters which make up the lift equation.
The dynamic pressure, half rho v squared, which is displayed to the pilot in the form of the indicated airspeed.
The dynamic pressure is often given the symbol Q.
The coefficient of lift, its angle of attack changed by the pilot operating controls.
And S, the surface area of the control.
In the lift equation, note that Q and a half rho v squared are used interchangeably as they are two
ways of referring to the same thing, that is, dynamic pressure.
The size or surface area of the flying control is set during construction.
This leaves two variables, the value of Q, which changes with indicated airspeed, and the value of the coefficient of
lift, altered by deflection of the flight control.
Consider an aircraft flying at 80 knots and the pilot wishes to pitch the aircraft nose up.
A value of aerodynamic force will be produced and a turning moment generated about the centre of gravity.
If the speed is now increased to 100 knots, dynamic pressure will increase and for the same control deflection a
greater aerodynamic force will be produced, generating a larger nose up pitching moment.
If the pilot wishes to maintain the same change in pitch at the higher speed, the degree of control deflection
would have to be reduced.
At low airspeeds, to obtain the required control response, large movements of the flight controls are generally required.
As airspeed increases, the controls will become more effective and less movement of the control surface is required to manoeuvre
the aircraft.
At low speed, the terms sloppy, sluggish and mushy are often used to describe less effective operation of the controls.
However, at high speed the controls can be so effective that they can generate forces beyond those which the aircraft
was designed to withstand.
These limitations are covered in Lesson 7.
An aircraft on the approach to land will be flying at low speed. This would imply less effective, sloppy controls.
With flaps and undergarage selected, drag will be high, requiring high levels of thrust to maintain balanced flight.
The increased propeller slipstream will energise the boundary layer across those control surfaces over which it flows.
On most single engine propeller aircraft, this means that the rudder and elevator or stabilator have increased control effectiveness under
these flight conditions.
The ailerons however, which are outside the influence of the slipstream, would still be sloppy.
The forces generated by the flying control surfaces are of sufficient magnitude to manoeuvre the aircraft around the sky.
By the nature of the design of the flying control surface, the aerodynamic force acting from its centre of pressure
will try to rotate the control around its hinge point in the direction of the force.
This is called the hinge moment.
To move the flying control surface to the required angular displacement and maintain it in that position, the pilot has
to overcome and then balance the hinge moment by applying a force to the cockpit control.
This is called stick force.
Note that this term is still used, even if talking about the rudder control.
If the hinge moment increases, stick force will increase.
This relationship between stick force and the hinge moment gives the pilot a feel for the aerodynamic forces being produced
on the aircraft.
At high speed, the stick force should be relatively high, indicating to the pilot that if he persists in trying
to displace the control to its full deflection, something might break.
Larger and faster aircraft can produce hinge moments that are too powerful for the pilot to easily overcome and balance.
The pilot will require assistance to move the controls.
This can be done either by using powered flying control systems or by aerodynamic balancing.
Modern large aircraft have flight control systems that are normally operated by hydraulic power.
The aircraft you're likely to fly will have a mechanical control system where you, the pilot, provide the motive force.
Although the force required to move the controls of a light aircraft is generally small, the effort required by the
pilot may become fatiguing during long flights at high cruise speeds, especially in turbulent weather conditions.
For this reason, controls are often balanced.
The most common and simplest principle of balancing applied to light aircraft, which helps the pilot move the controls but
still leaves him with a reasonable amount of feel, is the principle of aerodynamic balance.
Aerodynamic balance involves using the aerodynamic forces of the control surface to reduce the hinge moment.
It can be achieved in several ways.
The first method we shall look at is called the inset hinge.
Setting the hinge back in the control surface puts it nearer the centre of pressure, causing the hinge moment to
be reduced.
Furthermore, the airflow strikes the surface forward of the hinge, exerting a pressure force at F1 which opposes the hinge
moment and helps the pilot move the control.
You will appreciate that if the centre of pressure were to move forward of the hinge, there would be a
danger that all control feel would be removed, taking control out of the pilot's hands.
Consequently, not more than one fifth of the control surface may be forward of the hinge.
Similar in operation to the inset hinge is the horn balance.
The horn is part of the control surface and is forward of the hinge line.
In flight, when the control surface is displaced as shown in the diagram, aerodynamic forces will be generated both forward
and behind the hinge line.
This produces turning moments about the hinge, reducing the overall hinge moment and stick force.
Control effectiveness is unaltered.
Design of the horn balance will mean that during operation the horn extends above or below the surface of the
main aerofoil section.
This will cause an increase in drag.
The same principle is applied without the increase in drag by the use of an internal balance.
The internal balance is enclosed inside the rear of the main aerofoil section.
This forms a chamber that will feel the same changes in pressure as are produced on the control surface.
As shown in the graphic, the pressure differential inside the chamber will produce a moment in opposition to the hinge
moment.
The method of achieving aerodynamic balance discussed so far have had no effect on the force the control surface produces.
Stick force can also be reduced by small aerofoil tabs positioned at the rear of the control surface, but these
do alter the effectiveness of the flying control.
There are three main types of trailing edge tab device.
These are the balance tab, the anti-balance tab and the servo tab.
These tabs are small aerofoil sections hinged at the trailing edge of the flying control surface.
Their actual size will vary from aircraft to aircraft.
We'll first consider the balance tab.
The pilot has no direct control over tab movement.
The pilot's inputs to the flying control are transmitted by linkage to move the balance tab in the opposite direction
to the flying control surface.
The pilot moves the control surface, the control surface moves the tab.
The balance tab generates an aerodynamic force in the opposite direction to the flying control surface.
This reduces the hinge moment and stick force, but will also give some reduction in control effectiveness.
Aircraft fitted with large elevators or stabilators are capable of generating very powerful aerodynamic forces due to their large surface
area.
For relatively small control deflections, especially at high speeds, the forces produced could cause the pilot to over control.
If this were the case, an increase in stick force would reduce the possibility of over controlling.
This can be achieved by the use of an anti-balance tab.
You can see from the graphic the similarity between a balance and an anti-balance tab.
The difference is that the arrangement of the control linkage causes the anti-balance tab to move in the same
direction as the flying control.
It can be seen from the graphic that both the tab and the flying control surface force act in the
same direction.
This increases stick force and control effectiveness.
Anti-balance tabs are commonly used on all flying tail planes, otherwise known as stabilators.
The last type of tab control is the servo tab.
This differs from the other tab controls in that the pilot actually controls the tab and not the main flying
control surface.
Movement of the tab generates an aerodynamic force which moves the main flying control.
View the sequence of events in the graphic as the pilot selects a nose up pitch.
Note that to pitch nose up the tab moves down to cause the main flying control to move up.
This does not affect the way the pilot operates the cockpit controls.
Servo tabs tend to be used on larger aircraft, where the weight due to the physical size of the flight
controls, let alone their aerodynamic forces, would produce very large stick forces.
The various types of aerodynamic balance discussed can be used on any of the flying control surfaces.
They are more likely however to be found on the elevator or rudder of the aircraft.
Remember, these devices are used to reduce stick force to acceptable levels, yet still give the pilot feel for the
aerodynamic loads on the aircraft.
The aircraft designer ensures the stick force increases as the forces on the aircraft's flight controls increase.
This feel helps prevent the pilot trying to maneuver the aircraft beyond the structural limits of the airframe.
There is another form of balancing applicable to flying control surfaces, but this has nothing to do with assisting the
pilot to move the controls, or to reduce stick force.
It's called mass balancing, and it's used to prevent control surface flutter.
Flutter is a phenomenon associated with the higher speeds of the aircraft, and is an oscillation of the control surface
coupled with an oscillation in bending or twisting of the wing, bin, or tail plane.
Flutter can develop through poor maintenance of the aircraft's flight control system as a result of free play between the
cockpit control and the flying control surface.
Ignoring this factor, flutter is likely to occur if the position of the centre of gravity of the control surface
is behind the hinge line.
Consider the control surface in the graphic, being deflected down.
We know this will produce an aerodynamic force upwards.
This makes the main aerofoil rise, taking the hinge line up with it.
With the centre of gravity behind the hinge line, inertia, its reluctance to move, causes the control surface to lag
behind.
Eventually, the stiffness of the structure of the control surface causes it to flick upwards.
Its momentum carrying the control surface beyond the required position, and producing a downforce about the hinge.
The hinge moves down, with the control surface again lagging behind.
This cycle of oscillations continues, and can cause structural failure.
The problem can be resolved by moving the centre of gravity onto or just forward of the hinge line by
adding weights to the control.
This reduces the inertia moments about the hinge line.
Mass balancing can be applied to elevators, rudder, or ailerons.
Common methods of mass balancing are shown in the graphic.
This picture shows the mass balancing of the aircraft's ailerons.
An aeroplane is trimmed when it will maintain its selected attitude and speed without the pilot applying any force to
the cockpit controls.
Since no force is required to hold the controls, the stick force is zero.
If the aircraft can be flown in this way, it will be less fatiguing for the pilot, and allow him
time to concentrate on other tasks required during the flight.
The problem is that changes in speed, engine power, and centre of gravity position will all affect the control position
required to obtain trimmed flight.
The problem can be resolved by the use of a trim control system.
There are basically three methods of trimming a light aircraft and achieving zero stick forces.
This can be achieved by the use of fixed or adjustable tabs, a variable incidence trimming tailplane or springs.
The simplest method of achieving trim is with the use of a fixed trim tab.
This is a small metal plate attached to the trailing edge of the flying control surface.
Consider an aircraft where to maintain the wing's level in flight, the pilot must continually hold the control yoke or
stick slightly to the right.
To prevent this, once on the ground, the trim tab on the aileron is adjusted by bending it slightly to
apply an aerodynamic force to position the flying control surface in the desired position.
Ailerons may have a tab fitted to only one of them.
In this particular case, the fixed trim tab on the port wing is adjusted up so that a downforce is
applied to that aileron when the aircraft is flying at cruise speed.
This adjustment is by trial and error and can be checked on the next flight.
Fixed tabs are used on simple low performance aircraft and trimmed for only one speed. This speed is usually the
cruise.
The performance of many of today's light aircraft requires that trim be adjustable in flight throughout the aircraft's speed range
as stick forces can become unacceptable.
This can be achieved by the use of adjustable trim tabs.
The greatest changes in stick force with speed changes tend to occur in pitch.
The elevator or stabilator will therefore normally be fitted with an in-flight adjustable trimming tab.
The ailerons and rudder may also be fitted with adjustable tabs, but as their control forces tend to be lighter,
some aircraft may retain a fixed tab adjusted for cruise on these control surfaces.
The adjustable trim tab is moved either manually by a mechanical linkage or controlled electrically from the cockpit by the
pilot.
If the pilot were to pitch an aircraft nose up, to continue to fly at that new attitude a force
must be continually applied by the pilot to the controls to maintain the attitude.
This would soon become very tiring for the pilot.
Control of the trim tab is designed to be intuitive in its operation.
For example, if the pilot was pulling the flight control towards himself as in this example, the trim control would
be moved in the same direction.
The pilot's cockpit control could be a wheel, a lever or a switch spring loaded to the off position for
an electrically powered system.
Indication of the trim tab's position is shown to the pilot.
Movement of the tab slowly applies an increasing opposing hinge moment at the control surface until it balances the stick
force.
This makes the stick force value zero.
Note that the tab moves in the opposite direction to the flying control surface.
Generally, unlike the tabs used for aerodynamic balance, the trim tab does not move as the flying control surface moves.
It must be repositioned every time the control surface is displaced.
However, on some aircraft the tab can be designed to serve a dual role as both trim and aerodynamic balance
tab.
There are a number of disadvantages to using an adjustable tab type control to achieve trim.
These are a slight reduction in control effectiveness due to the tab force being in opposition to the main flying
control surface,
an increase in drag due to the deflection of the tab,
and reduced effective range of the control surface.
This last factor is of particular importance with respect to the control of pitch of the aircraft.
A variable incidence trimming tail plane does not have the disadvantages of the trim tab,
plus it has the ability to trim for larger changes in speed and centre of gravity position due to the
powerful forces it can produce.
The system is again controlled from the cockpit and can be either manual or power operated.
An example of the operation of a variable incidence trimming tail plane to achieve trim after change of attitude is
shown in this graphic.
The elevator is deflected to change the pitch attitude.
The angle of incidence of the tail plane is now altered to achieve trim.
The disadvantage of the variable incidence trimming tail plane is that it is more complex and heavy compared to a
tab system.
The last type of trim control system to be discussed is the spring bias type.
An adjustable spring force is used to decrease the stick force.
No tabs are required, so the spring system does not have the disadvantages associated with tabs,
nor is it complex and heavy like the variable incidence trimming tail plane.
The spring bias system differs from the other systems discussed so far
in that the pilot's cockpit trim control is connected not to the flying control surface but to its control linkage.
The trim control biases springs which apply a force to the control linkage to give zero stick force.
Prior to take-off, it's important that the pilot sets the trim to the correct position as stated in the
operating manual for the aircraft.
If the trim is not correctly set, an excessive rate of pitch of the aircraft could occur,
leading to the tail striking the ground at rotation or the pilot experiencing very heavy stick force.
The primary flight controls discussed in this lesson are those generally found on most aircraft.
There are, however, aircraft that have combined controls.
Elevens act as elevators and ailerons.
Tailerons for pitch and roll control.
And the elevator, or V-tail, butterfly tail as it's commonly known, combine the effects of rudder and elevator.
Knowledge of these control systems is not required with the private pilot's licence.
The term stall should not be new to you. It was first introduced in Lesson 3.
Let's recap the information covered so far with reference to the stall.
The coefficient of lift, CL, increases with an increase in angle of attack up to CL max, approximately 16 degrees
angle of attack.
Beyond this angle, CL reduces abruptly, but the coefficient of drag, CD, continues to increase.
This was defined as the point of stall.
An aircraft's wing will always stall at this angle of attack.
It's known as the stalling or critical angle of attack.
As the angle of attack increases, the pressure pattern alters around the wing and the centre of pressure moves forward.
It's at its most forward point on the chord, just before the stall.
As the centre of pressure moves forward, the boundary layer must flow against an increasing adverse pressure gradient.
This causes the airflow to thicken and become more turbulent and start to separate from the upper surface of the
wing.
The separation point moves forward as the angle of attack continues to increase and the wing area that generates a
pressure differential becomes smaller.
At angles of attack above approximately 16 degrees, the air has separated over most of the upper surface and insufficient
lift is produced to balance the weight of the aircraft.
It was also stated that the value of CL max corresponds to a minimum flight speed and this was defined
as the stall speed.
To be precise, it's the 1g stall speed where lift equals weight and the load factor is 1.
This occurs with the aircraft in steady straight and level flight.
This speed is an important point of reference and is abbreviated as VS1g.
It's important to remember that the definition of the angle of attack is the angle between the chord line and
the relative airflow.
Therefore, if the angle of attack is increased up to or beyond the critical angle, an aeroplane can be stalled
at any speed or flight attitude.
The pilot identifies when the aircraft is stalled, when the following symptoms occur individually or in combination.
If the pilot is unable to stop the nose of the aircraft pitching down, this is an indication that the
aircraft has stalled.
As the turbulent boundary layer separates from the wing at high angles of attack, it can strike the tail area
of the aircraft causing it to shake.
This is called buffet.
If the buffet is very severe, it would deter any further reduction in speed.
This identifies the fact that the aircraft is being stalled.
It should be noted that at slightly lower angles of attack than the critical angle, separation can cause light buffet
of the airframe which will warn the pilot that a continued increase in pitch attitude would stall the aircraft.
The last symptom identifying the stall is if the pitch control has reached its aft nose up pitch stop position
and no further increase in pitch attitude occurs with the control held in that position.
During your flying training for the private pilot's licence, you'll have to demonstrate the aircraft type being flown at the
stall.
Different aircraft will have different characteristics at the point of stall.
You will of course become familiar with those of the aircraft you are learning to fly.
We'll now look at how the design of the aircraft, especially the wing, affects those characteristics.
The stalling characteristics of an aircraft depend on many things.
The leading edge radius, the camber, and the thickness ratio of the aerofoil.
The wing planform, as well as the type of flow of the boundary layer, all contribute to the way the
aircraft stalls.
Comparing the graphical lift curves of two different profile aerofoils, it can be seen that the curve at the point
of stall for the thinner aerofoil is much sharper than that of the thicker section.
The reality of this is that an aircraft with this type of profile aerofoil would tend to stall suddenly with
a very rapid loss of lift.
As a student pilot, you'll be pleased to know that most aircraft used for flight training have thicker cambered aerofoils,
the stalling characteristics of which are very gentle and predictable.
The wing of an aircraft is designed to stall progressively from the root to the tip.
There are three reasons for doing this.
First, to induce early light buffet over the tail surface of the aircraft.
Then, to retain aileron control effectiveness up to the critical angle of attack.
And finally, to avoid wing drop, which is a large rolling moment which would occur if one wingtip were to
stall before the other.
The wing planform used for an aircraft has a major influence on where the stall first occurs.
A rectangular wing will usually stall as desired at the root, but you recall this wing planform suffers from high
induced drag.
A tapered wing reduces induced drag, but has a tendency to stall at the tip.
Again, compromise must be made, and the wing modified to achieve a stall at the root.
Induced drag is caused by the wingtip vortices that reduce the effective angle of attack.
On a rectangular wing, these vortices are very strong, and the reduction in the effective angle of attack at the
tips ensures that the wingroot stalls first.
The tapered wing has much weaker vortices at the wingtip, and therefore may have a greater effective angle of attack
than the root.
This would cause the wing to stall at the tip first.
Because of the reasons stated, tip stall is not a desirable feature.
We'll now consider ways to modify a wing to ensure that it will stall at the root first.
There are a number of methods used to prevent wingtip stall.
The most common design features are washout, stall strips, changes of aerofoil section, vortex generators, and slats or slots.
You may recall that washout was also used for reducing induced drag.
The reduction in the angle of incidence of the wing at its tip will result in the root reaching its
critical angle of attack before the wingtip.
It has been stated that an aerofoil with a sharper leading edge radius will stall before a more rounded section,
but its overall efficiency will reduce.
The same result can be accomplished by attaching stall strips to the root wing leading edge of an efficient section.
The stall strip is a small triangular section that at higher angles of attack promotes separation, but will not affect
wing efficiency in the cruise.
The designer may actually opt to use a section at the wing root that has a sharper leading edge and
little camber.
If the section at the wingtip is cambered with a large radius leading edge, the wing is more likely to
stall at the root first.
Vortex generators are rows of small, thin, aerofoil shaped blades which project about 2.5 cm vertically from the upper
surface of the wing.
Each blade produces a small vortex within the boundary layer. Higher pressure air from outside of the boundary layer, called
the free stream flow, is induced by the vortices.
This thickens the boundary layer, increasing its mass.
Although it's more turbulent and it increases drag, the greater mass increases the kinetic energy of the boundary layer, which
delays separation and retains aileron control effectiveness.
The use of slats and slots to energise the boundary layer and delay separation at the stall was covered in
Lesson 4.
Their use on the outboard wing leading edge gives a useful local increase in the coefficient of lift at the
tip and again retains aileron effectiveness.
An aircraft will always stall if its angle of attack is increased to or beyond the critical value.
Many large jet aircraft are fitted with angle of attack indicators, but the aircraft you will be flying are not
so sophisticated.
Since CL max corresponds to a minimum flight speed, it's easy to reference the stall of the aircraft to this
speed.
However, this is only true for when the aircraft is flying straight and level and the speed of the aircraft
is reduced very slowly.
It cannot be overemphasised that stalling really has nothing to do with the speed of the aircraft. The critical angle
of attack can be exceeded at any speed.
By rearranging the lift formula at CL max used in Lesson 3, the speed V can be made the subject.
This shows that the speed V is equal to the square root of the lift force divided by a half
multiplied by the density of the air multiplied by the maximum coefficient of lift multiplied by the wing surface area.
Don't press the escape button just yet. We're not about to do any mathematics. Well, not a lot anyway.
The formula will only be used to show by example facts you already know that will affect the speed.
For practical purposes, the density, rho, and the wing surface area, S, can be considered as constant for a given
altitude and configuration.
Since lift must be equal to weight to maintain level flight, load factor 1G, the value of lift in the
formula will therefore be the same as the value of weight.
At CL max, the speed V that is found by the formula was defined as the stall speed at 1G.
We'll now simply refer to it as the stall speed, VS.
The stalling speed for a particular aircraft is found during its initial flight testing by the manufacturer and the figure
is published in the pilot's flight manual.
The stall speed is an indicated airspeed and does not vary with altitude.
There are, however, a number of factors that will alter the stall speed during normal operation of the aircraft.
For the following section of this lesson, we'll assume that an aircraft has a clean stall speed of 60 knots
indicated airspeed.
The first factor we'll consider is that of weight.
If the weight of the aircraft were to be increased, lift would have to be increased to maintain level flight.
If the lift weight figures are increased in a formula, half row CL max S will divide into it more
times.
This will give an increase in VS.
Stalling speed increases as weight increases.
A 20% increase in weight will increase the stall speed by approximately 10%.
As fuel is used in flight, the weight of the aircraft will of course decrease and the stall speed decreases.
Again, a 20% reduction in weight reduces stall speed by approximately 10%.
The stall speed published in the flight manual is normally quoted for the maximum take-off mass or weight of
the aircraft.
The stall speed at 1g should not exceed the quoted value.
It's the principle of the effect of change in weight on stall speed that is important.
The centre of gravity position in addition to the weight also affects the stall speed.
If the centre of gravity is on its forward position limit, to balance the moment between lift and weight the
tail plane must produce a downforce.
This effectively increases the weight of the aircraft as both the tail plane force and the weight are acting downwards.
More lift must be generated to balance the downforce.
Increasing weight, increasing stall speed.
If the centre of gravity is moved aft, the stalling speed will decrease as effectively weight and lift reduce.
During flight testing, the stall speed is calculated at the maximum weight and most forward centre of gravity position.
Lesson 4 covered lift augmentation, where the use of flaps and slats increased CLmax.
If the value of CLmax is increased, the value of half row CLmax S will also increase.
This bigger value would divide into the lift or weight fewer times and consequently VS will decrease.
Of course, fowler flaps would also increase the value of S, the surface area, and this further reduces the stalling
speed.
So far, we have always considered that lift is equal to the weight in our calculations of the stall.
In Lesson 3, it was shown that as an aircraft banks, to maintain level flight it must increase lift.
This increases the load factor.
Load factor for the banked aircraft was found by the formula 1 divided by the cosine of the bank angle.
In this example, that is 1 divided by the cosine of 60 degrees or 1 divided by a half equals
2.
The aircraft is in a 2G manoeuvre.
To increase lift in banked flight, the angle of attack must be increased.
If the aircraft was already at CLmax straightened level, it would be impossible to bank the aircraft.
If a bank was started at an indicated airspeed above the stall speed, at some bank angle CLmax would be
reached.
The stall speed increases as bank angle increases.
The increase in the stalling speed is a function of the bank angle only and can be found from the
formula
the original level stall speed VS multiplied by the square root of the load factor.
In the example shown, the aircraft is banked at 60 degrees.
Taking its stall speed in straightened level flight to be 60 knots, the aircraft will stall at 84.8 knots.
This equates to a 41% increase in stall speed.
Select the different angles of bank and note the increase in stall speed compared to that for straightened level flight.
In all cases, the aircraft's stall speed in straightened level flight is 60 knots.
The stall speed at any altitude of flight can be increased significantly if the wing is contaminated.
You'll recall that snow, frost or heavy rain can reduce the value of CLmax by as much as 30%.
The contaminant will also increase the weight of the aircraft.
When establishing the stall speed at CLmax, the engines must be set to develop zero thrust.
It's assumed that all of the weight of the aircraft is entirely supported by lift.
If thrust is applied close to the stall, the stalling speed of the aircraft will be reduced.
The reduction in stall speed is achieved in two ways.
Firstly, with the nose-high attitude of the aircraft, at CLmax the vertical component of the thrust force assists lift
in supporting the weight.
The secondary cause of reduction in stall speed comes from the propeller slipstream energizing the boundary layer of the wings.
At low flight speeds and high power, the dynamic pressure within the propeller slipstream is much greater than that due
to forward speed.
Much greater lift is generated by the wing sections when they are subjected to the slipstream than at zero thrust.
This should be remembered as potentially a heavy landing could occur if power is reduced suddenly just above the runway.
A judicious blast of power at this stage could well help to avoid that happening.
PILOT SPEED
Flying at low speed close to the stall requires the pilot to be aware of certain changes in the behaviour
of the aircraft.
Having just considered the effects of power at the stall, we'll start by looking at how power affects the aircraft.
In Lesson 2, the effect of propeller slipstream producing a turning moment was discussed.
This is most apparent at low speed.
If power is applied, the aircraft will tend to yaw in the opposite direction to propeller rotation.
If power is reduced, the aircraft will tend to yaw in the direction of rotation.
Yaw of the aircraft will increase drag and if the speed is already low, it will reduce still further.
The other problem associated with yaw is of course that the aircraft will tend to roll.
The advancing wing will generate more lift than the retreating wing and a roll moment is produced.
Roll is controlled by the ailerons.
At low speeds close to the stall, the reduced dynamic pressure requires greater control deflections to achieve the required control
response.
Trying to lift a dropping wing due to roll with aileron will increase its angle of attack.
At speeds close to the stall, aileron must be used with caution.
The preferred method of lifting a wing at these speeds is to use rudder to yaw the aircraft just enough
to increase the speed of the dropping wing and increase its lift.
Wing drop at the stall should be avoided as it can quickly lead to the spin, which will be discussed
later.
The stall is not a manoeuvre that a pilot performs frequently.
Stalls are however practised during initial and continuation training and during the skills test for the private pilot's licence, you
must normally demonstrate that you can correctly recover the aircraft from a fully developed stall.
Since the stall occurs because the angle of attack has been increased to the critical angle or beyond, it should
come as no surprise that if the angle of attack is decreased to below the critical angle, the wing will
no longer be stalled.
It's not the purpose of this Principles of Flight course to teach you how to fly. Your flying instructor will
teach you the standard stall recovery method.
However, it is worth considering here the practical implications of all you've just learned about the stall in this lesson.
First of all, when an aircraft stalls the lift force generated by the wings decreases abruptly and there is a
sharp rise in drag.
Consequently, the aircraft will lose height quickly.
When practising stall recovery at altitude, mindful of your minimum recovery height and having carried out your checks, this rapid
loss of height is not a problem.
But, even if you apply the correct stall recovery method, a loss of height of several hundred feet is not
uncommon.
As you'll appreciate, therefore, an unintentional stall close to the ground can have catastrophic consequences.
It is, thus, extremely important for you as a pilot to ensure that you never allow your aircraft to stall
unintentionally.
During your flying training, you must, therefore, learn to recognise all the signs of an approaching stall so that you
can prevent the stall from occurring.
Your flying instructor will teach you how to do this. The wings of most aircraft actually assist in the recovery
of the stall.
At the point of stall, the centre of pressure of the wing moves rapidly rearwards and produces a nose-down
pitching moment of the aircraft.
This decreases the angle of attack below the critical angle.
During all manoeuvres, the pilot should operate the flight controls smoothly and progressively.
If very sudden changes in pitch are made, the aircraft can stall at speeds much higher than the published stall
speed.
This is known as an accelerated stall.
As stated, the pilot does not normally fly the aircraft in a stall condition.
Having established the stall speed of the aircraft clean and with flap, an aircraft must give a clear and distinctive
warning sufficiently in advance of the stall so that the stall itself is avoided.
This should be not less than 5 knots or 5% above the indicated stall speed, whichever is the greater.
This requirement must be met for all flight manoeuvres and changes in configuration of the aircraft.
The warning may be provided either through the inherent aerodynamic qualities of the airplane, by this we mean buffet, or
the automatic operation of leading edge slats.
Buffet, caused by the separation of turbulent air from the wing striking the tail of the aircraft, can cause the
airframe to shake sufficiently to warn the pilot that the aircraft is approaching the stall.
Automatic leading edge slats operate as the centre of pressure of the slat moves forward with increasing angle of attack
until the pressure difference between the upper and lower surfaces pulls the slat out.
This is designed to occur sufficiently before the wing reaches its critical angle of attack.
If the aircraft does not have inherent aerodynamic qualities to provide a warning, an artificial warning device must be fitted.
A visual warning device requiring the crew's attention is not acceptable by itself, and usually an audible warning is provided.
The graphic shown demonstrates the operation of a typical artificial stall warning device.
The movement of stagnation pressure around the aerofoil as angles of attack increase is used to operate a flapper switch
that makes an electric circuit to a buzzer in the cockpit.
There are various designs of artificial stall warning device, but they are all designed to prevent the pilot from inadvertently
stalling the aircraft.
If however the aircraft is stalled, and the correct recovery action is not carried out, as stated earlier, the aircraft
could then enter into a spin.
For the aircraft to enter a spin, both wings of the aircraft must have stalled, but one wing is at
a greater angle of attack than the other wing and is said to be more stalled.
This can occur if the aircraft inadvertently drops a wing at the stall.
The component of airflow caused as the wing moves down alters the direction of the relative airflow and increases the
angle of attack further beyond the critical angle.
The change in airflow on the up-going wing causes a reduction in its angle of attack, but it's still
above the critical angle and therefore stalled.
Comparing the lift and drag curves for an aerofoil, it can be seen that beyond the stall the down-going
wing has less lift and more drag than that of the up-going wing.
The reduction in lift at the stall cannot balance the force of weight. The aircraft pitches down and the direction
of flight of the aircraft is primarily downwards.
Change in drag between the up-going and down-going wings causes a turning moment that also induces yaw of
the aircraft in the direction of the down-going wing.
The aircraft is therefore pitching, rolling and yawing. The direction of the spin is determined by the yaw.
It's the induced yaw of the aircraft at the stall which causes it to spin.
Any recovery action will naturally require the pilot to be aware of the direction of the spin.
Since this is determined by the direction of the yaw which provoked the spin, the pilot must know in which
direction the aircraft is yawing.
But the spin can be very disorientating and a pilot's senses may not immediately give him the complete picture of
what the aircraft is doing.
Furthermore, the artificial horizon will almost certainly have toppled.
Here are two ways of determining the direction of spin.
In good visual flight conditions, the pilot can look along the nose of the aircraft at the ground and note
the direction of rotation of the aircraft.
Alternatively, he may refer to the turn needle of the turn and slip indicator. In a spin, the turn needle
will be fully deflected in the direction of spin.
The spin is one of the most complex of all flight manoeuvres. It may be defined as an aggravated stall
resulting in autorotation.
This means that the aircraft will stabilise in the spin with both wings remaining unequally stalled, continuing in its downward
spiral path until either the pilot recovers the aircraft or the ground beats him to it.
Spinning is not an objective in the private pilot's licensed flight training syllabus. The explanation of the spin given previously
is to make you aware of what could happen if you get it wrong.
Spinning is intentionally carried out by many pilots trained to do so. It's an exhilarating experience enjoyed by many pilots.
Emphasis in your flight training, however, is on avoiding the spin.
The spin consists of three separate phases. The incipient spin, the fully developed spin, and finally the recovery from the
spin.
As we've learned, during a spin, the aircraft's wings are stalled. Consequently, spin avoidance is achieved primarily by ensuring that
the aircraft does not stall.
If a stall does occur, application of the correct stall recovery procedure will ensure that a spin is avoided.
The most common cause of entering the spin is through poor coordination of the flight controls at low speeds and
during stall recovery.
For instance, at the stall, a pilot must never attempt to correct any wing drop using aileron control.
This will merely aggravate the stall on the downgoing wing and greatly increase the risk of a spin ensuing.
The interrelationship between aileron roll and rudder yawing was discussed in Lesson 5.
Distractions inside or outside of the cockpit in flight can often diminish the pilot's ability to fly coordinated manoeuvres.
System malfunctions, preoccupation with equipment, avoidance of other aircraft or obstacles are examples of these distractions.
A spin may also develop if forces on the aircraft are laterally unbalanced. Fuel is often carried in the wings
of the aircraft.
Poor management of the fuel system by the pilot can cause the centre of gravity to be laterally displaced and
the aircraft will have a tendency to drop a wing.
When flying at low speed and low altitude during take-off and landing, the importance of recognition of the stall
and the prompt abdication of the recovery procedure, so preventing the aircraft entering a spin, cannot be overemphasised.
In Lesson 2 we discussed the four forces acting on an aircraft and how the airframe must be capable of
withstanding those forces.
The graphical representation of these forces was for straight and level flight.
When the aircraft is manoeuvred, there is a change in the forces. We'll first consider the aircraft in a steady
climb.
Starting from straight and level flight, we consider here a simplified arrangement of the four forces acting on an aircraft
in flight.
As you can see, during straight and level flight, lift supports all of the aircraft's weight.
If we now consider the aircraft climbing vertically, clearly lift is not supporting any of the aircraft's weight.
In this example, it can be seen that drag acts in the same direction as the weight, adding to the
downward acting forces.
For the aircraft to actually fly in this condition, the force of thrust would have to balance not only the
drag, but also the weight of the aircraft.
The ability of the aircraft to be able to climb vertically would be solely dependent on it being able to
produce sufficient thrust.
Currently, barring some exceptional types, aircraft are unable to maintain a steady vertical climb. They're simply unable to produce the
magnitude of thrust required.
A full explanation of the forces acting in the climb is beyond the scope of this course, but if we
assume that the climb path of the aircraft is in the same direction as the line of thrust, the forces
will be arranged as shown on the screen.
If we now resolve the forces parallel and at right angles to the direction of flight, and assuming a climb
angle of 10 degrees, we obtain two equations.
Lift equals weight times the cosine of 10 degrees and thrust equals drag plus weight times the sine of 10
degrees.
We've assumed that the weight of the aircraft is 9,810 Newtons.
Now take a little time to look at the calculations in the text box.
What our two equations tell us in non-mathematical language is that, firstly in a climb, lift is less than
weight.
And secondly, in a climb, the thrust needed is greater than the drag and will increase with the steepness of
the climb.
The conclusion that lift is less than weight in a climb is interesting, but not contradictory.
If you examine the diagram closely, you'll see that the weight of the aircraft is balanced by lift and thrust
acting together.
Lift then is aided by thrust to counter weight.
Consequently, lift by itself is less than weight.
The largest angle of climb for an aircraft is achieved when the excess thrust available from the engine is at
its greatest compared to the total drag on the aircraft.
This will occur at a particular speed which is called Vx.
This speed is known as the best angle of climb speed and it's found at a point on the graph
where the distance between the thrust and drag curves is at its greatest.
The best angle of climb speed is also referred to as the maximum climb gradient speed.
Climb gradient is the ratio of the change in vertical height to the horizontal distance travelled.
There are other speeds used to achieve the best rate of climb and cruise climb. These are discussed in the
performance lessons on Disc 6 of this series.
Most people have no problem in accepting that when an aircraft is descending, lift is less than weight.
The force of weight on the aircraft when descending is split into two components.
One is perpendicular to the flight path and balanced by the lift.
The other acts forward along the flight path, adding to the value of thrust to balance the drag.
This achieves a balanced descent at a steady speed.
If the pilot closes the throttle and if we neglect residual thrust from the engine at idle RPM, the only
force remaining to propel the aircraft forward is that of the component of weight acting along the flight path.
Viewing the graphic, it can be seen that the drag vector is now greater than the component of weight acting
forward.
To maintain a constant indicated airspeed, the nose must be lowered to increase the component of weight acting forward. The
angle of descent will increase.
The aircraft is now in a glide.
To give the minimum descent angle and therefore the greatest gliding distance, drag must be at a minimum.
Minimum drag is achieved at a particular value of indicated airspeed. It's called VMD.
In the clean configuration, this is also the speed for maximum lift-drag ratio and corresponds to an angle of
attack of approximately 4 degrees.
The climb and descent of the aircraft discussed has been in balanced, unaccelerated flight. This produces no additional loads on
the airframe structure.
However, this is not the case when the aircraft is turning.
When an aircraft is flying, it tries constantly to obey Sir Isaac Newton's first law of motion.
Part of this law states that a body will continue to move in a straight line unless acted upon by
an external force.
The largest aerodynamic force produced on the aircraft is lift, and it's this force that is used to turn the
aircraft.
The speed of the aircraft and its mass will determine the value of force required to make it turn.
The turn is achieved by banking the aircraft.
The inclined lift force now has to both support the weight of the aircraft and provide the force to change
the aircraft's direction of flight.
Note that lift must be increased if its vertical component is to be large enough to maintain level flight, and
its horizontal component large enough to produce the required turning force.
The horizontal component of lift provides a centripetal, or centre seeking force, which during a turning manoeuvre is constantly deflecting
the path of the aircraft towards the centre of the turn.
Centripetal force is balanced by centrifugal force, which is its equal and opposite reaction.
The effect of these two forces can be seen if we consider a model aircraft attached to a piece of
string and being swung around in a circle as shown in the graphic.
In this example it's the string that provides the centripetal force.
Centrifugal force will be equal and opposite to it, and the model maintains a constant curved path.
The model is in a balanced, coordinated turn.
The effects on an aircraft of adverse aileronure and propeller slipstream give the possibility for the turn to be uncoordinated
and unbalanced.
This condition exists when the aircraft is sideslipping.
This will cause either a slipping or skidding turn.
The path taken by the aircraft in a slipping and skidding turn is shown in this graphic.
Indications of sideslip are displayed to the pilot by the inclinometer portion of the turn coordinator, or turn and slip
indicator instrument in the cockpit.
The inclinometer consists of a glass tube with a ball placed inside it.
This graphic shows a coordinated, balanced turn to the right.
The ball in the inclinometer is centered.
There is no sideslip.
If the aircraft were in a slipping turn, the indications would be as shown in this graphic.
The ball is displaced in the direction of the turn.
The ball can be centered in two ways.
Decreasing the bank angle will center the ball, but the following method I'm sure will be drummed into you by
your flight instructor.
Step on the ball is a simple rule used to center the ball by the use of rudder.
Ball right, apply right rudder.
Step on the ball.
In a skidding turn, the ball is displaced in the opposite direction to the turn.
This would require an increase in bank angle to center the ball, or, yes you guessed it, step on the
ball.
Apply left rudder in this example.
To produce the necessary force to support and turn the aircraft, the lift force is increased above that of weight
and load factor increases in a turn.
A 60 degree banked turn, such as the one shown here on the screen, induces a load factor of 2.
We say, popularly, that a 60 degree banked turn is a 2G turn.
The pilot feels this increased load factor as a force pushing him down into his seat.
In a 60 degree banked 2G turn, he will feel that his weight has doubled.
The load factor in a turn is a function of bank angle only.
It's not only in turning flight that the aircraft can experience increased load factors, or G as they are popularly
referred to.
If the pilot were to suddenly pitch the aircraft nose up whilst in a dive, the increasing angle of attack
at high speed would generate an increase in lift above that of the weight.
This maneuver, called a pull-up, would thus cause an increase in load factor.
Consider in this example the load factor has increased to a value of 2.
If the aircraft has a weight force of 9,810 Newtons, lift will be twice this value at 19,620
Newtons.
The airframe structure will have a total force acting on it equal to twice its own weight,
and will accelerate in the direction of the force at two times greater than that due to gravity.
This is a positive G maneuver.
The pilot will also be subjected to the increase in load factor.
But he'll feel it as the inertial reaction to the force produced by the wing, acting in the opposite direction,
that is, in a pull-up, through the pilot's seat downwards in the direction of weight force.
At a load factor of two then, the pilot will feel as if he's being pushed further down into his
seat, as his weight is effectively doubled.
This is what is known as positive G.
The aircraft can also be subjected to negative G.
Consider the aircraft on its curved flight path after the pull-up.
If the pilot were to suddenly pitch the aircraft nose down,
momentarily the angle of attack of the wing becomes negative, and the lift force will act downwards.
This produces an inertial reaction to the lift force acting in the opposite direction, that is, upwards.
This is what is popularly known as negative G force. It acts on both airframe and pilot.
It's important that a pilot be able to recognize the difference between this negative G maneuver and a stall.
You've probably actually experienced negative G in your everyday life.
Traveling in a car too fast over a humpback bridge causes you to rise up from your seat.
That's negative G.
You'll experience the same in an aircraft during negative G maneuvers.
The effects of G force on the body are covered in Disc 2 of this series.
This lesson is concerned with its effects on the aircraft's structure.
Aircraft structures are designed to cope with the level of forces imposed by the pilot during maneuvering as appropriate for
the aircraft type.
Regulations state the G forces which an aircraft must be able to withstand. These are for both positive and negative
G.
The positive G limits are normally greater than for negative.
For example, currently the JAA specify a minimum requirement of plus 6 and minus 3G for aerobatic aircraft.
The need for minimal structural weight dictates that safety factors are kept to a minimum.
G forces in excess of the approved values will cause permanent damage to the structure.
Forces in excess of 150% of the design values will almost certainly result in structural failure.
As well as the forces imposed on the aircraft by maneuvering, flight in rough weather can impose very significant G
loadings.
And in particular, flight through turbulence at high speeds can produce forces sufficient to cause structural failure.
Consider the aircraft in the graphic flying at 100 knots in smooth air.
If turbulence was to produce a vertical gust of wind at 20 knots, the relative airflow would alter in the
direction shown in the graphic.
This will of course change the angle of attack and the value of lift produced.
By simple mathematics, it can be shown that in this example the angle of attack would change by about 11
degrees.
This would greatly increase the lift and load factor.
The mathematics is irrelevant, but the principle must be understood.
The maximum load factors which must be allowed for during maneuvers are shown in an envelope of speed, V, against
load factor, N.
Each aircraft type has its own particular V-N diagram, with specific Vs and Ns.
This diagram shows a typical maneuver envelope or V-N diagram.
The curve OA represents the increasing G loading on the aircraft as the speed is increased with full backwards deflection
of the control column.
The maximum positive lift capability is initially limited, because the aircraft will stall before the load limit factor is reached.
At point A, the maximum positive lift capability will cause the loading on the structure to reach the positive limit
load factor.
The airspeed at this point is the minimum airspeed at which the positive limit load factor can be developed aerodynamically.
The usual term to describe this speed is V-A, or maneuvering speed.
If the aircraft is flown at speeds above that represented by point A, it must be borne in mind that
any full or sudden deflection of the elevators
will increase the load factor to a point where permanent deformation of the aircraft's primary structure may occur.
If aircraft speed is increased still further, it will eventually reach the point where structural damage or failure will occur.
This is the maximum speed which has to be considered when assessing the strength of the aircraft.
The usual term used to describe this speed is the design dive speed, or V-D.
In general, V-M-E, the never exceed speed, which we will define later, is set at about 10%
less than the design dive speed.
Similar stresses will be developed in the aircraft structure during negative lift flight.
However, the negative load factor limits are set lower than those for positive lift flight.
We now superimpose upon the maneuver envelope the significant speeds and speed ranges with which, as a pilot, you will
have to be very familiar.
The speeds and speed ranges will be given in the aircraft manual as well as indicated by coloured arcs on
the aircraft's airspeed indicator, or ASI.
We'll look at the ASI indications a little later on in this lesson.
Note that Vₙₒ, the higher limit of the normal operating range, is of a higher value than V
-A, the maneuvering speed.
V-A is the maximum speed at which full or abrupt control deflections will result in a stall without the
risk of causing structural damage.
The aircraft can be operated quite safely above this speed as long as the maneuvering envelope is fully understood.
Above V-A the pilot should not use full, abrupt control deflections.
The aircraft should not be flown in turbulent air at speeds above Vₙₒ.
The speed range between Vₙₒ and Vₙₑ is the caution range and should be entered only
in smooth air.
As we've said, Vₙₑ is the speed never to exceed.
Between Vₙₑ and V-D there is a danger of control surface flutter, which could lead to airframe
structural damage or failure.
The limit load factors will depend on the design category of the aircraft.
The Joint Aviation Regulations state that
This is so that structural weight can be kept to an absolute minimum so that a manufacturer will not design
an aircraft to be stronger than the minimum required by the regulations.
For utility aircraft the positive limit load factor is 4.4G.
For aerobatic category aircraft the positive limit load factor is 6G.
The negative limit load factor may not be less than minus 1G for normal category aircraft.
Minus 1.76G for utility category aircraft.
Minus 3G for aerobatic category aircraft.
On the airspeed indicator a green arc indicates the normal operating speed range.
Vₙₒ is the upper limit of the green arc.
Vₙₒ is the normal operating cruise speed limit.
Vₙₑ is the never exceed speed.
Although the aircraft can be flown safely in still air conditions at this speed, it's good airmanship to allow a
margin between Vₙₑ and the speed at which you are flying, just in case you encounter gusty conditions.
Gusts encountered while flying at Vₙₑ may increase the loads on the aircraft structure to above the positive
load limit factor.
Vₙₑ is shown by a radial red line on the airspeed indicator at the high speed end of
the yellow arc.
From Vₙₒ to Vₙₑ there will be a yellow arc, which is the caution range.
You may fly at speeds within the yellow arc only in smooth air, and then only with caution.
Some ASIs have two sets of coloured arcs to show the limiting speeds and speed ranges above and below a
given operating mass.
Flaps are designed to reduce take-off and landing distances.
The flaps and their operating mechanism and attachment points to the structure are not designed to withstand the loads which
would be applied at high airspeeds.
As you can see from the screen, with flaps deployed, the aircraft's flight envelope is significantly reduced.
A moment's study of the flaps deployed envelope will show us that the load factor limits are lower.
This means that we must set a much lower maximum permissible speed.
The critical speed never to exceed with flaps extended is termed V-F-E, Velocity Flaps Extended.
To indicate to the pilot the speeds at which he can fly without overstressing the aircraft,
a white arc on the airspeed indicator dial extends from the aircraft's power-off landing configuration stalling speed to V
-F-E.
V-F-E, the flaps extended speed, is the maximum airspeed at which the aircraft should be flown with the
flaps in a prescribed extended position.
The landing gear will normally be retracted as soon as possible after take-off, to reduce drag and increase the
climb gradient.
There is no normal requirement for the gear to be operated at high indicated airspeed,
so the retract and extend mechanism, together with the attachment points to the structure, are sized for the required task.
To design the gear for operation at high indicated airspeeds would unnecessarily increase structural weight.
Thus, either one or two speed limitations will be specified for your aircraft, depending on system design.
V-L-O is the Landing Gear Operating Speed. It's the speed at which it's safe both to extend and
to retract the landing gear.
If the extension speed is not the same as the retraction speed, the two speeds must be designated as V
-L-O Extend and V-L-O Retract.
When the gear is retracted or extended, the landing gear doors must open first.
The doors merely streamline the undercarriage bay and are not designed to take the aerodynamic loads which would be placed
on them at high indicated airspeeds.
Consequently, V-L-O is usually lower than V-L-E.
V-L-E is the Landing Gear Extended Speed.
There may be occasions when it's necessary to ferry the aircraft with the gear down, and to do this a
higher permissible speed would be convenient.
V-L-E is the speed at which it's safe to fly the aircraft with the landing gear secured in
the fully extended position.
Because the landing gear doors are closed, V-L-E is normally higher than V-L-O.
It's not only in the air that loads are applied to the aircraft.
During taxiing, take-off and landing, it's the undercarriage which must take the applied loads.
The undercarriage is primarily designed to withstand fore and aft loads, that is, loads applied on the undercarriage in the
direction of the aircraft's longitudinal axis.
These loads normally arise during take-off, landing and braking, so the undercarriage has to be strong.
Nevertheless, weight has to be kept to a minimum, for, as we know, in the air the undercarriage is just
dead, drag-generating weight.
Consequently, designers must ensure that margins of safety in the required load-bearing strength of the undercarriage are kept as
low as possible.
Excessive side loads on the aircraft can easily cause the undercarriage structure to fail.
This can be caused by taxiing too fast, or attempting to turn the aircraft through too tight an angle.
But more often, it's during take-off and landing, particularly with a crosswind, that the undercarriage is subjected to these
excessive loads.
Limits are set on the maximum value of crosswind acceptable for different aircraft.
Initially, during a crosswind take-off, the weight of the aircraft on the undercarriage will help resist sideways movement.
As speed increases and the wings generate more lift, the interwind wing can start to lift, putting a side load
on the undercarriage.
This is prevented by applying interwind aileron.
The aircraft is held on the ground for longer, maintaining direction with rudder, before lifting off.
In addition to any crosswind, during take-off the aircraft is also subjected to a turning moment from the propeller.
This is generally referred to as swing on take-off.
The effects of the propeller are covered in the engine's general section of this disc.
For a number of reasons, the aircraft will yaw in the opposite direction to the propeller rotation.
It must be controlled with rudder.
In this picture, if there was a crosswind from the left, the yawing moment caused by the propeller's rotation would
be aggravated.
Crosswind from the right would tend to cancel out the yaw due to propeller rotation.
Throughout this lesson we've generally showed pictures of what are commonly called tricycle undercarriage aircraft.
This type has two main wheels and another wheel at the front of the aircraft.
The tailwheel, or tail dragger aircraft as it's often referred to, is still widely used.
In the air it flies no differently from the tricycle undercarriage aircraft, but on the ground it handles very differently.
The differences in ground handling occur because the tailwheel aircraft has its centre of gravity behind the main wheels.
This tends to make the aircraft unstable during movement on the ground.
If, during the take-off roll, a tailwheel aircraft were to start to swing, the amplitude of the swing would
tend to increase because the aircraft's centre of gravity is behind the main wheels.
Unless checked, this directional instability could lead to a ground loop.
Tailwheel pilots soon learn that any swing on take-off must be quickly checked with rudder.
At the point of touchdown on landing, the throttle is closed and the propeller has little or no effect on
the airframe.
It's landing with a crosswind that can cause excessive sidelows on the undercarriage.
As with take-off, crosswind limits are set for a particular aircraft.
You'll be taught various methods of landing in a crosswind during your flight training.
It's important that as the aircraft touches down, any sideloads on the undercarriage are minimised.
This concludes the lessons on the principles of flight.
You should now be able to answer in detail the question,
How does something as big and heavy as an aircraft fly?, which was posed at the start of Lesson 1.
But next time you hear the sound of an aircraft engine, you may not be looking up at it wondering
at the marbles of flight,
but sitting within an aircraft and experiencing the exhilaration of flight.