The regulations, and good airmanship, require that during any phase of operation the loading, mass and centre of gravity of
the aircraft comply with the limitations specified in the approved aircraft flight manual, or the operations manual if it is
more restrictive.
It is the responsibility of the commander of the aircraft to satisfy himself that this requirement is met.
More detailed explanations of mass and centre of gravity limitations are given in the lesson following this one.
Limitations on mass are set to ensure adequate margins of strength and performance, and limitations on centre of gravity position
are set to ensure adequate stability and control of the aircraft in flight.
If the limiting mass of an aircraft is exceeded
Performance is reduced, take-off and landing distances will be increased, rate of climb and ceiling will be reduced, range
and endurance will be reduced, maximum speed will be reduced.
Also, stalling speed will be increased, manoeuvrability will be reduced, wear on tyres and brakes will be increased, and structural
safety margins will be reduced.
The maximum take-off mass is defined as the maximum permissible total aircraft mass at the beginning of the take
-off run.
The maximum landing mass is defined as the maximum permissible total aircraft mass upon landing, under normal circumstances.
The maximum zero fuel mass is the maximum permissible mass of an aircraft with no usable fuel.
If fuel must be retained in particular tanks for purposes of ballast or for centre of gravity reasons, it must
be included in the zero fuel mass when it is specified in the airframe limitations.
In service, an aircraft must observe certain speed limitations. These may be maximum speeds or minimum speeds, but in each
case they are set to give safe operation in the prevailing conditions.
The limits may be set by various considerations, the main ones being the strength of the aircraft structure, that is
the ability of the structure to withstand a load, the stiffness of the aircraft structure, which is the ability of
the structure to withstand deformation, that adequate control of the aircraft be maintained at all times.
The limit load is the maximum load to be expected in service. The ultimate load is the load at which
the structure will fail. The safety factor is the ratio of the ultimate load to the limit load.
For aircraft structures the safety factor is 1.5. The safety factor on aircraft structures is much lower than the
safety factors used in other forms of engineering because of the extreme importance of minimum weight in aircraft structures.
To keep the weight as low as possible, the safety factor must be kept to a minimum.
Because of this, it is extremely important not to exceed the limitations set on the operation of the aircraft, as
the safety margin can easily be exceeded and structural damage occur.
The airframe structure must obviously be strong enough to take the loads acting upon it in normal level flight, that
is the forces due to lift, drag, thrust and weight.
However, the aircraft is also required to manoeuvre and to fly in turbulent air.
Under these conditions the loads on the aircraft will be increased, so it must also be strong enough to perform
whatever manoeuvres are specified for the aircraft and to withstand the turbulence which the aircraft may reasonably be expected to
encounter.
The maximum load factors which must be allowed for during manoeuvres 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
manoeuvre 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 Vᴀ, or manoeuvring 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 Vᴅ.
In general, Vₙₑ, 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 manoeuvre 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 will 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 manoeuvring speed.
Vᴀ 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 manoeuvring envelope is fully understood.
Above Vᴀ 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 have said, Vₙₑ is the speed never to exceed.
Between Vₙₑ and Vᴅ there is a danger of controlled 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 JAR regulations state that for normal category aircraft the positive limit load factor may not be less than
2.5 g and need not be more than 3.8 g.
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 4.4 g.
For aerobatic category aircraft the positive limit load factor is 6 g.
The negative limit load factor may not be less than minus 1 g for normal category aircraft.
Minus 1.76 g for utility category aircraft.
Minus 3 g 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 air speeds.
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 Vꜰᴇ.
Vꜰᴇ, the flap's 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.
Vʟᴏ 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
Vʟᴏ extend and Vʟᴏ 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, Vʟᴏ is usually lower than Vʟᴇ.
Vʟᴇ 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.
Vʟᴇ 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, Vʟᴇ is normally higher than Vʟᴏ.
For flight in turbulence, an airspeed must be chosen to give protection against two possibilities, stalling and overstressing the aircraft
structure.
Because the rough airspeed has to satisfy these two requirements, it must be a compromise.
If you find yourself flying in turbulent conditions, and there is no specified Vʀᴀ for your aircraft, then
fly at the design manoeuvre speed, Vᴀ, for the aircraft's instantaneous all-up weight.
Remember that these speeds are targets. The turbulent conditions will cause your indicated airspeed to fluctuate, and by chasing a
speed you may overstress the aircraft.
Remember also that the turbulence may cause your altitude to fluctuate wildly. Don't chase it.
As long as you are in no danger of hitting the ground or other aircraft, allow the aircraft to ride
the gusts.
In any discussion on this subject, you will often hear the word weight substituted for the word mass.
In this particular context, it doesn't really matter, but it might help to clarify the situation if we define both
mass and weight.
Mass is the amount of matter in a body.
Weight, on the other hand, is the force produced when mass is subject to gravity.
A body will always have mass, but if there is no gravity, then it has no weight.
For the purpose of this subject, the two terms are interchangeable.
Aircraft are weighed in a draft rehanger.
On each and every occasion of weighing, a weighing schedule is compiled that states the basic equipment installed at the
time,
together with the mass and the position of the centre of gravity of the aircraft, after completion of the weighing
procedure.
The basic empty mass and centre of gravity of a new aircraft are determined by the manufacturer, using one of
a number of weighing procedures.
Light aircraft generally use the basic empty mass and centre of gravity as the start point for loading calculations.
The main type of equipment used to weigh light aircraft is electronic.
This equipment uses strain gauges, which vary their electrical resistance with load.
The centre of gravity is defined as the point through which the force of gravity is said to act on
a mass.
For an aircraft, this is the point on the aircraft through which the total mass is said to act in
a vertical downward direction, giving the aircraft its weight.
The centre of gravity is also the point of balance, and as such it affects the stability of the aircraft,
both on the ground and in the air.
The centre of gravity limits are the most forward and most aft positions that the centre of gravity can be
at which the aircraft is permitted to fly.
The centre of gravity limits are set by the manufacturer, defined in the aircraft flight manual and are mandatory.
They are defined relative to the centre of gravity datum.
The centre of gravity datum is a point designated by the manufacturer, from which all measurements are taken in the
calculation of the centre of gravity.
The datum can be anywhere along the longitudinal axis of the aircraft, even if the axis is extended in front
of or behind the aircraft.
The arm is the distance from the centre of gravity datum to the point at which the mass of a
component acts, the centre of gravity of the component.
By convention, all arms aft or behind the datum are positive, and all arms forward of the datum are negative.
The moment is the turning effect of a mass around the datum. It's the product of the mass multiplied by
the arm.
Since the arm may be positive or negative, a positive arm will give a positive moment and a negative arm
will give a negative moment.
The basic empty mass is the mass of the aircraft with all its basic equipment plus a declared quantity of
unusable fuel and oil.
The traffic load is the total mass of passengers, baggage and freight.
The dry operating mass is the total mass of the aircraft ready for a specific type of operation, excluding all
fuel and traffic load.
This mass includes items such as crew and crew baggage, catering and removable passenger service equipment.
The following procedure can be used to find the centre of gravity for the basic empty mass.
The datum for this aircraft is known to be level with the end of the spinner.
The aircraft is weighed, and the effect of the mass on each undercarriage leg is indicated by weighing devices.
If we use a table like the one shown here, we can fill in the fields with the appropriate information.
First, the mass indicated on the nosewheel weighing device and the arm, the distance between the datum and the nosewheel.
The mass indicated on the nosewheel weighing device is 100 pounds.
To obtain the figure for the moment column, we must multiply the mass by the arm.
Remember, in the other definitions part of this lesson we stated that whether the arm is aft or forward of
the datum is significant.
By convention, all arms aft of the datum are positive, and all arms forward of the datum are negative.
In this particular case the arm is 36 inches aft of the datum, so it will be positive.
36 times 100 gives us a moment of 3600 pound inches.
The figures required to fill in the left and right main wheel fields are identical because the aircraft is balanced
symmetrically.
The mass in both cases is 725 pounds, and the arm in both cases is positive 90 inches.
The resulting moment is plus 65,250 pound inches in each case.
Now we can add the figures in the mass and the moment columns to obtain our total mass of 1550
pounds and a total moment of 134,100 pound inches.
The centre of gravity position relative to datum is the total moment divided by the total mass.
So, substituting in the formula, we have the centre of gravity position relative to datum is 134,100 pound inches
divided by 1550.
Therefore, the centre of gravity position relative to datum equals 86.5 inches aft of the datum.
When the centre of gravity of the basic empty mass has been determined, the centre of gravity for the loaded
mass, both for take-off and landing, may be calculated using the same principle of moments.
For convenience, the data may be tabulated on the mass and balance documentation, more commonly called the load sheet.
This picture shows a simple table which will allow us to work out the position of the centre of gravity
and the mass of the aircraft when it's loaded.
Notice that the table already has the arm for each item shown, except that for the aircraft empty mass.
This is because this table is particular to the type of aircraft that we are going to fly.
We are required to fill in the details of the aircraft's registration, the date and the pilot's name.
The aircraft's empty mass, arm and moment can be obtained from the weight and centre of gravity schedule of our
aircraft.
If we now insert the mass of the front seat passenger, 190 pounds, into the correct field and multiply it
by the arm given, we get a moment of 15,295 pound inches.
Inserting the mass of the pilot, the other occupant of the front seats, into the designated field and multiplying it
by the arm given, gives us a moment of 16,100 pound inches.
Now we have to do the same for the sole passenger of the rear seats, whose mass of 100 pounds
is multiplied by the given arm of 118.1 to obtain a moment of 11,810 pound inches.
The luggage for this trip weighs 65 pounds and it's stowed in the baggage hold of the aircraft which is
to the rear of the passenger seats.
Multiplying the baggage weight by the appropriate arm we get a moment of 9,282.
The required fuel for this trip is 35 US gallons, which at 6 pounds per gallon will weigh 210 pounds.
Multiplying this by the given arm of 95 will give us 19,950 pound inches.
If we add all of the figures in the mass column we get a total of 2,315 pounds.
Adding the figures in the moment column we get a total of 206,512 pound inches.
Substituting these two totals in the formula shown at the bottom of the table, we get a centre of gravity
position which is 89.21 inches aft of the datum.
Each aircraft type will have a mass versus centre of gravity envelope.
The one shown here is for a light aircraft and so is relatively straightforward.
The green shaded portion of the envelope denotes the safe part of the envelope, the upper extent being limited by
the maximum loaded all up mass of the aircraft.
Now we must plot the centre of gravity position on the top and bottom scales and extend a line between
these positions.
And then intersect it with a line placed at a position which corresponds to the loaded all up mass.
The intersecting point of these two lines shows firstly that our aircraft loaded all up mass is below the maximum
allowed.
And secondly that the position of the centre of gravity is well inside the envelope.
Having determined the centre of gravity and the mass of the aircraft when it's loaded and ready for the flight
we are about to undertake,
it would be prudent to determine whether the use of fuel during the flight will move the centre of gravity
to a point outside the centre of gravity envelope.
To do this we must find the moment of the total fuel which is planned to be used on the
flight by multiplying its mass by its arm.
The mass of fuel to be used in this case is 150 pounds and the arm is 95 inches aft
of datum and therefore positive.
Multiplying the two together we get 14,250 pound inches.
From the take off mass we subtract the mass of the fuel used to obtain the landing mass of 2
,165 pounds.
From the take off moment we subtract the fuel used moment to obtain the landing moment of 192,262 pound
inches.
The landing centre of gravity can be determined by dividing the landing moment by the landing mass which gives us
88.8 inches aft of the datum.
This puts our weight and centre of gravity position here on the load sheet well inside the envelope.
The centre of gravity must lie between forward and aft limits.
If the centre of gravity is out of limits, control forces, stability, manoeuvrability and performance will all be affected.
The following are some of the effects that might occur if the centre of gravity is out of limits.
Imagine that these are the forward and aft limits of the centre of gravity.
Now imagine that the aircraft are somehow being loaded grossly nose heavy.
The centre of gravity will move to a position well in front of the forward centre of gravity limit.
If the centre of gravity is outside the forward limit then a large balancing down load is required from the
tail plane, requiring large elevator deflections.
Large elevator deflections will result in increased drag from the control surfaces.
And because of the increased wing lift required to balance the tail plane down load, increased induced drag.
This causes reduced aircraft performance.
Stalling speed will be increased because of the increased wing lift required.
Longitudinal stability is increased. In other words, the aircraft will be slow to respond to pitch inputs.
This leads to higher stick forces being required in pitch.
Range and endurance are decreased due to the increased trim drag as the elevators must be used to trim the
aircraft.
The ability to apply nose up pitch is decreased because the elevators, having been used to trim the aircraft's straightened
level, have less up range remaining.
Imagine once again that these are the forward and aft limits of the centre of gravity.
This time though, imagine that the aircraft has been loaded grossly tail heavy.
The centre of gravity will move to a position well behind the aft centre of gravity limit.
If the centre of gravity is outside the aft limit, then a large balancing upload is required from the tail
plane, requiring large elevator deflections.
Longitudinal stability is reduced. In other words, the aircraft becomes very sensitive to pitch control.
If the centre of gravity is too far aft, the aircraft will become unstable.
Stick forces in pitch will be light, leading to the possibility of overstressing the aircraft by applying excessive g.
Recovering from a spin may be more difficult because a flat spin may develop.
Range and endurance will usually decrease due to the extra drag.
Glide angle will be more difficult to sustain because of the tendency for the aircraft to pitch up.
In the lesson on airworthiness on Disc 5, we learnt that aircraft are categorised according to the purposes to which
they are put.
We also learnt that aircraft are further categorised according to the manoeuvres that they are allowed to perform.
These are
Normal category, below 5,700 kg and non-aerobatic. Manoeuvres are limited to stalls and steep turns of 60°.
Typical limit load factors are plus 2.5 g and minus 1 g.
Utility category
As for normal category plus limited aerobatics.
Typical limit load factors are plus 4.5 g and minus 1.8 g.
Aerobatic category
Fully aerobatic
Typical limit load factors are plus 6 g and minus 3 g.
Some types of aircraft may be allowed to operate in more than one category.
For instance, the normal and the utility category.
However, there will probably be different mass and centre of gravity limits which are applied when the aircraft is operating
in each category.
An example is shown here.
This is the mass versus centre of gravity envelope that we saw earlier when we were working out the centre
of gravity for the loaded mass.
The green shaded area is the envelope for the aircraft operating in the normal category.
The same aircraft can be operated in utility category, but only if the all-up mass and centre of gravity
limits are changed.
You can see here how the green shaded area has reduced considerably.
The maximum all-up mass has been lowered from 2,325 pounds to 2,020 pounds.
Whereas previously this aircraft could only operate in the normal category within limit load factors of between plus 2.5
g and minus 1 g,
within the new utility envelope it can operate within limit load factors of between plus 4.5 g and minus
1.8 g.
This section on flight performance deals with the factors that control and determine the performance of an aircraft throughout the
different stages of its flight.
We will start with the take-off, then proceed to the climb, cruise, descent and lastly the landing phase.
This first lesson deals with the take-off.
The take-off stage of flight is defined as being from brake release until the aircraft reaches a defined screen
height.
In light general aviation aircraft this screen height is 50 feet, as shown.
As the aircraft accelerates down the runway, rotation will be initiated at a specific speed.
The aircraft pitch will increase and eventually the aircraft will lift off and climb to the screen height.
A specific speed needs to be gained before the screen height.
This is called the take-off safety speed and it must be at least 20% more than the stalling
speed of the aircraft, in other words more than 1.2 vs.
This speed is called V2 in large twin-engine aircraft.
The take-off can be divided into two sections.
The first is the take-off roll or take-off run, which is the distance travelled whilst still on the
ground, in other words from brake release to lift-off.
The second section is the initial climb, which is the distance covered from lift-off until it reaches the screen
height.
The combined length of the take-off run and initial climb is known as the take-off distance.
As pilots we need to make sure that our take-off distance does not exceed the available distance at the
airfield we are using.
This may seem obvious, but there are many incidents to suggest that this important consideration is often forgotten.
Let's now look at the available distances at any given airfield.
The first of these is known as the take-off run available, or TORA for short.
This is defined as being the length of runway which is declared available by the appropriate authority and suitable for
the ground run of an aircraft taking off.
Essentially, this is the normal length of the runway from threshold to threshold.
However, at some airfields a clear way can be found.
This is defined as an area beyond the end of the runway in the direction of take-off under the
airport's authority,
which is clear of any significant obstacles and suitable for an aircraft to make its initial climb to the screen
height.
With the addition of the clear way we have what is called the take-off distance available, or TORA for
short.
This is therefore TORA plus any length of clear way.
There is one more additional section that can be found at the end of the runway in the direction of
take-off.
This section is only to be used to help bring the aircraft to a stop in the event of an
abandoned take-off.
It cannot be used for any other purpose.
This extra section is known as a stopway.
If we take the TORA plus any length of stopway we have what is called the accelerate stop distance available,
or ASDA for short.
This is sometimes also known as the emergency braking distance available.
What do these distances mean to us as pilots and how should we use them?
If your airfield has only a take-off run available then you must complete your take-off, including of course
the climb to the screen height, before the end of the runway as shown here.
However, if your airfield has a take-off distance available then you can use the TORA to complete your ground
run and do the initial climb within the clear way.
Notice that this time there is a greater amount of distance for you to use.
But be careful, not all airfields have this clear way and therefore the take-off run available is the same
as the take-off distance available.
Next we will examine how the take-off distance is calculated and displayed to the pilot.
In order to do this let us examine in detail the formula that is used to calculate the take-off
distance.
D is the distance required from rest to reach a specified speed, which is V, with a given acceleration force,
A.
Acceleration itself is force divided by mass.
The force is thrust minus aerodynamic drag and minus the wheel drag.
Having seen the components involved in calculating the take-off distance, let's now look at how these are changed and
more importantly how the take-off distance changes given different factors.
The first of these is the effect of increasing mass.
Increasing mass has several significant effects on the take-off.
The first is that a greater amount of mass for a given force will reduce the acceleration because it is
harder to move a heavier object.
This is the effect of inertia.
We can see this effect in this part of the formula.
So therefore the take-off distance will increase.
A second effect of mass is that there will be more load on the wheels and therefore wheel drag will
increase.
Again this effect will increase the take-off distance.
The third effect is not as apparent as the others.
A heavier aircraft will require more lift in order to get airborne.
To get this lift the aircraft will need to travel faster and will require a higher take-off speed.
Getting to this higher speed will require a greater distance and therefore the take-off distance will increase.
The last effect is that a higher mass will reduce the aircraft's angle in the initial climb.
This decreased angle will increase the airborne distance to get to 50 feet as shown.
A rough calculation to help you quickly quantify the effect is to know that a 10% increase in mass
will increase the take-off distance by approximately 20% or a factor of 1.2.
The next factor to discuss that will affect the take-off is the air density.
This has three effects.
The first is that a reduced density will reduce the thrust.
This is because not as much air and therefore not as much oxygen enters the engine for combustion.
This will reduce the acceleration force and as a result increase the take-off distance.
The second effect of a reduced density is to do with the take-off speed.
A reduction in density will increase the true airspeed for any given indicated airspeed.
To attain the same indicated airspeed, which is a function of dynamic pressure, the aircraft needs to travel faster through
the air.
This means that the aircraft's true take-off speed is higher and as a result the take-off distance will
be greater.
The third effect that density has on the take-off distance is that it will reduce the angle of the
initial climb
and as already mentioned this will increase the airborne distance to get to 50 feet.
We must be aware of what causes low density and therefore low aircraft performance.
A useful statement to memorise is that high, hot or humid conditions all decrease air density and aircraft performance.
A rough calculation you can use is to remember that for every 1000 feet increase in altitude or 10°C increase
in temperature,
the take-off distance will increase by approximately 10% or a factor of 1.1.
All aircraft will get airborne at a certain indicated airspeed or dynamic pressure.
However, if there is a headwind of 20 knots, then a stationary aircraft will already be experiencing dynamic pressure
and register an indicated airspeed of 20 knots, yet the aircraft ground speed is zero.
So the aircraft is already closer to its flying speed even though it is not yet moving.
So, headwinds will reduce the ground speed for any given indicated airspeed and consequently reduce your take-off distance.
Tailwinds will increase your ground speed for any given indicated airspeed and increase your take-off distance.
An additional effect that wind will have is on the angle of the initial climb.
Headwinds will tend to increase the climb angle and reduce the take-off distance,
whereas tailwinds will reduce the angle and increase the take-off distance.
When using winds it is recommended that you do not use the actual wind that is given,
just in case when you actually get airborne the wind changes and thus your take-off distance changes.
When calculating take-off distances, the recommendation is that you use 50% of the headwind component
and 150% of the tailwind component.
As an example, if the wind were reported to be 10 knots of headwind, only use 50% of this,
which is 5 knots.
Most performance graphs that calculate the take-off distance have this recommendation already applied.
The example shown here includes this.
You will notice it because of the different slopes of the headwind and tailwind lines.
Tailwinds have a greater effect on the distance than headwinds.
Slope has an impact on the take-off distance because it affects how the component of weight acts along the
longitudinal axis of the aircraft.
Notice here that a stationary aircraft on a level runway has weight acting straight down towards the centre of the
earth.
However, when the aircraft is on an upslope, the weight still acts towards the centre of the earth,
but there is a component of weight, shown here, which acts in the direction of drag, that is backwards along
the longitudinal axis.
This will reduce the acceleration and increase your take-off distance.
Conversely, when the aircraft is on a downslope, the weight component now acts in the direction of thrust.
This will increase the acceleration and decrease your take-off distance.
A rough calculation to help you quickly quantify the effect of slopes is to assume that for every 1%
slope,
the take-off distance is affected by 5% or a factor of 1.05.
Most take-off performance graphs assume a paved, hard surface.
If conditions of the runway are not like this, then the effect on take-off distance needs to be understood
and corrections applied.
A lot of small airfields have grass runways.
The grass will increase the drag on the wheels, or in other words, increase the rolling resistance.
This will reduce the acceleration and as a result increase the take-off distance.
A wet runway, whether grass or paved, will have a similar effect and increase the drag on the wheels,
and hence increase the take-off distance.
There are approximate calculations to help you calculate the effect.
Dry grass up to 20 cm will increase the take-off distance by 20% or a factor of 1
.2.
However, wet grass will have an even greater effect and it will increase the take-off distance by 30%
or by a factor of 1.3.
If the ground is very soft or there is snow, then the take-off distance will increase by approximately 25%,
which is a factor of 1.25.
Do not attempt to take off in grass more than 25 cm long or if there is slush on the
runway.
The reason is that the wheel drag may be so high that the acceleration will be insufficient to get the
aircraft to a flying speed.
Flaps are trailing edge devices used to increase the camber of the wing and generate more lift, thereby reducing the
stalling speed.
However, a lot of flap will dramatically increase drag as well.
Let's take a look at how different flap angles can affect the take-off distance.
With no flap, the aircraft will accelerate and then climb to the screen height as shown.
However, with some flaps selected, more lift will be generated and the aircraft will get airborne at a slower speed.
This can be as much as 10 knots less than without any flaps.
However, with flaps a small penalty is paid in the fact that the initial climb angle is shallower.
Nevertheless, the overall effect is that the take-off distance reduces.
If too much flap is used in the take-off, then the wing will generate more drag than lift.
The result of that would be that the acceleration is reduced and the take-off distance increases compared to a
no flap take-off.
Too much flap would also significantly reduce the initial climb angle, exacerbating the effect.
However, the effect of flap varies significantly from one aircraft to another.
Some aircraft require the use of flaps for take-off, whilst some don't.
Familiarise yourself with your aircraft and use the flap settings laid down by the manufacturer in the aircraft manual.
There are many other factors which need consideration when calculating your take-off distance.
These include the use of differential braking when trying to maintain directional control in strong crosswinds.
Other considerations include wind shear and turbulence, as well as contamination on the aircraft itself,
such as ice, snow or rain, and low tyre pressure which will increase rolling resistance.
The majority of modern light aeroplanes were originally classified as Performance Group E for the purposes of public transport.
The performance information contained in the manuals is unfactored, or otherwise known as gross performance.
This means that the data represents the performance achieved by the manufacturer using a test pilot and in ideal conditions.
This level of performance is unlikely to be met by the average pilot because of the variations in the actual
flying techniques described in the manual.
It is therefore prudent to add safety factors to the data in order to take account of this performance loss.
This gives us what is known as net performance.
After the application of the corrections for the variables, it has been recommended by the authority that the take-off
distances from the aircraft manual be increased by 33% or a factor of 1.33.
Having seen all the different variables in the take-off, let us now take a look at a typical take
-off distance graph that one might see in the aircraft manual.
Firstly, the title of the graph shows you that this is for calculating the take-off with no flaps.
For other flap settings, you may have to consult another graph.
Next, have a look at the associated conditions, paying particular notice to the techniques used for the power and undercarriage
as shown here.
Also notice that this graph is assuming a runway that is paved, level and dry.
If our runway is not like this, then corrections will need to be made to the values this graph will
give.
We have already discussed those correction factors.
Also notice this box here which highlights the rotation and screen height speeds for different weights.
Adhere to these accurately because these are the speeds which have been used to construct this graph.
Let's now examine the graph itself.
The left-hand carpet of the graph involves the variations of temperature and pressure altitude which correct the take-off
distance for density changes.
In the example, we will use a temperature of 15°C at an airfield 4000 feet above mean sea level.
To use the graph then, travel upwards from 15°C until you have reached the 4000 feet pressure altitude line.
Then move right until you meet the reference line shown here.
The next variable is weight. In our example, let us assume 3400 pounds.
From the reference line, we must move down along the sloping guidelines until we reach 3400 pounds as shown here.
From this point, we go horizontally to the right until the next reference line.
The next variable is wind. In our example, we will assume a 10 knot headwind.
Notice though that the slopes of the headwind and tailwind lines are different, suggesting that the 150% tailwind and
50% headwind rules we mentioned earlier have already been applied for us.
Travel down the headwind line until you reach 10 knots headwind and then move horizontally right once more until the
last reference line.
Remember that the take-off is not complete until we reach 50 feet.
So from this point on the graph, we must travel up the guidelines until the end.
The take-off distance in our example is approximately 2300 feet from brake release to the 50 feet screen.
However, let us now assume that the runway is grass and that it's wet.
Remember that both these variables cause an increase in the take-off distance.
Since the graph only assumes a paved dry runway, we will have to make a correction to the take-off
distance.
You may recall that wet grass increases the distance by 30% or by a factor of 1.3.
Our calculated take-off distance was 2300 feet, but correcting it for wet grass would now make the take-off
distance 2990 feet.
Let us now assume the runway has an upslope of 1%.
You may recall this would increase the take-off distance by 5% or by a factor of 1.05.
This now makes our take-off distance 3140 feet.
Finally, it is highly recommended that we add an additional safety factor of 1.33.
This now makes our assumed take-off distance 4176 feet.
You must ensure that the distance available at the airfield is more than this to ensure a safe take-off.
We have seen only one way in which take-off data may be presented, namely in a graph.
There are many others, some even in tabulated format.
Familiarise yourself with the format for the aircraft you are flying,
and correctly adjust the calculated take-off distance for any variable that is not accounted for in the data.
Do not forget to add the safety factor once you have applied all the variables.
An understanding of the effects of various conditions and factors that control the take-off
will better equip you to ensure that every take-off is well within the limits of the runway available and
thus minimise any accidents.
The CLIMB section of aircraft performance deals with the analysis of that stage of flight from take-off to the
en route altitude or cruising altitude.
There are primarily two ways in which we can examine the performance of an aircraft in the climb.
Firstly, by considering the angle or gradient at which the aircraft climbs,
and secondly, by looking at the rate at which it climbs.
We have assumed for the purpose of this lesson that the lift force acts through the centre of gravity and
directly counterbalances weight.
This is an oversimplification of the true nature of lift forces. These will be dealt with in detail in principles
of flight.
Notice that the forward acting force, thrust, balances the rearward acting force, drag.
However, when the aircraft attempts to climb, as shown here,
there is a component of weight which acts backwards along the longitudinal axis in the direction of drag.
This component is known as weight apparent drag.
You'll now notice that the amount of rearward acting force is more than the forward acting force.
If the situation were left unchecked, the aircraft would decelerate and possibly stall.
To maintain a steady speed along the flight path,
an additional amount of thrust is needed to balance the amount of weight apparent drag.
This is known as excess thrust available.
The excess thrust available can be calculated by taking the total thrust available and subtracting the total aerodynamic drag.
If there is a large amount of excess thrust available, then it's possible to balance a larger amount of weight
apparent drag.
In other words, the aircraft can afford to climb at a steep angle.
If the aircraft only has a small amount of excess thrust available, then notice now that there is too much
weight apparent drag
and that the angle will need to be reduced so that the excess thrust available balances the weight apparent drag.
From what we've just seen, for any given weight, the angle or gradient of climb depends on the amount of
excess thrust available.
We show here the formula to calculate the gradient of climb.
Remember that total thrust available minus total aerodynamic drag gave us excess thrust available.
For any given weight, the larger the excess thrust available, the larger the climb angle.
The less the excess thrust available, the smaller the climb angle.
In order to maximise the aircraft's climb angle, we need to maximise the excess thrust available.
Shown here is a graph of total thrust available and total drag for a typical light general aviation aircraft.
The maximum amount of excess thrust will be found when the distance between the thrust and the drag lines is
at its greatest, as shown here.
Notice that it occurs at this speed.
At any other speed, the excess thrust available is less.
This speed is known as VX.
It's defined as the speed for the best angle or gradient of climb.
If there are obstacles close to the end of the runway, you'll need the best possible climb angle to clear
the obstacles by the greatest margin.
Let's now examine what factors can influence the angle of climb.
Changes in weight can significantly alter the angle of climb.
Shown here is an aircraft in a steady climb.
However, if we increase the mass of the aircraft, notice that now the aircraft has more weight apparent drag.
The forward and rearward forces are not balanced, and the aircraft will decelerate and possibly stall.
In order to maintain the speed, the angle will need to be reduced so that the aircraft has less weight
apparent drag.
Now you can see that the forces are once again balanced, but the climb angle is less.
Weight has another effect though.
An increase in weight will require an increase in lift.
If lift is increased, then lift induced drag will increase, which in turn will increase total drag.
Having more drag will require yet more excess thrust.
If this is not available, then climb angle will reduce.
The graph here demonstrates this.
This first picture shows the drag curve and the excess thrust available for a light aircraft.
At higher weights, total drag is high, and the drag curve moves up and right.
Notice there is less excess thrust available.
However, what is important to see is that the speed for maximum excess thrust available is no longer the same.
It's higher.
So with higher weight, Vx is higher, and the climb angle is shallower.
The next factor that affects the climb angle is the configuration of the aircraft.
By this we refer to the influence of flaps and undercarriage.
If the undercarriage and flaps are deployed, then the profile drag of the aircraft will increase.
This of course increases total drag, and the drag curve moves upwards and to the left.
You can see that there is a reduction in excess thrust available as shown.
You may recall that a reduction in excess thrust available reduces your climb angle.
However, what is important to see is that the speed for maximum excess thrust available is no longer the same.
So with the undercarriage and or flaps deployed, Vx is lower and the climb angle is shallower.
It therefore seems prudent not to use flaps so that the climb angle is as large as possible.
But you may recall that flaps helped improve the take-off performance.
If you have used flap for take-off, then when it's safe to do so, remove the flap in stages.
Consult your aircraft manual for precise instructions.
Density is another important factor that affects the angle of climb.
Density has a direct influence on the thrust produced by the engine.
The graph shows the thrust and drag lines for high density.
Notice the large amount of excess thrust available and therefore large climb angle.
However, with a reduction in density the thrust is less and the thrust line moves downwards,
reducing the excess thrust available and therefore the climb angle.
You may recall that reduced density can be the result of high altitude, high temperature or high humidity.
Next we'll look at the effect of wind on our angle of climb.
This diagram shows a typical climb angle achieved in still air.
If there's a headwind, notice that the aircraft relative to the still air path is being pushed backwards.
The result is that the climb angle increases and the distance covered for the climb to an equivalent height is
reduced.
However, with a tailwind the aircraft is being pushed forward relative to the still air path,
with the result that the climb angle is decreased and the climb distance increased.
You may recall that there were two primary ways of analysing the performance of the aircraft in a climb,
namely the angle of climb and the rate of climb.
Let's now focus on the rate of climb.
The rate of climb is the vertical speed of an aircraft measured in feet per minute
and shown in the cockpit by a vertical speed indicator or VSI.
Another way to consider rate of climb is to see it as the velocity of a gradient.
Let's assume that two identical aircraft have the same angle of climb.
The one on the right is travelling faster along that gradient and will therefore have a better rate of climb.
Equally though, if two identical aircraft were to have the same speed,
but the aircraft on the right travelled along a steeper angle, then it would have a higher rate of climb.
Therefore, rate of climb is a function of both speed and gradient.
Let's return to the formula for the gradient illustrated here.
All that's needed now is to add the velocity function as shown.
This is now the formula for the rate of climb.
This adds a different perspective to the formula.
We now have a formula for power.
Velocity multiplied by thrust available gives power available
and velocity multiplied by drag gives power required.
The formula now reads rate of climb equals power available minus power required,
which gives us excess power, divided by weight.
In order to maximise the rate of climb, we need to maximise excess power for a given weight.
Shown here is a graph of power available and power required for a typical light aircraft.
The maximum amount of excess power available will be found where the distance between the curves is the greatest.
As shown here. Notice that it occurs at this speed.
In fact, this is the tangent point of the power required curve.
At any other speed, the excess power is less and the rate of climb less.
The speed indicated here is known as Vy.
It's defined as the speed for the best rate of climb.
There is another speed which is found at the same point as Vy.
That is the tangent of the power required curve. This is Vmd.
The velocity for minimum drag.
Vy is the airspeed to use to climb to the cruise or en route altitude in the shortest possible time,
and it is most commonly given as the climb speed.
However, some aircraft manuals suggest using a faster speed.
This causes a greater flow of air through the engine to help reduce engine temperatures with such high power settings.
Another reason is that flying faster means a lower nose attitude which improves forward visibility.
Let's now examine what factors can influence the rate of climb.
Changes in weight can significantly alter the rate of climb.
You may recall that an increase in weight creates more weight apparent drag, which reduces the angle of climb.
For any given airspeed, if the angle of climb reduces, then so will the rate of climb, because they are
fundamentally linked.
Weight has a further effect though. An increase in weight will require an increase in lift.
If lift is increased, then lift induced drag will increase which in turn will increase total drag.
Notice though that power required is a velocity multiplied by drag.
So if drag increases, so will the power required.
This results in less excess power available and will therefore reduce our rate of climb.
The graph here demonstrates this.
This first picture shows the power required curve and the excess power available of a light aircraft.
At higher weights, the power required is more and the curve moves up and right.
Notice there is less excess power available.
However, what is important to see is that the speed for maximum excess power available is no longer the same.
It's now higher.
So with higher weight, Vy is also higher and the rate of climb is lower.
The next factor that affects the rate of climb is the configuration of the aircraft, in other words the use
of flaps and undercarriage.
If the undercarriage and flaps are deployed, then the profile drag of the aircraft will increase.
This increases total drag and the drag curve moves upwards and to the left.
The power required curve will follow the same movement as the drag curve.
Notice the reduction in excess power available as shown.
You may recall that a reduction in excess power reduces your rate of climb.
However, the important effect here is that the speed for maximum excess power is no longer the same.
It's now lower.
So with the undercarriage and or flaps deployed, Vy is lower and so is the rate of climb.
If you use flaps for take-off, remove them in stages when you've attained a positive stable climb.
Density is another important factor that affects the rate of climb.
A reduction in density will cause the power available to reduce and power required to increase.
You'll notice that there is less excess power available and this causes a reduction in the rate of climb.
Notice that true airspeed for maximum excess power is higher.
However, the rate at which Vy, expressed as true airspeed, increases with altitude is fairly small.
This means that Vy, expressed as indicated airspeed, is actually reducing.
So, in summary, reduced density lowers the indicated airspeed of Vy and reduces the rate of climb.
The horizontal movement of air, otherwise known as wind, has no effect on the vertical velocity of an aircraft,
but it does affect the horizontal velocity and consequently the angle of climb.
We see here an aircraft climbing at 400 feet per minute with no wind.
If there was a headwind blowing the aircraft would reach the same vertical distance, or height, in one minute,
but not cover the same horizontal distance, because a headwind reduces ground speed.
In the case of a tailwind the aircraft would reach the same vertical distance, or height, in one minute,
but this time over a greater horizontal distance.
This completes the theoretical study of angle and rate of climb.
Most aircraft manuals have a graph similar to the one shown here.
These graphs are used to help calculate the time, distance and fuel required to attain a given cruising altitude in
various conditions.
Let's work through an example.
Let's assume an airfield elevation of 4,000 feet with an outside air temperature of 20 degrees Celsius.
An aircraft with an initial mass of 3,650 pounds climbs to an altitude of 10,000 feet where the
temperature is 0 degrees Celsius.
Firstly, start off by selecting the temperature at your cruise altitude.
Then move upwards until you reach the 10,000 foot line.
Now move right until you reach the weight line which is 3,650 pounds.
Move straight downwards to find the time, fuel and distance to climb.
However, the graph assumes the climb was from sea level to 10,000 feet, which of course it wasn't.
However, the airfield has an elevation of 4,000 feet, so we need to subtract the part that was from
sea level to 4,000 feet.
To do this, firstly calculate the time, fuel and distance required to climb from sea level to 4,000 feet
using the same method that you've just been shown.
Now subtract this information from the previous calculations and the figures now show the climb information from 4,000 feet
to 10,000 feet.
This completes the lesson on climb performance.
Ensure you familiarise yourself with how various factors and environmental conditions affect your climb performance in terms of both the
angle and rate of climb.
Practise using the climb graphs in your aircraft manual so that you can obtain the most accurate information as possible
on the climb performance of your aircraft.
But always remember, it's unlikely you will achieve the performance that the manual suggests, so make allowances and be prepared
in advance.
In this lesson we'll concentrate on the performance of an aircraft in the en route phase of flight.
The en route phase includes both the cruise and the descent.
Performance in the en route phase can be measured using the aircraft's range and endurance.
These will be discussed in detail later on in the lesson.
Let's firstly examine the forces acting on an aircraft in the cruise.
They're split into couples or pairs.
The first of these couples is produced by lift and weight.
Weight acts through the centre of gravity of the aircraft directly towards the centre of the earth.
Lift balances weight and acts through the centre of pressure shown here.
The effect of this couple on the aircraft is causing a nose-down pitching moment.
The lift-weight couple is comparatively strong and thus the nose-down pitching moment is also strong.
The other couple is produced by the forces of thrust and drag.
The effect of this couple is to cause a nose-up pitching moment.
Notice though that this couple is far weaker than the lift-weight couple,
and the nose-up pitching moment does not balance the stronger nose-down pitching moment.
As a result there is still a nose-down tendency.
Thus to maintain level flight we need somehow to augment the nose-up pitching moment.
This is achieved by the tailplane.
For instance the elevators can be deflected to create a downforce on the tail,
which will cause a nose-up pitching moment.
Notice that now the nose-up and nose-down pitching moments are balanced and level flight is possible.
However, though it's essential for level flight,
the tailplane downforce, or tail-load as it's sometimes referred to, has two adverse effects on aircraft performance.
Notice that the tailplane downforce acts in the same direction as weight.
It therefore increases the effective weight of the aircraft.
The second adverse effect is its contribution to drag.
The tailplane is an aerodynamic surface, able to generate both up and down forces.
It will therefore produce induced drag.
Thus it follows that the greater the amount of tailplane downforce, the greater the induced drag.
To some extent the amount of tailplane downforce required for level flight can be manipulated by changing the centre of
gravity.
If the centre of gravity is more forwards than normal, the strength of the lift-weight couple is increased because
the arm of the two forces is now longer.
This increases the pitching-down moment.
To balance this, more tailplane downforce is needed to maintain level flight.
This has a detrimental effect on the aircraft's performance, reducing both its range and endurance capability.
However, if the centre of gravity is moved aft, the strength of the lift-weight couple is decreased because the
arm of the two forces is now shorter.
This decreases the pitching-down moment.
To balance this, less tailplane downforce is needed to maintain level flight.
This has a beneficial effect on the aircraft's performance, this time increasing both its range and endurance capability.
But remember, before you fly, it's essential to always carry out a weight and balance check to ensure the centre
of gravity is still within the limits published in the aircraft manual.
Before we discuss the details of aircraft range and endurance, let's just look again at the Power Available and Power
Required graph, as shown.
In straight, unaccelerated level flight, thrust must balance drag.
Thrust is produced by the propeller's use of engine power.
At low speeds, induced drag is high and therefore more power is required to generate the thrust to overcome this,
as shown here.
The same applies for high speed.
At high speeds, drag is also high because of parasite drag, and therefore high power is required to generate enough
thrust to balance the drag, as shown here.
Between these speeds the drag is much less, and therefore the power required is also much less.
The speed for minimum power is called Vmp, and is found at the bottom of the power required curve, as
shown.
At various air speeds the engine of the aircraft can produce a different amount of maximum power. This is shown
by the Power Available curve.
Let's now examine in detail the endurance of an aircraft in the cruise.
Endurance is the time an aircraft can remain airborne on a given quantity of fuel, or, to put it another
way, the minimum fuel used for a given airborne time.
For maximum endurance the aircraft needs to fly at the lowest possible fuel consumption rate.
Since fuel consumption is a function of the power applied, then minimum fuel consumption in straightened level flight will occur
at the airspeed where power is at a minimum.
This is Vmp. So the speed for maximum endurance is Vmp.
The two major factors that can affect endurance and the speed for maximum endurance are mass and configuration.
Increasing the weight of the aircraft will require a greater amount of lift and consequently cause greater induced drag. This
will require more power.
If the aircraft uses more power then the fuel consumption will increase reducing endurance.
The effect of increasing just the induced drag causes the power required curve to move up and right.
Notice that the speed for maximum endurance, Vmp, is now higher.
Aircraft configuration changes, such as the deployment of flaps and undercarriage, will cause an increase in parasite drag.
This configuration requires more power and consequently will increase fuel consumption and decrease endurance.
The power required curve this time moves up and left.
Notice that the speed for max endurance, Vmp, is now lower.
Whenever you need to stay aloft for as long as possible, maybe due to a delay in a landing clearance,
it's advisable to set the best endurance speed.
A more useful performance parameter, and one that aircraft designers continually try to improve, is that of aircraft range.
Whereas endurance was about maximizing airborne time, range is more concerned with maximizing distance covered.
Maximum range can be defined as being the maximum distance an aircraft can fly with a given fuel quantity.
This fuel quantity being dictated by the distance you wish to fly.
It may seem logical to assume that if the aircraft were to fly at the speed for minimum power required,
that is, best endurance speed, using the least amount of fuel, then it would also cover the greatest distance.
Unfortunately, this is not the case.
The concern with range is not just about minimum fuel consumption, but also with speed.
The formula for range is shown here. It's a function of true airspeed and fuel flow.
Fuel flow is determined by the power required. From the formula, you can see that to maximize range we need
the best velocity power ratio. In other words, we need velocity to be high and power to be low.
We can resolve one of the issues already.
If we fly at VMP, power required will be at a minimum as shown here.
However, notice that the bottom of the power required curve is relatively flat.
This means that if speed were to increase from VMP by this margin, the penalty paid by an increase in
the power required is fairly minimal.
This means that the aircraft can travel significantly faster than VMP with only a small increase in the power required
and fuel consumption.
This faster speed will increase the range of the aircraft.
The tangent to the curve is where the velocity power required ratio is at a maximum. In other words, range
is at a maximum.
You may recall that the tangent to the power required curve gave us a speed called VMD, so the maximum
range speed is at VMD.
There are three main factors which can affect the range and the best range speed of an aircraft. These are
mass, configuration and wind.
Increasing the weight of the aircraft will require a greater amount of lift and consequently cause greater induced drag. This
will require more power.
If the aircraft uses more power, then the fuel consumption will increase, reducing the range.
The effect of increasing just the induced drag causes the power required curve to move up and right.
Notice that the speed for maximum range, VMD, is now higher.
Aircraft configuration changes, such as the deployment of flaps and undercarriage, cause an increase in parasite drag.
This configuration requires more power and will consequently increase fuel consumption and decrease range.
However, the power required curve this time moves up and left. Notice that the speed for maximum range, VMD, is
now lower.
A point always to bear in mind in the cruise is that any drag will be detrimental to range and
endurance. Such drag could be caused by damaged or contaminated aerofoils.
Be especially aware of icing. Icing will not only change the shape of the wing, making it less efficient at
producing lift, but it will also increase weight and drag. All this is detrimental to aircraft performance.
Wind is the last of the major factors affecting the range of an aircraft.
Headwinds will cause the aircraft to travel slower over the ground and therefore cover less distance for a given amount
of fuel. Thus in headwinds range is reduced.
In order to minimise this effect the speed of the aircraft is increased by a margin slightly less than the
amount of the headwind.
The increase in speed will increase the power required and therefore the fuel consumption, but the aircraft will be exposed
to the headwind for a shorter time period if it's flying faster.
However, a tailwind will increase the ground speed and increase the distance covered for a given amount of fuel consumed,
thereby increasing range.
For maximum range with a tailwind the speed for best range should be decreased slightly so as to reduce the
power required and fuel consumption.
The reduction in speed is slightly less than the amount of tailwind being experienced.
In consideration of the effect of headwinds and tailwinds you should always try and select an altitude which gives the
best wind conditions to help maximise your range.
In the aircraft manual there are graphs and tables which help you calculate the endurance and range of your aircraft
under various conditions. Shown here is a typical endurance graph for a light aircraft.
Let's assume a cruise altitude of 7,000 feet with an outside air temperature of 7 degrees Celsius at 65
% power.
The endurance of the aircraft allowing 45 minutes of reserve fuel is 6.6 hours or 7.4 hours allowing
for no reserves.
The graph shown here is for calculating the aircraft range and it works in exactly the same way as the
endurance graph except that instead of working out airborne time distance flown is shown.
To achieve these range and endurance figures be careful to follow the techniques described in the manual, especially with regard
to correctly leaning the mixture.
Another aspect of the en route phase of flight is the descent. There are two ways of assessing the descent
performance of an aircraft.
Either by angle of descent, sometimes called descent range, or rate of descent, sometimes called descent endurance.
Firstly let's consider angle of descent. Shown here is an aircraft in level flight with the forward forces balancing the
rearward forces.
In order to initiate a steady descent, thrust is normally reduced. Notice that the rearward forces exceed the forward forces
and the aircraft would slow down.
In order to balance the forces and keep the speed, the nose is lowered until the weight apparent thrust provides
enough forward force to balance the drag force.
Now the aircraft will maintain this steady descent angle at a constant speed.
If thrust were reduced even more, then more weight apparent thrust is needed to maintain the speed and the nose
must be lowered even more which will increase the descent angle.
In fact, the greater the margin of drag force over the thrust force, the greater the descent angle because the
weight apparent thrust must make up the difference.
This difference is known as excess drag.
Shown here is the formula for the angle of descent. In fact, it's the exact opposite of that for the
angle of climb formula.
Now let's consider the aircraft in a glide.
If the engine were to fail, then the objective would be to find a suitable forced landing field within the
gliding range of the aircraft.
To give the greatest chance of success, we would ideally want the greatest glide distance.
This will be achieved with the minimum possible angle of descent.
A large descent angle, as shown, decreases the descent range, leaving a smaller area in which to find suitable terrain
for a forced landing.
To achieve the minimum angle of descent and maximise your descent range, fly at the speed for minimum drag, VMD,
and keep the aircraft as clean as possible to make drag as small as possible.
This means that only a comparatively small amount of weight apparent thrust is needed and the nose need only be
lowered by a small amount.
Therefore, the minimum glide angle, or maximum descent range, is achieved by flying at VMD.
VMD is also the speed for the best lift drag ratio, which in fact is what determines the glide range.
When the lift drag ratio is greatest, the angle between the lift and the drag vectors is the smallest possible
angle between these vectors.
This angle is equal to the minimum glide angle in still air.
Because in a steady glide all the forces are equal, weight equals the vertical component of lift, and weight apparent
thrust equals drag.
The angle between weight and weight apparent thrust is thus also theta, the angle of glide.
It's customary to use this depiction of theta as the angle of glide.
However, if there is a lot of drag, as there would be with flaps lowered for instance, then this needs
to be balanced with a lot of weight apparent thrust if we are to maintain airspeed.
In this case, because there is so much drag, the nose would have to be lowered substantially to obtain the
necessary weight apparent thrust.
Having a steep glide angle can be useful, especially when making an approach to land and if you find yourself
too high.
If this was the case, increase the drag by increasing the flap angle, which will help you to steepen the
angle.
Side slipping can provide even greater drag and results in an even steeper descent angle.
Be sure not to try a side slip unless your aircraft is cleared for it and you have been cleared
for the manoeuvre by your instructor.
A change in mass has no effect on the glide angle.
For an aircraft of greater mass or weight, there will be more weight apparent thrust for the same nose down
pitch attitude.
The forward forces are greater than the rearward forces, and the aircraft will accelerate.
At this higher speed, more lift will be generated, which balances the greater weight.
As shown by the lift formula on the screen, L equals a half rho V squared C L S, an
increase in velocity will give an increase in lift.
Therefore, the lift-drag ratio remains the same as for the lighter aircraft.
As it's only the lift-drag ratio which determines the angle of glide, the glide angle of the heavier aircraft
will remain the same.
However, the speed at which this glide angle is achieved will be higher for the heavier aircraft.
We may sum up by saying that the minimum glide angle for a given aircraft will remain constant whatever its
payload.
But at the higher mass or weight, the speed for minimum glide angle will be higher.
An important consideration to the glide angle and descent range is the effect of wind.
The diagram shows the descent angle and descent range with no wind.
In a headwind, the horizontal distance covered by the aircraft during the descent will be less and the glide angle
consequently steeper.
Conversely, if there is a tailwind blowing, the aircraft will move forwards relative to the no-wind diagram.
This decreases the glide angle and increases the descent range.
In a tailwind, ensure you decrease your glide speed to maximise the effect of this.
You may recall we mentioned that there were two ways of assessing the descent performance.
We've covered the angle of descent or descent range, now let's consider the rate of descent or descent endurance.
Rate of descent is the opposite consideration from rate of climb.
This is the formula for the rate of climb.
But now, to work out rate of descent, instead of power available minus power required, it will read power required
minus power available.
In our example, we'll assume a glide and therefore no thrust.
This means that there is no power available.
In order to achieve the minimum rate of descent, we must fly at the speed for minimum power required for
straight and level flight for any given weight.
Therefore, in order to achieve the minimum rate of descent following engine failure or with the engine throttled right back,
you should fly the aircraft at Vmp.
This completes the lesson on enroute performance.
To calculate your range and endurance, use the performance graphs or tables published in the aircraft manual and correctly follow
the techniques listed.
Especially the speeds.
Always bear in mind the effect of wind on your range and glide distance in order to ensure you can
still get to your destination and or chosen forced landing field.
The landing distance is defined as beginning from a point where the aircraft is 50 feet above the landing surface
and ending where the aircraft comes to a full stop.
The 50 foot marker is referred to as the landing screen height, which is similar to the take-off screen
height.
You may remember that there was a specific speed that had to be reached by the screen height in take
-off.
Similarly for landing, a specific speed needs to be obtained before the landing screen height.
This is called the landing reference speed or Vref for short.
Vref must be more than 1.3 times the stalling speed in the landing configuration.
This stalling speed is known as Vso.
For example, if the stall speed in the landing configuration was 50 knots, then Vref would be 30% more
than this, that is 65 knots.
The objective when landing is to minimise the landing distance.
This is achieved by landing without thrust and using the maximum practical braking action.
The use of landing flaps will greatly increase the drag on the aircraft, helping to slow the aircraft down and
reduce the stall speed.
Obviously, the landing distance of the aircraft must not exceed the landing distance available.
The landing distance available or LDA at an airfield is defined as being the length of runway suitable for landing,
taking into account any obstacles.
Let's now look at the various factors that can influence the landing performance.
To begin with, there is the effect of increasing mass.
Increasing mass has several significant effects on the landing.
The first of these is that a greater amount of mass makes the aircraft harder to stop due to its
momentum.
So therefore the landing distance will increase.
The second effect is that a heavier aircraft will have a higher stall speed.
If the stall speed is higher, then Vref is higher, meaning the aircraft is approaching to land at a higher
speed.
A faster aircraft will be harder to stop.
So therefore the landing distance will increase.
The final effect of mass is actually beneficial to the landing performance.
A greater weight will put a greater load on the wheel and increase friction in the wheel bearing and between
the tyres and the runway.
This extra load on the wheels will allow for harder braking action.
In conclusion, the overall effect of increased aircraft mass is that the landing distance required is also increased.
A rough calculation to help quickly quantify the effect is that a 10% increase in mass will increase the
landing distance by approximately 10% or a factor of 1.1.
The next factor affecting the landing distance is the air density.
A reduction in air density will increase the true airspeed for any given indicated airspeed.
This means that if the indicated approach speed, Vref, is 65 knots, the true speed of the aircraft through the
air is more than this.
Therefore the aircraft is actually landing with a higher true speed.
This will increase the landing distance required.
What we must be aware of is what causes low density and therefore reduced aircraft performance.
You may recall from earlier lessons that high, hot or humid conditions all decrease air density and aircraft performance.
A rough calculation to remember is that for every 1,000ft increase in altitude or 10°C increase in temperature, the
landing distance will increase by approximately 5% or a factor of 1.05.
Wind affects the landing because of its influences on the ground speed of the aircraft.
Headwinds will reduce the ground speed for any given indicated airspeed and consequently reduce your landing distance required.
Tailwinds will increase your ground speed for a given indicated airspeed and increase your landing distance required.
It's therefore not advisable to land with tailwinds.
Select the runway which has the greatest headwind component.
When using winds to calculate the landing distance required, it's recommended that you do not use the actual wind that
is given.
The recommendation is that you use only 50% of the headwind component and 150% of the tailwind component.
Most performance graphs from which pilots calculate the landing distance have this recommendation already applied, as shown in this example.
The different coefficients applied to headwinds and tailwinds is noticeable because of the different slopes of the headwind and tailwind
lines.
Tailwinds have a greater influence on the distance than headwinds.
The slope of a runway has an impact on the landing distance because of its effect on the component of
weight which acts along the longitudinal axis of the aircraft.
Notice here that a stationary aircraft on a level runway has weight acting straight down towards the centre of the
earth.
However, when the aircraft is on an upslope, the greater part of the weight still acts towards the centre of
the earth, but there is a component of weight, shown here, which acts backwards along the longitudinal axis.
This will increase the deceleration and decrease your landing distance required.
Conversely, when the aircraft is on a downslope, a component of weight now acts in the direction of thrust. This
will decrease the deceleration and increase your landing distance required.
A rough calculation to help quickly quantify the effect of slopes is to assume that for every 1% of
slope the landing distance is increased or reduced by 5%, or a factor of 1.05.
Most landing performance graphs assume a paved hard surface.
If the runway surface is not paved, an understanding of surface effects is required, and then corrections applied to the
landing distance.
A lot of small airfields have grass runways. The grass will increase the drag on the wheels, but more importantly
will reduce the friction between the tyre and the runway.
This means that the brakes cannot be applied as hard as on a paved surface, otherwise the wheels would slip.
The result is an increased landing distance required.
A wet runway will reduce the friction between the wheel and runway, preventing full braking force being applied.
This will increase the landing distance required.
The following are approximate figures to help you calculate the effect.
Dry grass of up to 20cm in height will increase the landing distance by 20%, or a factor of 1
.2.
However, wet grass will increase the landing distance by 30%, or by a factor of 1.3.
If the ground is very soft, or if there is surface snow, then the landing distance will increase by approximately
25%, which is a factor of 1.25.
Flaps increase drag, particularly at high flap settings.
Let's take a look at how different flap angles can affect the landing distance.
Shown on screen is the landing distance required with no flap.
The main point to note about no flap landings is that the aircraft floats quite a long way before touching
down.
If a small flap angle is selected, the stall speed and therefore Vref will reduce, but also drag will increase.
Both of these effects will reduce the landing distance required.
If a large flap angle is selected, say landing flaps, the stall speed and therefore Vref will reduce even more
and drag will increase even more.
Landing flap settings will further reduce the landing distance required.
It's therefore advisable to land using the recommended landing flap setting for the aircraft, since this offers the greatest reduction
in landing distance.
However, the effect of flap varies significantly from one aircraft to another.
Familiarise yourself with your aircraft and use the flap settings laid down by the manufacturer in the aircraft manual.
There are many other factors which need consideration when assessing landing performance.
Be very wary of wind shear on the approach. Shown here is an aircraft on the approach with a constant
headwind.
However, the wind speed nearer the surface is often reduced because of the friction between the wind and the surface,
especially in high winds.
This can cause airspeed to decay causing the aircraft to sink and undershoot the aiming point on the runway.
Check the weather for wind shear warnings, increase your approach speed or, if necessary, divert to an alternative airport.
Other considerations include the effect of leaving the carb heat on.
Although it will not affect the landing distance, it will affect your ability to go around in the event of
an aborted landing.
This is because carb heat puts hot air into the engine, which reduces the density of the air and therefore
reduces the thrust the engine can produce.
On short finals, ensure you have selected carb heat off, but once again consult your aircraft manual and follow the
manufacturer's recommended actions.
The majority of modern light airplanes were originally classified as Performance Group E for the purposes of public transport.
The performance information contained in the manuals is unfactored or otherwise known as gross performance.
This means that the data represents the performance achieved by the manufacturer using a test pilot and in ideal conditions.
This level of performance is unlikely to be achieved when the aircraft is being flown by the average pilot, because
of the differences in the actual flying techniques described in the manual.
It is therefore prudent to add safety factors to the data in order to take account of this performance loss.
This is called net performance.
After the application of the corrections for variables, it has been recommended by the authority that the landing distance be
increased by 43%, or a factor of 1.43.
Having seen all the different possible variables affecting the landing performance, let's now take a look at a typical landing
distance graph that one might see in the aircraft manual.
Firstly, let's look at the associated conditions and assumptions.
Notice that we're assuming a rate of descent of 900 feet per minute, that the flaps are down, and that
the runway is paved, level, and dry.
If the runway surface conditions are not as assumed here, then corrections will have to be made to the figures
obtained from this graph.
We've already discussed the correction factors.
Also notice this box here, which highlights Vref, which is the speed to be attained at the 50 foot landing
screen height.
Adhere to these speeds accurately, because these are the speeds which the graph's information is based on.
Let's now examine the landing graph and calculate landing distances in various conditions.
In the example, we'll assume a temperature of 15 degrees Celsius at an airfield lying at 4000 feet above mean
sea level.
To use the graph then, travel upwards from 15 degrees Celsius until you reach the 4000 foot pressure altitude line.
Then move right until you meet the reference line shown here.
The next variable is weight. In our example, let's assume 3400 pounds.
From the reference line, we must move down along the sloping guidelines until we reach 3400 pounds as shown here.
From this point, we go horizontally to the right until the next reference line. The next variable is wind.
In our example, we'll assume a 10 knot headwind. Notice though that the slopes of the headwind and tailwind lines
are different,
suggesting that the 150% tailwind and 50% headwind rules we mentioned earlier have already been applied for us.
Travel down the headwind line until you reach 10 knots headwind and then move horizontally right once more until the
last reference line.
Remember that the landing distance begins at the 50 foot landing screen.
So from this point on the graph, we must travel up the guidelines until the end.
The landing distance in our example is approximately 1450 feet.
Now, let's assume that the runway is grass and that it is wet. Remember that both these variables cause an
increase in the landing distance.
Since the graph only assumes a paved, dry runway, we'll have to make correction to the landing distance.
You may recall that wet grass increases the distance by 30% or a factor of 1.3.
Our calculated landing distance was 1450 feet, but correcting it for wet grass would now make the landing distance 1885
feet.
Let's now assume the runway has a downslope of 1%. You may recall that this would increase the landing distance
by 5% or a factor of 1.05.
This now makes our landing distance 1979 feet.
Finally, you may recall that there was an additional safety factor of 1.43, which the Aviation Authority highly recommends
you apply.
This now makes the landing distance required 2830 feet.
You must ensure that the landing distance available at the airfield is more than the landing distance required.
We've seen only one way in which landing data may be presented, namely in a graph.
There are many others, some even in tabulated format.
Familiarise yourself with the format for the aircraft you are flying, and correctly adjust the calculated landing distance for any
variable that is not accounted for in the data.
Do not forget to add the safety factor once you've applied all variables.
An understanding of the effects of various conditions and factors that control the landing will better equip you to ensure
that all your landings are well within the limits of the runway available,
and thus help you avoid any accidents or mishaps on landing.