Related titles on the JAR syllabus
Ground Studies for Pilots series
Radio Aids
R.B. Underdown and David Cockburn
Sixth Edition
R.B. Underdown and Tony Palmer
Sixth Edition
Flight Planning
P.J. Swatton
Sixth Edition
Flight Instruments and Automatic Flight Control Systems
David Harris
Sixth Edition
R.B. Underdown and John Standen
Aviation Law for Pilots
R.B. Underdown and Tony Palmer
Tenth Edition
Human Performance and Limitations in Aviation
R.D. Campbell and M. Bagshaw
Third Edition
Sixth Edition
David Harris
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List of Abbreviations
Air Data Instruments
Gyroscopic Instruments and Compasses
Inertial Navigation Systems
Electronic Instrumentation
Automatic Flight Control
In-Flight Protection Systems
Powerplant and System Monitoring Instruments
Answers to Sample Questions
Since the publication of Roy Underdown's Fifth Edition of Volume 3 of
Ground Studies for Pilots, the format of the JAR-FCL examinations has become
established. Among other things, it has become clear that students need the
materials assembled in such a way as to facilitate study for individual papers
and that it is no longer practical to combine Instruments with Navigation
General, for example. The scope of what was once called the Instruments
paper (now paper number 022 in the ATPL syllabus) is now so broad that it
justifies a volume in its own right. Whilst still covering the air data and
gyroscopic instruments, inertial navigation systems and electronic navigation instruments, it has now been extended to cover the additional syllabus
requirements of engine and systems monitoring instruments and flight
warning systems.
As more and more students find the need to study subjects in the
groupings as they appear in the examinations, it makes sense to group
those subjects accordingly in the Ground Studies for Pilots series of
books. This new Sixth Edition addresses all the subjects listed in the
JAR Learning Objectives for the Instruments and Automatic Flight
paper number 022. I have tried to make the book both readable and
instructive, with the intention that it should be useful to those seeking
general information as well as to the examination student. Whilst it is
aimed principally at pilots studying for the JAR ATPL ground examinations, it will also be helpful to pilots at all professional and private
Just a few years ago, automatic flight systems and electronic instrument
systems were almost exclusively the preserve of large passenger transport
aircraft. In more recent times, however, it has become increasingly common
for smaller short-haul and executive-type jet and turbo-prop aircraft to be
equipped with such systems. Consequently, there is a need for those pilots
intending to progress from light and general aviation into commercial flying
to have knowledge of at least the basic principles of these systems and
instruments, even though they may not be immediately intending to sit the
professional examinations. The text and diagrams in this volume have been
deliberately designed to be understandable without preknowledge of the
The assistance of Roger Henshaw, Peter Swatton and David Webb, all of
Ground Training Services Ltd at Bournemouth International Airport, is
gratefully acknowledged for their advice and for scrutinising the manuscript
of this book for errors. Their professional knowledge of the JAR-FCL
examination requirements has helped to ensure that it will be extremely
useful to pilots undertaking ground studies.
David Harris
List of Abbreviations
alternating current
aircraft communications addressing and reporting system
airborne collision avoidance system
air data computer
attitude director indicator
above ground level
aircraft integrated data system
above mean sea level
Aircraft Operating Manual
airspeed indicator
airspeed indicator reading
built-in test equipment
central air data computer
calibrated airspeed
control and display unit
compressibility error correction
centre of gravity
cylinder head temperature
corrected outside air temperature
cathode ray tube
constant speed drive unit
control wheel steering
direct current
density error correction
directional gyro
decision height
electronic attitude and direction indicator
equivalent airspeed
Engine Centralised Aircraft Monitoring
Electronic Flight Instrument System
exhaust gas temperature
electronic horizontal situation indicator
Engine Indicating and Crew Alerting System
engine pressure ratio
x List of Abbreviations
in Hg
Full Flight Regime Autothrottle System
Flight Levels
flight management computer
flight management system
fibre optic gyro
Ground Proximity Warning System
horizontal situation indicator
indicated airspeed
International Civil Aviation Organisation
instrument error
instrument error correction
inches of mercury
Inertial Navigation System
Inertial Reference System
International Standard Atmosphere
instantaneous vertical speed indicator
Joint Aviation Authority
Joint Aviation Regulations
degrees kelvin
liquid crystal diode
light emitting diode
lateral navigation
local speed of sound
linear voltage displacement transmitter
mach number
manifold air pressure
multi-purpose control and display unit
critical mach number
mean sea level
nautical miles
outside air temperature
pitot pressure
position error
pressure error correction
dynamic pressure
height above a chosen ground datum
height above mean sea level
resolution advisory
rectified airspeed
List of Abbreviations
rate of climb/descent indicator
ring laser gyro
radio magnetic indicator
revolutions per minute
static pressure
static air temperature
secondary surveillance radar
traffic advisory
true airspeed
total air temperature
Terrain Avoidance Warning System
Traffic Collision Avoidance System
touch control steering
vertical navigation
VHF omnidirectional ranging
vertical speed indicator
Chapter 1
Air Data Instruments
Two of the most important pieces of information for a safe flight are height
and speed. Almost from the beginning of powered flight these have been
provided to the pilot by instruments that utilise the ambient atmospheric
pressure by means of a pitot/static system.
Pitot and static systems
Static pressure
The ambient atmospheric pressure at any location is known as the static
pressure. This pressure, in a standard atmosphere, decreases by 1 hectopascal (hPa) for each 27 feet (ft) increase in altitude at mean sea level. For
simplicity this figure is usually approximated to 1 hPa per 30 ft gain in
altitude. The rate of change of pressure with height is fundamental to the
operation of the pressure altimeter, the vertical speed indicator and the mach
meter. Each of these instruments uses static pressure to measure aircraft
altitude, or rate of change of altitude.
Static pressure, that is the pressure of the stationary air surrounding an
aircraft, irrespective of its height or speed, is sensed through a set of small
holes situated at a point on the aircraft unaffected by turbulence. This sensing point is known as the static source. It is typically on the side of the
fuselage or on the side of a tube projecting into the airstream.
Pitot pressure
As an aircraft moves through the air it displaces the surrounding air. As it
moves forward it compresses the air and there is a pressure increase on the
forward-facing parts of the aircraft. This pressure is known as dynamic
Suppose a cup were to be placed on the front of an aircraft, with its open
end facing forward. When the aircraft is stationary the pressure inside the
cup will be the same as the surrounding air pressure. In other words it will
be static pressure. When the aircraft begins to move forward the air inside
the cup will be compressed and dynamic pressure will be added to the static
2 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
pressure. The faster the aircraft moves, the greater the dynamic pressure will
become, but static pressure will always also be present.
The pressure measured on the forward-facing surfaces of an aircraft will
be the sum of static pressure and dynamic pressure. This is known as pitot
pressure, or total pressure, which is sensed by a forward-facing, open-ended
tube called a pitot tube, or pitot head. Figure 1.1 is a simplified diagram of a
pitot head.
Figure 1.1 Pitot head.
The pitot head comprises an aerodynamically shaped casing, usually
mounted beneath one wing, or on the side of the forward fuselage, clear of
any turbulent airflow. Within the casing is a tube, the rear of which is connected to the pitot system, which conveys pitot pressure to the pilot's
instruments. An electrical heating element is fitted within the tube to prevent the formation of ice, which could otherwise block the tube and render it
useless. Drain holes are provided in the bottom of the tube to allow water to
The dynamic element of pitot pressure is required to operate those air data
instruments that display speed relative to the surrounding air, the airspeed
indicator and the mach meter.
Pitot and static pressure is supplied to the air data instruments through a
system of tubes known as the pitot/static system. A schematic layout of the
pitot/static system for a light aircraft is shown in Figure 1.2. The static source
is duplicated on either side of the rear fuselage. This is to compensate for
false readings that would occur if the aircraft were side-slipping or in a
crosswind. The pitot and static pressure supplies are connected to a duplicate set of instruments in aircraft that have a pilot and co-pilot. In many
aircraft the pitot and static sources are combined in the pitot head, as
illustrated in Figure 1.3.
The static source consists of a number of small holes in the side of the pitot
head, connected to an annular chamber surrounding the pitot tube. This
chamber is connected to the static system, which conveys static pressure to
the pilot's instruments. A separate pipe connects the pitot tube to the pitot
Air Data Instruments
pitot system
static system
Figure 1.2
Pitot/static system for a light aircraft.
static source
static pressure
pitot pressure
drain holes
heater element
direction of flight
Figure 1.3
Pitot/static head.
system. As with the pitot head shown in Figure 1.1, an electrical heating
element is fitted to prevent blockage of the pitot and static sources due to
icing and water drain holes are provided in the bottom of the casing. In some
aircraft this type of pitot head is mounted on the fuselage near the nose and it
may be duplicated, one each side, to compensate for crosswind effects.
Static pressure error
The static system of air pressure measurement will be incorrect if the airflow
is turbulent, if there is a crosswind, or if the aircraft is side-slipping. The
effect of turbulence is minimised by locating the static source clear of protuberances and disturbed airflow. To eliminate the effect of crosswind or
side-slip the static source is duplicated and this is known as static balancing.
Despite all the measures taken by the aircraft designer there is often some
small error in the sensed static pressure, but this can usually be measured
and compensated for by a correction card or table. In many aircraft the
correction values will differ according to the position of flaps and/or
landing gear.
4 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
Correct and reliable indications from the various air data instruments can
only be achieved if the pitot and static sources are kept clear of any
blockages and the pitot/static system within the aircraft remains undamaged and pressure-tight.
Blockage of the pitot or static source may occur due to icing, insects, dirt or
dust. It is also not unknown for aircraft painters to forget to remove
masking tape from the perforated discs that form the static vents on the
fuselage sides. Icing can be prevented by the use of heaters, but this may
affect the sensed pressure to some small extent. The effect of blockages is
to render the instruments dangerously inaccurate or useless.
Blockage of the static sources will cause the altimeter reading to remain
constant regardless of changing aircraft altitude, the vertical speed indicator will not indicate rate of change of height and the airspeed indicator
will be dangerously inaccurate.
Blockage of the pitot source will not affect the altimeter or vertical speed
indicator, but it will render the airspeed indicator useless and the mach
meter grossly inaccurate.
Leakage in the piping of the pitot/static system will also seriously affect
the accuracy and usefulness of the air data instruments. Loss of pitot
pressure due to leakage in the pitot pressure system will cause the airspeed indicator to underread.
Leakage in the static pressure system within the cabin of a pressurised
aircraft is a serious problem, since the altimeter will register an altitude
equivalent to cabin altitude, which will almost certainly be much lower
than aircraft altitude. The vertical speed indicator will not function at all
and the airspeed indicator will be inaccurate.
In unpressurised aircraft the effect of a static system leak is less serious,
since internal pressure is much the same as external. However, it may
change at a slightly slower rate when the aircraft is climbing or descending and this would clearly affect the accuracy of the pressure
instruments during height changes.
Alternate static source
Blockage of the static source is a more probable hazard in flight and for this
reason many aircraft are fitted with an alternate static source. This may take
pressure from within the cabin in the case of unpressurised aircraft, or from
a separate external source. In either case there is likely to be a slight difference in pressure compared with that from the normal source. The aircraft
flight manual usually contains correction values to be used when the alter-
Air Data Instruments
nate static source has been selected. Changeover is made by means of a
selector cock easily accessible to the pilot.
In light aircraft, not fitted with an alternate static source, if the static source
is blocked an alternative source can be obtained by breaking the glass of the
vertical speed indicator (VSI).
The pressure altimeter
The function of the pressure altimeter is to indicate the aircraft height above
a given pressure datum. It operates on the principle of decreasing atmospheric pressure with increasing height and is, in fact, simply an aneroid
barometer that is calibrated to read pressure in terms of height. To do this,
the manufacturer assumes that air pressure changes at a given rate with
change of height. The International Standard Atmosphere (ISA) values are
the data used for this assumption.
In the ISA the temperature at mean sea level is +158C and the air pressure
is 1013.25 hectopascals (hPa). The temperature lapse rate (the rate at which
the temperature will decrease with increase of height) is 1.988C per 1000 ft
(6.58C per kilometre) up to a height of 36 090 ft. Above that height the temperature is assumed to remain constant at 756.58C up to a height of 65 600 ft.
Air is a fluid and it has mass, and therefore density. If we consider a
column of air, its mass exerts pressure at the base of the column; the taller the
column the greater the pressure exerted at the base. At any given height in
the column the pressure exerted is proportional to the mass of air above that
point and is known as hydrostatic pressure. Atmospheric pressure is
assumed to decrease at a rate of 1 hPa per 27 ft gain in height at sea level, this
rate decreases as height increases, so that at a height of say 5500 metres
(18 000 ft) the same 1 hPa change is equivalent to approximately 15 metres
(50 ft) change in height.
The pressure altimeter is calibrated to read height above a selected pressure datum for any specific atmospheric pressure.
The element of the pressure altimeter that measures atmospheric pressure
changes is a sealed capsule made from thin metal sheet. The capsule is
partially evacuated, so that the surrounding atmospheric pressure tends to
compress the capsule. However, a leaf spring attached to the capsule prevents this. The capsule may be of the diaphragm or the bellows type, as
illustrated in Figure 1.4.
The simple altimeter
The capsule is mounted in a sealed casing, connected to the static source.
Increased static pressure will cause the capsule to be compressed against the
restraining force of the leaf spring, decreased static pressure will allow the
6 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
bellows type
Figure 1.4 Aneroid capsule types.
leaf spring to expand the capsule. The spring force ensures that the extent of
compression or expansion is proportional to the static pressure being measured. This compression or expansion of the capsule is converted into rotary
motion of a pointer against a calibrated scale by a system of linkages and
gears. A schematic diagram of a simple pressure altimeter is shown in Figure
1.5. Expansion of the aneroid capsule will cause a lever to pivot about its
attachment to the instrument casing. This lever is connected to a drum by
means of a chain and its pivoting motion causes the drum to rotate. The
drum is attached to a pointer, which will consequently rotate against a
calibrated card scale.
Figure 1.5 Simple pressure altimeter.
Air Data Instruments
The setting of the height-indicating pointer can be adjusted by means of
the pointer setting knob. The purpose of this is to allow the pilot to set the
altimeter so that it displays height above a chosen datum. For example, if the
pointer is set to read zero with the aircraft on the airfield the altimeter will
thereafter indicate height above the airfield. Alternatively if, when the aircraft is on the airfield, the pointer is set to read height above mean sea level it
will thereafter show height above mean sea level. In both these examples the
surface pressure is assumed to remain constant.
The pressure altimeter is calibrated to read height in feet above the ISA
mean sea level pressure of 1013.25 hPa. Some altimeters of US manufacture
use inches of mercury (in Hg) as the unit of calibration and the equivalent of
1013.25 hPa is 29.92 in Hg. To convert inches Hg to hPa it is necessary to
multiply by 33.8639. Conversely, to convert hPa to inches Hg it is necessary
to multiply by 0.02953.
Some simple altimeters are calibrated on the assumption that static
pressure reduces with increasing height at a constant rate of 1 hPa per 30 ft.
Whilst this is a reasonable approximation up to about 10 000 ft, the height
increase represented by 1 hPa decrease in pressure becomes progressively
greater above this altitude. This is indicated in Table 1.1.
Table 1.1
Rate of pressure change with altitude.
Height in ISA (ft)
Mean sea level
Static pressure (hPa)
Height change (ft)
represented by 1 hPa
change in static pressure
10 000
20 000
40 000
60 000
In order to have a linear scale on the altimeter display it is necessary to
compensate for this non-linear distribution of atmospheric pressure. This is
achieved during manufacture by careful design of the capsules and the
mechanical linkages.
The simple altimeter is rarely used in aircraft because it lacks sensitivity,
furthermore the altitude scale is compressed and causes confusion. More
accurate height measurement is achieved with the sensitive altimeter.
The sensitive altimeter
The principle of operation of the sensitive altimeter is essentially the same as
8 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
that of the simple altimeter. Sensitivity of response to static pressure changes
is improved by incorporating two or three aneroid capsules connected in
series with each other. In a typical instrument the movement of the capsules
as altitude changes is transmitted to gearing via a rocking shaft. The gearing
in turn operates the height-indicating pointers, of which there are two or
Clearly, as an aircraft climbs the air temperature will fall. The altimeter is
calibrated to be accurate when operating at a sea level air temperature of
+158C (ISA mean sea level temperature) and without compensation it will
become increasingly inaccurate as instrument temperature deviates from
that datum value. In the sensitive altimeter, temperature compensation is
effected by a bimetallic strip inserted between the aneroid capsules and the
transmission shaft. Note that this temperature compensation has nothing to
do with altitude temperature error, which is discussed in sub-paragraph (6)
under Altimeter errors below.
Figure 1.6 illustrates the display of a sensitive altimeter with three
It is perhaps easiest to think of the display as being similar to a clock face,
in which one revolution of the long hand takes one hour and one revolution
of the shorter hand takes 12 hours. In the altimeter display one revolution of
the longest pointer represents a height change of 1000 ft, thus each numbered division on the instrument scale represents a height increment of 100 ft
when read against this pointer.
One revolution of the medium length pointer represents a height change
of 10 000 ft, thus each numbered division on the instrument scale represents
a height increment of 1000 ft when read against this pointer. The height scale
is divided into ten equal parts, so for every complete revolution of the 1000-ft
pointer the 10 000-ft pointer moves one-tenth of a revolution.
Figure 1.6 Sensitive altimeter display.
Air Data Instruments
The shortest pointer shows increments of 10 000 ft for each numbered
division on the instrument scale. The sensitive altimeter display in Figure 1.6
is therefore indicating a pressure altitude of 12 500 ft.
Whilst the sensitive altimeter is usable to greater altitudes than the simple
altimeter it becomes inaccurate at high altitudes where the pressure change
becomes small for a given increase in height. Its greatest disadvantage,
however, is its multi-pointer display, which is very open to misinterpretation.
Subscale settings
Sensitive altimeters incorporate a pressure datum adjustment, so that height
above any desired datum pressure will be indicated. The selected datum
pressure is indicated on a subscale calibrated in hPa or in Hg, and this is
usually referred to as the subscale setting.
If airfield datum pressure is set on the subscale the altimeter will indicate
height above the airfield once the aircraft is airborne and zero when on the
ground at that airfield. This setting is assigned the ICAO code of QFE for
communication purposes.
If mean sea level pressure is set on the subscale the altimeter will indicate
height above mean sea level. Thus, when the aircraft is on the airfield the
altimeter will indicate airfield elevation above mean sea level. This setting is
assigned the ICAO code of QNH for communication purposes.
Alternatively, the subscale may be set to 1013 hPa (29.92 in Hg), in which
case the altimeter will indicate pressure altitude, that is the altitude above
this pressure datum. This setting is used for aircraft flying Flight Levels (FL).
At this point it is appropriate to define the terms height and altitude with
reference to altimeter subscale settings.
Height is the vertical distance above a specified datum with known elevation, such as an airfield. Hence, if QFE is set on an altimeter it will indicate
height above that airfield.
Altitude is the vertical distance above mean sea level and it is therefore
altitude that is indicated by an altimeter with QNH set. Pressure altitude, as
we have already seen, is the altitude indicated on an altimeter with 1013 hPa,
or 29.92 in Hg, set on its subscale.
10 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
Density altitude is the height in ISA conditions at which a given air density
will occur and it is temperature dependent. If the local air temperature is ISA
temperature, then density altitude will be the same as pressure altitude. If
local air temperature is higher than ISA, then density altitude will be higher
than pressure altitude and vice versa.
True altitude is the exact vertical distance above any point on the surface.
A pressure altimeter cannot be relied upon to show true altitude, even with
QFE set, since its calibration assumes a rate of pressure and temperature
change with height that may not exist in the prevailing atmospheric
The servo-assisted altimeter
Whereas in the simple and sensitive altimeter the movement of the aneroid
capsule directly moves the altimeter pointer, in the servo-assisted altimeter
the capsule movement is transmitted to the indicator by an electromechanical system. When the capsules expand or contract their movement
causes an electro-magnetic transducer to transmit an electrical signal to a
motor, which drives the altimeter's single pointer and its height counters.
The servo-assisted altimeter has several advantages. Since the only
mechanical transmission is from the capsules to the transducer there is
considerably less resistance to motion and therefore less lag (see Lag error
below). It is much more sensitive to small pressure changes and therefore
remains accurate to greater altitude, where the pressure change is less for a
given change of height. Because its output is electrical it can readily be used
in conjunction with remote displays, altitude alerting and height encoding
systems. The servo-assisted altimeter display with its height counters, as
shown in Figure 1.7, is less likely to be misinterpreted than that of the sensitive altimeter.
12 5
Figure 1.7 Servo-assisted altimeter display.
pointer completes
1 revolution
per 1000 feet
Air Data Instruments
Altimeter errors
Pressure altimeters are prone to a number of errors, as listed below.
Lag error. There is inevitably a delay between an atmospheric pressure
change and the response of the aneroid capsules, with the result that the
movement of the instrument pointer lags behind a change in altitude.
The magnitude of the error depends upon the rate of change of altitude
and is clearly unacceptable. In sensitive altimeters it is often reduced by
means of a vibrator, or knocking, mechanism, which has the same effect
as tapping a barometer to make the pointer move after a small pressure
change. In servo-assisted altimeters lag error is virtually eliminated by
the reduction of mechanical resistance.
Blockage of static source. If the static source becomes blocked the
altimeter will cease to indicate changes of static pressure, and therefore
of altitude. The effect will be that the altimeter will continue to indicate
the reading at which the blockage occurred.
Instrument error (IE). As with any mechanical device, the pressure
altimeter is manufactured with small tolerances in its moving parts and
these give rise to small inaccuracies in its performance. They are usually
insignificant, but in some cases a correction table may be supplied with
the instrument.
Position error (PE). The static source is positioned at a point on the
airframe where disturbance to the airflow is minimal, so that the static
pressure measured is as close as possible to the undisturbed ambient
static pressure. However, there is usually some small error due to the
positioning of the static source. Use of the alternate static source may
also cause pressure error, since its siting is different to that of the
normal source and thus subject to different effects during manoeuvring. The PE for an aircraft type is determined during its initial flight
tests and is supplied in tabular or graphical format so that the pilot may
make the appropriate corrections. In servo-assisted altimeters the correction is usually incorporated into the transducer system.
Corrections to be applied to the altimeter reading for position error
may be listed in the Aircraft Operating Manual (AOM). These are
usually small and take account of the effects of aircraft speed, weight,
attitude and configuration, since all of these factors have some effect on
airflow over the static source. Corrections are also usually given for
both the normal and alternate static source, since position error may not
be the same for both. The correction is applied as indicated in the AOM
tables. For example, suppose the tables state that the correction to be
applied with flaps at the landing setting of 458 is +25 ft, then 25 ft must
be added to the altimeter reading.
12 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
Pressure error. Also known as barometric error and subscale-setting
error, this occurs when the actual datum pressure differs from that
selected on the subscale setting of the altimeter. Suppose, for example,
the subscale has been set for a regional QNH of 1020 hPa, but the aircraft is now operating in an area where the mean sea level (msl) pressure is 1000 hPa. Let us assume that the aircraft is at an altitude where
the ambient pressure is 920 hPa, so the altimeter will indicate 3000 feet:
(1020 7 920 = 100 hPa 6 30 ft/hPa = 3000 ft).
However, because of the lower msl pressure the actual height of the
aircraft above msl is:
1000 7 920 = 80 hPa 6 30 ft/hPa = 2400 feet.
Temperature error. The altimeter is calibrated to assume a Standard
Atmosphere temperature lapse rate of 1.988C per 1000 ft. Actual temperatures at any given altitude usually differ from this assumption and
the altimeter will be in error. In cold air the density is greater than in
warm air and a given pressure will occur at a lower altitude than in
warmer air. Consequently, the altimeter will overread and, with QNH
set, the aircraft will be lower than indicated, with the error increasing
from zero at msl to a significant amount at altitude. The AOM contains
a table of corrections for this situation. These corrections must be added
to published or calculated heights/altitudes when temperature is less
than ISA.
A `rule of thumb' calculation of altitude temperature error is that the
error will be approximately 4% of indicated altitude for every 108C
temperature deviation from ISA.
Altimeter tolerances
For altimeters with a test range of 0 to 9000 metres (0 to 30 000 ft) the required
tolerance is + 20 metres or + 60 ft. For altimeters with a test range of 0 to
15 000 metres (0 to 50 000 ft) the required tolerance is + 25 metres or + 80 ft.
The test is required to be carried out by the flight crew prior to flight, with
QNH or QFE set and the altimeter vibrated either manually or by the
mechanical vibration mechanism.
The airspeed indicator (ASI)
It is essential for the pilot of an aircraft to know its airspeed, because many
critical factors depend upon the speed of flight. For example, the pilot needs
Air Data Instruments
to know when the aircraft is moving fast enough for take-off, when it is
flying close to the stalling speed, when it has accelerated to the speed at
which landing gear and flaps must be raised and when it is approaching the
maximum safe flying speed, to name but a few. This critical information is
provided by the airspeed indicator (ASI).
The aircraft's speed relative to the surrounding air is proportional to the
dynamic pressure that results from the air being brought to rest on forward
facing parts of the airframe.
The airspeed indicator measures dynamic pressure and converts this to an
indication of airspeed. We have already seen that pitot pressure (P), as
measured in the pitot tube, is a combination of dynamic pressure (D) and
static pressure (S), i.e. P = D + S. It therefore follows that dynamic pressure is
pitot pressure less static pressure, i.e. D = P 7 S. Thus, the function of the
ASI is to remove the static pressure element of pitot pressure and use the
resulting dynamic pressure to move a pointer around a graduated scale.
The instrument comprises a sealed case connected to the static source and
containing a capsule supplied with pitot pressure. Hence, the static pressure
element of pitot pressure inside the capsule is balanced by static pressure
surrounding the capsule. Consequently, the capsule will respond only to
changes in the dynamic pressure element of pitot pressure. The faster the
aircraft flies through the atmosphere, the greater will be the resultant
dynamic pressure and the capsule will expand. This expansion is transmitted to a pointer by means of gearing and linkages. The principle is
illustrated schematically in Figure 1.8.
A simple ASI typically uses a single pointer that moves around a scale
calibrated in knots. More complex instruments may be used in high-speed
aircraft, incorporating an angle-of-attack indicator or a mach meter.
Figure 1.8
Principle of operation of the airspeed indicator.
14 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
Clearly the airspeed indicator must be calibrated to some fixed datum, for
the pilot must be confident that an indication of, say, 90 knots is aerodynamically the same in all conditions. The datum used is the air density
when ISA mean sea level conditions prevail, that is the density when the air
temperature is +158C and the air pressure is 1013 hPa. Only when these
conditions prevail will the ASI indicate the true airspeed.
Any point where air is brought to rest, such as in the pitot tube, is known
as a stagnation point. At this point the kinetic energy of the air is converted
to pressure energy. The pressure resulting is, of course, dynamic pressure
and is denoted by the symbol Q. It can be shown mathematically that Q =
2rV , where r is the air density and V is the airspeed.
As the aircraft climbs the air density decreases and the airspeed indicator
reading will be lower than the true speed of the aircraft through the surrounding atmosphere. A simplistic way of thinking of this is to consider that
as the aircraft moves through the air it is colliding with the air molecules; the
faster it flies the more molecules it strikes in a given time period.
The ASI indication of airspeed is based on the dynamic pressure
measured, assuming an ISA mean sea level value of air density. Therefore,
when the air density is lower, as with an increase of altitude, a given
dynamic pressure (and therefore indicated airspeed) will only be achieved at
a higher true airspeed.
Let us assume that it is flying in ISA msl conditions and is colliding with
air molecules at a rate that causes the ASI to indicate 90 knots flight speed.
When the aircraft climbs to a greater altitude the air density is less and so the
molecules are further apart. In order for the ASI to continue to read 90 knots
the aircraft must fly faster relative to the surrounding air in order to strike
the same number of molecules within a given time period. Thus the true air
speed, which is the speed of the aircraft relative to the surrounding atmosphere, increases.
Airspeed may be quoted in a number of ways and these are listed below:
ASIR: Airspeed indicator reading.
IAS: Indicated airspeed (IAS) is the speed indicated by the simple airspeed
indicator reading (ASIR) corrected for errors due to manufacturing tolerances in the instrument (instrument error), but not corrected for static
pressure errors occurring at the static source (pressure error).
CAS: Calibrated airspeed (CAS) is the airspeed obtained when the corrections for instrument error (IEC) and pressure error (PEC) have been applied
to IAS. These correction values are usually found in the Aircraft Operating
Manual and may be reproduced on a reference card kept in the cockpit. CAS
Calibrated airspeed used to be referred to as rectified airspeed (RAS).
Air Data Instruments
EAS: Because air is a compressible fluid it tends to become compressed as it
is brought to rest. At low to medium airspeeds the effect of compression is
negligible, but above about 300 knots TAS the compression in the pitot tube
is sufficient to cause the airspeed indicator to overread significantly. The
greater the airspeed above this threshold, the greater the error due to compression. The effect is also increasingly noticeable at high altitude, where the
lower density air is more easily compressed. The error produced by this
effect is known as compressibility error.
Compressibility error correction (CEC) can be calculated and when
applied to the calibrated airspeed the result is known as equivalent airspeed
TAS: The airspeed indicator only indicates true airspeed (TAS) when ISA
mean sea level conditions prevail; any change of air density from those
conditions will cause the indicated airspeed to differ from true airspeed. The
greater the altitude, the lower will be air density and therefore IAS (and
consequently EAS) will be progressively lower than TAS.
The error due to the difference in density from ISA msl density can be
calculated. When this density error correction (DEC) is applied to EAS the
result is the aircraft's true airspeed (TAS). TAS = EAS + DEC. The compressibility error and density error corrections are embodied in the circular
slide rule (navigation computer), from which TAS can be found using CAS
and the appropriate altitude, speed and temperature settings.
Square law compensation
Given that dynamic pressure increases as the square of airspeed it follows
that expansion of the capsule in the ASI must do likewise. Thus, at low
airspeeds the amount of expansion for a given increase in speed will be
small, whereas at higher airspeeds the amount of expansion will be relatively large for the same speed increase. This is known as square law
expansion and, transmitted to the instrument pointer, will require an
expanding scale on the face of the ASI, as illustrated at Figure 1.9. Such a
scale makes accurate interpretation of airspeed difficult at low speeds and
limits the upper speed range that can be displayed.
Most ASIs incorporate an internal compensation device that permits a
linear scale on the dial, i.e. equal spacing of the speed divisions over the
whole range, as depicted in Figure 1.10.
The effects of temperature variations on the sensitivity of the capsules and
their associated linkages are typically compensated by the inclusion of
bimetallic strips in the lever system. The expansion or contraction of these
strips varies the degree of lever movement in the system of linkages.
16 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
Figure 1.9 Square law calibration of ASI dial.
Figure 1.10 Typical ASI presentation.
The ASI scale is often marked with coloured arcs and radial lines. The arcs
indicate operating speed ranges and the radial lines indicate limiting speeds.
These speed ranges and limiting speeds are as follows:
Vne This is the `never exceed' speed, beyond which structural damage may
occur. It is indicated on the ASI presentation by a red radial line.
Vno This is the maximum speed under normal operating conditions with
the aircraft `clean' (i.e. flaps and landing gear retracted).
Vs1 This is the stalling speed with the aircraft `clean'. A green arc extends
around the ASI scale from Vs1 to Vno.
This is the stalling speed with flaps and landing gear fully extended.
Vfe This is the maximum permitted speed with flaps extended. A white arc
extends around the ASI scale from Vso to Vfe.
Air Data Instruments
Vmc This is the minimum control speed with one engine inoperative on a
multi-engine aircraft. It is indicated by a red radial line on the ASI scale.
Vyse This is the best rate-of-climb speed with one engine inoperative on a
multi-engine aircraft. It is indicated by a blue radial line on the ASI scale.
Vmo/Mmo Some ASIs incorporate a red and white striped pointer showing
the maximum allowable airspeed (Vmo) which is often the mach-limiting
airspeed, at which the airflow over parts of the airframe will be approaching
the critical mach number (Mcrit) for the aircraft. The pointer is actuated by a
static pressure capsule and specially calibrated mechanism. This allows the
Vmo pointer reading to increase progressively up to about 25 000 ft, as the air
becomes less dense. Above this altitude the Vmo pointer reading is progressively decreased, since Mcrit will be reached at progressively lower
values of indicated airspeed.
Vlo This is the maximum speed at which the landing gear may be safely
extended or retracted.
Vle This is the maximum speed at which the aircraft may be flown with
landing gear extended.
Note: Mach number and critical mach number will be explained in detail in
the section dealing with the mach meter.
ASI errors
Instrument error. This is effectively the same as for the pressure altimeter. It is the error between the airspeed that the ASI should indicate
and that which it actually does, due to manufacturing tolerances and
friction within the instrument.
Position or pressure error. This is the error caused by pressure fluctuations at the static source. These may be due to the position of the
static source, hence the term position error. The error can be determined
over the speed range of the aircraft and recorded on a correction card.
Pressure error may also occur when the aircraft is at an unusual attitude
or when flaps or landing gear are extended, in which case it is known as
manoeuvring error. An example of ASI values with a normal static
source in use, and the corrections to be applied when the alternate
source is in use, is shown in Table 1.2 below. These data are normally
found in the AOM.
Compressibility error. At true airspeeds above about 300 knots the air
brought to rest in the pitot tube is compressed to a pressure greater than
dynamic pressure, causing the ASI to overread. The effect increases
with altitude, since less dense air is more readily compressed. As pre-
18 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
Table 1.2 Airspeed calibration card.
Indicated airspeed (knots)
Flaps up
50 60
51 61
90 100
91 101
Flaps 108
50 60
51 61
90 100
90 100
Flaps 408
50 60
50 60
viously stated, the compressibility error correction can be found using
the circular slide rule (navigation computer).
Blocked pitot tube. A blocked pitot tube will mean that the pressure in
the pitot system will be unaffected by changes in airspeed. In level
flight the ASI will indicate a false airspeed, dependent upon the pressure locked in the system, and will not indicate changes in airspeed. In
the climb the decreasing static pressure will cause the ASI capsule to
expand and the instrument will falsely indicate an increasing airspeed.
In the descent the reverse will be the case.
Blocked static source. A blocked static source will cause the ASI to
underread as the aircraft climbs above the altitude at which the
blockage occurred, since the pressure trapped in the instrument case
will be progressively greater than ambient static pressure. In the descent the reverse will happen and a hazardous condition arises, since the
progressively overreading ASI may cause the pilot to reduce power and
airspeed may fall below stalling speed. A blockage to the static source
during level flight will not be readily apparent, since the ASI indication
will not change.
ASI tolerance
Typical required accuracy for the airspeed indicator is +3 knots.
The mach meter
The speed at which sound travels through the air is known as the speed of
sound, or sonic speed. This speed varies with air temperature and therefore
with location. The speed of sound at any specific location is known as the
local speed of sound (LSS).
Aircraft that are not designed to fly at supersonic speeds usually suf-
Air Data Instruments
fer both control and structural problems when the airflow over the airframe, particularly over the wings, approaches the LSS. Consequently it
is essential that pilots of such aircraft be aware of the aircraft's speed relative to the LSS. This is especially important at high altitude, since the
speed of sound decreases with temperature, and air temperature decreases with altitude in a normal atmosphere. Hence, the greater the altitude, the lower the LSS.
The aircraft's speed relative to the LSS is measured against a scale in
which the LSS is assigned a value of 1. The limiting speed above which
control problems may be encountered is known as the critical mach speed
and is assigned a value known as the critical mach number (Mcrit).
Supposing that, for a particular aircraft, this speed happens to be 70% of the
LSS, then the critical mach number would be 0.7.
The mach number (M) is the ratio of the aircraft's true airspeed (TAS) to
the LSS. This may be represented by the equation:
Thus, if an aircraft is flying at a TAS of 385 knots at an altitude where the LSS
is 550 knots, the aircraft's mach number is 385 7 550 = 0.7. Clearly, if the
aircraft were flying at a TAS of 550 knots in these conditions its mach
number would be 1.0 and it would be flying at sonic speed.
Since mach number is the ratio of TAS to LSS it follows that it is proportional to the ratio of EAS, CAS or RAS to LSS, but the requisite corrections
must be applied. Once again, the navigation computer, or circular slide rule,
is equipped to do this.
The speed of sound in air is entirely dependent upon the air temperature.
The lower the air temperature, the lower the LSS, therefore LSS decreases
with increasing altitude in a normal atmosphere. The LSS can be calculated
using the formula:
LSS ˆ 38:94 TK
where TK is the local air temperature in degrees kelvin.
The kelvin, or absolute temperature, scale is based upon absolute zero,
which is equal to 72738C. Thus, 08C is equal to 273K and an ISA mean sea
level temperature of +158C is equivalent to 273 + 15 = 288K. From this it can
be shown that the local speed of sound at ISA mean sea level is:
38:94 288 ˆ 661 knots
However, at 36 090 ft in the standard atmosphere, where the ambient air
temperature is 756.58C:
20 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
LSS ˆ 38:94 273 ÿ 56:5
ˆ 38:94 216:5
ˆ 573 knots
Since mach number is the ratio of TAS to LSS it follows that mach number is
also dependent upon local air temperature. At any given true airspeed the
aircraft's mach number varies inversely with temperature. Thus, an aircraft
climbing in standard atmospheric conditions at a true airspeed of 573 knots
would reach mach 1 at an altitude of 36 090 feet, whereas at sea level its mach
number would have been 573 7 661 = 0.87.
The mach meter is essentially a pressure altimeter and an ASI combined in
one instrument. Its purpose is to indicate the aircraft's airspeed relative to
the LSS.
We have seen that the LSS decreases with air temperature and, therefore,
with altitude in a normal atmosphere. The mach meter contains an altitude
capsule similar to that in the pressure altimeter. Static pressure is supplied to
the sealed instrument casing and the aneroid altitude capsule will expand
against a light spring as static pressure decreases with increasing altitude.
Figure 1.11 is a schematic diagram showing the principle of operation of
the mach meter. The instrument also contains an airspeed capsule, the
inside of which is connected to pitot pressure. As airspeed, and therefore
dynamic pressure, increases the airspeed capsule will expand, exactly as in
the ASI.
altitude capsule
ratio arm
ranging arm
airspeed capsule
Figure 1.11 Mach meter principle of operation.
The movement of the altitude and airspeed capsules is transmitted to the
mach meter pointer through a system of mechanical links and gears. The
pointer rotates against a dial calibrated to show the aircraft speed in terms of
mach number.
As a pressure instrument the mach meter cannot measure the ratio of TAS
to LSS, but it satisfies the requirement by measuring the ratio of dynamic
pressure (pitot 7 static) to static pressure. This can be expressed as:
Air Data Instruments
…p ÿ s†
From the foregoing it follows that an increase in airspeed will raise the
mach number, bringing the aircraft's speed closer to the LSS. An increase in
altitude will reduce the LSS, also bringing the aircraft's speed closer to the
LSS, raising the mach number. Let us examine the operation of the mach
meter under these circumstances.
Increased airspeed at constant altitude
Movement of the airspeed capsule is transmitted to the instrument pointer
through a ratio arm, a ranging arm and gearing. Let us assume there is an
increase of airspeed in level flight. The consequent increase in dynamic
pressure causes the airspeed capsule to expand and the ratio arm rotates to
bear against the ranging arm. This in turn rotates a quadrant and pinion gear
system, which is connected to the instrument pointer. Thus, the expansion of
the airspeed capsule causes the pointer to move against a calibrated scale,
indicating an increased mach number.
A decrease in airspeed will cause the airspeed capsule to contract and the
above sequence will be reversed, resulting in a decreased mach number
Increased altitude at constant airspeed
An increase in altitude will cause the altitude capsule to expand and the
linear movement of the ratio arm against the ranging arm will again rotate
the ranging arm and its associated quadrant and pinion to rotate the mach
meter pointer, indicating an increased mach number.
A decrease in altitude will cause the altitude capsule to contract and the
above sequence will be reversed, resulting in a decreased mach number
Figure 1.12 shows a typical mach meter display.
Mach/TAS calculations
Calculation of mach number given TAS and LSS has already been demonstrated. Clearly, by transposition of the formula, it is possible to calculate
TAS given mach number (M) and LSS.
LSS can be calculated if the ambient air temperature is known and converted
to kelvin.
22 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
mach number
Figure 1.12 Typical mach meter display.
Mach meter errors
Instrument error. In common with all pressure instruments the manufacturing tolerances inevitably lead to small errors of measurement
due to friction and lost motion in the linkages and gearing. These are
typically of the order of +0.1 M over a range of 0.5 M to 1.0 M.
Instrument error increases slightly with increasing speed and altitude.
Pressure error. The mach meter is more sensitive than the ASI to errors
in static pressure measurement, since it uses the ratio of dynamic
pressure (p7s) to static pressure, whereas the ASI uses the ratio of pitot
pressure to static pressure.
Blockages. Blockage of either pressure source will cause the measured
ratio to be incorrect. If the static source is blocked, changes in altitude
will not be sensed and the instrument will underread in the climb. If the
pitot source becomes blocked the instrument will not respond to speed
changes. The exact effect of a pitot blockage depends largely upon what
the aircraft is doing at the time and is therefore difficult to predict.
The vertical speed indicator (VSI)
The vertical speed of an aircraft is otherwise known as its rate of climb or
descent and the VSI is alternatively known as the rate of climb/descent
indicator (RCDI). The purpose of the VSI is to indicate to the pilot the aircraft's rate of climb or descent, typically in feet per minute.
Since it is rate of change of height being indicated it is necessary to create a
pressure difference within the instrument whilst a height change is occurring, and to arrange that the magnitude of the pressure difference is proportional to the rate of change of height. This is achieved by a metering unit
within the instrument, which is illustrated in schematic form in Figure 1.13.
Air Data Instruments
Figure 1.13
Vertical speed indicator principle of operation.
Static pressure is led directly to the inside of a capsule and also to the
inside of the sealed instrument casing via a metering orifice. The capsule is
connected to a pointer through linkages and a quadrant and pinion gear.
Whilst the aircraft is in level flight the pressure in the capsule is the same
as that in the casing and the pointer is in the horizontal, nine o'clock position
indicating zero. When the aircraft enters a climb, static pressure begins to fall
and this is sensed virtually immediately within the capsule. Pressure in the
instrument casing is now greater than that in the capsule, since air can only
escape at a controlled rate through the restricted orifice of the metering unit.
The pressure difference causes the capsule to contract, driving the pointer in
a clockwise direction to indicate a rate of climb. The faster the aircraft's rate
of climb, the greater will be the pressure difference between capsule and
casing and the greater the capsule compression, driving the pointer further
around the scale.
During a descent the increasing static pressure will be felt virtually
instantly within the capsule, but the pressure in the instrument casing will
rise at a slower rate due to the effect of the metering unit. Hence the capsule
will expand, against a spring, driving the pointer in an anti-clockwise
direction to indicate a rate of descent proportional to the pressure difference,
which is in turn proportional to the rate of descent.
When the aircraft levels out at a new altitude the pressure in the instrument casing will equalise with that in the capsule and the pointer will return
to zero. A typical VSI presentation is shown in Figure 1.14. It should be noted
that the scale graduation is logarithmic, having greater spacing at lower rates
of climb. This is deliberate, to facilitate easy interpretation when small
changes of height are being made.
24 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
Figure 1.14 VSI presentation.
Instantaneous vertical speed indicator (IVSI)
A disadvantage of the basic VSI described above is that there is an inherent
time lag in displaying climb or descent when either is initiated. This is due to
the hysteresis of the capsule, which needs a significant pressure differential
before it begins to expand or contract. In aircraft of moderate performance
this is unimportant, but in higher performance aircraft an instantaneous
response to height change is necessary.
This is achieved by introducing a small cylinder connected to the static
pressure supply to the capsule, containing a lightly spring-loaded free piston. This device is known as a dashpot accelerometer. When a climb is
initiated, inertia causes the piston to move down in the cylinder, creating an
instantaneous, but temporary, pressure drop in the capsule. The capsule
immediately contracts, causing the instrument to indicate immediately the
commencement of a climb. Thereafter, the decreasing static pressure due to
the climb causes an indicated rate of climb as previously described.
When a descent is initiated, inertia causes the dashpot piston to rise in its
cylinder, creating a small instantaneous pressure rise in the capsule, sufficient to expand it and indicate the commencement of descent.
A minor disadvantage of the IVSI is that, on entering a turn in level flight,
the centrifugal acceleration force is liable to displace the dashpot piston and
cause a temporary false indication of climb. Upon exiting the turn the
reverse will be the case. The IVSI is also more sensitive to turbulence and is
liable to give a fluctuating indication in such conditions. This is usually
damped out by the inclusion of a restriction in the static connection to the
capsule and the metering unit.
An adjusting screw is provided below the face of the instrument, by which
the pointer can be set to read zero. The range of adjustment available is,
typically, between +200 ft per minute and +400 ft per minute, depending
upon the scale range of the instrument.
Air Data Instruments
Instrument error. Manufacturing tolerances lead to small errors in the
internal mechanism. When these cause displacement of the pointer
from the zero position with the aircraft stationary on the ground it can
be corrected with an adjusting screw on the front of the instrument.
Pressure error. Disturbances at the static source may cause the instrument to display an incorrect rate of change of height.
Lag. A delay of a few seconds before a rate of change of height is
indicated is normal with the basic VSI. The IVSI virtually eliminates lag.
Transonic jump. A transonic shock wave passing over the static source
will cause the VSI briefly to give a false indication.
Blockage. Blockage of the static source will render the VSI useless, since
it will permanently give a zero indication.
Failure of the instrument will usually result in a fixed indication of zero rate
of change of height.
Dynamic vane-type VSI
Sailplanes and a few very light aircraft often use a different type of VSI
known as a variometer. Figure 1.15 shows a schematic illustration.
An enclosed casing shaped rather like a shallow tin can contains a pivoted
vane, held in a central position by a light leaf spring. The casing on one side
of the vane is connected to the static pressure source and on the other side of
the vane it is connected to an enclosed air reservoir. In level flight air can leak
past the vane sufficiently for the pressure in the reservoir to equalise with
from static source
to reservoir
Figure 1.15
26 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
static air pressure. Under these conditions the leaf spring will centralise the
vane, which is connected to a pointer indicating zero rate of climb/descent.
When a climb is initiated the static pressure will begin to fall below the
pressure stored in the air reservoir. The differential pressure across the vane
will deflect it and the attached pointer will indicate a rate of climb. The faster
the rate of climb, the greater the differential pressure and the more the vane/
pointer will be deflected. When the aircraft levels off, air will leak past the
vane to equalise the reservoir with static pressure once more and the leaf
spring will centralise the vane/pointer to the zero position. In a descent the
sequence of events will be the opposite of that described above.
The vane-type variometer is incapable of displaying large rates of height
change and is therefore unsuitable for most powered aircraft. It does,
however, have the advantage that it suffers less from time lag (i.e. initial
response) than the basic VSI.
There is also a type of variometer used specifically in sailplanes that
responds to height gain or loss not initiated by the pilot, as in a thermal. This
type is known as the total energy variometer.
The air data computer (ADC)
In addition to the four instruments already covered there are numerous
systems in modern aircraft that require air data inputs in terms of static
pressure, pitot pressure and air temperature. Since most, if not all, of these
systems are electronic in operation it is logical to supply such data in electronic form. The air data computer receives pitot and static pressure from the
normal and alternate sources and converts these into electrical signals and
transmits them to the various indicators and systems. As far as indicators are
concerned, this removes the need for bulky instruments that take up significant panel space and means that the information can be presented in
digital form if necessary.
The centralised air data computer can also be programmed to apply the
necessary corrections for pressure error, barometric pressure changes and
compressibility effects. With the addition of air temperature data inputs,
true airspeed can be calculated by the computer.
Analogue and digital ADCs
Air data computers are usually of the digital type; that is, they transmit data
in digital format which is compatible with other computer-based systems.
Analogue air data computers, which transmit their output data to servooperated devices, are less common, although a few are still in existence.
Figure 1.16 is a block diagram showing the data inputs and outputs of a
typical ADC. It will be seen that the data inputs are pitot and static pressure
Air Data Instruments
static pressure
pitot pressure
total air temperature
systems using adc outputs
flight director
automatic thrust control
automatic flight control cabin pressurisation
altitude reporting
flight management
flight recorder
stall warning
Figure 1.16
Data inputs and outputs of an air data computer.
and total air temperature (TAT). From these, electrical signals are generated
and transmitted as electronic data to operate the pilots' air data instrument
displays, plus TAS, TAT and SAT (static air temperature) displays. Additionally, these signals are transmitted to the various flight management
control systems, listed in Figure 1.16. Loss of air data input activates a
warning logic circuit within the ADC, which causes warning flags to appear
on the associated indicators and annunciators to illuminate on the computer
control panel.
At this point it is perhaps appropriate to define the various forms of air
temperature measurement.
28 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
SAT (static air temperature). This is the temperature the air at the
surface of the aircraft would be at if there were no compression effects
due to the aircraft's movement. At very low airspeeds these effects are
negligible, but for most transport aircraft normal flight speeds are such
that the direct measurement of SAT is virtually impossible. SAT is also
known as outside air temperature (OAT).
TAT (total air temperature). Total air temperature is the temperature of
the air when it has been brought completely to rest, as in the pitot tube.
The ram rise, that is the temperature increase due to compression, can
then be subtracted from TAT to give corrected outside air temperature
(COAT). The value of ram rise can be calculated for any given mach
number, so clearly the air data computer can be programmed to make
this correction.
Temperature measurement probes
Aircraft that operate at low airspeeds, such as some helicopters and light
aircraft, usually employ a simple bimetallic thermometer, which operates a
rotary pointer against a temperature scale to indicate what is essentially
static air temperature.
TAT sensors are more complex because they must, as far as possible,
recover the temperature rise due to adiabatic compression of the air as it is
brought to rest; what is known as the `ram rise'. To do this it is necessary to
use sophisticated probes to capture and slow the air and then convert the air
temperature to an electrical signal for transmission to the ADC.
The sensitivity of the probe in terms of the extent to which it truly
achieves, or recovers, the full ram rise effect is known as its recovery factor.
For example, a probe that senses SAT plus 85% of the ram rise in temperature
would be said to have a recovery factor of 0.85.
TAT probes are typically contained within an aerodynamically shaped
strut with an air intake mounted on the outer end, clear of any boundary
layer air. Air is drawn into the hollow strut, through the air intake, where it
is brought virtually to rest. Its temperature is sensed by a platinum resistance-type element mounted within the strut, which produces an electrical
signal proportional to the temperature. Most modern TAT probes have a
very high recovery factor, usually very close to unity (1.0).
An alternative type of probe uses engine bleed air to create a reduction of
pressure within the casing of the probe. This has the effect of drawing air
into the hollow strut at a higher rate, so that the de-icing heating element
within the strut cannot affect the sensed temperature of the indrawn air.
Figure 1.17 is a diagram of an air temperature measurement probe.
Air Data Instruments
air flow
air flow
bleed air
Figure 1.17
bleed air
TAT measuring probe.
Pressure transducers
Devices called transducers convert pitot and static pressure into suitable
electrical signals for transmission to the air data computer. In the case of
analogue ADCs the transducer is often of the electro-magnetic type, the
amplified output of which drives a servomotor and operates a synchro
system, which in turn operates the analogue instrument displays.
Digital ADCs more commonly utilise piezoelectric transducers that form
part of a solid-state circuit. Some crystalline materials, such as quartz, can be
made to generate varying electrical signals when subjected to pressure. A
diaphragm composed of thin quartz discs impregnated with metallic particles is subjected to pitot or static pressure and the subsequent flexing of the
diaphragm creates an electrical charge in the discs, the polarity of which is
dependent upon the direction of flexing. Thus a signal proportional to
increasing or decreasing pressure is generated.
Sample questions
1. At lower altitudes, near to sea level, the change of atmospheric pressure
with height is approximately:
1 hPa per 50 ft?
1 hPa per 27 ft?
27 hPa per 1 ft?
1 hPa per 10 ft?
30 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
2. ISA mean sea level pressure is:
1013 hPa?
1025 hPa?
29.02 in Hg?
1030 hPa?
3. An aircraft is at 600 ft above an airfield, the elevation of which is 250 ft
amsl. With QFE set, the altimeter will read:
850 ft?
350 ft?
600 ft?
250 ft?
4. An aircraft altimeter has been set for a regional QNH of 1010 hPa, but it
is operating in an area where msl pressure is 1000 hPa. The altimeter is
indicating 2000 ft; actual height amsl is:
2300 ft?
1730 ft?
1940 ft?
1970 ft?
5. Which of the following is correct?
Static pressure = pitot pressure + dynamic pressure?
Dynamic pressure = pitot pressure + static pressure?
Pitot pressure = dynamic pressure 7 static pressure?
Dynamic pressure = pitot pressure 7 static pressure?
6. An aircraft is flying at constant IAS:
As altitude
As altitude
As altitude
As altitude
decreases EAS will increase?
increases CAS will increase?
increases TAS will increase?
decreases TAS will increase?
7. The stalling speed with flaps and landing gear fully extended is assigned
the notation:
Air Data Instruments
8. An aircraft is flying at a TAS of 400 knots at an altitude where the LSS is
550 knots. The aircraft's mach number is:
9. The white arc on an ASI scale:
safe speed range with flaps retracted?
safe speed range with flaps fully extended?
safe speed range with one engine inoperative?
speed range from Vno to Vne?
10. The error suffered by air data instruments that is caused by manufacturing tolerances is known as:
Pressure error?
Position error?
Instrument error?
11. The type of VSI commonly used in sailplanes is:
Simple VSI?
Dynamic vane type?
Variation meter?
12. The data inputs to an ADC are:
TAT, pitot and static?
SAT, pitot and static?
COAT, dynamic and static?
OAT, pitot and static?
Chapter 2
Gyroscopic Instruments and
Gyro fundamentals
Gyroscopic instruments are of great importance in aircraft navigation
because of their ability to maintain a constant spatial reference and thereby
provide indication of the aircraft's attitude. The principal instruments that
use the properties of the gyroscope are the directional gyro, the artificial
horizon or attitude indicator and the turn and bank indicator.
Gyroscopic properties
The gyroscope used in these instruments comprises a rotor, or wheel,
spinning at high speed about an axis passing through its centre of mass and
known as the spin axis. A simple gyro rotor is illustrated in Figure 2.1. When
a rotor such as that in Figure 2.1 is rotating at high speed it exhibits two basic
properties, known as rigidity and precession. It is these properties that are
utilised to give gyroscopic instruments their unique features.
of r
Figure 2.1 Gyro rotor.
Gyroscopic Instruments and Compasses
The spinning rotor of the gyro has rotational velocity and therefore, if we
consider any point on the rotor, that point has angular velocity as indicated
by the arrow A in Figure 2.1. Since the rotor has mass, that angular velocity
produces angular momentum, which is the product of angular velocity and
mass. As stated in Newton's First Law of Motion, any moving body tends to
continue its motion in a straight line and this is known as inertia. In the case
of the spinning gyroscope there is a moment of inertia about the spin axis,
which tends to maintain the plane of rotation of the gyro. Consequently, the
spin axis of a gyroscope will maintain a fixed direction unless acted upon by
an external force. This property is known as rigidity. Another way of putting
this is that the spin axis of the gyro will remain pointing toward a fixed point
in space unless it is physically forced to move.
Since rigidity is the product of angular velocity and mass it follows that
the rigidity of a gyroscope may be increased by increasing either its angular
velocity, or its mass, or both. Increasing the speed of rotation of the rotor, or
its diameter, will increase angular velocity and therefore angular momentum. The rotor diameter is constrained by the need to keep the instrument as
compact as possible and so the gyro rotor is made to spin at very high speed.
Similarly, the mass of the rotor is constrained by its size limitations, but
angular momentum is improved if the mass is concentrated at the rim of the
rotor, as seen in Figure 2.1.
Precession is defined as the angular change in direction of the spin axis when
acted upon by an applied force. Let us suppose that the axis of the gyro rotor
in Figure 2.2 has a force applied to it as shown in (a). Application of a force to
the spin axis as shown is exactly the same as if the force had been applied at
point X on the rotor. If the rotor were stationary then it, and its spin axis,
would tilt as shown at Figure 2.2(b). However, when the rotor is rotating it
has not only the applied force acting upon it, but also the angular momentum previously described. The combination of the two displaces the effect of
the applied force through 908 in the direction of rotation, as shown in Figure
2.2(c). Thus, the spin axis of the gyroscope will precess as shown in Figure
2.2(d) in response to the force applied in Figure 2.2(a).
The rate at which a gyro precesses is dependent upon the magnitude of the
applied force and the rigidity of the rotor. The greater the applied force, the
greater the rate of precession. However, the greater the rigidity of the rotor
the slower the rate of precession for a given applied force.
A gyro will continue to precess so long as the applied force is maintained,
or until the applied force is in the same plane as the gyro plane of rotation, as
34 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
gyro rotor
applied force
applied force
Figure 2.2 Gyroscopic precession.
shown in Figure 2.2(e). If the applied force is removed, precession will
immediately cease.
Free gyro
Clearly the rotor of the gyroscope must be contained within a supporting
structure. The rotor spindle is mounted within a ring known as a gimbal and
this is in turn mounted within a framework, the design of which depends
upon the gyro function. All gyroscopes must have freedom for the rotor to
rotate and to precess. The gyroscope cannot precess about the axis of rotation,
but precession may take place about either of the two axes at right angles to the
plane of rotation. A gyroscope that has freedom to precess about both these
axes is known as a free gyro, and is said two have two degrees of freedom of
precession. Such a gyroscope is illustrated in Figure 2.3. The number of
Gyroscopic Instruments and Compasses
Figure 2.3
Free gyro.
degrees of freedom of precession of any gyroscope is the same as its number of
gimbals. It will be seen that the gyro rotor spindle is mounted in bearings
within a ring, or gimbal, known as the inner gimbal. This is in turn mounted in
bearings that are attached to a second gimbal ring, known as the outer gimbal.
Thus, the gyro rotor is free to spin about spin axis XX and it also has freedom of
movement about the inner gimbal axis YY. The outer gimbal is mounted in
bearings attached to the frame of the assembly and therefore has freedom of
movement about the third axis, ZZ.
If the frame of the free gyro were to be fixed to the instrument panel of an
aircraft, the aircraft could be pitched, rolled or even inverted and the spin
axis of the spinning gyroscope would remain aligned with the same fixed
point in space. In point of fact the free gyro has no practical application in
aircraft, but gyroscopes having freedom of precession about two axes,
known as tied gyros, are extremely useful.
Tied gyro
The aircraft instruments that employ gyros provide a fixed reference, about
which aircraft movement is indicated to the pilot. The directional gyro
provides the pilot with aircraft heading information and so its reference
axis is the aircraft's vertical axis and the gyro rotor must be sensitive to
movement about that axis and no other. The function of the attitude indicator is to provide the pilot with indications of aircraft attitude with reference to the pitch and roll axes of the aircraft and so its gyro must be
sensitive to aircraft movement about these axes. The turn indicator is
36 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
required to indicate rate of turn and so it must be sensitive to aircraft
movement about the vertical axis.
A gyroscope is not sensitive to movement about its spin axis, so its rotor
must be maintained at right angles to the required axis for maximum sensitivity. Consider the situations depicted in the illustration in Figure 2.4.
Figure 2.4 Gyroscope spin axis alignment.
Axis AA is the aircraft's vertical, or yaw, axis. Any movement about this
axis involves a change in aircraft heading and so we require the directional
gyro to be sensitive to movement about this axis. Since a gyro is not sensitive
to movement about its spin axis it is clear that a gyro with its spin axis
vertical (gyro D in Figure 2.4) would not be suitable, but that the spin axis of
a directional gyro must be maintained horizontal.
The attitude indicator is required to indicate aircraft attitude with reference to the aircraft pitch and roll axes, BB and CC respectively in Figure 2.4.
Clearly its spin axis must not be aligned with either of these aircraft axes.
Consequently gyro E would not be suitable, because its spin axis is aligned
with the aircraft's pitch axis, and gyro F would not be suitable, because its
spin axis is aligned with the aircraft's roll axis. Thus, the spin axis of the
attitude indicator gyro must be maintained vertical as in gyro D, and not just
aircraft vertical, but earth vertical.
The turn indicator is required to indicate rate of turn, that is the rate at
which the aircraft is turning about its vertical axis. The turn indicator gyro
must therefore be sensitive to movement about the aircraft vertical axis AA
and so its spin axis must be aligned with either axis BB or CC. For practical
reasons that will become apparent when we study this instrument in detail,
it is aligned with the aircraft's lateral axis BB, as for gyro E. The gyro of the
turn indicator is known as a rate gyro.
Gyroscopic Instruments and Compasses
In each case we require the spin axis of the gyroscope to be tied to a particular direction, e.g. earth vertical or aircraft horizontal. Such a gyroscope is
known as a tied gyro. A tied gyro that is controlled by the earth's gravity is
also known as an earth gyro; this is the case with the attitude indicator.
Another type of gyro, which we will meet later when we study the gyrostabilised platform of an inertial navigation system, is the rate integrating
gyro. This is a single degree of freedom gyro, sensing rate of movement
about one axis only, which is integrated to give change of distance.
Gyroscope drift (wander) and topple
Earlier in this dialogue it was stated that the spin axis of a gyroscope remains
aligned with some point in space, as opposed to alignment with any earth
reference such as true or magnetic north. Any deviation of a horizontally
aligned spin axis from its point of reference is known as gyro drift, or
wander. Gyro drift is of two types, real drift and apparent drift. Deviation of
a vertically aligned spin axis from its reference is known as gyro topple.
Real drift
As we have seen, the gyro comprises a spinning rotor mounted in a gimbal,
which is in turn pivoted to either another gimbal or a frame. If the rotor, its
spindle or a gimbal is not perfectly balanced the imbalance will apply a force
to the rotor. This force will cause precession, which will cause the spin axis of
the gyro to deviate from its spatial reference. The same effect can arise due to
friction or wear in the rotor spindle bearings. The drift due to the spatial
deviation is known as real or random drift; it is usually very small and it
cannot be calculated, so it is impossible to produce correction charts for real
Apparent drift
Let us now consider the case of a gyroscope with its spin axis tied to horizontal, as in the case of the directional gyro. Imagine this gyro is at the true
north pole, where all directions are south. The spin axis of the gyroscope has
been aligned with the Greenwich meridian, 08 of longitude. Now remember
that a gyroscope alignment is really with some point in space and it is to this
unknown point that the spin axis is truly pointing. The earth rotates at 158
per hour and so, assuming that our gyro is perfect and does not suffer from
any real drift, after one hour its spin axis will still be aligned with the same
point in space. However, to the earthbound observer it will no longer be
aligned with 08 of longitude, but will appear to have drifted by 158. This is
known as apparent drift due to earth rotation.
38 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
If the same gyro were to be taken to any point on the equator and aligned
with true north it would not suffer at all from apparent drift due to earth
rotation, because the earth reference point and the space reference point are
in alignment and remain so. Thus, the rate of apparent drift due to earth
rotation varies with latitude and can be calculated, since it varies as the sine
of the latitude. Apparent drift due to earth rotation is given as 15 6 sin
latitude8/hr. Apparent drift is illustrated in Figure 2.5.
Axis referenced to point in space.
Gyro appears to drift at 15°/hr
Axis referenced to point in space, but aligned with
local meridian. Gyro does not appear to drift
Figure 2.5 Apparent drift due to earth rotation.
Transport drift
Since the earth rotates about its north±south axis at 158 per hour it follows
that an aircraft flying on a westerly heading, against the direction of earth
rotation, will experience a greater rate of longitude change and one flying on
an easterly heading a lesser rate. To an observer on the aircraft, the spin axis
of a horizontally aligned gyro would appear to drift even though no change
of latitude has occurred. This apparent drift is known as transport drift.
Apparent topple
A vertical axis gyro will also suffer apparent wander, which is conventionally known as topple. Suppose a vertical axis gyro is taken to the true
north pole. Its spin axis will be aligned with the earth's spin axis and
pointing toward some point in space. The earth reference and the space
reference will remain in alignment and there will be no gyro topple. Suppose
now the same gyro were to be taken to a point on the equator and started
spinning with the spin axis perpendicular to the earth's surface, i.e. earth
Gyroscopic Instruments and Compasses
vertical. After one hour the earth will have rotated 158. The gyro spin axis
will have maintained its spatial alignment and will appear to have toppled
by 158. As with apparent drift, the rate of topple is dependent upon the
latitude at which the gyro is located, but in this case it varies as the cosine of
the latitude. Apparent topple is given as 15 6 cos latitude8/hr. Apparent
topple is illustrated in Figure 2.6.
Axis aligned with point in space and with
earth axis. Gyro does not appear to topple
Axis aligned with point in space
does not remain perpendicular
to earth's surface. Gyro appears
to topple at 15 /hr.
Figure 2.6
Apparent topple due to earth rotation.
Ring laser gyro
Unlike the conventional gyroscopes described above, the ring laser gyro is a
solid state device that does not have any moving parts. A simplified diagram
of a ring laser gyro is shown in Figure 2.7.
The device is made from a block of very expensive glass, within which
there is a triangular cavity of exact dimensions, filled with a suitable lasing
medium, such as helium±neon. Each side of the triangular cavity is exactly
laser path
Figure 2.7
Ring laser gyro.
40 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
the same length and at each of the three junctions is a mirror, one of which is
partially transmitting. At the mid-point of one side of the triangular cavity is
a cathode and in the other two sides an anode is positioned at exactly the
same distance from the mirrors. Laser beams travelling between the cathode
and each anode will take exactly the same length of time to travel exactly the
same distance.
However, if the ring laser gyro is rotated about the axis perpendicular to
the laser path, one laser beam will arrive at one anode slightly before the
other beam arrives at the other anode, and the time difference will be proportional to the rate of rotation. The direction of rotation will determine
which laser has the shorter distance to travel. The time difference is
measured and used to produce a digital readout of rate and direction of
Ring laser gyros, although very expensive to produce, have the advantage
of being much more reliable than conventional gyros, because there are no
moving parts subject to wear. Also they are available for immediate use
when switched on, whereas conventional gyros take some time to spin up
and stabilise.
Gyro drives
The rotors of gyroscopic instruments must spin at high speed to give the
degree of rigidity needed and the motive power for them is either pneumatic
or electrical.
Pneumatic drive. Air-driven gyro rotors are typically powered from
the aircraft's vacuum system, air being drawn through the instrument
by an engine-driven vacuum pump that maintains approximately 4 in
Hg of vacuum in the system. A schematic diagram of a typical light
aircraft vacuum system is shown in Figure 2.8. Pneumatically driven
gyros in aircraft that operate at high altitude are usually supplied with
air pressure rather than vacuum, because of the difficulty in producing
the requisite vacuum in a low-pressure environment. In either case the
air entering the instrument is directed onto bucket-shaped indentations
in the rim of the gyro rotor, driving it as a simple turbine.
In some light aircraft the vacuum is produced by means of a venturi
tube placed in the airflow. Because the device only operates in flight at
speeds in excess of 100 knots and is susceptible to icing, it is unsuitable
for use in aircraft where instrument flight may be required.
Electrical drive. Alternating current (a.c.) or direct current (d.c.) motors
are also used to drive gyroscopic instrument rotors, using power from
the aircraft electrical systems. As a general rule, a.c. motors are preferred for attitude indicators and d.c. motors for turn indicators. Simple
Gyroscopic Instruments and Compasses
inlet filter
Figure 2.8
Typical light aircraft vacuum system.
direction indicators are usually air-driven, but those forming part of a
magnetic heading reference system, such as the slaved gyro compass,
are normally driven by electric motors. In some aircraft the main panel
instruments may be electrically driven and the standby instruments airdriven.
Electrically powered gyros are necessary in aircraft intended for high
altitude flight. Because they are capable of much higher rotational
speeds than pneumatically powered instruments they offer increased
stability and lighter construction.
In either case it is clear that interruption of the power supply will render
the gyros unserviceable and it is important that the pilot should be immediately aware of this. Loss of power, whether pneumatic or electrical, will be
indicated by a warning flag on the face of the instruments.
Directional gyro (DG)
The function of the directional gyro is to indicate the aircraft heading, utilising the rigidity of a spinning gyroscope so to do. The gyro spin axis is
maintained horizontal and it can be set so that it is referenced to either
magnetic north or true north. It will then hold this reference whilst the aircraft heading changes. A compass scale is attached to the outer gimbal of the
gyroscope. The instrument casing, which is of course attached to the aircraft,
moves around the fixed reference scale card as the aircraft changes heading.
As a two-gimbal gyro, it has two degrees of freedom of precession.
Figure 2.9 shows the operating principle of the directional gyro. The rotor of
the directional gyro is mounted in an inner gimbal ring with its spin axis
42 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
inner gimbal
Figure 2.9 Directional gyro principle of operation.
horizontal, so that it is free to rotate in the vertical plane. The inner gimbal is in
turn pivoted to an outer gimbal, so that it is free to move about a horizontal
axis at right angles to the spin axis. The outer gimbal is pivoted to the case of
the instrument and is free to rotate about the vertical axis. The compass scale
card is attached to the outer gimbal and is typically marked from 08 to 3608.
The rotor is normally driven by air drawn in through the aircraft's vacuum
system and directed by a nozzle onto buckets machined in the rim of the
rotor. The rotor typically rotates at about 12 000 rpm.
Adjustment procedure
Before the start of a flight the directional gyro must be set up so that the rotor
is spinning with its spin axis horizontal and the heading indication agrees
with the aircraft compass reading. This is achieved with a caging mechanism
and an adjustment knob on the face of the instrument.
Depressing the adjustment knob engages a caging mechanism that locks
the inner gimbal in a horizontal position. It also engages a pinion with a
bevel gear attached to the outer gimbal of the instrument. Rotating the
adjustment knob will rotate the outer gimbal and its attached compass card
and this is done until the lubber line on the face of the instrument is aligned
with the required heading. Once this is satisfactory the adjustment knob is
pulled out to disengage the caging mechanism and bevel gear, leaving the
gyro spin axis free to maintain its fixed reference. This adjustment can also
be made in flight, but must be done with the aircraft flying straight and level.
The instrument should also be caged during violent manoeuvres to prevent
the gyro from toppling.
Gyroscopic Instruments and Compasses
Erection system
During a change of heading an aircraft is turning about its vertical, or yaw,
axis. Whilst it is doing so the aircraft is, of course, banked and so the spin axis
of the directional gyro must also be tilted to keep it aircraft horizontal. As we
know, to move the spin axis of a gyro away from its fixed reference it is
necessary to precess the gyro, and this is achieved through the design of the
nozzle that directs air onto the rim of the rotor.
In later designs of air-driven directional gyros the air from the rotor
is exhausted onto a wedge attached to the outer gimbal, as shown in Figure 2.10. Whilst the spin axis of the rotor remains aircraft horizontal the
spin axis and the outer gimbal are mutually perpendicular and the
exhaust air strikes both sides of the wedge equally, as seen in Figure
air in
air out
Figure 2.10
air out
Rotor erection system ± directional gyro.
When the aircraft begins to bank in a turn the outer gimbal banks
with it and the rotor axis is no longer at right angles to the outer gimbal. Exhaust air now strikes one side of the wedge more than the other.
This applies a force to the outer gimbal that is tending to rotate it about
the vertical axis, which is the same as applying a force to one side of
the gyro rotor. That force is precessed by the rotor through 908 in the
direction of rotation, tilting it to keep its spin axis aircraft horizontal.
Any tendency of the rotor to move from the aircraft horizontal reference
will be corrected by this device.
Earlier DGs used a split, or bifurcated, air nozzle to achieve the same
44 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
Gimbal error
Gimbal error is when the gimbals of a gyroscope are not mutually perpendicular and the gyroscope itself is displaced. Briefly, the reason for these
errors is because the spin axis of the DG is aligned with east±west on the
instrument compass card. Thus, with the DG properly set, the rotor spin axis
will be at right angles to the aircraft longitudinal axis (and therefore the
outer gimbal) when the aircraft is on an east or west heading and aligned
with it on a north or south heading. On these headings the DG will only
suffer gimbal error if the aircraft is pitched and rolled simultaneously. On
other headings it may occur during either pitch or roll attitude changes. The
effect of gimbal error is that the instrument will give a false heading indication whilst the manoeuvre is in progress, but it will normally indicate
correctly once the aircraft is returned to straight and level flight.
Drift calculations
The apparent drift due to earth rotation for a gyro that has no random (real)
drift and that is stationary on the ground (i.e. not affected by transport drift)
can be calculated given the latitude at which the gyro is located and the
hemisphere, north or south.
In the northern hemisphere a gyro that has been aligned with north will
appear to drift at the rate of 15 6 sin latitude 8/hr and the DG indication will
decrease at that rate. Reference to Figure 2.11(a) will show why this is the case.
Let us assume that the aircraft in which the DG is installed is stationary on
the ground at latitude 508N, that it is on a westerly heading, and that the DG
has been aligned with the local meridian and is indicating a heading of 2708.
This is the situation at point A in Figure 2.11(a). The gyro drift rate will be 15
6 sin 50 8/hr, which is 15 6 0.766 = 11.58/hr. After one hour has elapsed the
gyro has not moved, but the earth has rotated and the situation will be as at
point B. The gyro will appear to have drifted by 11.58 and its reading will
have decreased by that amount, because its space reference is now 11.58 to
the east of the local meridian. Consequently, it will now be indicating 258.58.
Without adjustment the gyro indication would continue to decrease at the
rate of 711.58/hr.
In the southern hemisphere a similar situation occurs and this is illustrated in Figure 2.11(b). With the aircraft on the ground at 508S and the gyro
indicating 2708 at point C the apparent drift rate due to earth rotation will
again be 11.58/hr, but now the readings will be increasing at that rate. This is
because, after one hour has elapsed, the gyro's space reference will lie 11.58
to the west of the local meridian, in position D, and so the DG will indicate a
heading of 281.58.
Gyroscopic Instruments and Compasses
direction of earth rotation
Figure 2.11
DG drift due to earth rotation.
Drift compensation
Clearly a heading indicator that was incapable of maintaining an accurate
heading indication would be of no use and it will come as no surprise to
learn that the directional gyro contains a compensation device. This is illustrated in Figure 2.12. Attached to the inner gimbal of the gyro is a threaded
spindle with a nut attached. The gimbal is manufactured with a slight
imbalance such that it is perfectly balanced with the nut, known as a latitude
nut, in approximately a mid-position on its spindle. If the nut is screwed
outward on the spindle it will apply a downward force to the gyro rotor,
which will be precessed 908 in the direction of rotation to cause the gyroscope to precess about the vertical axis. Screwing the latitude nut inward
will allow the slight imbalance of the inner gimbal to apply a force in the
opposite direction.
Figure 2.12
Latitude nut.
46 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
For the northern hemisphere situation described above, the latitude nut
would be adjusted to precess the gyro at the rate of +11.58/hr, thus exactly
compensating for the apparent drift rate of 711.58/hr. In the southern
hemisphere the adjustment would be in the opposite direction.
Comparison with magnetic compass
The latitude nut of the directional gyro provides compensation only at the
latitude for which it is set. This setting can only properly be made with the
instrument in a workshop. If the instrument is transported north or south of
the set latitude it will begin to suffer from apparent drift due to earth rotation, and the further it is moved the greater will be its error rate. Consequently, the DG must always be referenced to the aircraft magnetic compass
and this can be readily done in level flight or on the ground with the
adjustment knob.
An advantage of the directional gyro is that it does not suffer the turning
and acceleration errors of the magnetic compass and so its heading information tends to be more accurate, especially in a steady level turn.
In the event of failure of the gyroscope the rotor will almost inevitably
topple and its indication will be useless. Under these circumstances a
warning flag will appear, to obscure the display.
Effect of friction
Friction or wear in the bearings of the gyro rotor will cause real, or random,
drift. In a properly maintained DG this should be insignificant.
Attitude indicator (artificial horizon)
The attitude indicator uses a vertical earth gyro that has freedom of movement about all three axes. The gyro spin axis is maintained earth vertical,
using the force of gravity to keep it aligned with the earth's centre. Attitude
indicators may be pneumatically or electrically driven. The purpose of the
instrument is to provide the pilot with an indication of the aircraft attitude in
both pitch and roll.
Construction and principle of operation
Figure 2.13 illustrates the principle of operation of a pneumatic attitude
indicator. The vertical gyro rotates at about 15 000 rpm and is contained
within an inner gimbal. It is maintained earth vertical, thus spinning in the
earth horizontal plane, by utilising gravity. The inner gimbal is pivoted to an
outer gimbal with the pivot axis lying parallel to the aircraft lateral axis. The
Gyroscopic Instruments and Compasses
horizon bar
actuating pin
bank angle
Figure 2.13
XX vertical spin axis
YY lateral pitch axis
ZZ longitudinal roll axis
Attitude indicator ± principle of operation.
outer gimbal is in turn pivoted to the instrument casing with the pivot axis
lying parallel to the aircraft longitudinal axis.
Since the instrument casing is attached to the airframe it follows that any
change in aircraft attitude must take place about the vertically referenced
gyro. Thus, if the pitch attitude changes, the outer gimbal will pitch up or
down relative to the gyro spin axis. If the roll attitude changes the outer
gimbal will roll left or right relative to the gyro spin axis.
Attached to the outer gimbal is a sky plate, which is viewed through the
face of the instrument. The upper half of the plate is typically coloured pale
blue to represent the sky and the lower half black to represent the earth, the
two divided horizontally to represent the earth's surface. Also attached to
the outer gimbal by a pivoted spindle is a bar, which extends across the front
of the sky plate parallel to the dividing line. This bar, known as the horizon
bar, is driven by a spindle attached to the inner gimbal. Printed on, or
attached to, the glass cover of the instrument is a fixed symbol representing
the aircraft.
With the aircraft flying straight and level, the gyro spin axis will be perpendicular to both the lateral and the longitudinal aircraft axes and the
horizon bar and aircraft symbol will appear in the mid-position, as shown in
Figure 2.14(a). If the aircraft is pitched nose-up, the outer gimbal will be
pitched up with it, raising the front of the gimbal relative to the gyro spin
axis, which remains earth vertical.
Because the horizon bar is pivoted to the forward end of the outer gimbal,
that end of the bar will rise and pivot about the actuating pin protruding
from the inner gimbal. This causes the horizon bar to move down relative to
the aircraft symbol, indicating a climb. In a descent the reverse happens and
the horizon bar moves up relative to the aircraft symbol. These indications
are represented in Figures 2.14(b) and 2.14(c).
When the aircraft is rolled to left or right the bank indication is given by
the position of the horizon bar relative to the aircraft symbol. This is because
48 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
horizon bar
bank scale
Figure 2.14 Attitude indicator presentations.
the aircraft symbol, on the glass face of the instrument, will have rolled with
the aircraft about the vertically referenced gyro spin axis, which has maintained the outer gimbal, and therefore the horizon bar, earth horizontal.
Typically, a pointer attached to the outer gimbal will indicate bank angle
against a scale printed on the glass face of the instrument. Figure 2.14(d)
shows the indication with the aircraft banked to the left.
Some attitude indicators incorporate an adjustment knob that can be used
to raise or lower the aircraft symbol, so that it may be positioned against the
horizon bar when the aircraft is flying straight and level, but with the
fuselage pitched up or down. This is particularly useful in helicopters, which
frequently fly level in a pitched attitude.
Erection mechanism
The vertical gyro of the pneumatic attitude indicator is maintained earth
vertical by means of an erection unit beneath the inner gimbal. The air that
has driven the gyro rotor is exhausted through four equally spaced ports
machined in the sides of the unit, two in the lateral (athwartships) axis and
two in the longitudinal (fore and aft) axis. When the gyro spin axis is earth
vertical, each port is partly covered by a freely pivoted pendulous vane and
the exhaust air escapes equally from each port, as illustrated in Figure
Gyroscopic Instruments and Compasses
Figure 2.15
Erection system ± pneumatic attitude indicator.
If the spin axis, and with it the inner gimbal, tilts away from earth vertical,
the vanes, because they are pendulous, will continue to hang vertically.
Suppose the gyro has tilted as shown in Figure 2.15(b). On one side of the
erection unit the vane has fully uncovered its port, whilst on the other the
vane will have fully covered its port. The front and rear vanes will not have
moved relative to their ports, and so these ports will remain half open.
Consequently, air will be exhausted from one side of the erection unit only
and there will be a reaction force, in the opposite direction, applied to the
gyro rotor. The gyro will precess this force 908 in the direction of rotation,
which will serve to re-erect the gyroscope. When it is once again earth
vertical, all four ports will be equally uncovered and the erection forces will
once again be in balance.
Acceleration errors
The erection mechanism of the pneumatic attitude indicator is the cause of
false attitude indications during aircraft acceleration. When the aircraft
accelerates in a level attitude, such as during the take-off run, the pendulous
vanes tend to swing rearward due to inertia. This does not affect the front
and rear vanes. However, by referring to Figure 2.15 it will be seen that this
will result in the right side port becoming uncovered more than the left side
port. Because the gyro rotor spins anti-clockwise when viewed from above,
the reaction to this sideways imbalance of force will apply a force to the rotor
which, when precessed, will tilt the rotor to give a false climb indication.
The erection unit itself is also pendulous, suspended as it is beneath the
50 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
gyro rotor. Consequently, during a rapid acceleration inertia tends to swing
it rearward, thereby applying a rearward force to the rotor. This is precessed
908 in the anti-clockwise direction of rotation to tilt the rotor to the right,
giving a false indication of right bank.
Thus, the overall effect of aircraft acceleration is to give a false indication
of a climbing right turn.
Because of these errors, pneumatic attitude indicators are usually only
fitted to light and general aviation aircraft of low performance which have
limited electrical power available. Wherever the electrical power supplies
are adequate it is usual to fit electrically driven attitude indicators, which are
less susceptible to acceleration error.
Turning errors
During a turn, centrifugal force acts to swing the fore and aft pendulous
vanes outward from the centre of the turn. The resultant reaction force on the
erection unit, when precessed, tends to tilt the gyro to give a false indication
of bank. Furthermore, the pendulosity of the erection unit, suspended
beneath the inner gimbal, applies a sideways force to the gyro that, when
precessed, gives a false indication of pitch. In general the turning errors are
less serious than the acceleration error.
Electrically driven attitude indicator
The principle of operation of the electrically driven attitude indicator is
essentially the same as that of the pneumatic instrument. The gyro unit is an
electric motor that rotates at considerably higher speed, typically around
22 000 rpm, and therefore has greater rigidity.
The erection system is quite different, however. Instead of the pendulous
system of the pneumatic type, it employs two torque motors mounted on the
outer gimbal and operated by mercury-filled levelling switches. Figure 2.16
is a diagram showing the operating principle of the system.
The torque motors are a.c. induction machines with their stators mounted
on the outer gimbal in line with its lateral and longitudinal axes. When
current is supplied to the stator a rotating magnetic field is set up, which
tends to rotate the rotor surrounding the stator. This tendency is opposed by
the rigidity of the gyroscope, resulting in a torque reaction acting about the
axis of the motor, and therefore about the pitch or roll axis of the outer
The mercury-filled levelling switches are small tubes partly filled with
mercury, mounted on the inner gimbal of the gyroscope. When the gyroscope is running at normal speed an electrical current is supplied to a central
contact in the tube. Contacts at either end of the tube are connected to the
Gyroscopic Instruments and Compasses
roll torque
principle of
levelling switches
outer gimbal
Figure 2.16
Erection system ± electrical attitude indicator.
associated torque motor stator field windings. Whilst the inner gimbal
remains earth horizontal the mercury in the switch is centralised and there is
no conducting path between the central, supply contact and either of the
end, output contacts. If the switch is tilted, due to the outer gimbal tilting, the
mercury runs to one end of the tube and connects the electrical supply to one
output. The direction of tilt will determine the direction of the torque
applied by the appropriate torque motor. The principle is illustrated in the
scrap view in Figure 2.16.
Let us suppose that the inner gimbal has begun to topple rearward, that is
anti-clockwise about the pitch axis as viewed in Figure 2.16. The roll levelling switch, aligned with the pitch axis of the instrument, will not be affected
and its mercury will remain centralised. The pitch levelling switch, aligned
with the instrument roll axis, will be tilted and its mercury will run to the
rear end of the tube, completing the electrical supply circuit to the pitch
torque motor and causing it to apply a torque force anti-clockwise about the
roll axis, as viewed in Figure 2.16. The torque produced will be applied to the
vertical gyro and precessed through 908 in the direction of rotation. This will
result in a clockwise torque force about the longitudinal axis, as viewed in
Figure 2.16, acting upon the gyro to re-erect it. Once it is restored to earth
vertical the levelling switches will both be in the neutral (mercury centralised) positioned and supply to the torque motors isolated.
Acceleration and turning errors
The mercury switches are susceptible to acceleration, during which inertia
will force the mercury to one end of the tube and make the contacts to supply
power to one or both torque motors. This would, of course, lead to false
indications similar to those described for the pneumatic instrument. However, it is a relatively simple matter to incorporate a cut-out system in the
electrical circuitry, which will detect acceleration and distinguish it from
52 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
topple. This will cut off supply to the switches during detected acceleration
and prevent false climb or bank indication.
Warning indications
In the event of failure of the vacuum system the pneumatic attitude indicator
will normally display a warning flag on the face of the instrument. The
electrical attitude indicator typically displays an OFF flag when its power
supply is disconnected.
Erection speed
Pneumatic attitude indicators typically have an erection speed of about 88
per minute, which means that they usually take in excess of 5 minutes to
erect from start-up. Electrical attitude indicators have erection rates, of the
order of 38 to 58 per minute. Many have fast erection systems for use during
start-up, giving an erection time of less than one minute. Some pneumatic
instruments are equipped with a caging system, similar to that described for
the directional gyro, which shortens the start-up erection process.
Turn and bank indicator (rate gyro)
The purpose of the turn and bank indicator is to measure and display the
aircraft rate of turn and to indicate whether the aircraft is correctly banked
for a co-ordinated turn with no slip or skid. To measure the rate of turn, i.e.
rate of movement about the yaw axis, the instrument employs a rate gyro
that is sensitive to movement about the aircraft yaw axis only. The bank
indication is a separate device using a combination of gravitational and
centrifugal force.
Rate gyroscope
Since the rate gyroscope is required to be sensitive to movement about the
yaw axis it follows that its spin axis must be perpendicular to that axis, i.e.
horizontal. The gyro rotor is mounted in a gimbal with its spin axis aligned
with the lateral (athwartships) axis of the aircraft. The single gimbal is
pivoted fore and aft in the instrument casing, in line with the aircraft
longitudinal axis. The gyro rotor spins up and away from the pilot. The
general arrangement showing the principle of operation is shown in Figure
2.17. It will be seen that the gyro has freedom of movement about two axes
only, the lateral spin axis and the longitudinal precession axis.
When the aircraft yaws about the vertical axis this applies a force to the
gyro rotor at the front, in line with the spin axis. Let us suppose that the
Gyroscopic Instruments and Compasses
Figure 2.17
Turn and bank indicator ± principle of operation.
aircraft is turning to the left. This applies a torque force about the yaw axis in
an anti-clockwise direction viewed from above. This is as though a linear
force were applied to the front of the gyro rotor on the right side in line with
the spin axis, as illustrated at Figure 2.17.
The gyro will precess this force 908 in the direction of rotation, so that it
becomes torque acting in a clockwise direction about the longitudinal axis,
precessing the gyro so that the gimbal begins to tilt to the right. The extent to
which the gimbal tilts is limited by a spring connecting the gimbal to the
instrument casing. As the spring is stretched it exerts a force on the gimbal
opposing the precession. When the two are in balance the gimbal is held at a
tilt angle that is proportional to the rate of turn, because the precession is
equal to the rate of turn and the angular momentum of the gyroscope.
Thus, the greater the rate of turn, the greater the tilt of the gimbal. The
gimbal actuates a pointer, which moves against a calibrated scale on the
face of the instrument to indicate rate of turn. The actuation is such that
when the gimbal tilts to the right the pointer moves to the left and vice
The speed of rotation of the turn indicator gyro is relatively low, typically
about 4500 rpm. It is critical that its speed is maintained constant, since this is
a vital factor in ensuring that precession remains constant relative to rate of
54 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
turn. A warning flag will appear on the face of the instrument when the gyro
rotational speed is outside limits.
Rate of turn is classified numerically, where rate 1 equals 1808 per minute,
rate 2 equals 3608 per minute, rate 3 equals 5408 per minute and rate 4 equals
7208 per minute. These may also be quoted as 38, 68, 98 and 128 per second,
respectively. An aircraft maintaining a rate 1 turn for 2 minutes will therefore turn through 3608.
Bank indication
The bank indication given by the turn and bank indicator displays to the
pilot whether or not the aircraft is correctly banked for the turn being made.
If the aircraft is banked excessively it will tend to slip toward the centre of
the turn, whereas if it is underbanked it will skid outwards, away from the
centre of the turn. Hence the name by which this instrument was once
commonly known, the turn and slip indicator.
The display is provided by a device quite separate from the rate gyroscope
of the turn indicator, and typically comprises a curved glass tube filled with
liquid and containing a ball. When the aircraft is in level flight, gravity
ensures that the ball lies in the centre of the curved tube, as shown in Figure
2.18(a). When the pilot is making a properly co-ordinated banked turn the
glass tube, which is attached to the instrument, will be banked with the
aircraft and the resultant of centrifugal force and gravitational force will
keep the ball in the centre, as shown at Figure 2.18(b). Suppose now that the
aircraft is turning, but that the bank angle is greater than it should be, i.e. the
aircraft is overbanked. The centrifugal force acting on the ball is less than the
gravitational force and the ball falls into the lower part of the tube, as shown
in Figure 2.18(c). This indicates to the pilot that the aircraft is slipping into
the turn. If the aircraft is underbanked the centrifugal force acting on the ball
is greater than the gravitational force and the ball will be moved into the
upper part of the tube, indicating that the aircraft is skidding out of the turn.
This is shown in Figure 2.18(d). Figure 2.18(e) shows the turn and bank
indications during a properly co-ordinated 2 minute (rate 1) standard turn.
Turn co-ordinator
Light aircraft are often fitted with a variation of the turn and bank indicator, known as a turn co-ordinator. The purpose of the instrument is to
present the pilot with a display that makes co-ordination of bank angle
and turn rate as simple as possible. The display is as shown in Figure 2.19.
When the pilot banks the aircraft to initiate a turn the aircraft symbol on
the display banks in the appropriate direction, since it is actuated by the
gyro precession exactly as previously described. Provided that the aircraft
Gyroscopic Instruments and Compasses
Figure 2.18
Turn and bank indications.
symbol is aligned with the left or right bank indication on the display, and
the ball is in the centre, the aircraft will be making a properly banked rate
1 (2-minute) turn.
The main constructional difference between this and the turn and bank
indicator is that the longitudinal axis of the gyro gimbal is inclined at 308 to
the horizontal, so that the gyro will respond to banking as well as turning
input force.
The movement of the gimbal ring of all indicators is damped to control the
rate of precession. Among other effects, this will limit the instrument bank
indication when turning during ground taxiing.
56 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
Figure 2.19 Turn co-ordinator display.
Magnetic compass
The magnetic compass works on the principle that a freely suspended
magnet will align itself with the earth's magnetic field such that one end will
point toward the north magnetic pole. It is mandatory that all civil aircraft
must carry such a compass and in all but light general aviation aircraft it
serves as a standby compass.
To be effective a magnetic compass must meet three basic requirements.
The compass card must lie horizontal, since it is direction parallel to the
earth's surface that it is required to indicate. This is not quite as simple as it
might sound, since the earth's magnetic field dips toward the poles and is
only parallel to the surface near the equator. This is illustrated in Figure 2.20.
For simplicity, the magnetic force is resolved into its horizontal and vertical components, H and Z respectively. Thus, the nearer one is to either
terrestrial magnetic pole, the greater the Z force and the weaker the H force.
This causes the compass magnet to dip toward the earth's surface and,
clearly, the greater the angle of dip, the less accurate the compass becomes.
The amount by which the compass magnet system dips can be reduced
significantly by suspending it so that the centre of gravity of the magnet
system is well below the pivot point of the circular plate to which the
magnets and the compass card are attached. The principle is illustrated in
Figure 2.21, from which it can be seen that the tilt caused by the Z force effect
is countered by the CG becoming offset from the vertical pivot line and the
angle of dip is the resultant of the Z force and gravity.
Figure 2.21 shows the effect of a lowered CG on dip angle in the northern
hemisphere, where the magnet system dips toward the north magnetic pole.
In the UK the dip angle is about 28 to 38, but the further north the compass is
Gyroscopic Instruments and Compasses
TN = true north pole
MN = magnetic north pole
nal ax
magnetic eq
Figure 2.20
Terrestrial magnetism.
taken the greater this will become until, at about 708N the compass becomes
unusable. In the southern hemisphere the magnet system dips toward the
south magnetic pole and the compass again becomes useless above about
708S. Hence, in the northern hemisphere the magnet system CG lies to the
south of the pivot point and in the southern hemisphere it lies to the north of
the pivot point. This is of significance when we come to consider the effects
of turning and acceleration on the compass.
N.B.: Figure 2.21 is descriptive only. It should be noted that in reality the
suspended magnets would not appear as shown, but would be aligned with
the E±W cardinal points of the compass card.
It is essential that the magnet system of the compass shall point firmly along
the magnetic meridian toward the north magnetic pole, and this is achieved
by using magnets of sufficient pole strength. To ensure that they continue so
to do when the aircraft heading changes, friction at the pivot point is
minimised by using a jewelled bearing with an iridium pivot. Furthermore,
compass card
Figure 2.21
Magnet suspension system.
58 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
the compass system is suspended in a clear liquid, which lubricates the
When a compass is deflected away from its north seeking direction, it is
desirable that it will return to that direction as quickly as possible. If it tends
to oscillate about north for a significant period of time before once again
coming to rest it is said to be periodic. Ideally, it should come to rest with no
oscillations, when it would be said to be aperiodic. Aperiodicity is achieved in
aircraft magnetic compasses by three measures:
The interior of the compass case, known as the bowl, is filled with a clear
liquid, typically silicone or methyl alcohol. The compass card has filaments attached to it and these provide a damping effect in the liquid to
minimise oscillation.
Instead of a single powerful magnet, several small powerful magnets are
positioned close to the pivot point and this reduces the moment of inertia
of the magnet system.
The suspension of the magnet system in fluid helps support the apparent
weight of the magnet system, further reducing the moment of inertia.
Provision is made for the expansion and contraction of the compass liquid
by fitting a bellows, or in some cases a flexible diaphragm, inside the compass bowl. The construction of a typical aircraft standby compass is shown in
Figure 2.22.
compass card
Figure 2.22 Typical standby compass.
Turning and acceleration errors
We have already seen how the magnet system of the compass dips at
northern or southern latitudes, displacing the CG of the system from the
pivot point. This displacement has a profound effect upon the compass
magnet system during aircraft turns, accelerations and decelerations. Let us
Gyroscopic Instruments and Compasses
first consider the effect of acceleration and deceleration on a compass in the
northern hemisphere, using Figure 2.23 as a reference.
Acceleration error
Figure 2.23(a) shows the compass viewed from above, with the CG displaced
southward of the pivot point and the compass card markings simplified to
show only the cardinal points. In Figure 2.23(b) the aircraft is accelerating on
a westerly heading. Inertia causes the suspended magnet system to lag and,
because its CG is displaced from the pivot point the moment of inertia causes
the compass card to rotate anti-clockwise. To the pilot this presents a false
indication of a turn toward north. If the aircraft were to decelerate, the lag
would be in the opposite direction and the compass would swing to give a
false indication of a turn toward south.
In Figure 2.23(c) the aircraft is accelerating on an easterly heading and the
lag of the displaced CG causes the compass card to swing in a clockwise
direction. Once again, the effect as far as the pilot is concerned is that the
compass is falsely indicating a turn toward north. Deceleration on the same
heading would produce a false indication of a turn toward south.
Acceleration or deceleration on a northerly or southerly heading will not
cause false indications, since the CG and pivot point are aligned.
Summarising, in the northern hemisphere acceleration on an easterly or
westerly heading will produce a false indication of a turn toward north;
deceleration on those headings will produce a false indication of a turn
toward south.
In the southern hemisphere the errors are the opposite of the above.
Turning errors
Turning errors due to the displacement of the CG are maximum on a
northerly or southerly heading and are significant up to 358 on either side of
those headings. Let us consider the situation with an aircraft in the northern
Figure 2.23
Acceleration error.
60 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
hemisphere and turning through a northerly heading. Figure 2.24(a) illustrates the situation when a turn toward east is initiated.
As the aircraft turns, the centrifugal force acting upon the CG of the
magnet system creates a turning force on the compass card in the same
direction as the turn. The card should not turn at all, of course, but should
remain aligned with the magnetic meridian. The effect of the turning card is
to cause the compass to underread during the turn. A turn through north
toward west will have the same effect, the card turning in the same direction
as the aircraft turn and causing the compass to underread the turn. This is
illustrated in Figure 2.24(b).
Figure 2.24(c) shows the situation with the aircraft turning toward east
through a southerly heading. Now the centrifugal force acting on the displaced CG turns the compass card in the opposite direction to the turn and
the compass will overread during the turn. Exactly the same will happen
when a turn toward west is made through a southerly heading, as illustrated
in Figure 2.24(d).
aircraft turn
aircraft turn
aircraft turn
aircraft turn
Figure 2.24 Turning errors.
Gyroscopic Instruments and Compasses
In the southern hemisphere the turning errors are opposite to those
described above. The overall effects are summarised in Table 2.1.
Table 2.1
Turning errors.
Turn through
Turn toward
Compass error
Compass error
When turning through east or west there is no turning error, since the CG
and pivot point are in alignment. When turning with the magnetic compass
as reference it is necessary to roll out of the turn early (i.e. before the new
desired heading is reached) when it is underreading and late when it is
Variation and deviation
Referring back to Figure 2.20, it will be seen that the earth's true and magnetic north poles are not co-incident and there is, in fact, some considerable
geographic distance between the two. At most locations there will be an
angular difference between the magnetic north and true north, and this
difference is known as magnetic variation. On the shortest distance line
joining the true and magnetic poles, variation at any point is theoretically
1808, whereas elsewhere on a line joining the two, known as an agonic line,
variation is zero.
At locations where magnetic north lies to the east of true north, variation is
said to be easterly. At points where magnetic north lies to the west of true
north, variation is said to be westerly. The concept is illustrated in Figure
Magnetic variation at any point on the earth's surface can be plotted and is
shown on charts as a series of lines joining points of equal variation, known
as isogonals. The earth's magnetic field undergoes various changes which, in
the long term, cause the location of the magnetic poles to move. This
movement is reasonably predictable and usually appears on charts as annual
changes to variation, e.g. annual change 5'W.
62 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
Figure 2.25 Magnetic variation.
A compass needle will indicate the direction of magnetic north, provided
that it is only influenced by the earth's magnetic field. In an aircraft there are
many influencing local magnetic fields caused by hard or soft iron or electrical circuits. These will deflect the compass needle away from the direction
of magnetic north. The angular difference between compass north and
magnetic north is referred to as deviation.
The strength of the aircraft's magnetic field is usually reasonably constant
and its effect on the compass reading can be determined during a procedure
known as compass swinging. The aircraft is placed on a number of known
magnetic headings and the compass reading is compared with the aircraft
heading, to find the angular difference between compass north and magnetic north. The compass reading can be adjusted by means of a compensation device within the instrument until it agrees as nearly as possible with
the aircraft heading. It is usually impossible to adjust the compass so that it
exactly agrees on all headings and the remaining small angular difference is
known as residual compass deviation.
Upon completion of the compass swing the residual compass deviation is
recorded on a compass correction card, in either tabular or graphic form,
which is mounted in the cockpit. The correction card shows the deviation as
positive (+) or negative (7), indicating how it must be applied to the compass reading to obtain the correct magnetic heading.
When compass north lies to the east of magnetic north, deviation is said to
be easterly; when it lies to the west of magnetic north, deviation is said to be
westerly. Deviation easterly is positive and deviation westerly is negative.
Thus, if the correction card states that the compass has 718 residual
deviation on a heading of 1808(M), then the pilot must steer 1818(C) for the
aircraft to actually be on a heading of 1808(M). Alternatively, when the
Gyroscopic Instruments and Compasses
compass reads 1808(C) the aircraft's magnetic heading will be 1798(M). The
concept is illustrated in Figure 2.26.
The maximum permissible value for residual deviation of a standby
compass is +108.
Figure 2.26
Compass deviation.
Slaved gyro compass
The slaved gyro compass, also known as a remote indicating compass or
gyro-magnetic compass, embodies the best features of the directional gyro
and the direct reading compass.
The direct reading compass suffers from a number of disadvantages, in
particular its susceptibility to acceleration and turning errors. Because it
must be located at the position where it is to be read, e.g. the cockpit, it is
susceptible to aircraft magnetism and suffers from deviation.
The directional gyro is subject to apparent drift when it is at any location
other than that for which it is adjusted and consequently requires constant
The slaved gyro compass utilises the earth's magnetic field to sense
magnetic north and this magnetic flux is used to provide the constant
correction required by the gyroscopic element of the compass system to
maintain a magnetic north reference. The gyroscopic stabilisation reduces
the turning and acceleration errors inherent in the magnetic sensing ele-
64 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
ment and the device has the further advantage of being capable of
operating any number of remote compass indicators.
The fluxvalve
The magnetic detecting element of the slaved gyro compass is a device
known as a fluxvalve. This senses the direction of the earth's magnetic field
relative to the aircraft heading and converts this into an electrical signal. A
simple fluxvalve comprises two bars of highly permeable magnetic material
that are laid parallel to each other, each surrounded by a coil of wire, called a
primary coil, supplied with alternating current and connected in series.
Surrounding these is a secondary, pick-up coil. The general arrangement is
illustrated in Figure 2.27.
primary winding
secondary winding
Figure 2.27 Simple fluxvalve.
When a current is passed through a coil of wire a magnetic field is set up
around the coil and, if the current is alternating current, the field will be of
continuously varying strength and polarity. Because the primary coils are
wound in opposite directions, the fields surrounding them are of opposite
polarity and the permeable fluxvalve bars are magnetised with opposite
polarity by the primary fields. The alternating current is of sufficient
strength that, at its peak value, both bars are magnetically saturated.
A variable magnetic flux field cutting through the windings of the secondary coil would normally induce a current flow in them, but because the
primary fields are of opposite polarity and therefore self-cancelling, no
current is induced in the secondary coil.
However, if the fluxvalve is placed parallel to the earth's surface it will be
subjected not only to the electro-magnetic field due to the primary current
flow, but also to the horizontal (H) component of the earth's magnetic field.
This flux due to the earth's magnetic field will saturate the bars of the
fluxvalve before the alternating current supply to the primary coils reaches
Gyroscopic Instruments and Compasses
peak value and will cause a varying strength of flux around the primary
As this varying flux cuts through the windings of the secondary coil, it will
induce a current flow in the secondary windings. The strength of this current
is directly proportional to the strength of flux in the fluxvalve bars due to
earth magnetism. This will depend upon the direction of the bars relative to
the earth's magnetic field, which will be greatest when the bars lie parallel to
the earth's (H) field and weakest when they are at right angles to it. The
output of the secondary coil can be used to drive a remote indicating compass or, in the case of the slaved gyro compass, to apply corrections to the
gyro unit.
Detector unit
The detector unit contains the fluxvalve that senses the direction of the
earth's magnetic field relative to the aircraft heading. However, a simple
fluxvalve of the type described above would be inadequate, since it is prone
to ambiguity in that it will produce a signal of identical strength and polarity
on different headings. In order to overcome this anomaly an aircraft sensing
unit consists of three simple fluxvalves connected at a central point and
spaced at 1208 intervals, as shown in Figure 2.28. Each fluxvalve has a collector horn attached to its outer end to improve collection of the relatively
weak earth flux. The assembly is pendulously suspended so that it will
remain horizontal regardless of aircraft attitude and is mounted where it will
be least affected by aircraft magnetism, typically at the wingtip or near the
top of the fin (vertical stabiliser).
collector horn
Figure 2.28
Detector unit.
Remote compass indication
The detector unit is mounted such that it remains essentially earth horizontal, in order to best sense the H component of earth magnetism, and
fixed so that one `spoke' of the detector is permanently aligned with the
66 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
aircraft's longitudinal axis. Each spoke has its primary and secondary
windings, and the secondary windings are connected to the stator windings
of a device called a signal selsyn, as illustrated in Figure 2.29. The signal
selsyn comprises three stator windings, mounted at 1208 to each other,
surrounding a rotor upon which is wound a further coil.
earth magnetic
null-seeking rotor
(a) remote indicating compass before alignment
earth magnetic
(b) compass aligned ± rotor in null position
Figure 2.29 Remote compass indication.
The following is a description of the principle of operation of a remote
indicating compass system. Let us suppose that the aircraft is on a heading of
0008M. Coil A of the detector unit is aligned with the earth H field and
therefore maximum current is induced in it. Coils B and C have weaker
strength current induced in them and, because of the direction of their
windings, it is of opposite polarity. These currents are supplied to coils A, B
and C of the signal selsyn and the combination of their electro-magnetic
fields reproduces a flux field identical to that sensed by the detector unit.
Assume for the moment that the rotor of the signal selsyn is aligned with
this reproduced field. The lines of flux cutting through the windings of the
rotor coil will induce maximum voltage and current flow within the coil.
Gyroscopic Instruments and Compasses
This current flow creates an electro-magnetic field, which will seek to align
with the reproduced `earth' magnetic, causing the rotor to rotate until it is at
right angles to the reproduced field. When this is reached there will be no
voltage induced in the rotor coil and the motor will cease turning, with the
compass pointer indicating the heading as sensed by the detector unit. This
is known as the null point of the signal selsyn rotor. Attached to the rotor is a
compass pointer, indicating the aircraft magnetic heading.
Because the fluxvalves of the detector unit are at 1208 to each other their
combined secondary coil outputs are unique on every heading. This ensures
that the null-seeking rotor will always be at right angles to the selsyn
reproduced `earth' magnetic field.
As aircraft heading changes, the relationship of the detector unit fluxvalves to the earth's magnetic field will change with it. This field will be
reproduced in the signal selsyn and the rotor will turn to maintain itself at
right angles to the field, moving the compass pointer so that the heading
indication changes with the heading change.
In the case of the slaved gyro compass a signal from the null-seeking rotor
is used to precess a gyroscope to the correct magnetic heading reference. A
schematic diagram of a simple slaved gyro compass system is in Figure 2.30.
detector unit
gyro rotor
bevel gears
mechanical transmission
Figure 2.30
Slaved gyro compass system schematic.
68 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
Gyro compass operation
The gyroscope of the slaved gyro compass is basically similar to that of the
directional gyro. Its horizontal spin axis is mounted within an inner gimbal,
which is in turn pivoted to an outer gimbal that has freedom of movement
about the vertical axis. Attached to the inner gimbal is a permanent magnet,
as shown in Figure 2.30.
A precession coil is wound around the permanent magnet and is supplied
with direct current (d.c.) from the null-seeking rotor coil. When current
flows through the precession coil a magnetic field will be produced, the
polarity of which will depend upon the direction of the current flow. This
magnetic field will react with that of the permanent magnet, applying a force
to it that tends to tilt the spin axis of the gyroscope.
However, the gyroscope will precess this force through 908 in the direction
of rotation, which will cause the gyro to rotate about its vertical axis. It will
continue to do this until the current flow to the precession coil ceases, which
will happen when the null-seeking rotor is at right angles to the selsyn
reproduced magnetic field. Thus, aircraft heading changes will generate a
current flow from the null-seeking rotor, precessing the gyro to align itself to
the new heading.
The gyro is mechanically connected through bevel gears to the compass
display pointer and to the null-seeking rotor. This latter ensures that any
tendency of the gyro to drift will move the rotor from its null position,
causing a current to be generated in the rotor coil. This will be transmitted to
the gyro precession coil, precessing the gyro to keep it aligned with the
aircraft magnetic heading. Apparent drift due to earth rotation or transport
wander is thereby eliminated.
As with the directional gyro, the inner gimbal of the gyro compass gyro
must be maintained horizontal and to achieve this an erection system is
required. In this case the erection system consists of a torque motor mounted
on the outer gimbal, activated by a levelling switch mounted on the inner
Gyro compass errors
Since the detector unit of the slaved gyro compass is pendulously mounted,
it follows that during aircraft accelerations the unit will tilt and sense the
vertical (Z) component of earth magnetism. In order to minimise the errors
that would otherwise arise during turning and acceleration, the precession
rate of the gyroscope is deliberately kept slow. Thus, during an aircraft
acceleration the heading reference is maintained by the gyro, although the
rotor coil may be transmitting a false signal to the precession coil. Errors due
to detector unit tilt during a turn only become significant if a slow rate of
Gyroscopic Instruments and Compasses
turn is maintained for a significant period of time. During a normal turn the
error is very small due to the slow precession rate of the gyro. Deviation
errors are small in the slaved gyro compass because the detector unit is
mounted as far as possible from deviating magnetic influences. Compensation for those small remaining influences may be made by means of a
compensator unit that produces small electro-magnetic fields to oppose
those causing deviation.
Gyro compass outputs
Output from the slaved gyro compass may be used to supply magnetic
heading information to the radio magnetic indicator (RMI), the horizontal
situation indicator (HSI) of a flight director system, the autopilot system and
navigation systems such as INS and Doppler.
Sample questions
1. The two basic properties of a gyroscope are rigidity and . . . . . . . . . . .:
2. Concentrating the mass of a gyro rotor at the rim:
Increases the rate of precession?
Decreases its rigidity?
Decreases its rate of precession?
Reduces its angular momentum?
3. A free gyro is one that:
two gimbals and freedom of movement about three axes?
one gimbal and freedom of movement about two axes?
two gimbals and freedom of movement about two axes?
three gimbals and freedom of movement about three axes?
4. When the spin axis of a gyroscope deviates from its fixed reference due
to manufacturing imperfections this is known as:
Apparent drift?
Apparent wander?
Apparent topple?
Real wander?
70 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
5. The rate of apparent drift due to earth rotation is given as:
15 6 sin latitude 8/hr?
15 6 cos latitude 8/hr?
Sin latitude 8/hr?
6. An advantage of the ring laser gyro over a conventional gyro is that it:
Is cheaper to produce?
Has greater freedom of movement?
Has higher spin speed?
Is immediately available for use when switched on?
7. A directional gyro is most likely to suffer from gimbal error:
manoeuvring on headings other than cardinal headings?
flying straight and level on cardinal headings?
manoeuvring on cardinal headings?
in a steady climb or descent on a non-cardinal heading?
8. An aircraft is stationary on the ground at latitude 558N. The apparent
drift rate suffered by its directional gyro will be:
12.38/hr increasing?
8.68/hr increasing?
12.38/hr decreasing?
8.68/hr decreasing?
9. Compensation for apparent drift in the directional gyro is made by:
Eliminating manufacturing imperfections?
Adjustment of the latitude nut?
Imbalance of the outer gimbal?
Increasing the rotational speed?
10. The attitude indicator uses:
An earth gyro spinning at about 4500 rpm?
A free gyro spinning at about 12 000 rpm?
A tied gyro spinning at about 25 000 rpm?
An earth gyro spinning at about 15 000 rpm?
11. During acceleration a pneumatic attitude indicator will give a false
indication of:
a. Descending left turn?
Gyroscopic Instruments and Compasses
b. Climbing left turn?
c. Climbing right turn?
a. Descending right turn?
12. The erection system of an electrical attitude indicator uses:
a. Mercury switches on the inner gimbal and torque motors on the
outer gimbal?
b. Mercury switches on the outer gimbal and torque motors on the
inner gimbal?
c. Mercury switches and torque motors on the inner gimbal?
d. Mercury switches and torque motors on the outer gimbal?
13. The turn and bank indicator uses:
An earth gyro?
A free gyro?
A vertical gyro?
A rate gyro?
14. An aircraft is making a rate 1 turn through 1808. The turn will take:
2 minutes?
1 minute?
112 minutes?
3 minutes?
15. The principal constructional difference between the turn co-ordinator
and the turn and bank indicator is that:
a. The turn co-ordinator does not require bank indication?
b. The turn co-ordinator has two gimbals whilst the turn and bank
indicator has only one?
c. The longitudinal axis of the gyro gimbal is inclined at 308 to the
d. The turn co-ordinator display does not use an aircraft symbol?
16. The three basic requirements of a direct-reading magnetic compass are:
Horizontality, rigidity and sensitivity?
Sensitivity, rigidity and aperiodicity?
Horizontality, sensitivity and aperiodicity?
Aperiodicity, sensitivity and rigidity?
17. An aircraft in the northern hemisphere is accelerating on an easterly
heading. Its standby magnetic compass will:
72 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
Indicate a turn toward north?
Indicate a turn toward south?
Be unaffected on this heading?
18. An aircraft flying in the southern hemisphere is making a turn toward
east through a southerly heading. Its standby magnetic compass will:
Overread during the turn?
Underread during the turn?
Indicate a turn toward north?
Be unaffected on this heading?
19. An aircraft compass deviation card shows that the residual deviation on
a heading of 1358 is 28 easterly. In order to fly a heading of 1358M, the
pilot must steer a compass heading of:
20. The maximum permissible residual deviation for a direct reading
standby magnetic compass is:
21. The output signal of a fluxvalve is generated in:
primary windings?
permeable metal bars?
secondary winding?
collector horns?
22. The detector unit of a slaved gyro compass:
Is pendulously mounted, but fixed in azimuth?
Is pendulously mounted and free to move in azimuth?
Is free to move in azimuth because it is north-seeking?
Is designed to sense the Z component of earth magnetism?
23. The output of the null-seeking rotor of a slaved gyro compass will be
maximum when it is:
Gyroscopic Instruments and Compasses
At right angles to the signal selsyn field?
Aligned with the earth magnetic field?
Aligned with the signal selsyn magnetic field?
Aligned with the gyro heading?
Chapter 3
Inertial Navigation Systems
Inertial navigation systems are computer-based self-contained systems that
provide aircraft geographic position information in terms of latitude and
longitude, together with aircraft speed, heading and tracking information.
When provided with a TAS input, the system also produces an output of
wind velocity and direction. They require no external information or reference other than the starting location of the aircraft.
The basis of the inertial navigation system lies in measurement of the
aircraft's acceleration in a known direction and this is accomplished with the
use of accelerometers. These are devices that measure acceleration along a
specific axis; normally one measures accelerations and decelerations along
the east±west axis and a second measures accelerations and decelerations
along the north±south axis. Acceleration may be defined as increase of
velocity per unit time and is usually expressed in terms of metres or feet per
second per second (m/s2 or ft/s2).
If a vehicle, such as an aircraft, accelerates from rest or steady speed at a
constant rate over a given period of time, its final velocity and the distance
travelled can be calculated from simple formulae:
v ˆ u ‡ at
s ˆ ut ‡ 12 at2
where: v
final velocity
initial velocity
distance travelled
Since aircraft accelerations and decelerations are seldom constant it
becomes necessary to integrate each acceleration with respect to time in
order to obtain velocity and then to integrate the result of that with time in
order to obtain distance travelled. To achieve this, the outputs from the
accelerometers are fed to two integrators in series, as shown in Figure 3.1.
Inertial Navigation Systems
2nd integrator
1st integrator
Figure 3.1
Calculation of speed and distance travelled.
In order to determine position from the factors known, speed, distance
travelled and start position, it is necessary to know the direction of travel.
This is determined by virtue of the two accelerometers being aligned east±
west and north±south. Suppose, for example, that the north±south accelerometer/integrator combination has recorded a distance travelled of 60
nautical miles (nm) and the east±west accelerometer/integrator combination
has recorded zero distance travelled. Clearly the aircraft is now 60 nm, or 18
of latitude, north of its previous position. If both the east±west and north±
south accelerometers have recorded speed and distance, then the aircraft is
at some point at a known distance and in a calculable direction from its start
In order for the system to work, the accelerometers must only measure
aircraft accelerations and, to do this, they must be maintained earth horizontal at all times so that they do not measure the acceleration due to
gravity (9.81 m/s2 or 32.2 ft/s2).
Accelerometers can be maintained physically horizontal to the earth on a
gyro-stabilised platform called an Inertial Navigation System (INS). Alternatively, the accelerometers can be fixed to the aircraft axes, in which case
the accelerations due to gravity and aircraft manoeuvres are removed
mathematically from the accelerometer outputs. This system, called a
strapdown inertial system, is the basis of an Inertial Reference System (IRS).
We shall first study the gyro stabilised platform inertial navigation
Since the accelerometers are the heart of the inertial navigation system it
should come as no surprise that they use inertia to measure acceleration.
There are a number of different types of accelerometer, but the one most
commonly used in aircraft inertial systems is the pendulous force balance
If a freely suspended pendulum is subjected to acceleration it will lag, due
to inertia, in a direction opposite to that of the acceleration. The accelerometer incorporates a pendulous mass that is constrained so that it
responds only to acceleration or deceleration along its sensitive axis.
76 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
In the example illustrated in Figure 3.2 the pendulous mass is suspended
by two light leaf springs between two pick-off transformers. Ferrite armatures are attached to the ends of the mass and the transformer primary coils
are supplied with low voltage alternating current at a frequency of about
12 kHz. Whilst the accelerometer is not subjected to acceleration, the suspended mass is positioned exactly midway between the two transformers.
Under this condition the secondary voltage is identical in both transformers
and there is consequently no current flow in the secondary circuit.
primary circuit
secondary circuit
permanent magnet
leaf spring
force coil
permanent magnet
a.c. supply
Figure 3.2 Accelerometer principle of operation.
When the accelerometer is subjected to an acceleration or deceleration
along the axis of the pendulous mass, the mass is deflected by inertia,
reducing the gap between one armature and its transformer and increasing
the gap at the other end. This causes the voltages induced in the transformer
secondary windings to be unbalanced, with a resultant current flow between
This current flow is amplified and fed back to the force coil surrounding
the pendulous mass, creating an electro-magnetic field around the mass.
This field reacts with the field produced by the permanent magnets on either
side of the mass, to return the mass to its mid, or null, position. The
Inertial Navigation Systems
secondary current required to achieve this is directly proportional to the
acceleration that caused the deflection and it therefore serves as the output
signal from the accelerometer.
As previously stated, the purpose of the gyro-stabilised platform is to
provide a mounting for the accelerometers that is earth horizontal at all
times and that remains aligned with true north. The system uses rate integrating gyroscopes to sense platform horizontality and alignment. The
platform is mounted on gimbals which are controlled by pick-off signals
from the rate integrating gyros. Servo motors drive the gimbals to keep the
platform level and aligned irrespective of aircraft maneouvres.
Rate integrating gyro
The rate integrating gyros typically used are single degree of freedom (i.e.
single gimbal) gyroscopes that are highly accurate because the friction in the
gimbal is reduced to virtually nil. The gyro rotor is mounted within a canshaped gimbal, which is in turn mounted within a can-shaped case filled
with viscous fluid. Transformers at either end of the cans create a magnetic
field, upon which the inner gimbal is suspended, eliminating the need for
mechanical bearings.
The gyro rotor is a two-phase synchronous motor rotating at about
24 000 rpm and, because it is a single degree of freedom gyro, it is sensitive to
movement about one axis only, known as the input axis. Any movement
about this axis will cause the gyro to precess, rotating the gimbal within
which it is mounted. The gimbal is `floating' within the outer casing and the
relative movement between gimbal and casing is sensed by pick-off coils,
generating an output signal. A schematic diagram of a rate integrating gyro
is shown in Figure 3.3.
The gyro-stabilised (gimballed) platform
Three gyroscopes are mounted on the stable platform, fixed so that their
sensitive axes are aligned with north±south, east±west and earth vertical,
respectively. The platform is pivoted to inner and outer gimbals, as shown in
Figure 3.4. Each gimbal axis is connected to a servo motor; the two horizontal-axis motors are known as the pitch and roll motors and the vertical
axis motor is known as the azimuth motor.
The platform is shown in Figure 3.4 with the aircraft on a northerly
heading and with the platform levelled (i.e. earth horizontal) and
aligned with north. Under these circumstances the north gyro sensitive
axis is aligned north±south and the east gyro sensitive axis is aligned
east±west. Similarly, the north±south accelerometer pendulous mass axis
is aligned north±south and the east±west accelerometer mass axis is
78 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
outer gimbal
inner gimbal
Figure 3.3 Rate integrating gyro schematic.
roll motor
pitch motor
to aircraft
Figure 3.4 Gyro-stabilised platform ± aircraft on northerly heading.
Inertial Navigation Systems
aligned east±west. The sensitive axis of the azimuth gyro is aligned
with earth vertical.
Let us assume that the aircraft is stationary on the ground and accept for
the moment that the system will maintain the platform both level and
aligned. If the pilot now taxies the aircraft away from its stand, still on a
northerly heading, the north±south accelerometer will sense the northerly
acceleration as the aircraft moves away and the integrators will convert this
to speed and distance for the computer and its cockpit display. The east±
west accelerometer senses no acceleration and so the system continuously
computes distance travelled north from the starting point.
On arrival at the eastern end of the east±west runway (very convenient
airfield, this), the pilot turns the aircraft left onto a heading of 2708T. The
frame of the stable platform, being attached to the airframe, will have turned
through 908 with the aircraft, as illustrated in plan view in Figure 3.5.
Figure 3.5
Gyro-stabilised platform ± aircraft on westerly heading.
Any tendency of the stable platform to turn away from its north±south
alignment will be immediately sensed by the azimuth gyro, which will
precess and send an error signal to the azimuth servo motor. The platform is
therefore turned relative to the airframe to maintain north±south alignment.
The amount by which it is turned gives the change of aircraft heading for the
INS computer.
As the aircraft accelerates along the east±west runway on the take-off roll,
the acceleration is sensed by the east±west accelerometer and integrated to
show increasing velocity and distance travelled westward. The aircraft nose
80 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
is pitched up and this nose-up pitch is maintained during the climb, which
for simplicity we will assume continues on a westerly heading.
Any tendency of the stable platform to tilt in pitch will be immediately
sensed by the north gyro, since on this heading the pitch axis is coincident
with the north±south axis. The north gyro will send an error signal to the
pitch servo motor to correct the tilt until the north gyro signal is nullified.
Finally in this example, let us assume that the aircraft levels out at the top
of the climb at constant speed (i.e. no acceleration) and turns left onto a
heading of 2258T. During the turn the aircraft will, of course, bank and any
tilting of the stable platform will be sensed by both north and east gyros,
which will signal the pitch and roll servo motors to maintain the platform
level. North alignment during the turn is again maintained by the azimuth
gyro and azimuth servo motor, turning the platform relative to the airframe
and generating the heading change for the INS computer.
Whenever the aircraft is on a non-cardinal heading, control of platform
levelling is shared by the north and east gyros and the pitch and roll servo
motors. North±south alignment is at all times maintained by the azimuth gyro
and its servo motor. The situation is illustrated in plan view in Figure 3.6.
m itc
ot h
m roll
Figure 3.6 Gyro stabilised platform ± aircraft on non-cardinal heading.
Inertial Navigation Systems
Position calculation
We saw in Figure 3.1 how two stages of integration convert acceleration into
speed and distance travelled. The ultimate function of an inertial navigation
system is to provide the pilot with navigational data such as track,
groundspeed and present position in terms of latitude and longitude. The
block diagram in Figure 3.7 illustrates how this is achieved.
gyro–stabilised platform
1st stage integration
great circle
1st stage integration
accn. N–S d.t
= speed N–S
track and groundspeed
accn. E–W d.t
= speed E–W
2nd stage integration
great circle
speed N–S d.t
= distance N–S
track and distance
start latitude plus
change of latitude
= current latitude
departure x sec lat
= change of longitude
2nd stage integration
speed E–W d.t
= distance E–W
start longitude plus
change of longitude
= current longitude
current position
Figure 3.7
INS calculations.
Track and speed
The track calculated by the INS is a great circle track. The speed calculated at
any instant is groundspeed, because the INS bases its calculations upon
accelerations of the aircraft over the surface of the earth.
82 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
Latitude and longitude
Distance travelled north±south in nautical miles converts directly to change
of latitude, since each nautical mile along a meridian equates to one minute
of latitude. Distance travelled east±west is known as departure and is calculated using the equation:
Departure E-W …nm† ˆ change of longitude cosine latitude
Since, in the case of the INS computation, departure is known the calculation
carried out by the INS computer to determine change of longitude is
Change of longitude ˆ departure …nm† secant latitude
INS self-alignment
The gyro-stabilising platform is self-levelling and self-aligning, but these
functions can only normally be carried out in non-military aircraft with the
aircraft stationary on the ground. In order to compute the speed, distance
travelled and position of the aircraft the INS must first be referenced to north
at the current aircraft position, which is fed into the computer by the pilot
from airfield information. The process of self-alignment is performed in two
distinct phases; first the platform must be levelled so that the accelerometers
are not influenced by gravity and then it must be aligned with true north so
that they only sense horizontal accelerations on the north±south and east±
west axes.
Levelling and alignment are initiated by switching the INS to STANDBY
and inserting the present position in terms of latitude and longitude, and
then selecting ALIGN mode.
Conventional systems reduce the time taken to level the platform by using
gravity to achieve coarse levelling, followed by fine levelling using the
Coarse levelling
During coarse levelling the pitch and roll gimbals are typically driven by the
servo motors until they are mutually perpendicular and the platform is
brought to within about 18 of level using either gravity or reference to the
Inertial Navigation Systems
Fine levelling
With the aircraft stationary and the platform level there should be no output
from the accelerometers. If there is an output from either or both accelerometers this indicates that the platform is not level and that acceleration
due to gravity is being sensed. The output signals from the accelerometers
are used to torque the north and east gyros, which in turn use the pitch and
roll servo motors to drive the platform about the pitch and roll axes until
output from the accelerometers is zero.
However, it will be appreciated that a platform levelled to earth horizontal
using spatial referenced gyroscopes will not remain horizontal over an earth
that is rotating. From Figure 3.8 it will be seen that a level platform above the
poles would suffer drift about the vertical axis at earth rate (15.048/hr), and
one at the equator would suffer topple about the north±south axis at the
same rate. At intermediate latitudes the platform would drift and topple at a
rate dependent upon the latitude. To correct for this the north gyro is biased
for topple at 15.04 6 cos lat8/hr and the azimuth gyro is biased for drift at
15.04 6 sin lat8/hr.
Consequently, the platform will only be level when the output of the east±
west accelerometer is zero and the output of the north±south accelerometer
is also zero with the correct bias applied. The INS computer is able to apply
the correct bias, because current latitude is inserted before the levelling
procedure begins. Fine levelling typically brings the platform to within 6
seconds of arc of earth horizontal and takes about 112 minutes.
polar axis
platform remains
earth horizontal, but
drifts at 15°/hr
platform topples
at 15°/hr
earth rotation
Figure 3.8
Effect of earth rotation on stable platform.
84 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
The process of aligning the platform with the local meridian is usually
referred to as gyro-compassing or azimuth alignment and it is the final stage
of the alignment procedure. Most modern north-referenced gyro-stabilised
platforms use a system called open loop gyro-compassing. Since both earth
rate and the local latitude are known, any misalignment between the platform north±south axis and the local meridian can be calculated. When the
inertial navigation system is switched from the alignment mode to the
navigation mode the platform is rotated by the azimuth servo motor through
the computed misalignment angle.
Many earlier systems use what is called closed loop gyro-compassing to
achieve the same result. In this, any gravitational acceleration due to tilt of
the platform sensed by the north accelerometer after fine levelling results in
an output signal that is used to torque the azimuth gyro until the signal is
nullified, which will only occur when the platform north±south axis is
aligned with the local meridian. During this process the fine levelling
process remains operative.
With either system the gyro-compassing phase takes between 6 and 10
minutes and typically achieves an alignment accuracy of about 6 minutes of
arc. The time taken for self-alignment and the accuracy achieved increases
with latitude. Close to the poles it becomes impossible to align a northreferenced platform; in UK latitudes self-alignment usually takes between 10
and 15 minutes.
Mode selector panel
The various modes of operation of the stable platform INS are selected by
means of a simple panel, similar to that illustrated in Figure 3.9. The rotary
switch shown has five positions and there are two illuminated annunciators.
STANDBY (SBY) is the mode selected to switch the system on. In this mode
power is supplied to the system and the gyros are warmed and spun up to
operating speed. Whilst this is taking place it is usual to insert the current
position to the nearest 6 seconds of arc.
ALIGN selects the alignment mode, during which the levelling and alignment procedure described above takes place. When this is completed the
READY NAV annunciator illuminates, indicating that the INS is ready for
NAV is the navigation mode, used throughout the period of ground
movement and flight and during which the INS will make all its navigation
Inertial Navigation Systems
Figure 3.9
Mode selector panel.
calculations and display them as required on the control and display unit, to
be described shortly.
ATT REF is the attitude reference mode and is only used when the INS
computer fails to provide its navigational information. In this mode the
stable platform is used to provide heading, pitch and roll information.
The INS has its own internal battery, which is capable of supplying power to
the system for a limited period, typically about 15 minutes, in the event of
loss of the normal power supply. The BATT annunciator will illuminate red
when the battery power falls to a predetermined level, warning the pilot that
the INS is about to fail.
INS error corrections
It has already been shown that the gyro-stabilised platform is subject to drift
and topple due to earth rotation and that this is continuously corrected by
computation of earth rate for the current latitude. In addition to this, the
system is affected by transport wander and by coriolis effect.
Transport wander
As the platform is transported around the earth it will, if uncorrected, drift
and topple. This is because it is spatially referenced and it is necessary for the
INS computer to bias the rate-integrating gyros by an amount equal and
opposite to the drift and topple rates in order to maintain the platform level
and north-aligned. A biasing signal is applied to a torque motor on each gyro
output axis, causing the gyros to send correcting signals to the servo motors.
Coriolis effect
Any vehicle moving over the face of the earth is subjected to coriolis effect.
Put simply, coriolis effect is the combination of the vehicle movement and
earth rotation and it produces acceleration to the right in the northern
hemisphere and to the left in the southern hemisphere. As we know, any
acceleration is sensed by the INS accelerometers and, if uncorrected, would
86 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
result in a false computation of track, groundspeed, etc. Given that the INS is
continuously computing latitude, it is able to calculate and apply the
necessary corrections.
The Schuler loop
The periodic rate of a pendulum, that is the rate at which it swings, is
dependent upon the length of the pendulum. The longer it is, the longer it
takes to complete a full cycle of swing from centre to left, to right and back to
centre. Some considerable time ago a mathematician by the name of Schuler
calculated that a pendulum the length of the earth's radius would have a
periodic rate of 84.4 minutes. No doubt he had his reasons.
The relevance of this to the gyro stabilised platform is that it behaves as
though it were attached to a pendulum from the centre of this earth, by
virtue of the fact that it is maintained earth horizontal. Consequently, if the
platform is subjected to a disturbance, such as shock or vibration, it is apt to
oscillate with a frequency the same as a Schuler pendulum. The oscillations
will, of course, move the platform away from earth horizontal and the
accelerometers will produce an output that will be integrated into false
speed and distance signals.
The oscillations will cause the platform to move from level to a maximum
tilt (usually very small) over a period of 21.1 minutes, back to horizontal after
42.2 minutes, maximum tilt in the opposite direction after 63.3 minutes and
back to horizontal again after 84.4 minutes. The effect on the speed integration, for example, might be an error of +3 knots at 21.1 minutes and 73
knots at 63.3 minutes. At 0, 42.2 and 84.4 minutes there would be no error
and clearly the mean error over the full Schuler cycle is zero. This is known
as a bounded error, since it does not increase with time. However, the
positional calculations made from speed and distance will increase in error
with time and are therefore known as unbounded errors.
Calculation of magnetic north
The INS computer is programmed with the values of magnetic variation at
all geographic locations and can therefore apply this to the INS aligned true
north to obtain magnetic north, so that track, heading, etc., may be displayed
as either true or magnetic, as required.
Control and display unit (CDU)
A typical INS control and display unit panel is illustrated in Figure 3.10. The
unit comprises an alphanumeric light-emitting diode (LED) display, a rotary
selector switch and a keyboard for the insertion of data. In addition, there is
Inertial Navigation Systems
Figure 3.10
INS control and display unit.
the facility to set up a route by inserting waypoint latitudes and longitudes at
which track changes are to be made. Waypoints are identified in numerical
sequence using the selector wheel, starting with waypoint 1 as the departure
airfield. Each waypoint is entered using the keyboard to set its latitude in the
left alphanumeric display and its longitude in the right alphanumeric display. Insertion is made by pressing the INSERT button. If a change of track is
authorised en-route, say direct from waypoint 5 to waypoint 8, this is made
using the TK CHG pushbutton.
CDU selections
The displays associated with the various rotary display selector switch
positions are listed and explained in the following paragraphs:
TK/GS (track and groundspeed)
The INS computed track, usually referenced to magnetic north, is displayed
to the nearest tenth of a degree in the left display and the groundspeed in
knots in the right display. For example, a current track of 1358M and a
groundspeed of 467 knots would appear as 135.08 and 0467.
HDG/DA (heading and drift angle)
The heading obtained from the angle between the platform frame and north
reference is displayed to the nearest tenth of a degree in the left display. The
angular difference between heading and track (drift angle) is displayed to
the nearest tenth of a degree in the right display, preceded by the letter R or L
to indicate whether drift is right or left. Thus, a heading of 1378M on a track
of 1358M would be presented as 137.08 and L 02.08.
88 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
XTK/TKE (cross track distance and track error angle)
Cross track distance is the distance by which the aircraft is displaced right or
left of the desired great circle track and is displayed in the left display to the
nearest tenth of a nautical mile. The track error angle is the angular difference, right or left, between the desired great circle track and the actual track
being made, to the nearest tenth of a degree. If the aircraft were displaced
112 nm to the left of the desired track of 1358M, the left display would read
L 01.5. If the track being made good happened to be 1308M, the right display
would read L 005.08.
POS (present position)
The aircraft's current latitude and longitude are shown to the nearest 6
minutes of arc in the left and right displays, respectively. Suppose it happens
to be at 51.15.7N, 04.23.6W, this would appear as 51815.7'N in the left display
and 04823.6'W in the right display. This is the situation illustrated in Figure
WPT (waypoint positions)
The position of each inserted waypoint is shown as latitude in the left display and longitude in the right display by selecting WPT on the rotary
selector switch and scrolling through the waypoint numbers with the
waypoint selector wheel.
DIS/TIME (distance and time to the next waypoint)
The distance in nautical miles from the present position to the next waypoint
is shown in the left display and the time at present groundspeed to the
nearest tenth of a minute in the right display.
WIND (wind speed)
The INS is able to compute wind direction and speed and these are displayed in the left and right displays, respectively, to the nearest degree of arc
and knot.
DSR TK/STS (desired track and status)
The great circle track from one waypoint to the next changes as the aircraft
progresses between the two and the INS computes the present desired
magnetic track based upon distance from the waypoints, magnetic variation
and the assumption that the aircraft is on track. This will appear in the left
display to the nearest tenth of a degree and the right display will be blank.
The status function is for use only whilst the INS is in ALIGN mode and it
shows a numerical display in the right window that indicates the status of
the alignment procedure. The display typically shows 99 at the start of
Inertial Navigation Systems
alignment and counts down to 0, when alignment is completed and READY
NAV is illuminated.
Other CDU controls
It will be noted from Figure 3.10 that there are additional controls and
annunciators and their functions are as follows:
The waypoint selection controls are situated immediately below the left
LED display. A thumbwheel is rotated to select the number of a waypoint
that is to be inserted or amended. An LED display indicates the current
from/to situation; the illustration depicts the display that would appear
when flying between waypoints 2 and 3. The track change (TK CHG) push
button is used when altering the pre-planned sequence of waypoints. Conventionally waypoint 0 is the aircraft's current position. Suppose ATC has
cleared you to fly direct to, say, waypoint 5 then the TK CHG button is
pressed until 0±5 appears in the display. The DIM thumbwheel adjusts the
brightness of the LED displays.
The AUTO/MAN/RMT rotary switch is used to select the type of flight
control to be used in flying from waypoint to waypoint. In AUTO the INS
will automatically change the from/to display as each waypoint is overflown and would normally be used in conjunction with automatic flight. In
MAN (manual) the pilot is required to enter the from/to display as each
waypoint is reached. The RMT (remote) position is used when two or more
INS are fitted and enables the waypoint information to be transferred from
one INS to the other(s).
The three annunciators situated below the right LED display serve to draw
attention to specific events. The ALERT annunciator illuminates amber as
the aircraft approaches the next waypoint, typically when about 2 minutes
short of it, and will continue to flash until either cancelled by the pilot or,
when in AUTO mode, by overflying the waypoint. The BATT (battery)
annunciator will illuminate amber when the INS is operating on battery
power, reminding the pilot that the system will only operate for a limited
time on internal power. The WARN annunciator illuminates red in the event
of a system failure. At the same time the status display will show a number
that cross refers in the system manual to the nature of the failure.
Wander angle systems
It was mentioned earlier that the north referenced stable platform system
becomes unacceptably inaccurate at high latitudes and is therefore useless
for trans-polar flight. Because of this an alternative system has been developed in which the azimuth gyro is allowed to wander and the INS computer
90 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
transforms the wandering azimuth reference into any required earth coordinate.
Strapped-down systems
The disadvantages of the gyro stabilised platform INS are that the gyroscopes are expensive to manufacture and they inevitably suffer to some
extent from random wander, however small, due to manufacturing imperfections. These, of course, lead to inaccuracies in the INS output. The
gyroscopes take some time to warm up and reach their operating speed, and
platform alignment is a relatively slow process.
With the vast improvements in computer technology and the introduction
of the ring laser gyro it became possible to develop inertial systems that do
not require a stabilised, earth horizontal platform, but which can be
mounted directly to the aircraft structure. Hence the term `strapped down'.
These systems typically use three ring laser gyros with their sensitive axes
aligned to the aircraft roll, pitch and yaw axes.
Since the accelerometers are also attached to the airframe and move with
it, it follows that they will sense gravitational accelerations. The INS computer must differentiate between the accelerations that occur in the earth
horizontal plane and those that occur in the aircraft horizontal plane in order
to eliminate gravitational accelerations.
Alignment is achieved during an alignment phase with the aircraft stationary on the ground. The INS computer is able to discriminate between the
accelerometer outputs due to gravity, since aircraft attitude is fixed, and
those due to earth rotation and to compute the angle between the aircraft
fore and aft (roll) axis and true north.
In the navigation mode the INS computer receives inputs of aircraft
manoeuvres from the ring laser gyros and uses these to identify accelerometer outputs due to aircraft movement in the earth horizontal plane.
These are then related to the north±south, east±west and azimuth axes to
compute speed and direction of movement, distance travelled, current
position, etc. Corrections for earth rate, transport wander and coriolis effect
are computed in much the same way as previously described, from the
calculated angular differences between aircraft and earth horizontal and
between true north and the aircraft longitudinal axis.
Some strapped-down systems use an alternative type of ring laser gyro
that has four sides as opposed to three and some military systems incorporate a third accelerometer to measure vertical acceleration.
Fibre optic gyro (FOG)
Some manufacturers employ fibre optic gyros (FOGs) instead of ring laser
Inertial Navigation Systems
gyros. In these the beam of light is carried in fibre optics instead of the much
more expensive Cervit block required to transport the laser beam of the RLG.
Tuned rotor gyro
The tuned rotor gyro has a rotor that is, in its simplest form, flexibly suspended on a drive shaft so that it has limited freedom of movement relative
to the shaft. In other words, the gyro rotor axis of rotation may be misaligned
with the drive shaft axis by as much as 28.
At a specific speed of rotation the centrifugal force of the spinning rotor,
acting on the spring that flexibly attaches the rotor to the drive shaft, creates
a `tuned' condition in which the rotor behaves as though it were freely
suspended, i.e. it behaves as a free gyro.
These gyroscopes are much less expensive to manufacture than the rate
integrating gyro previously described and are generally more robust.
Sample questions
1. The output of the INS accelerometers is integrated . . . . . . with time to
obtain . . . . . . . . . and . . . . . . . . :
2. A typical gyro-stabilised platform contains:
Two accelerometers and three rate-integrating gyros?
Three accelerometers and two rate-integrating gyros?
Two accelerometers and two rate-integrating gyros?
Three accelerometers and three rate-integrating gyros?
3. With the aircraft on a northerly heading and the gyro-stabilised platform
properly levelled and aligned, acceleration will be sensed by the:
East gyro?
North gyro?
East±west accelerometer?
North±south accelerometer?
4. Change of longitude in minutes is calculated by the INS computer by:
a. Multiplying departure E±W by cosine latitude?
b. Multiplying departure E±W by cosine longitude?
92 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
c. Multiplying departure E±W by secant latitude?
d. Multiplying departure E±W by sine latitude?
5. The ALIGN mode of the INS mode selector:
Is only used with the aircraft stationary on the ground?
Is used when inserting waypoint information?
May not be used until the READY NAV annunciator illuminates?
Is selected when the platform has levelled and aligned?
6. The effect of earth rotation on the stable platform is corrected by:
a. Sending a biasing signal to the accelerometers?
b. Sending biasing signals to the east±west accelerometer and the north
c. Sending biasing signals to the north gyro and the azimuth gyro?
d. Sending biasing signals to the east gyro and the azimuth gyro?
7. The process of aligning the stable platform with the local meridian is
known as:
Coarse levelling?
Gyro compassing?
Fine levelling?
Attitude referencing?
8. The periodicity of a Schuler pendulum is:
42.2 minutes?
63.3 minutes?
21.1 minutes?
84.4 minutes?
9. An aircraft is displaced 212 nm to the right of its desired track of 2658M, on
a heading of 2728M. With XTK/TKE selected on the INS CDU, the left
and right displays would show:
R 02.5 R 007.08?
R 02.5 L 007.08?
L 02.5 L 007.08?
L 02.5 R 007.08?
10. A gyro stabilised platform INS that is not referenced to true north is
known as:
a. A drift angle system?
Inertial Navigation Systems
b. A non-levelled system?
c. A wander angle system?
d. A non-aligned system?
11. A strapped-down INS typically uses:
a. Three ring laser gyros with their sensitive axes aligned north±south,
east±west and earth vertical?
b. Three ring laser gyros with their sensitive axes aligned with the
aircraft lateral, longitudinal and vertical axes?
c. Two ring laser gyros and two accelerometers?
d. Three rate-integrating gyros with their axes aligned north±south,
east±west and aircraft vertical?
12. The advantage of a tuned rotor gyro over a rate integrating gyro is that:
Its speed of rotation is unimportant?
Its alignment is unimportant?
It has no moving parts?
It is less expensive to manufacture?
Chapter 4
Electronic Instrumentation
As the operation of transport aircraft and their systems has become
increasingly automated and, at the same time, the number of flight deck
crew members has been decreased, it has become impossible for the crew to
monitor and control the automated processes using conventional instrument
displays. To overcome this situation the traditional form of instrument has
been almost entirely replaced by computer-generated displays projected
upon a few cathode ray tube (CRT) screens. These screens usually combine
the features of a number of the conventional instruments and, especially in
the case of engine and system displays, normally only show the more
essential information, with less critical information being selected by the
pilots only as required.
The principal electronic display systems in use are the Electronic Flight
Instrument System (EFIS) for the presentation and control of navigational
information and, for presentation of engines and systems information, either
the Engine Indicating and Crew Alerting System (EICAS) or Engine Centralised Aircraft Monitoring (ECAM). EICAS is generally used in aircraft of
American manufacture and ECAM in Airbus Industries aircraft.
Electronic Flight Instrument Systems (EFIS)
The EFIS comprises two identical systems supplying the captain and first
officer with navigational information on two display screens each, mounted
one above the other. The upper display screen is an electronic attitude and
direction indicator (EADI) and the lower screen display is an electronic
horizontal situation indicator (EHSI). Each pilot's display has its own control
panel, and a symbol generator from which the electronic representations on
the screens are generated. A third symbol generator acts as a standby unit
and may supply either of the pilot's displays in case of failure.
Each symbol generator receives inputs from all navigational sources, both
internal and external, and interfaces between these inputs and the display
screens to present the information in a standard format. In addition, the
symbol generators perform the monitoring and control functions of the EFIS.
A block diagram of the EFIS process is shown in Figure 4.1.
Electronic Instrumentation
light sensor
light sensor
control panel
control panel
TMC (ctr) (left)
Figure 4.1
(ctr) TMC
TMC (ctr) (right)
EFIS operating schematic.
It will be seen from Figure 4.1 that both the captain's and first officer's
displays are provided with a control panel and a remote light sensor unit.
The control panel is used to control the EADI and EHSI displays and the
remote light sensor automatically adjust the brightness of the screen displays according to the light level on the flight deck. A typical EFIS control
panel is illustrated in Figure 4.2.
The electronic attitude and direction indicator (EADI)
The EADI screen displays aircraft attitude in pitch and roll in the conventional format of an artificial horizon divided horizontally, with the upper
half coloured blue and the lower half coloured yellow (or sand). The source
data for the attitude indication are the aircraft inertial reference systems. The
display also includes flight director command bars, ILS glideslope and
localiser deviation indications, and deviation indication from a selected
Radio altitude, decision height and operating modes of the automatic
flight and autothrottle systems are also displayed on the EADI screen.
Between 1000 ft and 2500 ft, radio altitude is displayed in digital format only,
but below 1000 ft above ground level (agl) the display changes to include an
96 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
EHSI range
setting knob
80 160
decision height
reference indicator
EHSI manual
brightness con
decision height
setting knob
decision height
reset button
weather radar
display button
EADI brightness
control knob
EHSI map mode
selector buttons
(select features to be added to MAP display)
Figure 4.2
EFIS control panel.
analogue, circular scale display as well. As height agl decreases the white
circular scale segments are progressively removed in an anti-clockwise
direction. The decision height (DH) can be set by a control knob on the EFIS
control panel and the selected DH is digitally displayed on the EADI screen
and a magenta coloured marker appears at the selected height on the circular
scale. At 50 ft above DH an aural chime begins to sound and its frequency
increases until DH is reached. At this point the circular radio altitude scale
and the DH marker both change colour to amber (yellow) and flash for
several seconds. This alert can be cancelled by a push button on the EFIS
control panel.
The display on both this and the EHSI screen is in colour and the colours
used follow conventions laid down in JAR Ops 25 as listed below:
Display features should be colour coded as follows:
Flight envelopes and system limits
Cautions, abnormal sources
Engaged modes
ILS deviation pointer
Flight director bar
Electronic Instrumentation
Specified display features should be allocated colours from one of the following colour sets:
Fixed reference symbols
Current data, values
Armed modes
Selected data, values
Selected heading
Active route/flight plan
Set 1
Set 2
The extensive use of yellow for other than caution/abnormal information is
In colour set 1, magenta is intended to be associated with those analogue
parameters that constitute `fly to' or `keep centred' type information.
Precipitation and predicted turbulence areas should be colour coded as
Precipitation (mm/hr)
0 to 1
1 to 4
4 to 12
12 to 50
Above 50
White or magenta
A typical EADI display screen is shown in Figure 4.3.
The electronic horizontal situation indicator (EHSI)
The lower of the two EFIS screens, the EHSI, presents a display of flight
navigational information and progress in one of nine possible modes,
selected from the HSI section of the EFIS control panel. The modes available
are as follows:
MAP. The display used for en-route navigational information and the
one most commonly selected in cruise flight. The display shows features
ahead of the aircraft, with the aircraft symbol appearing at the bottom of
the display. This is illustrated in Figure 4.4.
CTR MAP. Essentially the same as MAP, but the display is centred upon
the aircraft current position, with an aircraft symbol in the centre of the
PLAN. This display shows the planned route with waypoints and is
principally used when entering waypoints into the flight management
98 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
roll pointer (W)
roll scale (W)
decision radio
height (G) altitude (W)
speed (W)
GS 350
scale (W)
DH 100
display below 1000ft
flight director
command bars (M)
aircraft symbol (W)
scale (W)
glideslope deviation
pointer (M)
pointer (M)
status (G)
mode (G)
armed pitch
mode (W)
pitch mode (G)
AFDS status (G)
armed roll
mode (W)
engaged roll mode
pointer (M)
slip indicator
scale (W)
W white
G green
M magenta
Figure 4.3 Typical EADI display.
system (FMS) computer before flight or when making changes to the
planned route. It is illustrated in Figure 4.5.
FULL VOR/FULL ILS. These displays are basically identical and show a
compass rose with heading and deviation indications that follow conventional formats. They are used when checking aircraft track against a
VOR bearing or ILS localiser. The display with ILS selected is shown in
Figure 4.6. With VOR selected the display would be essentially the same,
except that the in-use VOR would be indicated in the lower left corner,
where ILS appears in the diagram.
EXP VOR/ILS. In the expanded mode the information displayed is the
same as in the full mode, but is in semi-map format. Only the relevant
segment of the compass rose is displayed at the top of the screen, with a
heading pointer. The aircraft's current (instantaneous) track is shown as a
solid line extending from the aircraft symbol to the compass arc. The
bearing of the selected radio aid, ILS or VOR, is shown as a solid line
extending from the centre of the deviation scale to the compass arc. The
display with ILS selected is shown in Figure 4.7. Again, the display with
VOR selected is essentially the same. In either case the weather radar
picture can be superimposed upon the display, if required.
EXP NAV/FULL NAV. These two modes display lateral and vertical
navigational information in much the same format as a conventional HSI.
Electronic Instrumentation
distance to
next waypoint (W)
magnetic track (W)
28.5 NM
heading pointer (W)
ETA next
waypoint (W)
expanded compass
rose (W)
track (W)
scale (W)
set bug (M)
(G, Y, R, M)
pointer (M)
range to
altitude (G)
scale (W)
vector (W)
wind speed
and direction (W)
symbol (W)
selected ILS
frequency (G)
lateral deviation
scale (W)
Figure 4.4
lateral deviation
pointer (M)
airport (C)
radio navaid (C)
off-route waypoint (C)
waypoint (active M, inactive W)
C cyan
W white
G green
Y yellow
R red
M magenta
EHSI MAP mode display.
Expanded NAV mode shows a compass arc, whereas full NAV mode
displays a full compass rose and does not permit the weather radar display to be superimposed, exactly as with the expanded and full VOR/ILS
Map mode display
Between latitudes 658S and 738N the expanded compass rose may be
referenced to either magnetic or true north as required, above those latitudes
it may only be referenced to true north. Heading information is provided by
the aircraft inertial reference systems; the heading and track pointers will, of
course, only be aligned when there is no drift.
The vertical deviation scale and pointer indicates whether the aircraft is
above or below the planned flight path and the lateral deviation scale and
pointer whether it is to right or left of the planned flight path. Wind speed is
indicated digitally in the lower left corner of the display, with an arrow
indicating wind direction. The arrow is orientated to the map display, which
100 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
85 NM
Figure 4.5 EHSI PLAN mode display.
DME 34.0
Figure 4.6 EHSI FULL ILS display.
Electronic Instrumentation
DME 27.5
Figure 4.7
EHSI EXP ILS display.
is orientated to the aircraft track, so the wind direction is displayed relative
to track.
The instantaneous track appears as a solid line extending from the apex of
the triangular aircraft symbol to the compass rose. The selected range scale is
superimposed and, when a planned change of altitude is taking place, an arc
indicating range to the next altitude (at present rate of change of altitude)
appears against this scale.
During change of heading a curved, dashed trend vector appears, showing predicted heading at the end of 30, 60 and 90 seconds from the present
time. The planned flight path appears as a solid line from the apex of the
aircraft symbol to the next, and subsequent, waypoints.
Navigational ground features, such as radio navigational aids and airports, are positioned on the display by the flight management system (FMS)
from data obtained from VOR/DME stations and aircraft position is based
upon these inputs together with those from the aircraft INS/IRS. When out
of range of VOR/DME coverage the aircraft position is updated from INS/
IRS only, but can be manually updated if required.
The weather radar display may be superimposed upon the MAP display,
with intensity of radar returns indicated by green, yellow and red in order of
intensity. Some displays will also indicate predicted turbulence, based upon
very high intensity returns, in white or magenta.
102 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
PLAN mode display
The PLAN mode display, illustrated in Figure 4.5, is principally for use
when entering or amending the lateral flight plan with the insertion of
waypoints. The active route, from waypoint to waypoint, is displayed on the
lower part of the screen, with significant navigational features such as airports and radio beacons included in their locations relative to the planned
route. It should be noted that, in this mode, the display is orientated to true
north. Wind speed and direction and the weather radar display cannot be
superimposed whilst in PLAN mode. At the top of the display the expanded
compass rose with track and heading as in the MAP display is maintained,
together with distance and time to the next waypoint.
This display is particularly useful as a checking medium when inserting
waypoint co-ordinates, before entering them into the flight management
ILS mode displays
In ILS mode there are two possible screen displays, full and extended. These
displays are primarily for use during landing approach. In the full ILS mode
a complete compass rose fills the central part of the screen, with an aircraft
symbol, deviation scale and deviation pointer indicating aircraft position
relative to the ILS localiser beam superimposed, as illustrated in Figure 4.6.
The compass rose, driven by input from the inertial reference system,
rotates against a fixed heading pointer as aircraft heading changes. The
heading, magnetic or true, appears digitally within the pointer. Selected
DME range is displayed in the top left corner of the screen, with wind speed
and direction in the lower left corner. The selected type of radio beacon, in
this case ILS, is also shown here, with the beacon frequency shown in the
lower right corner. A vertical deviation scale and pointer, to indicate aircraft
position relative to the glide path, is on the right side of the display.
With expanded (EXP) ILS mode selected, an expanded compass arc is
displayed at the top of the screen, as in MAP and PLAN modes, with aircraft
heading displayed in digital and analogue form as before. Selected DME
range appears in the top left-hand corner of the screen, whilst wind speed
and direction, and a reminder that the display is based upon ILS transmissions, appear in the lower left-hand corner. The selected ILS frequency
appears in the lower right corner.
A semi-map display fills the lower part of the screen, with the triangular
aircraft symbol positioned near the bottom of the display. A lateral deviation
scale with deviation pointer intersects the apex of the aircraft symbol and a
solid line extends from the central bar of the lateral deviation scale to the
compass arc, indicating the bearing of the localiser transmitter. A second
solid line, with a range scale, extends from the apex of the aircraft symbol to
the compass arc, indicating the current (instantaneous) aircraft track. When
Electronic Instrumentation
a new set heading is selected, a dotted line appears briefly from the apex of
the aircraft symbol to the heading bug. In ILS mode a vertical deviation scale,
with a pointer indicating aircraft position relative to the glide path, is on the
right-hand side of the screen. Weather radar returns may be superimposed
upon the display if required. The expanded ILS display is illustrated in
Figure 4.7.
VOR mode displays
Full and expanded VOR mode displays are essentially the same as the ILS
displays described above. These displays are useful when checking the
aircraft track and heading relative to a selected VOR/DME and the principal
differences in display features are that they will identify the selected beacon
and indicate whether the aircraft is tracking toward (TO) or away from
(FROM) the beacon. In VOR modes there is, of course, no vertical deviation
Engine indicating and crew alerting system (EICAS)
The EICAS system is an electronic display consisting of two CRT screens
mounted vertically, one above the other, and usually positioned centrally on
the cockpit console, where they are easily visible to either pilot. The displays
are capable of presenting all the engine and system operating data traditionally displayed by a mass of dials at a flight engineer's station, with
facilities for displaying a great deal more information besides. The upper of
the two screens, known as the primary display, normally shows only
essential (i.e. primary) engine information such as engine pressure ratio
(EPR), turbine spool speed (N1) and exhaust gas temperature (EGT). The
lower of the two screens, known as the secondary display, may be used to
display less important (secondary) information and details of abnormal
engine or system operating conditions.
The EICAS displays are generated by two computers that are continuously
receiving operating data from the engines and the various aircraft systems.
At any given time only one computer is operating the system, whilst the
other functions as a standby. A display selection panel enables the pilots to
select one of two operating modes, operational or status. A third mode,
maintenance, is available on the ground, specifically for use by maintenance
personnel. A block schematic diagram of a typical EICAS is shown in Figure
Operational mode
This is the mode in which the system is used throughout flight. In this mode
the upper screen displays the primary engine information listed above and
the lower screen remains blank so long as all engine and system operating
104 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
Figure 4.8
display select panel
engine sensors
system sensors
other systems
hyd press
hyd qty
hyd syst temp
fg cont surfaces
elect sys (I, V, F)
csdu temp
ECS temps
brake temps
oil temp
oil press
oil qty
rad alt
EICAS block schematic.
parameters are normal. In the event of an abnormal condition developing an
alert message will appear on the upper screen and the lower screen will
display details of the abnormal condition in analogue and digital format.
Figure 4.9 shows the upper screen display with conditions normal.
The EICAS display is colour coded, following the JAR Ops 25 codes
previously listed for EFIS displays.
When an abnormal engine or system operating condition develops, an
appropriate warning or cautionary message will appear on the left side of
the upper screen together with a row of pointers directing attention to the
lower screen, where an analogue and digital display details the nature of the
failure or critical condition. The alert messages relating to abnormal conditions are prioritised by the EICAS computer so that they appear in order of
importance and degree of crew response required. Warning messages,
requiring immediate corrective action, appear at the top of the screen in red.
Electronic Instrumentation
Figure 4.9
EICAS primary display.
Cautionary messages, requiring immediate crew awareness and possible
remedial action, appear below the warning message(s) in yellow (amber).
Both warning and cautionary messages are accompanied by an aural alert,
such as a fire bell or a repeated tone. Advisory messages, which only require
crew awareness, are indented on the display and also appear in amber. No
aural alert accompanies these messages.
An EICAS display with primary alerts and secondary display is illustrated
in Figure 4.10.
Status mode
This mode is primarily for use during preparation of the aircraft for flight
and shows the status of aircraft systems and their readiness for flight. The
information is allied to the aircraft minimum equipment list. The display
appears on the lower screen of the EICAS and shows flying control surface
positions in analogue format, with system status information in digital
message format. The quantity of information available is too great for a
single display and is available by selecting successive `pages'. The number of
the page being viewed is displayed on the screen. A typical status mode
display is shown in Figure 4.11.
Maintenance mode
This mode is available to maintenance engineers for diagnosis of operating
faults. It contains records of engine and system operating conditions and is
only available with the aircraft on the ground. A separate control panel is
provided for the display of maintenance data.
106 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
primary panel
warning message (R)
cautionary message (Y)
advisory message (Y)
secondary panel
Figure 4.10 EICAS primary and secondary display.
system states
1.00 0.97 0.99
3015 2995
EGT 498 RPM 98
secondary panel
Figure 4.11 EICAS status mode display.
Electronic Instrumentation
Display select panel
This panel, illustrated in Figure 4.12, is used to select the type of EICAS
display required and is usually situated on the centre console adjacent to the
display screens. The function of the various controls is as follows:
Display push buttons. When the engine display push button is depressed
secondary information appears on the lower screen. Depression of the
status push button selects the status mode referred to above.
Event record push button. Engine or system malfunctions in flight are
recorded automatically and stored in the EICAS computer memory.
Should the flight crew have reason to suspect that a transient fault has
occurred, depression of the event record push button will highlight
relevant data in the computer stored records for subsequent investigation
by maintenance personnel. This latter is known as a manual event.
Computer rotary switch. This is used to select the in-use computer of the
system. In the AUTO position the left computer will normally be in use;
switching to the right computer will occur automatically in the event of
failure. The LEFT or RIGHT positions are used for manual selection of the
in-use computer.
Brightness control. This is a dual rotary switch. The inner knob controls
display intensity and the outer knob controls the brightness balance
between the two displays.
Thrust reference setting. This is also a dual rotary switch. The outer knob
is used to select an engine and the inner knob is pulled and rotated to
position a cursor on the EPR or N1 circular scale.
Maximum indicator reset. If a measured parameter, such as engine oil
temperature, exceeds a preset limit an alert will appear on the EICAS
display. The maximum indicator reset push button is depressed to clear
the alert when the excess condition has been rectified.
Figure 4.12
EICAS display select panel.
System failures
It has already been explained that, in the event of failure of one computer,
the standby computer will take over either automatically or by pilot selection. If the lower display screen should fail whilst secondary information is
108 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
being displayed, the information is transferred to the lower half of the upper
screen in abbreviated, digital format. This is known as a compact display
and is illustrated in Figure 4.13. Failure of one display screen inhibits use of
status mode. Should both display screens fail, a standby engine indicator
displays essential engine performance data in a liquid crystal diode (LCD)
display, as illustrated in Figure 4.14. The indicator has a two-position control
switch. With this switch in the AUTO position the standby indicator is
functioning, but does not display any data unless the CRT displays are not
functioning. With the switch in the ON position the unit displays continuously. The test switch has three positions and is used to test the alternative power supplies to the indicator.
TAT -20C
1.7N 2
Figure 4.13 EICAS compact display.
Figure 4.14 Standby engine indicator.
Electronic Instrumentation
Failure of the display select panel is indicated on the upper EICAS screen,
which continues to display primary engine information. Secondary information still automatically appears on the lower screen, but the panel control
switches are inoperative.
Electronic centralised aircraft monitoring (ECAM)
The ECAM system was developed for the Airbus A310 aircraft and the block
schematic diagram in Figure 4.15 illustrates that version of the system. It is
principally an aircraft systems display medium, with primary engine
information displayed on traditional instruments. The ECAM display
screens are mounted side-by-side and both are in use continuously. The left
screen displays information covering systems status, warnings and corrective actions required in check list format. The right screen shows associated
information in analogue displays.
left hand display
right hand display
warning light
display unit
control panel
analogue inputs
system data
Figure 4.15
flight warning
flight warning
aircraft systems
data inputs
aircraft systems
data inputs
ECAM block schematic diagram.
Control panel
The ECAM control panel is illustrated in Figure 4.16. The left and right
display control knobs are for switching on the displays and adjusting the
display brightness. The functions of the various push button switches are as
110 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
SGU select switches
synoptic display switches
Figure 4.16 ECAM control panel.
SGU select. In normal operation of the system both symbol generator
units (SGUs) are functional. In the event of a fault being detected by an
SGU self-test circuitry, a fault caption is illuminated on the appropriate
switch. Releasing the switch isolates the affected SGU and extinguishes
the fault caption, illuminating the OFF caption in its place.
Clear (CLR). This is a clear switch, which will illuminate whenever a
warning or status message appears on the left screen. Depressing the
switch clears the message.
Status (STS). Depressing this switch allows manual selection of aircraft
system status displays, provided that there is no warning message displayed.
Recall (RCL). If a warning message is cleared whilst its associated failure
condition is still existent, it may be recalled by depressing the RCL push
Synoptic displays. Synoptic diagrams of each of the 12 aircraft systems
are called up on the right screen, provided that there is no warning
message displayed, by depressing the appropriate synoptic display
Operating modes
The system has four operating modes, known as NORMAL, ADVISORY,
FAILURE and MANUAL. Apart from the last-named, the display modes are
automatically selected.
NORMAL mode. This mode is flight-related and is the mode in which the
system normally operates throughout the flight progress from pre-flight
through to post-flight checks. In this mode the left screen displays system
states in check list format and the right screen contains a relevant pictorial
Electronic Instrumentation
display. For example, during post engine start checks the left screen
would typically list the state of each of the aircraft's systems and the right
screen would display the various system states pictorially (e.g. hydraulic,
electrical, pressurisation, etc.). These displays are selected by depressing
the appropriate push button on the ECAM control panel, illustrated in
Figure 4.16.
ADVISORY mode. The display automatically switches to this mode
when the status of a system changes. For example, starting the APU will
cause a message to that effect to appear on the left screen. The right screen
will continue to display the selected diagram.
FAILURE mode. This mode takes precedence over all others and is
automatically selected by the ECAM system in the event of normal
operating parameters being exceeded in any of the aircraft systems. An
appropriate warning message appears on the left screen, accompanied
by an aural alert. Below this message, the corrective actions required by
the flight deck crew are listed. On the right screen a diagrammatic display of the affected system illustrates the fault. When the corrective
action has been taken the displays change to illustrate this. Examples of
these displays are shown in Figures 4.17 and 4.18. In the example shown
the No. 2 generator frequency is outside permitted limits and disconnection of the constant speed drive is called for, with the situation
FREQ......................DISCONNECT CSD 2 DRIVE
115 V 400 HZ
Figure 4.17
115 V 420 HZ
FAILURE mode display before corrective action.
115 V 400 HZ
Figure 4.18
FAILURE mode display after corrective action.
112 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
displayed pictorially on the right screen. When the corrective action has
been taken, the left screen display changes to show that No. 2 generator
is disconnected and the right screen display shows the new status of the
a.c. electrical system.
MANUAL mode. Provided that there are no warning messages displayed
on the left screen, diagrams related to the aircraft systems can be called up
on the right screen by depressing any of the 12 synoptic push buttons on
the control panel.
Sample questions
1. The EFIS display comprises:
a. Two screens mounted side-by-side on a central console?
b. A pair of screens mounted one above the other in front of each pilot?
c. An upper screen showing the horizontal situation and a lower screen
showing attitude and direction?
d. A left screen showing navigational information and a right screen
showing engine and systems displays?
2. Symbols coloured magenta on the EFIS display indicate:
Warning information?
Command information?
Cautionary information?
Current situation information?
3. Below 1000 feet agl, the radio altitude display on the EADI is:
a. A digital only display?
b. An analogue display of a pointer moving around a circular scale?
c. An analogue and digital display with the selected decision height
displayed in digital format only?
d. An analogue display consisting of a circular scale in which the segments disappear as height decreases, with height agl displayed
4. At 50 ft above decision height:
a. An aural chime alert sounds and increases in frequency until DH is
reached, at which point the circular scale and DH marker both
change colour to amber and flash for several seconds?
b. An aural chime alert sounds and increases in frequency until DH is
reached, at which point the circular scale and DH marker both
change colour to red and flash continuously until cancelled?
Electronic Instrumentation
c. An aural chime alert sounds and continues until cancelled. At DH
the circular scale disappears and the DH marker changes colour to
d. An aural chime alert sounds for several seconds. When DH is
reached, the circular scale and DH marker both change colour to
magenta and flash until touchdown?
5. In MAP mode, a curved dashed line extending from the apex of the
aircraft symbol indicates:
a. Rate of change of aircraft heading?
b. Range to selected altitude?
c. Predicted heading at the end of 30, 60 and 90 seconds from preset
d. Deviation from desired track?
6. Weather radar returns are available in the following EFIS EHSI modes
(answer a, b, c or d):
3, 4, 5, 6?
4, 6?
7. The expanded compass rose in EFIS EHSI modes may be referenced to:
Magnetic or true north between latitudes 658S and 738N?
True north only above latitudes 638N or S?
Magnetic or true north between latitudes 638S and 758N?
Magnetic or true north between latitudes 758S and 758N?
8. In PLAN mode the planned route display:
orientated to
orientated to
orientated to
orientated to
the present aircraft track?
true north?
the present aircraft heading?
magnetic north?
114 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
9. A circle coloured cyan on the EFIS MAP display indicates:
An off-route waypoint?
An active waypoint?
An airport?
A radio beacon?
10. In the engine indicating and crew alerting system:
a. There are two computers, both of which are operating at all times?
b. There are two computers, but both will only operate simultaneously
if so selected manually?
c. There are three computers, two in operation and one on standby at
any time?
d. There are two computers, normally one is operating and the other is
on standby?
11. EICAS has:
Three modes, all of which are available during flight?
Three modes, of which only two are available during flight?
Four modes, all of which are available during flight?
Four modes, three of which are available during flight and the fourth
on the ground only?
12. EICAS advisory messages:
a. Appear on the primary display in analogue form, coloured yellow?
b. Appear on the secondary display in digital and analogue form?
c. Appear on the primary display in digital form, indented and
coloured red?
d. Appear on the primary display in digital form, indented and
coloured yellow?
13. In the event of failure of an EICAS display screen:
a. Primary and secondary information will be displayed on the
remaining screen in compact form?
b. Primary engine information only will appear on the standby engine
c. Primary and limited secondary information will appear on the
standby engine indicator?
d. Primary engine information only will appear on the remaining CRT
Electronic Instrumentation
14. Failure of the EICAS display select panel will:
a. Render the entire system inoperative?
b. Reduce the display to the standby engine indicator only?
c. Not affect the primary and secondary displays, but will be indicated
on the primary display?
d. Reduce the display to a compact version on one screen only?
15. The ECAM system has four operating modes:
a. The mode principally used during flight is known as the operational
b. The manual mode may only be selected on the ground?
c. The advisory mode may only be selected during flight?
d. The failure mode takes precedence over all other modes?
16. Warning messages on CRT displays are required to be coloured:
17. An armed mode on an EFIS EADI display will appear as:
magenta symbol?
white analogue feature?
white alphanumeric message?
cyan digital message?
18. Predicted turbulence will appear on the EHSI weather radar display as
an area of:
Red or cyan?
Red or magenta?
White or magenta?
Red or white?
Chapter 5
Automatic Flight Control
Automatic control of an aircraft in flight has developed from relatively
simple autopilot systems, in which the aircraft was automatically maintained on a steady heading, to complex systems that automatically control all
aspects of aircraft flight in terms of lateral and vertical navigation (LNAV
and VNAV) and speed from immediately post take-off to the end of the
landing roll and beyond.
To achieve this, the autopilot requires inputs from all navigational sources, both internal and external to the aircraft, and the engine thrust must be
managed at all flight stages for optimum power at optimum economy. The
co-ordination of all these requirements is achieved in modern transport
aircraft by the flight management system (FMS).
Flight management system
Almost all modern passenger transport aircraft employ a computerised
flight management system, the purpose of which is to reduce crew workload, thereby enabling a reduction in crew numbers, and to achieve the best
possible fuel economy with the overall result that operating costs are
The system may function purely as an advisory unit, providing the flight
crew with advice on the control settings required to achieve optimum fuel
economy in each of the various flight conditions such as take-off, climb,
cruise, descent and approach. In most cases this is done by displaying the
necessary engine pressure ratio (EPR) or torque settings and the recommended climb/descent rates. In this type of flight management system all
the control adjustments are made manually by the crew in response to the
FMS advisory messages. Associated with this type of FMS will be a discrete
flight director system, providing the flight crew with command and advisory information regarding the vertical and lateral flight path of the aircraft
through an attitude director indicator (ADI) and horizontal situation indicator (HSI). The flight director system will be described later in this chapter.
More sophisticated flight management systems interface with the aircraft
automatic flight control and automatic thrust control systems to achieve
fully automatic control of all flight phases with the exception of take-off and
Automatic Flight Control
initial climb-out. In these systems the management of the aircraft planned
flight path is divided into lateral and vertical profiles. The FMS guides the
navigation of the aircraft vertically (VNAV) to achieve the planned altitude
at each stage (waypoint) of the planned flight and laterally (LNAV) to arrive
overhead each geographical turning point (waypoint) of the planned flight.
At each stage of the flight the FMS instructs the automatic thrust control
system as to the power setting necessary to achieve maximum fuel economy.
A typical flight management system profile is shown in Figure 5.1.
departure procedures
arrival procedures
(APPR, STAR, TRANS) approach
en-route procedures
lateral flight plan
vertical flight plan
Figure 5.1
Typical FMS flight profile.
In addition to the control functions performed by the FMS, it continuously
provides information to the flight deck displays such as EFIS, EICAS or
ECAM, as described in the preceding chapter. Flight director commands to
the flight crew, particularly necessary during take-off and initial climb-out,
are transmitted through the EADI and EHSI functions of the EFIS, as we
have already seen to some extent.
In order to perform its multitudinous functions, the FMS must be provided with navigational data from all the navigation systems such as INS/
IRS, DME, VOR and Doppler and from all the engine and associated systems
monitoring equipment.
Since the term IRS has appeared in this text, it is perhaps appropriate at
this stage to differentiate between an inertial navigation system (INS) and an
inertial reference system (IRS). The former, as described in Chapter 3, is a
stand-alone navigation system that does not interface with other systems.
The IRS is essentially the same insofar as it performs basically the same
118 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
navigational functions, but its computer system is considerably more
sophisticated to permit interfacing with the flight management system.
Figure 5.2 shows a block schematic diagram of the data inputs and outputs
of a typical flight management system.
control and
display unit
weight and
Figure 5.2 FMS data interfacing.
Multi-purpose control and display unit (MCDU)
The link between the flight crew and the FMS is the multi-purpose control
and display unit. It provides the crew with the means to make inputs to the
system in order to obtain required displays, or information to assist with
decision-making in respect of the aircraft's flight progress. The CDU display
is in paged, digital format on a small, typically 2 6 3 inch, CRT screen.
The MCDU is primarily used for long-term actions, such as monitoring
and revising the flight plan, selection of operating mode and insertion of
data such as aircraft weight, wind speed and direction, various temperatures
and performance data. It provides the flight management computer (FMC)
with readout capability, together with verification of the data entered into
the computer memory. Flight plan and advisory data are continuously
available for display on the MCDU.
The MCDU panel has a full alpha-numeric keyboard, along with mode,
function and data entry keys. The keyboard includes advisory annunciators,
display light sensors and a manual brightness control. A typical multipurpose control and display unit is illustrated in Figure 5.3.
Automatic Flight Control
left field
line select keys
line select keys
title field
right field
numeric keys
Figure 5.3
function and
mode keys
alpha keys
Typical MCDU.
Display screen
The display screen of the MCDU shown in Figure 5.3 has 14 lines with a total
of 24 characters per line. The page format of the screen is divided into four
areas, these are:
Title field, which contains the title of the page of subject data displayed
and the page number (e.g.1/2, meaning page 1 of 2).
Left and right fields, each containing six pairs of lines of 11 characters per
line. The pilot has access to one line of each pair through the line select
keys on either side of the unit.
Scratchpad, which forms the bottom line of the display screen. Scratchpad
entries may be pilot-inserted, unless an FMC originated message is displayed in this field, and they are independent of the page displayed.
Line select keys
Momentarily depressing a line select key affects the line adjacent to the key
on the respective side of the MCDU for entry, selection or deletion of data.
Brightness control
The light intensity of the MCDU display may be adjusted by rotating the
BRT knob. Brightness of the illuminated keys is automatically adjusted by a
remote flight deck control.
120 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
There are two annunciators on each side of the keyboard. These display the
following messages when appropriate:
DSPY. Display (upper left). The white display light illuminates when the
active lateral or vertical leg performance mode is not displayed on the
current MCDU page.
FAIL. Fail (lower left). The amber light illuminates when there is a fault in
the FMC.
MSG. Message (upper right). The white light illuminates to indicate to the
pilot that an FMC-generated message is displayed on the scratchpad, or is
waiting to be displayed when the scratchpad is cleared.
OFST. Offset (lower right). The white light is illuminated when lateral
navigation (LNAV) is based on a route parallel to, but offset from, the
active route.
Alpha-numeric keys
These keys allow the pilots to enter letters and numbers onto the scratchpad
successively from left to right. They include space (SP), delete (DEL), slash
(/), and plus/minus (+/7) keys.
Function keys
The function keys control the MCDU field displays, accomplished by the
execution of pilot inputs and requests. The purpose of each of the function
keys is briefly described below:
EXEC. The execute key is the command key for the FMCS. Whenever a
modification or activation is pending, a white light bar illuminates.
Depressing the key will activate the flight plan, change the active flight
plan or change the vertical profile, as appropriate.
NEXT PAGE. Depressing this key causes the display to page on to the
next higher page number in multi-page displays.
PREV PAGE. Depressing this key causes the display to page back to the
next lower page number in multi-page displays.
CLR. The clear key extinguishes the MSG annunciator light and clears any
message from the scratchpad. Where more than one message is displayed,
each momentary push clears a single message; multiple messages are
cleared by repetitive momentary pushes or a single long push.
DEL. Pressing this key inserts the word DELETE onto the scratchpad,
provided that the pad is clear. Line selection by means of a line selection
key deletes the entered data on that line, but is only available for specific
Automatic Flight Control
Mode keys
The mode keys control the type of page displayed on the MCDU and are
therefore the means by which the pilots gain operational access to the flight
management system. There are twelve keys, as described below:
INIT REF. (Initialisation/reference.) The initialisation/reference key
selects the first of a series of pages used to initialise the position of the
FMCS and the inertial reference system (IRS).
RTE. (Route.) The route key provides access to planned routes and selects
the page for entering or changing the point of departure, destination or
DEP ARR. (Departure/arrival.) Depressing this key calls up an index
listing all terminal area procedures.
ATC. (Air traffic control.) This key selects the ATC automatic dependent
surveillance status page.
VNAV. (Vertical navigation.) Depressing this key provides access to the
climb (CLB), cruise (CRZ) and descent (DES) pages for evaluation and
FIX. The fix key provides access to the fix information pages, which are
used for creation of waypoint fixes.
LEGS. The legs key provides a page for evaluating or modifying lateral or
vertical details of each route leg.
HOLD. The hold key calls up the page for entering, exiting or amending a
holding pattern.
FMC COMM. (Flight management computer communications.) In most
current systems this key is non-operational.
PROG. (Progress.) This key is used to select current dynamic flight and
navigation data, such as ETAs and fuel remaining at a given point (e.g.
next two waypoints, destination or alternate).
MENU. The menu key provides access to other aircraft subsystems and to
the alternate control for the EFIS and EICAS control panels in the event of
NAV RAD. (Navigational radio.) Depressing this key selects the page for
monitoring or modifying navigational radio tuning.
Figure 5.4 shows a typical MCDU page display.
The MCDU is duplicated so that each pilot has access to the system and
the two units are usually located on either side of the central console.
Operation of the FMS is fully described in the Aircraft Operating Manual for
each aircraft type.
122 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
Boeing Aircraft
Figure 5.4 Typical MCDU display.
Flight director systems
A flight director system (FDS) integrates the display of aircraft attitude, in
terms of roll and pitch, with radio navigational data to provide a complete
directional command function for both vertical and lateral navigation. The
components of a typical FDS are shown in block schematic form in Figure
5.5. The vertical gyro unit referred to in the diagram is similar to the attitude
indicator described in Chapter 2, but instead of providing attitude information directly to a display it transmits attitude-related electronic signals to
the flight director computer. In aircraft not equipped with a flight management system and EFIS the outputs from the computer are fed to an
attitude director indicator (ADI) and a horizontal situation indicator (HSI).
Attitude director indicator (ADI)
The attitude director indicator display closely resembles that of the gyro
attitude indicator, but with vertical and lateral deviation indicators and a
radio altitude readout added. A fixed aircraft symbol, typically in the form
of a flattened triangle, is positioned centrally against a background tape that
is able to scroll up or down to indicate aircraft pitch attitude. The tape is
coloured to represent sky and ground, with a horizontal dividing line
representing the horizon. The tape is driven by a servomotor, which receives
signals from the pitch channel of the vertical gyro unit.
The background tape and its roller drive are mounted upon a ring gear
and this is rotated by a second servomotor that receives signals from the roll
channel of the vertical gyro unit, to indicate roll attitude. The tape has +908
freedom of movement in pitch and 3608 in roll. Bank angle is indicated by a
pointer attached to the ring gear, which moves against a fixed scale on the
instrument casing.
Automatic Flight Control
Figure 5.5
magnetic heading
reference system
magnetic heading
reference system
Flight director system block diagram.
Deviation from the ILS localiser beam is indicated by a pointer and lateral
deviation scale at the bottom of the display. Below this is a conventional `ball
and tube' slip and skid indicator. Deviation from the glideslope is indicated
at the left-hand side of the display by a pointer and vertical deviation scale.
A radio altitude readout of height agl is displayed, typically during the last
200 ft of descent. This may take the form of a scrolling digital readout, as in
Figure 5.6, or a `rising runway' directly beneath the fixed aircraft symbol.
It will be noted that the ADI includes command bars. These are driven in
response to signals from the flight director computer and they indicate pitch
and roll commands to the pilot. The two bars are not physically connected to
each other, but they normally move as a pair and the aircraft must then be
flown to position the triangular aircraft symbol in the `vee' formed by the
bars, in order to satisfy the command. The principle is illustrated in Figures
5.7 and 5.8.
In Figure 5.7 the command bars have moved to demand `fly up' and `fly
left'. In order to satisfy the flight director command the pilot must therefore
pitch the aircraft up into a climb and bank it to the left. Figure 5.8 illustrates
124 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
bank indicator
and scale
background tape
horizon bar
glideslope deviation
indicator and scale
radio altitude
command bars
aircraft symbol
localiser deviation
indicator and scale
slip and skid indicator
Figure 5.6 Attitude director indicator.
Figure 5.7 Pitch and roll commands unsatisfied.
Automatic Flight Control
Figure 5.8
Pitch and roll commands satisfied.
the ADI display when these manoeuvres have been carried out and the
aircraft is satisfying the climbing left turn command.
Some ADIs are equipped with a pitch command facility with which the
pilot can select a given climb or descent in certain modes of operation of the
flight director system. This will position the command bars to indicate the
required pitch attitude.
Horizontal situation indicator (HSI)
The HSI presents a display of the lateral aircraft situation against a compass
rose. The compass rose is driven by a servomotor receiving signals from the
aircraft magnetic heading reference system. In aircraft fitted with an inertial
navigation system the servomotor may also receive signals from this,
enabling either magnetic or true heading to be selected. Aircraft heading is
indicated by a fixed lubber line on the instrument casing.
The aircraft symbol, in the form of a miniature aircraft, is fixed at the
centre of the display and points toward the aircraft heading lubber line.
Control knobs at the bottom of the display allow a VOR/localiser course and
a desired heading to be selected. Rotation of the course selection knob rotates
the course arrow to indicate the selected course on the compass rose. At the
same time, a digital course counter at the upper right of the display provides
a readout of the selected course. The centre section of the course arrow is a
movable lateral deviation bar, which is deflected left or right of the course
arrow to indicate deviation from the selected VOR radial or localiser centre
line. It thus indicates fly left or right to intercept the localiser beam or VOR
126 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
radial and the command bars of the ADI will move to direct roll in the
appropriate direction. A to/from arrow indicates whether the selected
course is to or from the VOR. The dots of the deviation scale represent
displacement of approximately 1148 and 2128 from the localiser centre line or
approximately 58 and 108 from the VOR radial. Once a course is set, the
course arrow rotates with the compass rose as aircraft heading changes. In
VOR mode of operation the deviation bar begins to indicate when the aircraft reaches approximately 168 from the radial; in ILS mode, movement of
the bar begins at approximately 48 from the localiser beam centre.
The heading select knob, at the lower left side of the display, is used to set
a desired heading. When it is rotated, a triangular heading bug moves
against the compass scale to indicate the heading selected. In the heading
mode of operation the ADI command bars will move to direct roll in the
appropriate direction until the desired heading is attained.
Vertical deviation from the selected ILS glideslope is indicated by a
pointer moving against a deviation scale. Range to a selected DME is displayed digitally at the upper left side of the HSI.
Figure 5.9 illustrates a typical HSI display.
lubber line
selected cour
DME range
heading bug
bar and
pointer and
Figure 5.9 Horizontal situation indicator.
Warning indications
In the event of a fault developing it is vital that the pilot should be aware that
the display may be in error. The command signals generated by the flight
director system are continuously monitored and, should they weaken to the
point that the information provided is unreliable, small red warning flags
appear at the relevant part of the ADI or HSI display.
Automatic Flight Control
The ADI has three warning flags to indicate failure of the ILS glideslope
receiver, the vertical gyro and the computer. These are labelled GS, GYRO
and COMPUTER, respectively. When the GS flag is activated it obscures the
glideslope pointer and deviation scale; the GYRO and COMPUTER flags
appear at the bottom of the tape display.
The HSI also has three warning flags, labelled GS, COMPASS and VOR/
LOC. The GS flag operates in the same manner as that on the ADI, whilst the
COMPASS flag is activated in the event of failure of signal from the magnetic
heading reference system (MHRS) and the VOR/LOC flag indicates failure
of the VOR or localiser signal.
Modes of operation
The flight director system may be operated in a number of different modes
and these vary with the various system manufacturers and aircraft types.
Similarly, the method of mode selection by the pilots varies between
systems, but is usually achieved through push button selector switches.
The basic operating modes common to most systems are described briefly
OFF. With the system switched off the command bars of the ADI are
retracted from view and the indicator is used purely for attitude reference.
HDG. With heading selected, the ADI command bars provide guidance in
roll to reach and maintain the compass heading selected with the heading
select control knob and indicated by the heading bug. Where a pitch
command facility is fitted this will be enabled in this mode.
VOR/NAV. The ADI command bars provide guidance in roll in order to
capture and maintain a selected VOR radial or ILS localiser beam and
lateral deviation is indicated on both ADI and HSI displays. Pitch command facility is also enabled in this mode.
GS. With glideslope selected the ADI command bars provide vertical and
lateral guidance to capture and maintain the ILS glideslope and localiser
beams. Lateral and vertical deviation indicators are activated on both
GS AUTO. This mode is basically the same as GS mode, except that
interception and capture of the glideslope is automatic once the localiser
beam has been captured.
ALT. In this mode the ADI command bars provide guidance in pitch to
maintain a preselected altitude.
APPR I. This mode is used when making an approach to a Category I ILS.
Lateral and vertical guidance for the capture and tracking of glideslope
and localiser beams is provided by the ADI command bars.
128 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
APPR II. Guidance is the same as for APPR I, but beam tracking is of a
higher standard to meet the requirements of Category II ILS.
GA. (Go-around.) After a missed approach using one of the approach
modes, this mode may be selected for the execution of a go-around procedure. The ADI command bars will order a wings level, pitch-up attitude. Once the required power and speed settings have been achieved,
HDG and IAS modes may be selected.
IAS. This mode is used when it is necessary to maintain a given airspeed
during climb-out or descent. The ADI command bars provide guidance in
VS. This mode is selected when guidance is required for a specific rate of
climb or descent (vertical speed). The ADI command bars provide
guidance in pitch.
MACH. This mode is for use at higher altitudes. Guidance is the same as
for the IAS mode.
Figure 5.10 shows an example of a mode selector panel.
Figure 5.10 Flight director mode selector panel.
Gain programme in approach mode
The system response is optimised during an ILS approach by means of a
programmer in the gain control section of the FD computer. Following
capture of the localiser and glideslope beams the pitch and roll signals and
the deviation signals are reduced to allow for the convergence of the beams
during descent. The programmer is selected automatically at predetermined
positions on the localiser and glideslope.
Lateral and vertical beam sensors
The task of these sensors is to supply input data to the FD computer to assist
with the task of stabilising and adjusting the aircraft attitude as necessary.
Disturbances about the aircraft's lateral and longitudinal axes are sensed
Automatic Flight Control
and signalled to the FD computer, which generates pitch and roll command
bar movements in response. The commands are computed to reflect the rate
of change of deviation due to disturbance.
Automatic flight control systems
The core of every automatic flight system is the autopilot. This is an autostabilisation system capable of maintaining the aircraft in stable flight about
one or more of the aircraft's three axes of roll, pitch and yaw. Early autopilot
systems were only designed to maintain wings level flight by controlling the
aircraft in roll, that is about its longitudinal axis, and such a system is still in
use today in some light aircraft. This is known as a single axis, or single
channel, autopilot providing lateral stabilisation.
Inner loop control
The single axis lateral stabilising autopilot employs a closed loop control
system in which the aircraft attitude in roll is controlled by operating the
ailerons through a servomotor, or actuator. Any change in roll attitude is
sensed by a rate gyro sensitive only to movement about the aircraft longitudinal axis. Such movement will cause the gyro to precess and this precession will be picked off and transmitted as an error signal to a controller,
which compares input signals with the `wings level' message programmed
in its memory. The error signal will indicate the rate and direction of
deviation from wings level and the controller will generate a correcting
output signal of corresponding amplitude and rate of change. This signal is
amplified and transmitted to the servomotor, which moves the ailerons in
the appropriate direction to arrest the roll and restore the aircraft to wings
level. The restoring movement is sensed by the roll gyro, with the result that
the error signal received by the controller diminishes as the wings approach
the level condition, until there is no error when the wings are once again
level, by which time the servomotor has returned the ailerons to their neutral
position. A simple block diagram of such a closed loop automatic control
system is shown in Figure 5.11.
(error sensing)
aileron position feedback
Figure 5.11
Single-axis autopilot system.
130 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
It will be appreciated that a system such as that described is only capable of maintaining the wings level, because the only information the controller has is the programmed desired condition, the actual condition as
sensed by the rate gyroscope and response feedback from the aileron
servomotor. It is incapable of maintaining a particular heading or
course, since there is no directional information input to the system.
Such a control loop is known as an inner loop, since it has no external
references. The `wings level' reference programmed into the controller
may be, and almost invariably is, adjustable by the pilot. By rotating a
knob, for example, on a control panel the desired roll attitude can be
biased to left or right and the controller will signal the servomotor to
move the ailerons and roll the aircraft until the required bank angle is
achieved. The autopilot loop will now maintain that bank angle until
the reference attitude is re-adjusted. Thus the pilot can control aircraft
heading through the autopilot system.
A second autopilot inner loop channel can be added, to provide automatic
control of the aircraft about the pitch, or lateral, axis. Once again, this is
purely an autostabilisation system that will maintain a preselected pitch
attitude. The reference pitch attitude may be adjusted by the pilot, as before,
in order that the autopilot maintains a selected aircraft attitude for level
flight, climb or descent.
An autopilot system that has both roll and pitch channels is known as a
two-axis autopilot and this is probably the most common form of autostabilisation system. A third, yaw channel is often provided in larger
transport aircraft, especially where stability about the yaw axis is a problem,
as is the case with many swept-wing aircraft.
Outer loop control
Few transport aircraft fitted with autopilots use them just for autostabilisation. Instead, the autopilot is supplied with inputs from external
sources, which direct it to control the aircraft laterally to maintain a preselected heading or course and to control it longitudinally to maintain a
given altitude or vertical speed. Once this is done, it is a relatively
straightforward progression to supply inputs, via the flight management
system, from the aircraft's navigational systems to direct the autopilot so
that it flies the aircraft along a preselected lateral and vertical flight path.
These external control inputs are known as outer loop controls and a block
diagram illustrating one channel of an autopilot system with inner and outer
loops is shown in Figure 5.12.
Listed below are some of the sources of outer loop control signals supplied
to the autopilot systems of a typical modern large transport aircraft:
Automatic Flight Control
inner loop
error sensing
pilot inputs
autopilot select
radio navigation
and other data
control surface
outer loop
aerodynamic feedback
Figure 5.12
Autopilot inner and outer control loops.
Central air data computer (CADC). Manometric data of airspeed, pressure altitude, vertical speed and mach number are supplied by the CADC
for control of the aircraft in pitch to select and maintain a given airspeed
(airspeed select and hold), select and maintain a given altitude (altitude
select and hold) or maintain a given vertical speed or mach number (V/S
or mach hold). These are all pitch modes.
Magnetic heading reference system. The aircraft magnetic heading is
selected and maintained by control about the roll axis, using inputs from
the MHRS. This is a roll mode, known as heading select and hold.
Radio navigation aids. The ADF, VOR and ILS localiser receivers provide
signals for lateral guidance of the aircraft. In these modes of operation the
autopilot uses control about the roll axis to achieve a required course and
they are consequently roll modes. The ILS glideslope receiver provides
signals for vertical navigation, through control about the pitch axis, during approach and landing phases. This is, of course, a pitch mode of
Flight management system. The lateral and vertical information programmed into the flight management computer, when supplied to the
autopilot systems, enables them to navigate the aircraft vertically and
laterally along a predetermined flight path. Through this source many
refinements are possible, such as pitch trim to compensate for fuel consumption to name but one. In this mode of operation (LNAV and VNAV)
the FMS will be coupled to the roll, pitch and (where fitted) yaw channels
of the autopilot.
132 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
Control surface actuation
Autopilot operation of the aircraft flying controls may be by hydraulic
power-operated actuators, as is usually the case in large aircraft, or by
servomotors otherwise known as servo-actuators. These are electrically
driven mechanical, hydraulic or pneumatic machines that move the control
surfaces by linear or rotary motion. In most cases the same servomotors
operate the control surfaces whether the signal controlling them is from the
autopilot or from the human pilot's controls.
In some aircraft control systems the servomotors are connected in series
with the pilot's controls. In this case, when the autopilot is engaged and the
flying control surfaces are moved, there is no corresponding movement of
the pilot's controls on the flight deck. Alternatively, the servomotors may be
connected in parallel with the pilot's controls and autopilot-initiated
movement of the control surfaces will be mirrored by the flight deck
Torque limiters
The stress loads imparted to the airframe when the control surfaces of a large
aircraft are deflected can be enormous. Consequently, large or rapid
deflections can conceivably induce structural loads beyond the design limitations of the aircraft, resulting in serious structural damage. In manual
control of flight this danger is largely averted by artificial `feel' devices,
which warn the pilot if excessive control deflection is attempted. Since the
autopilot system has no such `feel' it is necessary to introduce safeguards
into the control surface operation to prevent overloading of the airframe.
Torque limiting devices are inserted in the drive between servomotor and
control surface, which will either slip or disengage if the torque required to
achieve the attempted rate of deflection exceeds a preset limit. This device
has the added benefit of preventing a servomotor runaway (undemanded
operation of the servomotor) from more than slight deflection of the associated control surface, before disengagement occurs. The torque limiter
typically comprises a spring-loaded coupling and friction clutch.
Autopilot engagement
Before the autopilot is engaged, and control of the aeroplane is transferred
from manual to automatic, it is important that a number of conditions are
satisfied to ensure that the changeover occurs without hazard to flight. For
example, the trim of the aircraft must be set by the pilot to avoid any possibility of sudden attitude change when the autopilot is engaged. Similarly,
all power supplies to the autopilot system must be operational and a host of
Automatic Flight Control
operational parameters must be met. To ensure that it is impossible to
engage the autopilot until all requirements are satisfied a system of interlocks is interposed between the autopilot engage switch and its electrical
supply. These interlocks take the form of relays and switches that only close
when parameters are satisfactory. Since they are connected in series, they
must all close before the autopilot can be engaged.
Manual inputs
Mention has been made of the simplest form of manual manoeuvring input
to the autopilot, in the form of a rotary knob used by the pilot to bias the
inner loops to change the roll or pitch attitude of the aircraft. This type of
input control may still be found on the automatic flight system control
panel of many aircraft types. In modern transport aircraft, however, it is
much more common to apply roll and pitch manoeuvring inputs to the
autopilot by means of the control yoke. There are two methods of achieving
this, known as control wheel steering (CWS) and touch control steering
Control wheel steering. When the autopilot is engaged the pilots can
override it, without disengaging, by applying normal manoeuvring force
to the control wheel or column. Upon release of the control wheel the
autopilot will hold the aircraft at its new attitude and in some cases, if the
bank angle is less than 58, roll the aircraft wings level and hold the new
heading until a new automatic flight mode is set on the control panel.
Touch control steering. With this system a thumb switch on the control
column is depressed to disengage the autopilot whilst the pilot manoeuvres the aircraft. When the thumb switch is released the autopilot reengages to hold the aircraft at its new attitude until an automatic flight
mode is reselected.
JAR-OPS requirements for autopilots
Each approved autopilot system must be capable of quick and positive
disengagement by the pilots, so that it does not interfere with their control
of the aircraft.
Each system must have a means of indicating to the pilot the alignment of
actuating devices in relation to the controls they operate.
Manually operated control must be so positioned as to be readily accessible to the flight crew.
Each control wheel or yoke must have a quick release control on the side
opposite the throttles.
Attitude controls must operate in the sense and plane of motion to be
134 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
achieved and the direction of motion must be clearly indicated on, or
adjacent to, the respective control.
The autopilot system must not be capable of placing hazardous loads on
the aircraft or create hazardous deviations from the flight path under any
flight condition appropriate to its use.
When the autopilot system is integrated with other systems or auxiliary
controls there must be interlocks and sequencing to ensure proper
operation. There must also be protection against adverse interaction of
Armed and engaged modes of operation must be indicated to the flight
Automatic flight system
A typical automatic flight system for a passenger transport aircraft in the
medium to large range is made up of a number of component systems. These
include the flight management system, the flight director system and an
autothrottle system. The latter will be described in more detail later in this
chapter, but briefly its role is to maintain selected EPR and N1 conditions at
specific flight phases as directed by the flight management computer or as
set by the pilots on the autopilot flight director system (AFDS) mode control
panel. It will be recalled from Chapter 4 that the current automatic flight and
autothrottle status is displayed on the EFIS ADI and HSI display screens.
The AFDS mode control panel is usually situated on the cockpit coaming
beneath the windscreen and its function is to provide the pilots with control
of the autopilot systems, the flight director, autothrottle settings and altitude
alert settings. The design of the AFDS and its control panel will clearly vary
according to the size and performance of the aircraft in question. The type
described in the following paragraphs is typical of certain marks of Boeing
737, but its features and the general operating parameters are similar to those
of most passenger transport turbofan aircraft. This particular system uses
two completely independent autopilot systems to allow for system redundancy, but many modern aircraft employ three independent autopilots to
improve failure protection even further.
Figure 5.13 shows a diagram of the captain's half of the autopilot FDS
control panel.
The control panel
The system uses two independent flight control computers that, in automatic
flight, supply pitch and roll commands to the inner loops of the autopilot
systems. In manual flight control the computers position the command bars
on the captain's and first officer's ADI displays. Each pilot has a flight
director selector switch. When switched on, the ADI command bars will
arm switch
and indicator pushbutton
bank angle
select knob
altitude select
flight director
and indicator
Figure 5.13 AFDS control panel.
speed select
vertical speed autopilot engage
Automatic Flight Control
heading select
136 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
appear in certain command modes; when switched off, the command bars
will retract out of view. The various mode selector push button switches are
depressed for selection and will illuminate to indicate mode selection.
Depressing an illuminated switch will deselect that mode. The system will
only accept a new mode selection provided that it does not conflict with the
mode(s) currently in operation.
Engagement and disengagement of the autopilots is made with paddle switches, one for each autopilot. The paddles have three positions;
OFF disengages the respective autopilot, labelled A and B, CWS engages the autopilot but control of flight is by operation of the control
wheel and column and CMD is the position for full automatic flight control, enabling all the command modes and CWS operation as required.
In all flight phases other than approach (APP) only one autopilot may
be engaged at any one time, but approach mode requires both autopilots to be engaged for a fully automatic landing. Command modes may
only be armed or selected provided that at least one of the engage paddles is set to CMD and one or both flight directors are switched on. An
armed mode is one that has been selected, but will only engage when certain parameters are met.
Autopilot command modes
The following are brief descriptions of each of the command modes of the
system under discussion:
Vertical navigation (VNAV) mode. When this selector is depressed the
flight management computer commands the AFDS pitch control and
autothrottle to follow the selected vertical flight profile programmed into
the flight management system. The programmed climb and descent rates,
cruise altitudes, speeds and height limitations will be followed through
automatic selection of pitch attitude and thrust. With VNAV selected the
EFIS ADI will display VNAV PTH or VNAV SPD, depending upon the
phase of the planned flight and SPD, N1, RETARD or ARM for the current
autothrottle mode.
Lateral navigation (LNAV) mode. Engagement of LNAV mode causes
the flight management computer to command the AFDS roll control to
intercept and track the lateral route programmed into the flight management system from waypoint to waypoint. The programme includes all
flight procedures such as SIDs, STARS and ILS approach. LNAV will only
engage provided that there is a flight path programmed into the flight
management system computer. It will automatically disengage if the
planned track is not intercepted within certain criteria or if the HDG SEL
push button is depressed.
N1 mode. With N1 selected the autothrottle system positions the thrust
Automatic Flight Control
levers to maintain whatever limiting rpm is set on the flight management
computer for the current phase of flight.
Speed mode. With this mode selected the autothrottle system positions
the thrust levers to maintain the speed selected with the rotary speed
select knob and displayed on the AFDS control panel. The autothrottle
system will ensure that the selected speed is achieved without exceeding
N1 limits and will equalise N1 on both engines provided that it can do so
without exceeding 88 difference of thrust lever position.
Level change (LVL CHG) mode. In this mode automatic control of pitch
and thrust is co-ordinated for climb or descent to a preselected altitude at
preselected airspeed. Before engaging LVL CHG a new altitude is selected
with the rotary altitude select knob on the AFDS control panel and this is
displayed digitally in the appropriate window on the panel.
Heading select (HDG SEL) mode. A selected heading is made by rotating
the heading select knob on the AFDS control panel and is displayed
digitally in the HDG window. Depressing the HDG SEL push button will
send a roll command to the autopilot to intercept and hold the selected
heading. The bank angle during the turn can be controlled with the bank
angle select knob, which forms the outer perimeter of the heading select
Approach (APP) mode. With approach mode selected the AFDS is armed
to capture and hold the ILS localiser and glideslope. Only when this mode
is armed is it possible to engage both autopilots; at any other time moving
one autopilot paddle to CMD will automatically disengage the other. To
meet the requirements of a fail passive control system (to be explained
shortly), both autopilots must be engaged for completion of a fully
automatic landing sequence. In this mode the AFDS will command the
autopilots through the ILS descent, landing flare, touchdown and roll-out
phases. An autoland sequence is described later in this chapter.
Take-off/go-around (TO/GA) mode. The go-around mode is automatically armed when FLARE ARMED is annunciated on the flight mode
annunciator and/or EFIS display. Depressing the TO/GA selector push
button under these circumstances will engage go-around mode, whereupon the flight director will command a 158 pitch up attitude for a climb
on present track to a radio altitude of 400 ft. The autothrottle system will
simultaneously command the thrust levers to advance for go-around
N1 rpm. Once 400 ft radio altitude has been passed, other pitch and roll
modes may be engaged; prior to that both autopilots must be disengaged
if pitch or roll attitude is to be changed.
Altitude hold (ALT HOLD) mode. Selection of altitude hold mode will
either maintain the aircraft at the selected altitude or adjust the aircraft's
attitude until the selected altitude is attained, in either case by pitch
commands. If a new altitude is selected with ALT HOLD engaged, the
138 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
select push button will illuminate until the new altitude is reached.
Alternatively, the new altitude may be selected first and then ALT HOLD
engaged. With ALT HOLD engaged, LVL CHG, V/S and VNAV modes
are inhibited.
Vertical speed (V/S) mode. In this mode the flight director provides pitch
commands to maintain the selected rate of climb or descent and the
autothrottle system adjusts the thrust levers to maintain the selected
indicated airspeed. Engagement of V/S mode is annunciated on EFIS
and/or the flight mode annunciator and the present vertical speed is
displayed on the control panel, prefixed by + or 7to indicate rate of climb
or descent, respectively. The desired vertical speed is set by rotation of a
thumbwheel on the mode control panel.
Automatic landing (autoland)
For an automatic flight control system to be capable of automatic landing it
must meet certain criteria. As has already been stated, it must contain a
minimum of two independent autopilot systems and, in addition, it must
satisfy the following safety requirements:
The response of the system must be such that there will be no deviation
from the flight path in the event of external disturbance such as turbulence
or windshear.
Control system faults must be indicated to the pilot as a warning or alert.
Control system failures must not cause the aircraft to deviate from the
flight path.
The flight control system must have sufficient control authority to ensure
accurate maintenance of the flight path.
The effect of a servomotor runaway must be limited, such that safe
recovery by the pilot is not jeopardised.
The automatic flight control system must not prevent completion of the
intended landing manoeuvre in the event of a system failure.
The above criteria are met by incorporating redundancy in the flight control
system through duplication or triplication of the autopilot systems, so that a
single failure within the system has a minimal effect on the overall aircraft
performance during approach and landing. Depending upon the degree of
redundancy, the autoland system is classified as being either a fail passive
(fail soft) system or a fail operational (fail active) system.
Fail passive system
An automatic flight system is considered to be fail passive if there is no
significant deviation from the flight path, or out-of-trim condition, following
Automatic Flight Control
a failure within the system, but the landing cannot be completed under
automatic control. In simple terms it means that, if one of the autopilots fails,
the other will disengage (since two are required for completion of an automatic landing), but there will be nothing to prevent the pilot completing the
landing manually. It follows from this that an automatic flight control system incorporating two independent autopilots must be a fail passive system.
Furthermore, a self-monitoring system is essential to ensure that both
autopilots are in agreement at all times. These are the minimum requirements for the multiplex type of control system necessary to meet autoland
In the event of failure of either autopilot or the monitoring system during an
automatic approach, the approach will continue on one autopilot, but
automatic landing is no longer possible. The flight crew must take over
manual control and revert to category 1 minima for landing, either continuing
the landing or executing go-around procedures at decision height. The single
autopilot will disengage automatically at about 350 ft radio altitude.
Fail operational system
In order for a landing to be completed automatically, following a failure
within the system, it follows that there must be at least three independent
autopilots and two independent monitoring systems. A single failure in
either of these will render the system fail passive, but it still has sufficient
redundancy to meet the criteria for completion of an automatic landing. In a
fail operational system all the autopilots and self-monitoring systems must
be engaged for an automatic approach and landing.
The EFIS display indicates the number of engaged autopilots, with a
caption reading LAND 3 indicating three autopilots engaged and a fail
operational system, LAND 2 a fail passive system with two autopilots
engaged and LAND 1 a passive failure with automatic disengagement
pending and completion of the automatic landing impossible. At all other
flight phases only one autopilot may be engaged at a time.
In the case of fail operational systems there is a specified alert height,
determined by the performance of the aircraft and the automatic landing
system. Failure of a redundant autopilot or monitoring system above this
radio altitude will result in discontinuance of the automatic landing. If
failure occurs below alert height the automatic landing is continued on the
remaining autopilot, on the basis that manual reversion is more hazardous at
this late phase (typically below decision height for manual completion of
landing) than to continue in automatic control.
Automatic landing sequence
The sequence of events during an automatic landing is illustrated in Figure
5.14. The radio altitudes for the events during the final stages of the
Figure 5.14
AFDS remains
until disengaged
by flt crew
with selection
of reverse thrust
Automatic landing sequence.
flare mode
thrust lever
45 ft
2 ft/sec descent path
flare mode engaged
330 ft
nose-up trim
1500 ft
flare mode armed
off-line autopilots
glideslope and
localiser captured
140 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
Automatic Flight Control
approach to touchdown will vary according to aircraft size and performance, but the sequence is typical for most aircraft types.
During the descent from the cruise, approach mode is selected by
depressing the APP pushbutton and this arms the off-line autopilots; the
second in the case of a fail passive system and the second and third in the
case of a fail operational system. At the same time the ILS glideslope and
localiser channels become the armed pitch and roll modes.
The radio altimeter becomes effective at, typically, 2500 ft agl and provides
all height measurements for the automatic flight control system from then
until touchdown. At 1500 ft radio altitude, provided that the localiser and
glideslope beams have been captured, the off-line autopilots engage and
LAND 2 or LAND 3 is displayed on the autoland status annunciation,
depending on the number of engaged systems. The aircraft continues to be
flown by one autopilot, with the remainder performing a comparative
function, overseen by the monitoring system. If these sequences have been
satisfactorily completed, FLARE mode becomes armed and the glideslope
and localiser beams become the engaged pitch and roll command modes,
maintaining the aircraft on the glidepath centre line.
When the aircraft has descended to 330 ft radio altitude, the AFCS commands a nose-up trim adjustment, with pitch control being maintained
through the elevators. When the main landing gear is 45 ft above ground
level, as measured by the radio altimeter and adjusted to take account of the
height difference between the radio altimeter transceiver and the main gear,
FLARE mode engages and provides pitch commands. Roll commands are
still from the localiser, to keep the aircraft on the centre line of the glidepath.
The aircraft now follows a 2 ft per second descent path, rather than the
glideslope beam, and the autothrottle system begins retarding the thrust
levers to control airspeed for the touchdown.
Just prior to touchdown, at about 5 ft gear altitude, flare mode disengages
and touchdown and roll-out modes engage. At approximately 1 ft gear
altitude the AFCS commands a decrease in pitch attitude to 28 nose-up and,
at touchdown, the elevators are adjusted to lower the nose and bring the
nose wheels into contact with the runway. Selection of reverse thrust by the
pilot disengages the autothrottle system, but the AFCS remains in control of
the roll-out until disengaged by the flight crew.
Automatic thrust control (autothrottle)
The autothrottle system receives its commands from an autothrottle computer, which is linked to the flight management and flight control computers
and operates the thrust levers through servo-actuators. Its function is to
control the thrust in terms of engine pressure ratio (EPR), HP spool rpm (N1)
or the aircraft's flight speed. Its primary function is to operate in conjunction
142 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
with the automatic flight control system in its VNAV and approach modes,
to attain a required airspeed and to maintain the programmed vertical flight
path. The autothrottle system is armed by operation of a switch on the mode
control panel of the automatic flight control system and is controlled
through this panel during automatic flight. Figure 5.15 shows in block
schematic format the signal interfacing between the autothrottle computer
and other systems.
thrust lever
mode control
thrust lever
thrust lever
flight control
air data
landing gear
Figure 5.15 Autothrottle signal interfacing.
Operating modes
The autothrottle system operates in one of three possible modes: take-off,
speed control and go-around.
Take-off mode
Before commencing the take-off the flight management system is engaged
and its computer supplies the N1 limits for each stage of the flight profile,
together with a selected, or `target', N1 rpm. These values are displayed as
markers on the N1 indicators of the engine displays. Switching the autothrottle engage switch on the mode control panel to ARM will arm the
autothrottle system for take-off and this will be annunciated on the EFIS, or
other, display. The thrust lever servo-actuators are engaged by pressing
switches mounted on the thrust levers, known as take-off/go-around
Automatic Flight Control
switches. An example of these is illustrated in Figure 5.16. Once this has been
done, the servo-actuators advance the thrust levers at a preset rate in order to
reach the position for take-off N1 by the time a specific speed has been
reached on the take-off roll. For example, the advance rate for the thrust
levers might be 158 per second to ensure all engines have reached take-off N1
before the aircraft has reached a speed of 60 knots. When this target speed
has been exceeded by a preset amount, autothrottle movement of the thrust
levers is interrupted by a speed detection circuit and the levers are held at
their current position, a condition known as throttle hold (THR HOLD).
Should the speed detection circuit fail, a back-up system, activated by the
main landing gear `squat' microswitches, will operate to instate throttle hold
shortly after the aircraft lifts off. At a radio altitude of 400 ft the autothrottle
system arms to control N1 for the vertical profile of the remainder of the
flight and the automatic flight control system takes over control of the
autothrottle system.
select switches
Figure 5.16
Thrust lever go-around switches.
Speed control mode
Speed control mode is selected through the mode control panel of the
automatic flight control system, either by the pilot pressing the SPD push
button switch or automatically if the system is in other than a speed mode
(e.g. VNAV). In either case, the autothrottle system will command the thrust
lever actuators to adjust the levers until the IAS or mach No. selected has
been reached and held. The autothrottle system controls airspeed/mach to
maximum and minimum safe values, regardless of the selected airspeed/
mach, and it prevents the angle of attack (alpha angle) from exceeding a safe
value. Minimum safe airspeed and maximum safe alpha are computed from
144 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
data received from the flap position sensors and angle of attack sensors.
Under VNAV control mode the autothrottle system begins to retard the
thrust levers at the top of descent at, typically, 28 per second until they either
reach the idle stop or are arrested by pilot intervention. During retardation a
RETARD annunciation appears, followed by ARM when the retarding
movement ceases, to indicate that speed mode is armed. At glideslope
capture the AFCS mode changes from VNAV to approach (APP) and the
autothrottle engaged mode changes to speed, with the displayed speed
being that computed by the flight management system. The gain, or sensitivity, of the autothrottle system is increased for greater precision of speed
control during the approach. During the flare manoeuvre the thrust levers
are retarded at a rate computed to reach idle in 6 seconds. Immediately after
the landing gear microswitches have indicated touchdown, further thrust
lever retardation is initiated by the autothrottle system, until it automatically
disengages 2 seconds after touchdown.
Go-around mode
Depressing a take-off/go-around switch on one of the thrust levers with the
autothrottle system engaged and the aircraft below 2000 ft radio altitude will
initiate advance of the levers until they reach the position for reduced goaround thrust. The caption GA will immediately appear on the ADI and the
flight management system computer will calculate the full go-around thrust
rating determined by the present aircraft all-up weight and the density
altitude. A second depression of the thrust lever-mounted TO/GA switch
will now advance the thrust levers to increase engine thrust to the full goaround thrust rating. The AFCS will generate the pitch-up and wings level
commands necessary to establish the aircraft in the go-around climb-out.
Full flight regime autothrottle system (FFRATS)
This system performs all the functions described above and additionally
provides engine overboost protection and selection of variable engine rating.
The system monitors demands on the engines made by air conditioning
system and anti-icing system air bleeds and adjusts the engine pressure ratio
(EPR) limits to suit. Most large modern passenger transport aircraft are fitted
with an autothrottle system meeting the FFRATS specifications.
Thrust computation
In order to achieve maximum fuel economy and to prolong engine life,
advanced aircraft turbine engines utilise electronic engine control systems.
Full-authority electronic engine control systems receive data from the aircraft and engine systems to enable safe and efficient operation of the thrust
Automatic Flight Control
management system over the entire operating range of the engines. One
aspect of such a system is computation of the optimum and maximum thrust
requirement for every condition of flight. The computed total air temperature (TAT) and measured pressure altitude are used to compute the optimum and limiting engine pressure ratio (EPR) for the current flight phase.
EPR is the ratio of HP turbine exhaust pressure to LP compressor inlet
pressure and has been found to be directly proportional to the thrust
delivered by the engine.
The computed EPR for the current flight phase is presented on an indicator on the flight deck, which typically displays TAT, the current flight
mode (e.g. take-off, climb, cruise, etc.) and the EPR limit for that mode. The
actual EPR, with limit and target markers, continues to be indicated on the
engine monitoring display (e.g. EICAS). Examples of EPR indicators are
shown in Figure 5.17.
- 13
actual EPR
Figure 5.17
EPR indicators.
The flight mode for which EPR computation is required is selected by the
pilot through an EPR limit control panel and this is fed to the EPR computer.
Typical modes for EPR limit computation are climb (CLB), economy (CON),
cruise (CRZ), top-of-descent (TOD) and go-around (GA). When the automatic flight control system is in use, go-around EPR limit will automatically
display as the glideslope is captured. Additionally, the panel may contain
thrust rating selector switches, with which the pilot can command the
computer to calculate the EPR for specific engine performance ratings. The
system incorporates a test function for preflight testing. In the event of
system failure or electrical power loss, a warning flag obscures the EPR limit
146 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
Sample questions
1. The display screen of a flight management system multi-purpose control
and display unit:
a. Provides an analogue display of flight progress?
b. Provides an alpha-numeric display of flight plan data?
c. Is divided into two fields, the left field providing information in
alpha-numeric form and the right field providing the same information in analogue form?
d. Provides 24 lines of information, 14 characters per line?
2. The background tape of a flight director ADI normally has . . . . . . . . . . . .
freedom of movement in roll and . . . . . . . . . . . . freedom of movement in
3. The aircraft symbol on a flight director HSI:
Rotates with the compass rose?
Aligns with the selected VOR radial?
Aligns with the selected heading?
Is fixed at the centre of the display?
4. With VOR selected and the HSI lateral deviation bar displaced one dot to
the right, the aircraft is approximately . . . . . . . . . . . . from the radial:
5 nm?
5. The warning flags on a flight director ADI will indicate failure of
(answer a, b, c or d):
glideslope receiver
FD vertical gyro
FD computer
MHRS signal
VOR or localiser signal
a. 1, 2, 3, 4, 5?
b. 1, 3, 4, 5?
Automatic Flight Control
c. 1, 2, 3?
d. 3, 4, 5?
6. With the flight director in VOR/NAV mode, the ADI command bars
Guidance in pitch only?
Guidance in pitch and roll?
No guidance, since they are retracted in this mode?
Guidance in roll only?
7. A two-axis autopilot has:
A single inner loop and two outer loops?
Two inner loops and two outer loops?
Two inner loops?
One inner loop?
8. The data supplied to an autopilot system from the central air data
computer are known as:
Manual data?
Manometric data?
Monometric data?
Aerodynamic data?
9. The function of a torque limiter in the servo-drive of a flying control
surface is to:
Prevent excess rate of movement of the surface?
Prevent slip in the drive system?
Control the rate of movement of the surface?
Prevent over-torquing of the servo-motor?
10. Control wheel steering (CWS) is engaged by:
Rotating a knob on the AFDS control panel?
Applying normal manoeuvring force to the pilot's controls?
Operating a thumb switch on the control wheel?
Moving the autopilot engage paddles to OFF?
11. With the automatic flight director system in VNAV and LNAV mode,
engaging a second autopilot will:
a. Improve the sensitivity of the automatic flight system?
b. Automatically engage the autothrottle system?
148 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
c. Disengage the active autopilot?
d. Automatically change the control mode to approach mode?
12. LVL CHG, V/S and VNAV modes are inhibited when . . . . . . . . . . . . is
13. A fail operational automatic landing system:
a. Requires two independent autopilots and one monitoring system?
b. Is one which will not continue with an automatic landing in the event
of a single failure within the system?
c. Will abort the landing if a failure occurs below alert height?
d. Requires at least three independent autopilots and two independent
monitoring systems?
14. For an automatic landing to continue below 1500 ft, which of the following conditions must be satisfied (answer a, b, c or d):
Glideslope capture
Localiser capture
Radio altimeter serviceable
Off-line autopilots engaged
FLARE mode armed
1, 2,
1, 2,
1, 2,
1, 2,
3, 4?
4, 5?
3, 4, 5?
15. Autothrottle take-off mode is engaged by:
A push button switch on the AFDS mode control panel?
Press switches on the thrust levers?
A lever switch on the AFDS mode control panel?
A push button switch on the control yoke?
16. The factors necessary for computation of EPR limit are:
TAT, pressure altitude, flight mode?
TAT, pressure altitude?
N1, TAT?
N1, TAT, flight mode?
Chapter 6
In-Flight Protection Systems
Flight envelope protection
Every aircraft design is tested mathematically and in flight to determine the
limits of pitch, roll, yaw, angle of attack and `g' force that the airframe can
withstand in flight without suffering structural damage. These limits then
form what is known as the flight envelope for that particular design, within
which the aircraft can be safely operated. With a conventionally controlled
aircraft it is clearly possible to exceed the limits of the flight envelope by
applying excessive control movements.
As a means of eliminating the possibility of exceeding the limits of the
flight envelope through human error, the fly-by-wire system of flight control
has been developed. With such a system, the pilot's control demands are
transmitted to computers that are programmed to respond with signals to
the appropriate flying control servo-actuators which will limit their rate of
movement, thus ensuring that the aircraft response remains within the limits
of the flight envelope.
In the Airbus series of aircraft, beginning with the A320, the fly-by-wire
concept has been developed to the extent that the fly-by-wire computers
have complete control over each of the flying control surfaces, in response to
pilot demands from a small side-stick type of control. The response of the
computerised system to pilot inputs must be the same as in a conventional
direct control system, but the nature of the inputs is more complex because
the pilot can demand, for example, a rate of pitch or roll instead of a simple
control movement. This type of fly-by-wire system is known as an active
control system.
Given that there is no provision for reversion to manual control in these
aircraft, it is clearly vital that there must be a degree of redundancy in the flyby-wire control system sufficient to sustain failure of a computer without
degradation of aircraft control. This is achieved by employing a number of
computers in an active control system, such that no single computer can
command a control surface movement without being monitored by at least
one other. The A320 aircraft employs seven computers, connected by a data
bus, to control the elevators, ailerons, horizontal stabiliser, spoilers and
rudder. Two computers control the elevators, ailerons and the horizontal
150 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
stabiliser and are known as the elevator/aileron computers (ELAC). Three
computers control the spoilers, elevators and horizontal stabiliser and are
known as the spoiler/elevator computers (SEC). It can be seen that control of
the aircraft in pitch and roll is shared between the two computer systems so
that a fault in one system will not adversely affect the aircraft control. A third
pair of computers controls the aircraft in yaw, known as the flight augmentation computers (FAC).
Yaw damper
Swept-wing aircraft, to a greater or lesser extent, exhibit a tendency to
develop an oscillatory motion in flight, following a disturbance, which is a
combination of yawing and rolling and is known as `Dutch roll'. In many
cases the motion is damped out naturally by the `weathercocking' effect of
the vertical stabiliser and the aircraft quickly returns to steady flight.
However, swept-wing aircraft exhibit less natural damping because the
yawing motion initiates rolling and, in some cases, the oscillations increase if
unchecked, especially at lower flight speeds. The tendency can only be
checked by deflection of the rudder and to achieve this manually throughout
a long flight would place considerable strain on the pilot.
In aircraft that are susceptible to Dutch roll it is usual to install a yaw
damping system that automatically applies rudder deflection to control the
yawing tendency. The system comprises the third (yaw) axis of an autopilot
system and can be operated in either automatic or manual flight control. A
block schematic diagram of a yaw damping system is shown in Figure 6.1.
yaw rate
yaw damper
mechanical drive
Figure 6.1 Yaw damping system.
Motion about the aircraft's yaw axis is sensed by a rate gyroscope situated
in a coupler unit and powered from the aircraft's 115 V a.c. electrical system.
Output signals from the yaw rate gyro are amplified and filtered to remove
frequencies not associated with Dutch roll, and transmitted to an hydraulic
transfer valve in the rudder power control unit (PCU). Movement of this
valve directs hydraulic pressure to the yaw damper actuator. The resultant
movement of a piston in the yaw damper actuator operates a control valve in
the main rudder actuator, which moves the rudder in the required direction
to correct the yawing tendency sensed by the rate gyro. The yaw damper
In-Flight Protection Systems
piston motion is sensed by a transducer, known as a linear voltage displacement transmitter (LVDT), and fed back to the gyro unit. When the
actuator piston has moved by the amount demanded, this feedback of
rudder position cancels the gyro output and rudder movement is arrested.
When the Dutch roll oscillations have ceased, the LVDT signal is integrated
in the rate gyro coupler unit, to produce an output signal returning the
rudder to its neutral, centralised position.
Yaw damper indicator
On many aircraft equipped with a yaw damping system the operation of the
yaw damper is indicated on the EADI in conjunction with the rate of turn
indicator. This receives a signal from the yaw rate gyro. Whenever the gyro
precesses, the signal causes the rate of turn indicator to move away from its
neutral position. Rudder movement is displayed on a control position
Pilot operation of the rudder is by direct linkage to the main rudder
actuator and is therefore independent of the yaw damping system. Rudder
movements commanded by the yaw damping system are not transmitted
back to the rudder pedals.
Automatic pitch trim
In an aircraft equipped with a movable horizontal stabiliser (trimmable
stabiliser) and elevator for pitch control, pitch trim is normally adjusted by
first moving the elevators, followed by trimming the horizontal stabiliser
until the elevator is returned to the neutral, centralised position. The normal
action of an autopilot system in compensating for an out of trim condition in
pitch is to move the elevators until the condition is corrected.
The disadvantage of this system is that, once the elevators are deflected,
the amount of remaining movement in that direction is limited and control
authority in pitch is reduced. Furthermore, with the elevators deflected from
their centralised position, drag is increased, with the obvious adverse effects
upon fuel economy and, ultimately, range and endurance.
Consequently, it is not uncommon for aircraft with the stabiliser/elevator
configuration to incorporate a system additional to the automatic flight
system, which will automatically adjust the horizontal stabiliser until the
elevators are restored to the neutral position. Such a system is known as an
automatic stabiliser trim (AUTO STAB TRIM) system and it is usually
engaged automatically with autopilot engagement. It is a requirement of
autopilot engagement that the automatic stabiliser trim system must be
The degree of elevator deflection necessary will depend on airspeed and
152 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
the automatic stabiliser trim controls incorporate a feel unit which adjusts
the trimming signal according to sensed dynamic pressure.
Figure 6.2 illustrates schematically an automatic stabiliser trim system.
Pitch commands from the autopilot or from manual inputs are sent to the
powered control unit, which deflects the elevators through a screwjack,
either up or down depending upon the pitch attitude change required. The
automatic trim system will then move the horizontal stabiliser, through the
trim actuator and screwjack to apply the nose-up or nose-down trim
adjustment initially required. As the stabiliser takes up its new position, its
motion is mechanically transmitted to a feel and centring unit and a neutral
shift sensor. The deflection of the stabiliser removes the need for elevator
deflection and the neutral shift sensor sends a feedback signal to the elevator
PCU, removing elevator deflection as stabiliser deflection increases, until the
elevator and stabiliser are centralised.
feel and
horizontal stabiliser
trim system
Figure 6.2 Automatic stabiliser trim unit.
In aircraft with a fixed horizontal stabiliser, the pitch trim is achieved by
means of elevator trim tabs, which are deflected to assist the elevators, to
relieve the aerodynamic loads and some of the drag created by elevator
deflection. Automatic pitch trim control is accomplished by means of a
separate elevator trim tab servo-actuator coupled to the trim tabs and
working in parallel with the elevator servo-actuator. Figure 6.3 illustrates
schematically a system sometimes used in conjunction with small aircraft
automatic flight control systems.
A sliding bar on a mounting attached to the airframe is connected to a
capstan, positioned between the elevator `up' and `down' control cables. The
cables are lightly tensioned by pulleys so that, when the elevator is in the
In-Flight Protection Systems
d.c. supply
Figure 6.3
Automatic trim tab pitch trim system.
neutral (streamlined) position, the capstan and bar are centralised between
the cables. An electrical contact attached to the sliding bar is supplied from
the aircraft's d.c. bus bar. Adjustable contacts fixed to the mounting are
connected to the `up' and `down' field windings of a reversible d.c. motor,
which is the trim tab servo-motor.
Let us suppose that the autopilot has demanded a nose-down pitch. The
elevator actuator will deflect the elevator down, through the control cables,
tensioning the down cable and relieving tension on the up cable. The difference in cable tension will force the sliding bar upward and electrical
contact will be made with the trim tab `down' line, supplying the `down'
field coil of the trim tab servo-motor and driving the tab down to reduce the
aerodynamic force on the elevator. As the load on the elevator decreases, the
tension of the elevator cables will once again equalise and the sliding bar will
return to the centralised position, cutting off supply to the trim tab servomotor.
Automatic pitch trim systems normally include warnings and alerts in the
event of system failure. These typically take the form of warning lights or
captions and may include an aural alert should a runaway condition,
resulting in excessive trim input, occur. In the case of the automatic stabiliser
trim system, there is always a trim indicator on the flight deck.
Warnings general
Flight warning system
The function of a flight warning system is to alert the pilots to the existence
of an abnormal situation that requires action; it also identifies the nature and
location of the failure or condition. Warnings may be aural or visual, or a
154 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
combination of both. Aural warnings may be in the form of a klaxon alarm, a
bell or chime or, in some cases, a verbal message. Visual warnings and
cautions may be illuminated lights or captions, red for those requiring
immediate attention and amber for those requiring urgent attention, and
they may also appear as printed messages on the EFIS, EICAS or ECAM
displays. Certain hazardous situations, such as a fire warning, are indicated
by a general alert in the form of a red master warning light accompanied by
an aural alarm. The location of the incident will be indicated separately, for
example on an annunciator panel or the relevant electronic display.
Annunciator panels usually incorporate other attention-getting lights in
blue, green and white to notify the pilots of system availability and status, in
much the same manner as the EICAS and ECAM display systems.
In many large aircraft types the flight warning system is computerised and
an example of such a system is illustrated in block schematic form in Figure
flight warning
Figure 6.4 Flight warning system.
Altitude alert system
The function of the altitude alert system, as distinct from the ground
proximity warning system (GPWS), is to alert the pilots both aurally and
visually when the aircraft departs from or approaches a selected altitude on
the automatic flight director system. When the aircraft is approaching a
selected altitude, and it is within 900 ft of that altitude, the amber altitude
In-Flight Protection Systems
alert lights illuminate and an aural chime alert sounds for a 2-second period.
The alert lights remain steadily illuminated until the aircraft is within 300 ft
of the selected altitude.
Should the aircraft deviate from a selected altitude by 300 ft the amber
altitude alert lights will illuminate and flash repeatedly and the aural chime
alert will sound for 2 seconds. The alert lights will continue to flash until
the aircraft has deviated from the selected altitude by 900 ft, or has returned
to it. Deviation from the selected altitude will not be alerted when the
landing gear is extended, since this could lead to confusion with the GPWS
An altitude alert system is a JAR requirement for all turbojet passenger
aircraft and for turboprop aircraft over 5700 kg take-off weight and capable
of seating more than nine passengers.
Radio altimeter
The function of the radio altimeter is to measure and display the vertical
distance between the aircraft and the ground directly beneath it. It is
important to remember that, whilst it is very accurate, it only measures
vertical distance and is incapable of measuring terrain clearance ahead of the
Principle of operation
Radio altimeters for civil use operate in the SHF band within the frequency
range of 4200 MHz to 4400 MHz. A second frequency range of 1600 MHz to
1700 MHz, in the UHF band, is also reserved for radio altimeter operation,
but is not used by civil aircraft. The principle of operation is to continuously
transmit a variable frequency signal in a relatively narrow beam vertically
downward. The signal is reflected from the ground and received at the radio
altimeter receiver, located separately from the transmitter. Since it takes a
finite time for the signal to travel to the ground and return, and given that
the transmitted signal frequency is continuously changing, it follows that the
received frequency will differ from the transmitted frequency. The difference between the received and transmitted frequencies will vary as the
aircraft height varies, and the time taken for the signal to travel to the ground
and back varies. It is the frequency difference that is used to determine the
aircraft height above the ground at any instant, using the speed of propagation of the radio beam and the rate of change of transmitted frequency,
which are both known.
The transmitted signal is modulated to sweep over a frequency range of,
typically, 100 MHz around 500 times per second. This is a deliberately low
sweep rate, designed to avoid height ambiguity which might occur at a
156 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
graph A
re itte
ce d
ive sig
sig al
4300 MHz
4250 MHz
4200 MHz
time base
graph B
4250 MHz
4300 MHz
4200 MHz
time base
Figure 6.5 Radio altimeter principle of operation.
higher rate of frequency change. The concept of the radio altimeter principle
of operation is illustrated graphically in Figure 6.5.
In the example in Figure 6.5 the frequency is being modulated to sweep
over the range 4200 MHz to 4300 MHz at a constant rate. The solid line
represents the transmitted frequency and the dotted line the received frequency. At a particular point in time, represented by the vertical line
annotated 1 in graph A, it will be seen that there is a difference between the
two frequencies; this difference is directly proportional to the aircraft height
and is represented as such on the radio altimeter display. In graph B, the
vertical line annotated 2 represents an instant in time with the aircraft at a
greater height and consequently the difference between transmitted and
received frequencies is greater. The frequency difference is exaggerated in
these diagrams; in reality it is quite small, even at the usual maximum radio
altimeter operating height of 2500 ft.
In-Flight Protection Systems
System components
The principle components of a radio altimeter system are the transmitter,
the receiver and the display unit. However, as we have seen in previous
chapters, the display is incorporated in the ADI displays of EFIS or flight
director equipped aircraft. An example of the type of display unit found in
aircraft not so equipped is shown in Figure 6.6. It will be seen that the
instrument includes a decision height feature that allows the pilot to set the
decision height index bug on the face of the instrument. When the pointer
reaches the set height during a descent a visual and/or aural warning is
activated. The visual warning is usually in the form of a light and the aural
warning may be a chime alert or a recorded voice message. In the event of
failure of the system due to loss of power, a system or reception fault, a
prominent warning flag appears on the face of the instrument. Additionally, the pointer will be obscured on these occasions or when flying above
2500 ft. The pointer will take up a known position when the press-to-test
button is depressed. In some displays the instrument scale is logarithmic
for heights above 500 ft.
warning flag
decision hei
setting knob
Figure 6.6
Radio altimeter display.
The accuracy of the radio altimeter is given as +1 ft or +3% of the indicated
height, whichever is the greater. It can be subject to errors due to reflections
from parts of the aircraft structure, such as the landing gear, or to leakage of
signals between the transmitting and receiving aerials. The positioning of
the aerials is therefore very important and every effort is made by the
manufacturer to avoid these errors. It is also conceivable that the receiver
might pick up a signal that has been reflected from ground to aircraft more
than once, known as a multi-path signal. To a large extent this potentially
ambiguous situation is avoided by gain control in the receiver.
158 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
The principle use of radio altimeter information in large modern aircraft is
in conjunction with the automatic landing system and the ground proximity
warning system.
Ground proximity warning system (GPWS)
The purpose of a ground proximity warning system is to provide aural and
visual signals to the pilot when the aircraft is in danger of impacting with the
ground, unless corrective action is taken. Joint aviation requirements are that
all turbine-powered aircraft having a maximum certificated take-off weight
greater than 5700 kg and seating for more than nine passengers must be
equipped with a GPWS.
There are three types of GPWS currently in use: basic, advanced and
Basic GPWS
The basic GPWS has five modes of operation, which require the following
source inputs:
Radio altimeter. Accurate measurement of height above ground level is
provided by the radio altimeter.
Central air data computer. Barometric pressure is integrated by the
GPWS to compute descent rates.
ILS glidepath receiver. The GPWS is required to give warning of descent
below the glidepath during a landing approach.
Approach configuration. The landing gear and flaps positions are
necessary inputs to the system during the approach to land.
The GPWS must be active between 2450 ft and 50 ft above ground level.
Operating modes
Mode one, excessive descent rate
This mode is activated at and below 2450 ft radio altitude and is designed to
alert the pilot to the fact that the aircraft is descending at a rate that is
hazardous. The system measures the barometric rate of descent and gives an
aural warning in the form of a klaxon type `whoop-whoop' alarm followed
by a verbal `pull up' message. These are accompanied by a visual warning in
the form of a red light with the caption PULL UP prominently displayed
within the pilots' normal field of view. At the upper limit of operation of
2450 ft agl, the warning will be given if the descent rate exceeds 7350 ft per
In-Flight Protection Systems
minute. The lower the aircraft is, the more urgent the required response and
so the descent rate that will initiate a warning decreases with height. At
1000 ft agl, the warning is triggered by a descent rate of 3000 ft per minute. At
the minimum GPWS operating height of 50 ft agl, the trigger value is 1500 ft
per minute.
Mode two, excessive terrain closure rate
Mode two is designed to provide warning when the aircraft is in level flight
at or below 1800 ft radio altitude and the terrain beneath it is rising. In this
mode the input is from the radio altimeter, but the system also takes account
of flap position. Mode 2A is active above 790 ft agl and will only provide a
warning if the flaps are NOT in the landing configuration and the terrain
closure rate is in excess of 2063 ft per minute. Mode 2B becomes active below
790 ft agl and will give a warning regardless of flap position if the terrain
closure rate is 3000 ft per minute or greater. As with all five modes, the lower
limit of operation is 50 ft agl. The warnings given, in the event of the terrain
closure parameters being exceeded, are the same as in mode one.
Mode three, altitude loss after take-off or go-around
This mode is active between 50 ft and 700 ft agl as measured by the radio
altimeter. Its purpose is to alert the pilot to accumulated height loss immediately following take-off or during a go-around manoeuvre. The aural and
visual warnings will be initiated if a cumulative barometric height loss
occurs, the triggering value of which depends upon radio altitude. For
example, at 50 ft agl the warnings will be triggered if the cumulative barometric height loss exceeds 10 ft. At the upper level of 700 ft the warnings will
not be activated unless the cumulative barometric height loss is in excess of
70 ft. The trigger value increases linearly between the lower and upper limits
of operation.
Mode four, unsafe terrain clearance
The purpose of mode four is to warn the pilot if the terrain clearance is
inadequate with the aircraft NOT in the landing configuration. It is divided
into two parts, mode 4A and mode 4B. Mode 4A is active between 500 ft and
200 ft radio altitude if the landing gear is NOT down and locked. It is triggered by a barometric descent rate of 1900 ft per minute or more. Mode 4B is
active between 200 ft and 50 ft radio altitude and will be triggered if the flaps
are not in the landing configuration with the landing gear down and locked.
The warnings given are the same as for the previous modes.
Mode five, aircraft below the ILS glideslope
Unlike the previous modes, mode five gives an alert as opposed to a
warning. The definition of a warning is a command requiring immediate
160 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
response in the form of a maximum gradient climb to a safe altitude. An alert
is defined as a caution requiring immediate action to correct the flight path
or aircraft configuration. Mode five is active between 500 ft and 50 ft radio
altitude and will generate an alert if the aircraft is significantly below the ILS
glideslope with the landing gear extended. In broad terms, a significant
deviation is given as greater than 150 micro amps displacement from the
central null at the glideslope receiver. Mode five is inhibited when mode
three is active, to avoid confliction between the two. The alert given is visual
and takes the form of an amber light, with a `below glideslope' caption,
situated adjacent to the red `pull up' warning light. The mode five alert can
be inhibited by pressing the amber light.
An example of a GPWS control and display unit is shown in Figure 6.7.
System integrity is tested by pressing the red `pull up' light cover, whereupon the `pull up', `below glideslope' and `inop' lights should illuminate.
The override switch and light provide the facility to inform the GPWS when
the approach flap configuration is to be other than normal, as in an approach
with asymmetric power, for example.
Figure 6.7 GPWS control and display unit.
Advanced GPWS
A disadvantage of the basic GPWS is that it does not differentiate between
modes 1 to 4, because it gives the same warning in each case. It is, of course,
useful for the pilot to know the cause of the warning when responding to it.
Advanced GPWS overcomes this to a large extent by providing an identifying aural alert message to accompany the warning. The amber alert light,
which only illuminates in mode 5 of basic GPWS, will illuminate in all cases
and consequently carries the caption `ground proximity' instead of `below
The alert messages in modes one to three follow the aural `whoop-whoop',
`pull up' warning and are as follows:
In-Flight Protection Systems
Mode one. `Sink rate' repeated.
Mode two. `Terrain' repeated.
Mode three. `Don't sink' repeated.
Mode four. The warning message is `too low terrain', instead of `pull up'.
The alert message in mode 4A is `too low gear' repeated and in mode 4b it
is `too low flaps' repeated.
Mode five. As with basic GPWS, there is no warning message and the
alert message is the same, `glideslope'.
Mode six. Advanced GPWS has a sixth mode, which is alert only. It differs
from the previous five modes in that it alerts the pilot to the fact that the
aircraft has reached a certain point in the landing approach, such as
decision height, and that the approach may be continued or a missed
approach procedure executed, depending upon visibility criteria. The
alert message with this mode is `minimums' repeated.
Advanced GPWS has built-in test equipment (BITE), which continuously
monitors the system for faults and these are displayed automatically during
flight. The self-testing system display cannot be selected during flight, but is
normally part of the pre-flight checks. The ground test will activate all the
visual and aural warnings. In the event of unserviceability of the GPWS, the
aircraft may only be permitted to fly if the equipment cannot be repaired at
the location where the fault is first discovered. In such a case, the aircraft
may be flown to an airfield where the fault can be rectified, provided that the
journey requires no more than six segments.
Terrain avoidance warning systems (TAWS)
Both basic and advanced GPWS have shortcomings in that they take no
account of terrain ahead of the aircraft, the time available to respond to
warnings is very limited and the recovery advice is minimal. It would clearly
be of great value to the pilot to receive information that a potentially
hazardous situation could exist, as far in advance as possible.
With the amount of information stored in flight management computer
systems about ground locations of radio navaids, airports, etc., it is a relatively simple task to add a terrain database that covers all the world's major
air routes. This database is held in the GPWS computer memory and,
combined with the route and location data held in the flight management
system computer, gives the system the capability to warn of terrain hazards
in the projected flight path. The system currently in use, that complies with
TAWS requirements, is known as enhanced GPWS (EGPWS) manufactured
by Honeywell.
The inputs to the enhanced ground proximity warning system (EGPWS)
162 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
Radio altimeter. As in the basic and advanced systems, the radio altimeter
determines the vertical height of the aircraft above ground level.
Central air data computer (CADC). The aircraft's vertical speed is computed from rate of change of barometric pressure and its airspeed from
pitot and static pressure. In addition, the CADC provides the EGPWS
with static air temperature.
VHF receiver. ILS glideslope information is fed from the VHF navigation
radio receiver.
Navigation computer. The aircraft's lateral position and track are supplied continuously from the navigation computer, using radio navigational fixes and IRS computed position information.
Flight management system (FMS). The FMS provides heading, track,
attitude and groundspeed data as generated by the IRS. Where these
systems are not available, the information can be provided by a series of
The enhanced system displays surrounding ground features and is able to
give significantly greater warning of impact hazard than its predecessors.
Terrain details can be presented on the EFIS HSI display or the weather
radar display, with colour coding indicating the terrain height relative to the
aircraft. Green indicates that the terrain is below the aircraft's projected
flight path, terrain that extends above the projected flight path is indicated
by yellow and red indicates terrain that is well above the projected flight
path. Resolution increases with terrain height. Terrain extending above the
projected flight path and which is sufficiently close to generate an alert is
painted in solid yellow or red.
Warnings and alerts
When the projected flight path of the aircraft is obstructed by terrain at
ranges calculated to present a hazard, caution lights illuminate on the
EGPWS panel and a recorded spoken message gives an appropriate aural
alert such as `caution terrain'. If no action is taken to amend the flight path
the caution lights are replaced by a red warning light and the aural message
becomes `terrain, terrain' ± `pull up, pull up'.
The general terrain database of the system contains a model of the earth's
surface with contours stored in rough detail. Around well used routes and
airways the model is more precise and includes obstructions as well as
terrain features. In the vicinity of airports it is assumed that aircraft are likely
to be descending and so the terrain proximity warning distances are less
than those in areas where aircraft would be expected to be operating at
altitude. For the areas surrounding airports the database contains precise
and up-to-date information of all terrain and obstacles under the approach
In-Flight Protection Systems
EGPWS mode seven
EGPWS has a seventh mode, the function of which is to provide alerts in the
event of encountering windshear below a radio altitude of 1500 ft. The
system computer compares airspeed and groundspeed to detect significant
changes in windspeed and direction. In the event of an increasing headwind,
decreasing tailwind and/or severe updraught a caution light will illuminate
and the aural message `caution windshear' will be spoken. In the event of a
decreasing headwind, increasing tailwind and/or a severe downdraught, a
windshear warning will be given in the form of an illuminated red warning
light and the spoken aural message `windshear, windshear, windshear'.
Traffic Collision Avoidance System (TCAS II)
For many years the separation of aircraft in flight was dependent upon the
air traffic control services, using secondary surveillance radar (SSR), and the
alertness of the flight crews, using the mark 1 eyeball. As air traffic congestion increased, especially around large airports, and the speed of aircraft
made visual warning and avoidance more and more unlikely, the airline
operators became highly concerned at the number of near misses and the
increasing likelihood of a catastrophic mid-air collision. Fortunately, these
circumstances coincided with the rapid advance in computer technology
and miniaturisation. It became possible, in theory at least, to extend the SSR
principle of a ground station interrogating an airborne transponder and fit
aircraft with interrogation equipment as well. Thus, an aircraft in flight
would be able to continuously transmit interrogation signals that, when
received by another aircraft, would trigger its transponder to respond with
details of its altitude. If these, when computed, indicated a potential collision
course between the two aircraft, the airborne equipment would supply flight
directions to divert the interrogator and avoid conflict.
The operating requirements for a collision avoidance system using these
principles were stipulated by ICAO under the title Airborne Collision
Avoidance System (ACAS). To date only one system has been introduced
into general use that meets these requirements and it is known as the Traffic
Alert and Collision Avoidance System (TCAS), developed in the USA. There
are two versions of the system, TCAS I and TCAS II, of which only the latter
is approved for passenger transport aircraft in the USA and Europe. In
Europe the Joint Aviation Authority requires all fixed wing turbine-powered
aircraft with seats for more than 30 passengers to be equipped with TCAS II
and, by the beginning of 2005, this will be extended to include all aircraft
over 5700 kg take-off weight with seats for more than 19 passengers. In the
USA, the Federal Aviation Administration (FAA) requires all aircraft with
seating for more than 30 passengers to be equipped with serviceable TCAS
164 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
TCAS I, whilst not approved for use in fixed wing passenger transport
aircraft, has proved popular in aircraft that operate at generally lower speeds
and altitudes, especially helicopters. It provides the pilot with a visual display of the range and relative bearing of an aircraft on a potential collision
course, with 40 seconds in which to acquire the aircraft visually and take
avoiding action. Mode C equipped transponders provide altitude information concerning the conflicting traffic and render the system capable of
eliminating traffic that has adequate vertical separation from the TCAS
TCAS II principle of operation
The TCAS installation consists of a transponder that continuously transmits a
pulsed interrogating signal on the standard SSR frequency of 1030 MHz and
responds with a coded pulsed signal on the standard SSR frequency of
1090 MHz. The transmitter power required is relatively low, since the range
required is limited by the distance necessary to provide adequate warning of
potential confliction. The interrogating signal is sent in SSR mode S, which
includes mode A interrogation and mode C altitude information (SSR is
explained in detail in the Radio Aids volume of this series of books). A TCAS II
equipped aircraft's transponder receiving this signal will respond with a
coded pulsed signal at 1090 MHz containing details of the aircraft altitude.
The two receiving antennae are directional and are located one on the top of
the aircraft and one underneath. They enable the interrogating aircraft's
TCAS computer to calculate relative bearing, and the response time enables it
to compute range. From these data the computer can immediately determine
whether there is a potential confliction. By integrating the received altitude
with time, the interrogating aircraft's computer can calculate and display the
rate of change of altitude of the intruder aircraft. All information from the
TCAS computer concerning responding aircraft transponders is displayed on
an electronic vertical speed indicator (VSI) or on the EFIS display.
The TCAS II system provides a visual display of all responding air traffic.
Those responses that are not computed to currently present a threat of collision are known as traffic advisory (TA) messages. When the computer
calculates that a confliction is possible it generates resolution advisory (RA)
messages describing the necessary avoidance manoeuvre. The resolution
advisory only prescribes manoeuvres in the vertical plane, i.e. climb or
descend, since the accuracy of the directional antennae is insufficient for safe
lateral deviation directions. The resolution advisory messages are passed to
the pilot both aurally and visually. The aural messages are in the form of
recorded voice messages that indicate the urgency of the pilot response
necessary. For example, where an intruder aircraft presenting no current
threat exists the traffic advisory (TA) message will be `traffic, traffic',
In-Flight Protection Systems
drawing the pilot's attention to the visual display and alerting him to the
possibility of a resolution advisory (RA). Resolution advisories such as
`climb, climb' or `descend, descend' require the initiation of a 1500 ft per
minute climb or descent within 5 seconds. `Increase climb/descent' requires
the rate of change of vertical speed to be increased to 2500 ft per minute
within 2 to 3 seconds and `climb, climb NOW' or `descend, descend NOW'
requires an immediate reversal of the vertical flight direction.
Visual display
The electronic VSI display referred to previously comprises a `conventional'
VSI scale, albeit electronically generated, around its perimeter and calibrated
to show rates of climb in the upper half and rates of descent in the lower. The
lateral situation is indicated by an aircraft symbol and azimuth scale in the
lower centre of the display and symbols representing intruder aircraft about
the upper centre in the respective lateral positions of the transponding
intruder aircraft.
There are four alternative intruder symbols, indicating the threat potential
presented and the vertical movement of the intruder. Provided that the
intruder aircraft is responding with SSR mode C, the symbol will show the
relative altitude numerically in hundreds of feet. If the relative altitude of the
intruder is changing by more the 500 ft per minute the numeric annotation
will be preceded by a plus or minus sign, indicating that it is climbing or
descending. This is further emphasized by an accompanying arrow pointing
up or down, as appropriate. The shape and colour of the intruder symbol
indicates the nature of the advisory message, as shown in the summary
below and in Figure 6.8.
intruder – no threat
vertical separation
more than 1200 ft
range more than 6 nm
range less than 6 nm
1200 ft above
traffic advisory
resolution advisory
500 ft below
descending at
more than 500 ft/min
400 ft above
climbing at more
than 500 ft/min
Figure 6.8
proximate traffic – no threat
TCAS II symbology.
166 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
No threat. Traffic beyond 6 nm range or more than 1200 ft vertical
separation. An open diamond symbol in white or cyan, with no altitude
Proximate traffic. Traffic within 6 nm range, but not computed to present
a threat. A solid diamond symbol in cyan, with relative altitude annotation.
Traffic advisory. A solid circle in amber, with relative altitude annotation.
Resolution advisory. A solid square in red, with relative altitude annotation.
The circular scale has coloured arcs superimposed upon it to indicate safe
and unsafe climb and descent rate areas when a resolution advisory exists.
Unsafe areas are indicated by a narrow red arc; the advised rate of change of
altitude by a broad green arc. The range of the intruder is indicated digitally
outside the circular scale and in analogue form by the position of its symbol
relative to the range ring. A typical TCAS II electronic VSI display is shown
in Figure 6.9. In this, the equipment has issued a resolution advisory for an
intruder at 10 nm range, which is 600 ft above the interrogating aircraft and
descending at greater than 500 ft per minute. The broad green arc between
500 and 1000 ft per minute descent rate indicates the RA recommended safe
descent rate and the narrow (red) arc the unsafe area.
red arc
Figure 6.9 TCAS II VSI display.
In EFIS equipped aircraft the intruder symbols are displayed on the horizontal situation indicator in the PLAN and expanded modes and the RA
avoidance manoeuvre is shown by the command bars on the ADI display.
When both the interrogating and the intruder aircraft are equipped with
TCAS II and SSR mode S capability, the two TCAS computers are able to coordinate the resolution advisories in each aircraft to achieve optimum
In-Flight Protection Systems
separation, with the least disruption to either. This is designed to ensure that
both flight crews do not take the same avoiding action and worsen the
danger of collision.
Crew response
The system is so designed that, provided pilot response to an RA is taken
within the time limits of 5 seconds and 2 to 3 seconds referred to above, the
altitude change to avoid confliction should not exceed 500 ft. If the aircraft is
under Air Traffic control when a TCAS resolution advisory is received, the
pilot is required to obey the TCAS command and inform ATC `TCAS climb/
descent' as appropriate. Upon receipt of the TCAS message `clear of conflict',
the aircraft must be returned to the ATC assigned flight level. Should a
manoeuvre instruction be received from both TCAS and ATC simultaneously, the pilot is required to obey the TCAS instruction and advise ATC
At radio altitudes of less than 1000 ft the TCAS will not give a resolution
advisory involving a descent, and below 1800 ft agl it will not recommend an
increased rate of descent, since in either case the hazard to the aircraft in
terms of ground impact would be greater than the collision hazard. All
resolution advisories are inhibited at radio altitudes of less than 500 ft, and
traffic advisories at less than 400 ft.
In order to avoid confusion, TCAS warnings are co-ordinated and
prioritised with those of other in-flight protection systems. It is usual for
GPWS alerts and warnings to take precedence over TCAS, and a windshear
warning will be awarded the highest priority.
Collision warning systems are also described in the Radio Aids volume in
this Ground Studies for Pilots series of books.
Overspeed warning
The function of the overspeed warning system is to provide aural warning,
to supplement the visual warnings on the airspeed and mach indicators,
when the maximum operating speed Vmo/Mmo is exceeded. The aural
warning given is usually a `clacker' type of alarm, which can only be stopped
by reducing airspeed below the maximum operating limit.
Input data of airspeed and mach number for the overspeed system is
obtained from the central air data system. Since the maximum operating
airspeed is affected by criteria such as aircraft weight, flap and slat positions
and centre of gravity position, these are provided from the flight management system computer.
Visual displays of Vmo/Mmo are presented on the airspeed indicator,
typically in the form of a separate pointer, known as the maximum allowable
168 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
pointer, actuated by a static pressure diaphragm and a specially calibrated
mechanism. The pointer is usually distinctively marked and is often referred
to as the `barber's pole'. Vmo increases with altitude until Mmo is reached, so
the calibration of the maximum allowable pointer indicates the maximum
operating speed adjusted for altitude. Mmo is a set value (e.g. 0.84 M) and a
mach number operated switch triggers the clacker alarm if this value is
reached before Vmo, as would typically be the case at altitude. Failure of the
Vmo/Mmo system is indicated by a warning flag on the face of the
The overspeed warning system incorporates a test function to prove
serviceability during preflight checks, which sounds the aural warning.
Stall warning system
The function of a stall warning system is to alert the pilot to the fact that the
aircraft is approaching the stalling angle of attack in sufficient time for
corrective action to be taken. The actual nature of the warning takes several
different forms, depending upon the aircraft type and its behaviour at the
point of the stall and thereafter. Since the stall occurs at a known angle of
attack, the stall warning sensing device measures aircraft angle of attack and
the stall warning system activates an alert before the stalling angle is
Many aircraft provide some warning of the approach of a stall in the form
of buffeting, followed by a pitch-down change of attitude as the centre of
pressure moves rapidly aft on the wing. The buffeting is felt on the controls
in aircraft with controls linked directly to the control surfaces, but with
powered flying controls this is less likely to be the case and the natural stall
warning may be so slight that the pilot might miss it in conditions of high
work-load. In aircraft with the horizontal stabiliser mounted at or near the
top of the fin (so-called T-tail aircraft) a condition known as deep stall is
likely to develop if prompt recovery action is not taken immediately a stall
warning is received. In aircraft having this configuration the stall warning
has to be more direct than a simple aural alert; initially the system applies
vibration to the control column (stick shaking) and in some aircraft, if pilot
response is not immediate, it is followed by a forward pressure to the column (stick pushing).
Angle of attack sensing
In general aviation aircraft the angle of attack sensing device often used
consists of a vane mounted in the leading edge of a wing at the point where
stagnation occurs at normal flight attitudes. Under these circumstances the
vane occupies a neutral, mid-position with the air pressure approximately
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equal on both its upper and lower surface. As the aircraft attitude becomes
increasingly pitched nose-up, the stagnation point moves toward the lower
surface of the wing and the pressure beneath the vane becomes significantly
greater than that above it. The vane therefore moves upward and closes a
switch attached to it, completing an electrical circuit to the stall warning
horn in the cockpit. The mechanism is adjustable, so that it can be set to
activate the stall warning at a precise angle of attack, slightly lower than the
stalling angle. This type of sensor is illustrated in Figure 6.10.
to warning horn
Figure 6.10
Leading edge angle of attack vane.
Some light aircraft and training types use an even simpler stall warning
device known as a plenum chamber. An adjustable plate with a slot cut in it
is mounted on a wing leading edge and adjusted so that the slot aligns with
the stagnation point at normal flight attitudes. The slot is connected to a
chamber, which is in turn connected by tube to a reed-operated horn in the
cockpit. When the angle of attack reaches a preset value, the reduced pressure at the slot induces airflow through the horn, vibrating the reed to
provide an aural alert to the pilot.
In larger aircraft, where the stall warning may well be accompanied by
stick shaking or pushing, a more sophisticated type of angle of attack sensor
is typically used. This uses an aerodynamic vane that projects into the airstream at a location unaffected by disturbances, usually on the side of the
fuselage toward the nose of the aircraft. The vane is dynamically balanced
and is connected to the rotor of a synchro, which transmits an electrical
signal proportional to the angle taken by the vane, relative to a preset null
position. Since the vane has a symmetrical aerofoil section and is freely
balanced, it naturally aligns itself with the airstream when the aircraft is in
flight. Thus, as aircraft pitch attitude changes, the vane remains aligned with
the airstream and the aircraft moves in pitch relative to the vane. The relative
movement is sensed by the vane synchro and transmitted to the stall
warning system and the angle of attack indicator, where fitted. An angle of
170 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
attack vane is illustrated in Figure 6.11. Upon installation, the vane unit is
accurately aligned by locating pins on the fuselage that engage in holes in the
vane unit mounting flange.
Because the stalling angle of attack changes when flaps and slats are
deployed, it is necessary to bias the vane synchro with inputs from the flap
and slat position synchros to adjust the stall warning angle of attack
holes for locating pins
aerodynamic vane
indexing scale
Figure 6.11 Angle of attack sensor unit.
Stick shaker
In aircraft where an immediate response to a stall warning is required it is
usual to install a stick shaker. This consists of a d.c. motor that, when activated, applies vibration to the control column, immediately drawing the
pilot's attention to the need to apply forward stick to reduce the pitch attitude of the aircraft. This method is chosen because it is reminiscent of the
pilot's earliest flight training in stall recovery procedures in a training aircraft. The motor will continue to vibrate the control column until the angle of
attack sensor senses that the aircraft's angle of attack has reduced below the
warning threshold, whereupon the system cuts off d.c. supply to the motor.
Figure 6.12 shows a block diagram of a stick shaker system. Switching the
system control switch to TEST, with the aircraft on the ground, activates a
small motor to rotate the dial of an indicator, proving that the control circuit
is functional.
Stick pusher
Aircraft with poor stall recovery characteristics, such as T-tail configured
aircraft, may be fitted with a system that actually pushes the control column
forward if the pilot does not respond to the stick shaker. It was found with
the BAC 1-11 and Trident aircraft, both high T-tailed with aft-mounted
engines, that if stall recovery was not promptly executed, the aircraft was
prone to enter a deep-stall (super stall) condition from which recovery is
almost always impossible. The reason for this is that the tailplane (horizontal
In-Flight Protection Systems
angle of attack
sensing unit
flap position
a.c. supply
bias unit
Figure 6.12
Stall warning system block diagram.
stabiliser) becomes stalled by the turbulent airflow from the stalled wings
and all longitudinal control is lost. The aircraft enters a steep descent in a
nose-up, wings level attitude. Thus, it is essential that rapid corrective action
is taken in aircraft of this configuration, hence the stick pusher. The motive
force for pushing the control column is often of the linear actuator type and
the signal to it is passed through the elevator channel feel and centring unit.
Activation of the stick pusher automatically disengages the automatic flight
control system.
Flight data recorder
The Joint Aviation Authority (JAA) requires that all turbine powered aircraft
with a take-off weight greater than 5700 kg and with seating for more than
nine passengers shall be equipped with a flight data recorder. The device
must be capable of retaining data recorded during at least the last 25 hours of
aircraft operation, although this figure may be reduced to 10 hours for aircraft with a take-off weight of less than 5700 kg.
The data recorded must be sufficient to establish the following flight
Attitude in pitch and roll
Thrust or power on each engine
Configuration of lift or drag devices
Radio transmission keying
Use of automatic flight control systems
172 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
Angle of attack
Air temperature
For aircraft with a take-off weight in excess of 27 000 kg it is required that
additional data must be recorded in order to be able to establish the following parameters, as well as those listed above:
Primary flight control positions
Pitch trim
Primary navigation information as displayed to the flight crew
Flight deck warnings
Landing gear position
Radio altitude
The data recorded must be from essentially the same sources as those which
supply the information displayed to the flight crew and it must include any
parameters that are peculiar to the operating characteristics of the aircraft
The flight data recorder must automatically begin recording all the above
data before the aircraft is capable of moving under its own power and must
automatically cease recording after the aircraft is no longer capable of
moving under its own power. In practical terms, this usually means that
recording starts with start of the first engine and ceases at shut-down of the
last engine. The recorder must be contained within a container painted in a
distinctive orange or yellow colour and its recovery must be assisted by
reflective material and an underwater locating device that is automatically
activated upon immersion.
It must be so installed in the aircraft that the probability of damage to the
recorded data from shock, heat or fire is minimised. This is usually satisfied
by locating the flight data recorder as far aft as practicable, typically in the
vicinity of the rear pressure bulkhead.
The electrical supply to the recorder must be from a bus bar that can be
expected to provide power under all circumstances, without jeopardising
essential or emergency services. There must also be a pre-flight testing
facility to check the serviceability of the recorder.
Figure 6.13 shows a block diagram of the typical inputs to the flight data
recorder of a large transport aircraft.
Types of flight data recorder
Flight data recorders are classified according to the amount of information to
be retained and the length of aircraft operating time over which data are to
be recorded and stored. Recorders meeting the JAR-OPS requirements for
In-Flight Protection Systems
flight data
Figure 6.13
three axis
HF radio
Flight data recorder block diagram.
aircraft with a maximum take-off weight in excess of 27 000 kg must be
capable of recording at least 32 parameters and these are classified as Type I
flight data recorders. Type II recorders meet the JAR-OPS requirements for
smaller aircraft (take-off weight of 5700 kg) and these must be capable of
recording at least 15 parameters. Type IIA recorders only have a 30-minute
recording span, but must be capable of retaining data recorded during the
preceding take-off.
The minimum 32 parameters required of Type I flight data recorders are
listed below. Normally a Type II recorder would record the first 15 of these
parameters, although the parameters may vary according to aircraft type.
UTC or elapsed time
Pressure altitude
Indicated airspeed
Vertical acceleration
Pitch attitude
Roll attitude
Radio transmission keying
Power on each engine
Trailing edge flap position or control selection
Leading edge flap position or control selection
Thrust reverser position
Ground spoiler and/or speed brake position
Outside air temperature
Autopilot, autothrottle and automatic flight control system modes and
Longitudinal acceleration
Lateral acceleration
Primary control surface positions and/or pilot's control inputs
174 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
Pitch trim setting
Radio altitude
Glide path deviation
Localiser deviation
Marker beacon transit
Master warnings
Navigational radio frequencies
DME distances
Landing gear status from squat switch
Landing gear selector position
Angle of attack
Hydraulics systems pressures
Latitude and longitude, groundspeed and drift angle
System monitoring
The flight data recorder system has its own built-in test equipment (BITE)
and the serviceability of this and the recorder should be checked before the
first flight of the day. FDRs are subject to annual inspection and to calibration on a 5-year cycle. Dedicated airspeed and altitude sensing equipment is
subject to bi-annual inspection and calibration.
Aircraft integrated data system
Many of the larger transport aircraft types are equipped with data gathering
and retention systems for monitoring the health and performance of the
engines and aircraft systems. The system most commonly used is known as
the aircraft integrated data system (AIDS), which provides the option of a
real time display of current operating conditions, or downloading and printout of the data when the aircraft is on the ground. Some operators make use
of an extension to AIDS known as the aircraft communication addressing
and reporting system (ACARS), whereby the system can be interrogated
from the operator's ground base and technical data downloaded whilst the
aircraft is in flight. The data recorded and stored by AIDS can be interchanged with the flight data recorder and the FDR data can be printed out
during aircraft maintenance.
Cockpit voice recorder
The Joint Aviation Authority requires that all multi-engine turbine-powered
aircraft with a maximum take-off weight in excess of 5700 kg and with
seating for more than nine passengers shall be equipped with a cockpit voice
In-Flight Protection Systems
recorder. The voice recorder must be capable of retaining recorded information over the period of the last 2 hours of operation and the parameters
recorded must be as follows:
All radio voice communications received or transmitted from the flight
All sounds within the flight deck environment, including audio signals
received by each boom and mask microphone in use.
Voice communications between flight crew members on the interphone
All voice or audio signals identifying navigation or approach aids, as
received on crew headphones or speakers.
All announcements made by the flight crew on the public address system.
For aircraft with a maximum take-off weight of less than 5700 kg the
recording time may be limited to 30 minutes.
The cockpit voice recorder must automatically begin recording before the
aircraft first moves under its own power and continue until it is no longer
capable of moving under its own power. In practical terms, this is usually
from first engine start to last engine shut-down.
The voice recorder container must be easy to locate in a crash situation by
painting it a distinctive orange or yellow colour with reflective material
attached. It must also include an automatically activated underwater
detection device and it must be resistant to shock, heat and fire.
The recorder must be installed in a location where its recordings are least
likely to suffer damage. The site chosen is usually as far aft as practicable,
typically close to the rear pressure bulkhead. It must receive its electrical
power from a bus bar that can be relied upon to continue providing power
under all circumstances and that is separate from the aircraft's essential and
emergency services.
Figure 6.14
cockpit voice
Cockpit voice recorder block diagram.
176 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
Figure 6.14 shows a block diagram of a cockpit voice recorder system for a
large passenger transport aircraft.
There must be a means of preflight checking the cockpit voice recorder for
serviceability. An aircraft may only be despatched with an unserviceable
recorder provided that the means of repairing it are not available and that
the aircraft does not complete more than eight subsequent consecutive
flights with the device unserviceable.
Sample questions
1. The type of fly-by-wire system that converts pilot demands into rate of
pitch or roll is known as:
An active control system?
A passive control system?
A rate control system?
An integral control system?
2. A yaw damping system is necessary in aircraft:
To control the rate of yaw following a pilot rudder input?
That are susceptible to Dutch roll?
That are susceptible to phugoidal flight?
That do not have swept wings?
3. An automatic stabiliser trim system:
a. Adjusts the position of the elevators to align with the horizontal
b. Adjusts a trim tab on the elevators to reduce elevator deflection?
c. Adjusts the position of the horizontal stabiliser to centralise the
d. Can only be engaged after the autopilot system has been engaged?
4. The altitude alert system will provide:
a. An aural and visual alert that commences when the aircraft is within
900 ft of the selected altitude and continues until it is within 300 ft of
the selected altitude?
b. An aural and visual alert that commences when the aircraft is within
900 ft of the selected altitude and repeats when it is within 300 ft of
the selected altitude?
c. An aural alert that sounds for 2 seconds when the aircraft is within
900 ft of the selected altitude and a visual alert that remains illuminated until the aircraft is within 300 ft of the selected altitude?
In-Flight Protection Systems
d. An aural alert that sounds for 2 seconds when the aircraft is within
600 ft of the selected altitude and a visual alert that remains illuminated until the aircraft is within 300 ft of the selected altitude?
5. Radio altimeters operate in the . . . . . . . . . . . . band with a frequency range
of . . . . . . . . . . . .:
4200 MHz±4400 MHz?
1600 MHz±1700 MHz?
4200 MHz±4400 MHz?
4200 MHz±4400 MHz?
6. The sweep rate of a radio altimeter is, typically:
High, to avoid interference with other aircraft transmissions?
Low, to avoid height ambiguity?
Low, because the transmitter and receiver are not co-located?
Variable, to make each aircraft's transmitted signals unique?
7. The accuracy of a radio altimeter is given as:
+1 ft
+3 ft
+1 ft
+5 ft
8. A basic GPWS requires inputs from (answer a, b, c or d):
Radio altimeter
Central air data computer
ILS glidepath receiver
Approach configuration
Navigation computer
Flight management system
3, 4 only?
3, 4, 5 only?
3, 4, 5, 6?
3 only?
9. GPWS mode 1 gives warning of:
Excessive descent rate?
Excessive terrain closure rate?
Altitude loss after take-off or go-around?
Unsafe terrain clearance?
178 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
10. The GPWS mode that gives warning of altitude loss after take-off or goaround is active between:
50 ft
50 ft
50 ft
50 ft
and 1800 ft agl?
and 700 ft agl?
and 2450 ft agl?
and 500 ft agl?
11. GPWS mode 5 is inhibited when:
Mode 1 is active?
The aircraft is below the glideslope?
Mode 3 is active?
Flaps and landing gear are deployed?
12. A fundamental difference between advanced GPWS and basic GPWS is
that the former:
Has only five operating modes?
Gives alerts, rather than warnings?
Operates from a greater height above ground level?
Identifies the warning mode with an alert message?
13. The colour display generated by the enhanced GPWS shows terrain that
is below the aircraft's projected flight path:
shaded yellow?
solid yellow?
shaded red?
14. EGPWS mode 7 gives warning of:
Windshear, below a radio altitude of 1500 ft?
`Minimums', below a radio altitude of 1000 ft?
Terrain that is well above the projected flight path?
Clear air turbulence?
15. TCAS II transponders transmit an interrogating signal on a frequency of
. . . . . . . . . . . . and respond on a frequency of . . . . . . . . . . . .:
1090 MHz
4200 MHz
1030 MHz
1030 GHz
1030 MHz?
4400 MHz?
1090 MHz?
1090 GHz?
In-Flight Protection Systems
16. A TCAS II message that relates to a possible confliction and requires
avoiding action is known as a:
Traffic advisory?
Resolution advisory?
Manoeuvring advisory?
Vertical speed advisory?
17. A TCAS II symbol depicting a solid yellow circle indicates:
No threat?
Proximate traffic?
Traffic advisory?
Resolution advisory?
18. The minimum altitude at which the TCAS II system will issue a traffic
advisory is:
400 ft?
500 ft?
1000 ft?
1800 ft?
19. The aural warning usually associated with the overspeed warning is:
chime alert?
warbling tone?
20. A stall warning system is set to operate an alarm:
a speed just below stalling speed?
an angle of attack just below stalling angle?
a speed just above stalling speed?
an angle of attack just above stalling angle?
21. The minimum retained data period required by JAR-OPS for a flight
data recorder installed in an aircraft with a take-off weight in excess of
5700 kg and seating for more than 9 passengers is:
180 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
22. The minimum number of parameters required to be covered by a Type I
FDR is:
23. A cockpit voice recorder must automatically begin and cease operating:
a. At take-off and touchdown?
b. During voice transmissions by the flight crew?
c. Before the aircraft first moves under its own power and after it is no
longer capable of moving under its own power?
d. From start of take-off roll to end of landing roll?
Chapter 7
Powerplant and System
Monitoring Instruments
In Chapter 4 we saw how engine and system monitoring is presented to the
pilots by electronic displays such as EICAS and ECAM. In this chapter we
shall be examining the type of instruments often found in less sophisticated
aircraft, and the methods of measurement of pressure, temperature, fluid
flow and quantity, rotary speed, torque and vibration used in all aircraft.
Pressure gauge
In aircraft of British and American manufacture it has been conventional to
measure pressure in units of pounds per square inch (lb/in2 or psi), whereas
most European manufacturers preferred metric units and used kilograms
per square centimetre (kg/cm2). Recently the trend in European manufactured aircraft has been to measure pressure in atmospheric units known
as bars, one bar being equal to 14.7 psi. Low pressures, especially where
related to atmospheric pressure, are usually measured against absolute zero
in inches of mercury (in Hg).
The method of pressure measurement depends largely upon the value of
the pressure to be measured and how it is to be displayed. Not surprisingly,
high pressures such as those found in a hydraulic system, for example,
require more robust methods than the low pressures associated with piston
engine manifolds. In early aircraft it was normal to connect the pressure
instrument (gauge) on the pilot's instrument panel direct to the pressure
source, in which case the pressure measuring device is contained within the
instrument itself. Thus, for example, engine oil pressure is piped to a flexible
element within the pressure gauge and this is known as a direct-reading
instrument. The disadvantages of such a system are that a leak in the connecting pipe not only renders the pressure gauge useless, but also presents a
loss of vital engine lubricant and a potential fire hazard. Furthermore, the
weight of piping required to convey high pressure fluids is not inconsiderable.
To overcome these disadvantages it has become the usual practice in all
but the simplest cases to use remote reading instruments, in which the
182 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
pressure is measured at source and transmitted to the pilot's instruments by
mechanical or electrical means. A simple mechanical transmitter, now rarely
used, consisted of a cylinder containing a free piston. One side of the piston
was connected to the pressure source, say hydraulic system pressure, and
the other was connected to the instrument panel pressure gauge by an
enclosed pipe filled with non-flammable fluid. The force exerted on the
piston by the hydraulic system pressure was transmitted by the enclosed
system to the pressure gauge. Such methods have been replaced almost
entirely by electrical transmission of a signal proportional to the pressure
exerted on a transducer, the signal then operating a suitably calibrated
indicator in the cockpit.
Direct-reading pressure gauges
Where relatively high pressures are involved, such as engine oil pressure or
hydraulic systems, the type of pressure measuring element commonly used
in direct-reading pressure gauges is the Bourdon tube, as illustrated in
Figure 7.1. The device comprises a flattened tube formed into a semi-circle
and closed at one end. The other end of the tube is connected to the pressure
source. The tube is made of a flexible material so that, when pressure is
applied to the inside of it, it tends to straighten. This tendency is opposed by
a spring, or by the natural resistance of the tube itself, so the extent of the
`straightening' movement is proportional to the pressure applied. The closed
end of the tube is connected through gearing to the instrument pointer,
which moves against a calibrated scale to indicate system pressure in the
chosen units of measurement.
Where lower pressures are to be measured the Bourdon tube is not sufficiently sensitive. The pressure measuring elements used are typically
corrugated capsules of the type described in Chapter 1 under air data
instruments. Figure 7.2 illustrates an example of the use of these elements in
a piston engine manifold pressure gauge. Manifold air pressure (MAP)
to linkage
pressure in
Figure 7.1 Bourdon tube principle.
Powerplant and System Monitoring Instruments
Figure 7.2
Manifold air pressure gauge operating principle.
typically ranges from a value less than ambient atmospheric pressure to a
small amount (perhaps 1 or 2 bar) above, and so it must be measured against
absolute zero (i.e. the pressure in a total vacuum). To achieve this, two
sensing capsules are used, one of which is evacuated and spring loaded to
respond to ambient atmospheric pressure and the other of which is connected internally to the engine manifold by piping. The two capsules are
linked mechanically to the instrument pointer.
The manifold air pressure or boost gauge is usually calibrated to read
absolute pressure in inches of mercury (in Hg), thus at sea level with the
engine stopped it will indicate approximately 30 in Hg. When the engine is
running and the pistons are drawing air into the cylinders through the intake
manifold, a partial vacuum is created in the manifold and the gauge will
read less than ambient atmospheric pressure. If the engine is supercharged,
or `boosted', the supercharger or turbocharger will force air into the manifold at higher engine powers and the manifold pressure will be greater than
ambient atmospheric pressure. The MAP is a measure of the power being
developed by a piston engine, hence the reason for indicating it to the pilot.
In the system illustrated in Figure 7.2 the situation depicted is that which
would exist at sea level with the engine stopped. Ambient atmospheric
184 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
pressure is acting on the outside of the aneroid capsule, against the force of
its internal spring, and on the inside and outside of the manifold capsule.
The forces exerted are in balance and the gauge pointer is indicating 30 in Hg
against the calibrated scale. If the engine were to be started and run at low
rpm, the pistons would draw a partial vacuum in the intake manifold and
the force exerted by the manifold capsule would decrease. This is because
the ambient atmospheric pressure acting on the outside of the manifold
capsule tends to compress the capsule and this is transmitted to the gauge
pointer through mechanical linkage and gearing. The amount of compression is restricted by the atmospheric pressure acting on the outside of the
sealed, aneroid capsule and so the pointer movement is proportional to the
change in manifold air pressure.
At increased engine rpm and power the supercharger-boosted pressure in
the manifold becomes greater than atmospheric pressure and the manifold
capsule expands, moving the gauge pointer toward a higher value. The
extent of movement is limited by the opposing force of the spring surrounding the aneroid capsule.
Remote-reading pressure gauges
The use of direct-reading pressure gauges is mainly restricted to small aircraft with a limited number of gauging requirements. The more complex the
aircraft and its systems, the greater the number of pressure measurements
required and the less practical it becomes to pipe these to an instrument
array in the cockpit. Consequently, the various pressures are measured at
source and transmitted electrically to the pilot's instrument displays, which
may comprise individual electrically operated indicators or a computerised
electronic display.
The conversion of pressure into a proportional electrical signal and its
transmission to a calibrated indicating instrument necessarily involves the
conversion of mechanical movement into an electrical output at the measuring source and, in the case of a mechanical indicator on the flight deck, a
reversal of this conversion. There are various types of device for achieving
this, including the synchronous transmission, or synchro, system, the
inductive transmitter and the potentiometer system.
Synchronous transmission
The principle of synchronous transmission was described in Chapter 2
under the direct reading compass, but is repeated here as it applies to the
measurement and transmission of engine oil pressure, and as illustrated in
Figure 7.3.
In the example, oil pressure is sensed in a capsule that expands against a
spring to create linear movement proportional to the measured pressure.
Powerplant and System Monitoring Instruments
stator coils
24 v 400 Hz a.c.
oil pressure
Figure 7.3
pressure gauge
Synchronous transmission system.
This movement is transmitted mechanically to a rotor upon which is wound
a coil carrying alternating current. The rotor is positioned centrally within a
stator having three coils at 1208 spacing. The electro-magnetic field created
around the rotor induces a current flow in each of the stator coils, the
strength of the current in each coil being dependent upon the orientation of
the rotor. These three transmitter currents are fed to, and repeated in, an
identical receiver stator system located behind the cockpit instrument panel,
where they create a magnetic field identical to that in the transmitter. This
field interacts with the a.c. induced field surrounding the receiver rotor coil,
causing the receiver rotor to rotate and adopt an orientation corresponding
to that of the transmitter rotor. The receiver rotor is mechanically connected
to the pointer of a pressure-indicating instrument.
Induction transmitter
In this type of pressure transmission system the pressure to be measured is
led to a capsule inside the pressure transmitter, which is located as close as
possible to the pressure source. The capsule is mechanically connected to a
permanent magnet armature, and linear expansion or contraction of the
capsule moves the armature linearly against the opposition of a spring. The
armature is surrounded by two sets of coils, supplied with current and
186 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
connected to a moving coil indicator on the flight deck. As the armature
moves, its position relative to each of the coils differs, and the inductance of
the two coils will vary in direct proportion to the pressure being measured.
This will cause the output current from the coils to vary, positioning the
pointer of the moving coil indicator pressure gauge accordingly. The principle is illustrated in Figure 7.4.
coil 1
coil 2
coil 1
coil 2
induction transmitter
Figure 7.4 Induction transmitter and ratiometer.
The type of moving coil indicator typically associated with this type of
remote reading pressure gauge is the ratiometer, which is illustrated in
Figure 7.4 above. Current from the transmitter coils is supplied to two coils
wound around armatures on a spindle-mounted iron core. The core is
positioned eccentrically between the poles of a permanent magnet, so that
the air gap between the core and the magnet poles is greater on one side than
on the other. Where the gap is greatest, the strength of the permanent
magnetic field will be weakest and vice versa. Current flowing through the
core coils creates electro-magnetic fields that will interact with the permanent magnetic field. If the two current flows are equal, their magnetic field
strengths will be equal and of opposite polarity, thus cancelling each other
and the iron core will be held stationary. If, however, the current flow in coil
B is greater than that in coil A the stronger magnetic field surrounding coil B
will be attracted toward the larger air gap where the field strength is weaker,
rotating the iron core on its spindle and moving the attached gauge pointer
against a calibrated scale. At the same time, the weaker induced field surrounding coil A is moved into a narrowing air gap where the permanent
magnetic field is stronger. This will eventually arrest the rotation of the core
when the opposition of the permanent field matches the attraction of the
induced field and the gauge pointer will indicate the changed pressure that
caused the current imbalance. If the current flow in coil A is greater than that
in coil B, the effect will be the reverse of that described above.
Powerplant and System Monitoring Instruments
Potentiometer transmission
This type of pressure transmission system uses an inductance transmitter
similar or identical to that described above, but its output current is
amplified and used to drive an a.c. motor, which is connected to the pointer
of the pressure gauge and a potentiometer. In simplistic terms, the potentiometer is a variable resistance connected to an a.c. supply, the output of
which is fed back to the transmitter amplifier. As the motor drives the gauge
pointer, the potentiometer resistance varies until its output current and
phase balances the transmitter signal and supply to the motor ceases,
holding the gauge pointer at its new position.
Piezo-electric transmitters
In most large modern aircraft the transmission of low pressure utilises solidstate transmitters that operate on the piezo-electric principle. These comprise a thin stack of quartz discs impregnated with metallic deposits. When
acted upon by pressure the disc stack flexes and small electrical charges are
produced. The polarity of the induced charge depends upon the direction of
flexing, due to increased or decreased pressure, and the output is amplified
and used to actuate an electronic representation of a pressure indicator.
Pressure gauge indications
Traditionally, aircraft pressure gauges use a central pointer moving
against a circular calibrated scale, as illustrated in Figure 7.5. The gauge
scale is calibrated in the chosen units of pressure measurement and
coloured markings are added in many cases to indicate operating limits
Figure 7.5
Typical piston engine oil pressure gauge.
188 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
and ranges. In the case of the engine oil pressure gauge illustrated, maximum and minimum oil pressures are indicated by radial red lines and
the normal operating range of pressures by a green arc. If these markings are made on the glass cover, rather than the instrument face, it is
required that a white radial line, known as a slip indicator, must be painted on the glass and the adjoining casing to indicate movement of the
glass relative to the casing.
Piston engine manifold air pressure (MAP) gauges are also usually colour
coded to indicate operating power ranges and limits. A red radial line
indicates the maximum permissible MAP for take-off power and blue and
green arcs indicate the lean and rich mixture ranges, respectively. Normally
aspirated (unsupercharged) piston engines with fuel injection systems often
have a fuel pressure gauge with coloured arcs to indicate lean and rich
mixture ranges, since in these powerplants fuel pressure is directly proportional to engine power. In the cruise at reduced power a lean mixture
may be used for fuel economy, but at high power settings it is essential to use
a rich mixture to avoid detonation and engine damage. The coloured arcs are
the same as those on a MAP gauge.
Pressure operated switches
In many cases the pilot does not need to know the actual operating pressure
of a particular system, but merely that it is within acceptable limits. Examples of this are the constant speed drive unit (CSDU) oil pressure, where only
a low pressure warning is necessary, and similarly the hydraulic system
pressure in some light aircraft with very limited hydraulically operated
devices. In these cases it is usual to use a simple transmitter in the form of a
pressure-operated switch connected to a warning light in the cockpit. The
source pressure is applied to a small piston in the switch assembly, the
movement of which is opposed by a calibrated spring. When the source
pressure is within operating limits the spring is compressed and a switch
connected to the piston is held with its contacts open. If pressure falls below
a preset value, the spring overcomes the pressure acting on the piston and
moves the switch to close the contacts and connect supply current to the
warning light. Clearly, the same principle, but with suitably amended calibration, can be adapted to indicate excessive pressure. An example of this is
the oil filter bypass warning light, which will be activated as the filter
becomes clogged and the pressure differential across it increases, to warn the
pilot that the filter bypass valve will open unless the filter is changed at the
earliest opportunity.
Powerplant and System Monitoring Instruments
Temperature gauge
Temperature gauges are used in aircraft piston engines to monitor lubricating oil temperature and cylinder head temperature (CHT). Additionally,
carburettor air intake temperature is measured in some engines, to give
warning of carburettor icing, and exhaust gas temperature may also be
monitored, since this gives an indication of combustion efficiency and is
useful when adjusting mixture settings. In aircraft turbine engines lubricating oil temperature and exhaust gas temperature (EGT) are invariably
monitored. In the case of turbo-propeller engines, turbine inlet gas temperature may be measured instead of EGT.
There are fundamentally two methods of temperature sensing in common
use in aircraft engines and systems: the variable resistance method and the
thermocouple. The variable resistance type of sensing element makes use of
the tendency of metals to change their conductivity as temperature changes,
such that the conductivity decreases (i.e. resistance increases) with increasing temperature. The thermocouple operates on the principle that heat
energy can be converted into electrical energy due to what is known as the
Seebeck effect. The type of sensor used depends mainly upon the degree of
heat involved. The sensing elements that make use of resistive change with
temperature are not generally suitable for use with the high temperatures
associated with exhaust gas, but are ideal for the temperature ranges
experienced in engine lubricating oil systems. The electrical energy generated by the Seebeck effect is small, so this type of sensor is better suited for
measurement of high temperatures.
Resistive systems
A resistive temperature sensing system comprises a sensing element containing a resistance element supplied with low voltage electrical current and
connected in series with an indicator that will convert the electrical output of
the sensor into mechanical movement of an instrument pointer. The electrical supply is usually d.c., but in some cases single phase a.c. may be used.
The sensor element is contained within a closed tube, or `bulb', which is
immersed in the fluid to be measured. As fluid temperature increases, the
resistance of the element will increase and current flow to the indicator unit
will decrease in proportion. A typical sensing unit is illustrated in Figure 7.6.
Sensing unit
The d.c. supply is led to the sensing probe through a two-pin connector. The
resistive element is wound around a central core made of non-conducting
material and enclosed within a leakproof casing made of thin, heatconducting material, typically copper or aluminium. The tube is inserted
190 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
Figure 7.6 Resistive temperature sensing unit.
into the fluid system (e.g. the engine lubricating oil system) and held in place
by a threaded union nut. The calibrating coil shown in Figure 7.6 has a
resistance value that is set during manufacture to determine the temperature/resistance characteristic of the sensing probe for the temperature range
it is intended to measure.
Indicating unit
The temperature gauge associated with resistive temperature measurement
is typically a moving coil instrument, which may be operated by a Wheatstone bridge circuit. The principle of operation of the Wheatstone bridge is
illustrated in Figure 7.7.
24 v d.c. supply
Figure 7.7 Temperature probe and Wheatstone bridge circuit.
The Wheatstone bridge comprises two pairs of series resistances, connected in parallel. The resistances are connected to a low voltage circuit,
typically 24 V d.c. In the diagram above, resistances R1, R2 and R3 are of
identical value, whilst resistance R4 is the element of the temperature probe
Powerplant and System Monitoring Instruments
and varies with the temperature of the sensed fluid. Resistances R1 and R2
are connected in series and form one side of the bridge system, whilst
resistances R3 and R4 are also connected in series and form the other side. A
coil surrounding an iron core is connected from point A, between resistances
R1 and R2, to point B between resistances R3 and R4.
Let us assume for the moment that the temperature of the sensed fluid is
such that the resistance value of R4 is the same as the other three resistances.
Since the total resistance on each side of the parallel bridge circuit is the
same, it follows that the current flow will also be the same on each side.
Bearing Ohm's Law in mind, it therefore follows that the voltage at points A
and B will be identical and there will consequently be no current flow
through the armature coil, since current will only flow from a higher voltage
point to a lower one.
Suppose now the temperature of the sensed fluid increases. The resistance
of the probe element will increase and current flow through resistance R4
will decrease (Ohm's Law again, I = V/R), whilst the current flow through
resistance R2 remains constant. As a result, the voltage at point B will
increase (V = IR), whilst the voltage at point A remains constant; the voltage
difference will cause a current flow from B to A and this current flow will
induce a magnetic field about the coil, concentrated in the soft iron armature.
The armature is situated within a permanent magnetic field, and magnetic
attraction/repulsion will cause the armature to rotate upon its spindle. The
direction of rotation will depend upon the polarity of the armature field,
which is in turn dependent upon the direction of current flow in the
armature coil. The armature is connected mechanically to the pointer of the
temperature gauge.
If the temperature of the sensed fluid were to decrease, the resistance of
the probe element would decrease and current flow through resistance R4
would increase above that through resistance R2. The voltage at point A
would now be greater than that at point B and current flow through the
armature coil would be from A to B, reversing the polarity of the induced
magnetic field. The armature would consequently rotate in the opposite
direction, indicating the reduced temperature on the gauge.
If you find this concept hard to understand, consider the analogy of water
flowing through a similar canal system, as shown in Figure 7.8. If the flow is
restricted at R4, it is clear that it must divert through the path provided by
the interconnecting ditch. If the restriction at R2 and R4 is the same, there will
be no flow through the interconnecting ditch, and so forth.
Thermocouple sensors
Thermocouple temperature measuring sensors require no external electrical
supply, since they directly convert heat energy into electrical energy. They
192 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
Figure 7.8 Wheatstone bridge principle.
operate on the principle that, if two conductors made of dissimilar metals are
connected at either end, as illustrated in Figure 7.9, a potential difference will
exist between the two junctions provided that there is a temperature difference between the junctions. The value of the potential difference will be
directly proportional to the temperature difference and, since the two joined
conductors form a loop, will cause current to flow around the loop. Clearly,
the greater the temperature/potential difference the greater the current flow
and, when suitably amplified, the thermocouple current can be made to
operate the indicator of a temperature gauge.
Figure 7.9 Thermocouple principle.
For the measurement of piston engine cylinder head temperature, where
the temperature range is typically of the order of 4008C to 8508C, the metals
used are usually copper and copper±nickel alloy, or iron and copper±nickel
alloy for the higher end of the range. For gas turbine exhaust gas temperature measurement, where the maximum temperature may be as high as
11008C, nickel±aluminium and nickel±chromium alloy conductors are
usually used.
Cylinder head temperature thermocouples typically take the form of a
`washer' bolted to the cylinder head and forming the hot junction of the
thermocouple. The cold junction is at the amplifier of the temperature
Powerplant and System Monitoring Instruments
hot junction
cylinder head
temperature gauge
surface contact
Figure 7.10
Surface contact sensor.
indicator. This is known as a surface contact sensor and is illustrated in
Figure 7.10.
The measurement of gas temperature uses an immersion sensor which, as
its name suggests, consists of a probe immersed in the hot gas flow, containing the hot junction of the thermocouple. The cold junction is at the
indicator as before. An immersion probe is illustrated in Figure 7.11. In all
cases the indicator contains a compensating device that automatically allows
for variations in temperature at the indicator.
Figure 7.11
exhaust gas
Immersion sensor.
Turbine exhaust gas temperature is usually measured within the jet pipe
as close as possible to the turbine outlet. In order to allow for the harsh
conditions, in which probes might become damaged, it is usual for a number
of probes to be connected in parallel and positioned radially at intervals
around the perimeter of the jet pipe.
194 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
Air temperature measurement
The measurement of intake air temperature is important in piston engines to
provide warning of potential carburettor icing and in piston and gas turbine
engines for the determination of engine performance. In the case of gas
turbine performance, it is preferable that the measured temperature should
be the static air temperature (SAT) at the altitude at which the aircraft is
operating. However, as was discussed in Chapter 1 under air data computers, SAT cannot be directly measured because heating occurs due to compression and friction. The increased temperature due to this heating is
known as the ram rise and the resultant temperature as total air temperature
(TAT). The ram rise can be calculated and the air temperature gauge reading
can be corrected, either automatically or from a chart, to give SAT. The
ability of a temperature sensor to measure the full extent of the ram rise
effect is known as its recovery factor and, in most modern sensors, is close to
Meaning of coloured arcs
Temperature gauges, especially those used in conjunction with piston
engines, usually have coloured arcs and radial lines to indicate operating
temperature ranges and limits.
Engine oil temperature
The piston engine lubricating oil temperature gauge is usually marked with
a green arc, typically between 608C and 708C, to indicate the normal operating temperature range. Red radial lines indicate minimum and maximum
safe operating temperatures and these are typically 408C and 1008C
Carburettor intake temperature
The carburettor air intake temperature gauge typically has three arcs
coloured yellow, green and red. The yellow arc extends from 7108C to
+158C and indicates the temperature range within which a carburettor icing
hazard exists. The green arc extends from +158C to +408C and indicates the
normal operating temperature range. The red arc begins at +408C and
indicates intake temperatures that are liable to cause detonation within the
Exhaust gas temperature
Exhaust gas temperature gauges are usually marked with red coloured arcs
or a red radial line to indicate maximum temperature ranges or limits.
Powerplant and System Monitoring Instruments
Vapour pressure gauge
A few light aircraft still use a very simple form of temperature gauge that
operates on the Bourdon tube principle. The gauge is connected by capillary
tube to a bulb filled with a highly volatile liquid. The bulb is immersed in the
medium to be sensed (e.g. the engine oil system) and the heat of the medium
vaporises the liquid in the bulb. The pressure in the closed system of bulb
and tube increases in direct proportion to the medium temperature and acts
upon the gauge Bourdon tube to move a pointer against a scale graduated to
indicate temperature.
RPM indicator
The measurement of engine revolutions per minute (rpm) is important in
unsupercharged piston engines, since it is an indication of the power being
delivered to the propeller. Similarly, in gas turbine engines, rpm is related to
thrust, although it is more common to measure this in terms of engine
pressure ratio (EPR). In early single-engine aircraft the pilot's rpm indicator,
or tachometer as it is properly known, was usually driven directly from the
engine by means of a flexible drive and a system of flyweights. As aircraft
became more complex this method became impractical and electrical
transmission of the measured rpm to the pilot's instruments was developed.
Electrical tachometer
The method commonly used to achieve electrical transmission of rpm in
piston engines is illustrated in Figure 7.12. A small three-phase a.c. generator
is driven from the engine accessories gearbox and its output is used to drive
a synchronous motor, which operates the pilot's engine tachometer. The a.c.
frequency of the generator output will vary directly with engine rpm, and it
is the supply frequency that determines the speed of rotation of an a.c.
synchronous motor. Hence, the higher the engine rpm, the greater the rotary
speed of the tachometer motor.
a.c. generator
Figure 7.12
Electrical tachometer system.
196 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
The tachometer indicator is usually a conventional pointer moving
through an arc against a calibrated scale, so it is clear that the continuous
rotary motion of the synchronous motor must be converted into semi-rotary
movement of the tachometer pointer. This is achieved by means of a magnetic device called a drag cup. Mounted on the rotor shaft of the synchronous motor is a permanent magnet which rotates inside an aluminium cup.
The rotating field of the permanent magnet sets up eddy currents in the
aluminium drag cup. These create electro-magnetic forces that react with the
rotating permanent magnetic field and create a rotary force on the drag cup.
The drag cup rotary movement is restrained by a coil spring attached to the
shaft connecting it to the tachometer pointer. Thus, its amount of deflection
is dependent upon the electro-magnetic force acting on the drag cup, which
is in turn dependent upon the speed of rotation of the permanent magnet.
The degree of pointer movement is therefore directly proportional to the
engine rpm.
A typical piston engine rpm indicator is shown in Figure 7.13. It will be
noted that the actual engine rpm is indicated, as is normal with piston
engines. Gas turbine engine tachometers usually indicate rpm as a percentage, where 100% rpm is the optimum engine rotary speed. In the case of
piston engine tachometers, a green coloured arc may indicate the normal
operating rpm range, with a red radial line or arc to indicate maximum
permissible rpm and time-limited operating rpm ranges.
Electronic tachometer
A type of electronic tachometer sometimes used with piston engines
converts the impulses from the engine magneto into voltage, to drive an
X 100
0008 5
Figure 7.13 Typical piston engine rpm indicator.
Powerplant and System Monitoring Instruments
indicator pointer. Clearly, the higher the engine rpm, the more impulses per
minute from the magneto and the higher the voltage from the conversion
circuit. When supplied to a voltmeter calibrated to read rpm, the amount of
pointer deflection will be directly proportional to engine rpm. The advantage of both this type of tachometer and the electrical type is that they require
no external electrical supply and will continue to operate in the event of
failure of normal aircraft electrical services. There is a more complex type of
magneto-driven electronic tachometer that requires a transistorised amplifier circuit needing a 12 V supply from the aircraft electrical system.
Servo-operated tachometer
Some gas turbine engine tachometers use a variation of the electrical tachometer described above, in which the generator output is converted into a
square waveform by a solid-state circuit. A `square' pulse is formed each
half-cycle of the generator output, resulting in a pulse repetition frequency
that is twice the a.c. generator output frequency. The pulsed transmission
produces direct current (d.c.) to drive a d.c. motor, which operates the
tachometer indicator pointer. The d.c. voltage, and therefore the motor
speed, is dependent upon the pulse repetition frequency, which is in turn
dependent upon engine rpm. This system is not independent of the aircraft
electrical system, since it includes an overspeed pointer mechanism, which
requires an external 28 V d.c. supply to reset it.
Tacho-probe system
This method of rpm measurement, illustrated in Figure 7.14, is commonly
used with gas turbine engines as it has a number of significant advantages.
Not only is it electrically independent, but its output can be used to supply
flight data and autothrottle systems as well as to operate the rpm indicator.
phonic wheel teeth
co-incident with pole pieces
pole pieces
magnetic field cutting through coil
Figure 7.14
Tacho-probe system.
magnetic field collapsed
198 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
A toothed wheel is mounted on a shaft, the rpm of which is to be measured. In gas turbine engines this is usually the HP compressor/turbine
shaft, but in many turbo-fan engines the fan rpm is also measured. This
wheel is known as a phonic wheel and it clearly rotates at the same
speed as the HP shaft or fan shaft. Mounted on the engine casing adjacent to the phonic wheel is a probe unit comprising a permanent magnet, two pole pieces that are spaced to exactly match the spacing of the
phonic wheel teeth and sensing coils in which electrical current is generated.
When two of the phonic wheel teeth are exactly opposite the two pole
pieces of the probe the permanent magnetic field surrounding the coils is at
maximum strength, as in Figure 7.14(a). As the wheel rotates and the teeth
are no longer co-incident with the pole pieces, as in Figure 7.14(b), the
magnetic field strength surrounding the coils falls to near zero. This fluctuating field strength through the coil windings induces voltage in the coils
and an alternating current flows through the associated output circuit. The
frequency of the induced a.c. is directly proportional to the rotary speed of
the phonic wheel and is used to actuate the rpm indicator pointer to show
shaft speed as a percentage.
A typical gas turbine rpm indicator is shown in Figure 7.15. The large
pointer shows percentage rpm from 0% to 100% in 10% increments. The
smaller scale and pointer shows percentage rpm increments in unitary
values from 0% to 10%. In the display illustrated the gauge is indicating 94%
of optimum rpm.
14 12 10
Figure 7.15 Typical gas turbine tachometer.
Powerplant and System Monitoring Instruments
Flight hour meter
Figure 7.13 shows an hour meter incorporated in the piston engine tachometer display. It is important for routine servicing, inspection and component life to have a record of the number of flight hours accumulated by the
aircraft. In light general aviation aircraft this is often achieved by a meter
activated only when the engine is operating at cruise rpm. For example, if
the normal cruise rpm is 2200 it follows that the engine will complete 132 000
revolutions in one hour at cruising speed. Thus, for each 132 000 revolutions
the flight hour meter progresses by one unit. Such a meter suffers from the
obvious limitation that it does not record flight times at other than cruise
rpm. However, for aircraft that operate at this speed for all but take-off and
landing, this is adequate and has the advantage of being independent of
aircraft electrical power.
A slightly more sophisticated system is operated by an electric clock,
powered from the aircraft electrical system and activated by the landing gear
`weight-on' or `squat' switch. This system truly records flight time, since the
switch supplying power to the clock only closes when aircraft weight is off
the landing gear and the battery master switch is closed. This type of flight
hour recorder is known as the Hobbs meter.
Fuel consumption gauge
It is important for the flight crew to be aware of the rate at which fuel is being
consumed in flight in order to calculate range, endurance and economy. This
information is vital to the operation of automated thrust and flight control
systems. In large gas turbine powered transport aircraft the measurement of
fuel flow is made by relatively complex flow metering systems that are
capable of integrating the flow rate with time to compute and display both
rate of fuel flow and the total fuel consumed. In smaller, short range
transports, particularly those powered by piston engines, it is usual to
measure only fuel flow rate using less sophisticated devices.
Fuel flow is ideally measured in terms of mass flow rather than volumetric
flow, since it is the mass of fuel consumed that determines the power output
of an engine. Mass per unit volume varies with the density of the fuel, which
in turn varies with its temperature. Thus, if fuel flow is measured in terms of
volumetric rate (gallons or litres per hour) a further calculation is necessary,
taking temperature into account, in order to determine the mass flow rate. In
short range aircraft this is less important, but in long range transports the
flowmeter calibration usually takes fuel temperature into account and
computes mass flow rates. Mass flow is usually measured in pounds (lb) or
kilograms (kg) per hour.
200 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
Fuel flow and pressure
With a continuous flow fuel-injected piston engine the rate of fuel flow to the
injectors is proportional to the fuel supply pressure. In piston-engine aircraft
that are equipped with direct fuel injection systems it is not uncommon for
the fuel pressure gauge to be calibrated to indicate fuel flow rate as well as
fuel pressure. Since such aircraft are not usually long range types it is adequate for the fuel flow indication to be volumetric, in gallons per hour or
litres per hour. An example of a fuel pressure gauge calibrated to serve also
as a flowmeter is shown in Figure 7.16. It will be noted that the calibration
indicates the fuel flow rates for various power settings (e.g. cruise, take-off
and climb).
CLIMB 8000
OFF 4000
16 14
Figure 7.16 Fuel pressure gauge/flowmeter.
Vane-type flowmeter
The principal disadvantage of the combined pressure gauge/flowmeter is
that it will provide erroneous indications if the fuel pressure is artificially
high. Suppose, for instance, that an injector has become partly blocked. The
fuel being consumed by the engine will be less, because of the blockage,
whilst the fuel supply pressure will increase because of it. Consequently, the
fuel flow indication would be falsely high. To measure actual flow rates it is
necessary to place a device in the fuel supply line to the engine which will
convert fuel flow into mechanical movement and transmit a signal to a fuel
flow gauge in the cockpit. A simple type of flow measuring device, used in
some piston engine and smaller gas turbine engine aircraft, is shown in
Figure 7.17.
Fuel is directed through a volute chamber in which there is a springloaded vane. The action of the fuel pressing on the flat vane causes it to rotate
Powerplant and System Monitoring Instruments
by-pass valve
fuel in
fuel out
Figure 7.17
coil spring
Vane-type flowmeter.
against the force of a coil spring until the gap between the edge of the vane
and the inside of the volute chamber no longer impedes fuel flow. The
rotation of the vane is transmitted to a fuel flow gauge in the cockpit by
means of a synchro system. The greater the fuel flow, the more the vane will
be deflected against the force of the coil spring before the gap is sufficiently
wide. To protect against failure of the vane or other blockage in the volute
chamber, a lightly loaded by-pass valve will open in the event of the differential pressure across the chamber exceeding a preset value, thereby
maintaining fuel flow to the engine.
Integrated flowmeter
Larger, long-range, turbine-powered transports employ a more complex
type of fuel flowmeter that is much more accurate than the foregoing systems and which integrates the measured flow rate with time, either
mechanically or electronically, to compute and indicate fuel consumed as
well as fuel flow. An example of the fuel flow measuring system is shown in
Figure 7.18. The flowmeter is situated in the high pressure fuel supply to the
engine and contains an impeller driven at constant rotary speed by an a.c.
synchronous motor. Fuel entering the flowmeter passes through the
impeller, which imparts a swirling motion to the fuel flow. The flow then
enters a turbine where the force of the swirling fuel striking the turbine
vanes drives the turbine to rotate in the same direction as the impeller
rotation. The rotary motion of the turbine is restrained by a coil spring
attached to the turbine shaft. As flow rate increases, the force of the swirling
fuel acting on the turbine vanes increases, rotating the turbine further
against the spring.
The rotary motion of the turbine shaft is sensed by a low voltage trans-
202 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
a.c. supply
Figure 7.18 Integrated flowmeter.
former (LVDT), which produces an output voltage proportional to the
amount of rotation. This output voltage forms the signal transmitted to the
fuel consumption display, which is typically a digital readout in kilograms
or pounds of fuel consumed. The signal also operates a servo-motor, which
positions the pointer of an analogue fuel flow gauge, showing fuel flow in
kilograms or pounds per hour. The two displays are often on the same
instrument and an example is shown in Figure 7.19.
kg/hr x 1000
Figure 7.19 Combined fuel flow and consumption gauge.
Failure of the integrated flowmeter impeller or turbine will not impede the
fuel flow to the engine, but will obviously render the gauge readings useless.
Failure is typically indicated by a warning flag on the instrument display.
The flowmeter circuitry may also incorporate a low fuel flow warning.
Powerplant and System Monitoring Instruments
Fuel quantity measurement
In addition to knowing the rate at which fuel is being consumed, and the
cumulative total of fuel consumed, it is clearly essential that the pilots are
also aware of the quantity of fuel remaining in the aircraft tanks and so some
form of tank gauging is necessary. One of the simplest forms of aircraft fuel
tank gauge ever devised comprised a float-operated vertical dipstick, which
protruded through a hole in the top of the tank at a location visible from the
cockpit. The extent of dipstick protruding directly indicated the quantity of
fuel remaining in the tank.
Resistive fuel quantity measurement
Such a system as that described above is obviously impractical for a large
transport aircraft where the pilots and the automated systems need to know
the exact quantity of fuel in each tank. However, in smaller aircraft a system
of remote tank contents indication is often employed which does not differ
greatly in principle from the float-operated dipstick. A float in the tank
operates the wiper of a potentiometer, the output of which drives a moving
coil instrument. As fuel level in the tank varies, the float moves up or down
accordingly and operates the potentiometer, which is supplied with low
voltage from the aircraft electrical system. The potentiometer signal causes
the moving coil in the indicating gauge to position itself according to the
strength of the signal, positioning the gauge pointer against a calibrated
scale to show tank contents in gallons or litres. The principle of this method
of fuel quantity measurement is illustrated in Figure 7.20.
d.c. supply
Figure 7.20
Resistive tank contents measurement.
fuel level
204 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
This system of measuring the amount of fuel remaining in a tank suffers
from a number of potential inaccuracies and is only suitable for use in the
relatively small fuel tanks of light general aviation aircraft. Movement of fuel
in the tank, due to changes of aircraft attitude or acceleration, will cause the
float to move up or down and the system will falsely indicate the fuel
quantity in the tank. Furthermore, since the system can only measure the
level of fuel in the tank, it will give a false indication of quantity if the level
changes due to temperature change. As both of these causes of inaccuracy
are limited in the small tanks of light aircraft, they are generally acceptable.
Additionally, the system is only capable of indicating the volume of fuel
contained in the tank, computed from the fuel level and the known tank
dimensions, so the associated fuel tank contents gauge must be calibrated in
litres or gallons, rather than the preferred kilograms or pounds.
Capacitive fuel quantity measurement
The fuel tank contents in large aircraft, with a significant number of correspondingly large tanks, are measured by a system that uses the electrical
capacitance of the fuel to determine the exact quantity of fuel in each tank
and to indicate it in terms of mass rather than volume.
A capacitor is an electrical device consisting basically of two conducting
plates separated by a resistive medium known as a dielectric. Such a device
is capable of storing an electric charge and this property is known as capacitance. The principle of operation of a capacitor is shown in Figure 7.21.
When the two conducting plates of the capacitor are connected to an
electrical supply, a potential difference exists between the plates. In the
example in Figure 7.21 the switch in the circuit has been placed in position A
and a 12 V supply has been connected to the plates. Current cannot flow
between the plates since they are separated by a non-conducting dielectric,
and so the potential difference between the plates is 12 V. If the switch is now
moved to position B, the potential difference will continue to exist between
the plates, since there is nowhere for it to discharge to. If the switch is now
moved to position C the capacitor will discharge its stored electrons through
the circuit provided.
In this example direct current is used and it will be noted that the capacitor
only becomes charged, or discharges, when the switch is operated. If,
however, a capacitor is supplied with alternating current, the constantly
changing voltage and polarity of the supply acts in a similar manner to the
switch. When voltage is increasing, the capacitor is charging as the potential
difference between the plates increases. As the supply voltage decreases, the
potential difference between the plates becomes greater than the supply
voltage and the capacitor discharges. This process is continuously repeated
through each half cycle of the a.c. supply.
Powerplant and System Monitoring Instruments
12 V
Figure 7.21
Capacitor principle.
Capacitance, the ability to store an electric charge, depends upon the
surface area of the capacitor plates and the permittivity of the medium
separating the plates. Permittivity is also sometimes referred to as the
dielectric constant of the separating medium. It is usually measured as a
relative value, where air has a permittivity of 1.00, so other media are
assigned a relative permittivity indicating the capacitance they offer relative
to air. Aviation kerosene, for example, has a relative permittivity of 2.10.
The capacitance or `charge-holding capability' of a capacitor is the ratio
between the charge supplied and the potential difference between the two
plates and is measured in picofarads (pF). The strength of the discharge
current from a capacitor will depend upon its capacitance and the rate of
change of the supply voltage. The latter is a constant given a constant frequency a.c. supply and the capacitance, as we have seen, is dependent upon
the permittivity of the medium. Suppose two identical capacitors are supplied with 6 V a.c. at a frequency of 400 Hz and one is placed in air whilst the
other is placed in aviation kerosene. The capacitor with air separating its
plates will produce a lower discharge current than the one with fuel
separating its plates.
The capacitive probe inserted into an aircraft fuel tank is basically two
open-ended tubes, one inside the other. The tubes act as the capacitor
`plates', separated by air or fuel, depending upon the depth of fuel in the
tank. The concept is illustrated in Figure 7.22.
206 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
Figure 7.22 Effect of fuel level on capacitance.
In Figure 7.22(a) the tank is empty and the separating medium is air, with
a relative permittivity of 1.00. For the sake of simplicity, let us assume that
the capacitance of the probe is 100 pF. In Figure 7.22(b) the tank is filled with
aviation kerosene with a permittivity of 2.10, so the capacitance of the probe
will now be 210 pF, an increase of 110 pF, and its discharge current, when it is
supplied with constant frequency a.c., will be 2.10 times greater than when
the tank was empty. If the tank is half full, as shown in Figure 7.22(c), the
increase in capacitance over the `tank empty' value will be half as much, i.e.
55 pF. Thus, the capacitance of the probe will be 155 pF and the discharge
current will be 1.55 times the empty value. From this it can be seen that the
tank fuel level is accurately represented by the probe discharge current, and
this is used to operate the tank contents gauge.
Compensation for movement of fuel in the tank due to aircraft attitude or
acceleration changes is easily made by inserting several probes at different
locations within the tank and `averaging' their outputs to give a mean
reading. Smaller tanks typically have two probes, whereas larger ones may
have as many as six.
A simplified capacitive tank contents measuring system is shown in
Figure 7.23. The low voltage alternating current supply to the system is from
a power transformer to the fuel tank capacitor probe and to a reference
capacitor in parallel with it. Because the fuel tank should never be completely empty, the charge from the probe will always be greater than the
constant value of the reference capacitor. The higher the fuel level in the
tank, the greater the difference and it is this potential difference that actuates
the voltmeter, which is calibrated to read tank contents. A complete capacitive system is more complex than this, since it contains circuitry that allows
the indicator gauge to be `zeroed', that is the maximum and minimum
readings to be adjusted for purposes of calibration.
Powerplant and System Monitoring Instruments
115 V
400 Hz
Figure 7.23
Simplified capacitive gauging system.
Fuel quantity by weight
The observant reader will have noted that, so far, all references have been to
the measurement of fuel level in the tank, but it is ideally required that the
measurement of fuel contents should be by weight. We have already
established that fuel will expand as its temperature increases and the level of
fuel in the tank will increase, whilst the important factor, its weight, remains
the same. Therefore, if the gauges are calibrated to show quantity by weight,
one would expect them to be in error.
However, the increase in volume with no change in weight means, by
definition, that the fuel density has decreased. The reduction in density
reduces the permittivity of the fuel and so the capacitance of the probe is
reduced. This reduction almost exactly mirrors the increase in capacitance
due to the higher level of fuel and the two effectively cancel each other out.
Thus, the capacitance method of fuel tank measurement automatically
compensates for changes in fuel density due to temperature changes.
Torque meter
The function of the torque meter is to measure and indicate the power
developed by a turbo-propeller engine. The turning moment, or torque,
delivered to the propeller through the reduction gearing is proportional to
the horsepower developed, which is the product of the torque and the
propeller rpm.
In the example of a torque meter system shown in Figure 7.24, the
reduction gearbox between the engine and the propeller uses a type of
208 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
fixed ring gear
planet gear
Figure 7.24 Torque meter system.
gearing known as sun and planet, or epicyclic. A large gearwheel, with
helical teeth, drives smaller helical-toothed gearwheels surrounding it.
These are the sun and planet wheels from which the name of the system is
derived. The planet wheels rotate within a fixed ring gear and are carried in
a drum attached to the propeller shaft, thereby transmitting rotation to the
propeller shaft.
The torsional force of the engine turning the propeller is transmitted
through the planet gears, the helical teeth of which transfer some of that
force into an axial direction. The shafts of the planet gears form pistons that
fit into cylinders machined in the carrier drum. These cylinders are connected to an enclosed hydraulic system, supplied with oil from the engine
lubricating system and connected to a pressure gauge. Axial movement of
the sun gears is restrained by the hydraulic system, but the greater the
torsional force, and consequent axial force, the greater the hydraulic pressure created. The gauge reading is usually calibrated to read torque in ft lb
and can be used to calculate the brake horsepower (bhp) being delivered to
the propeller using the formula bhp = TNK, where T is the torque in ft lb, N
the engine rpm and K a constant (2p/33 000).
In some systems the gauge may also show negative torque, the undesirable condition that exists when the propeller is windmilling and tending to
drive the engine. The torque meter output is used in some turbo-propeller
systems to automatically operate the propeller feathering mechanism in the
event of engine failure and excessive negative torque.
Powerplant and System Monitoring Instruments
An alternative type of torque measuring system makes use of the fact that,
under torsional loading, the propeller drive shaft twists. Strain gauges
attached to a torque ring mounted on the engine/propeller drive shaft twist
and deform with the shaft as torque increases. These produce a small electrical signal that varies according to the deformation of the strain gauges and
is amplified to actuate the torque meter on the flight deck.
The torque meter gauge scale may incorporate coloured arcs. A green arc
indicates the normal operating range of positive torque, a yellow arc indicates negative torque and red radial lines or arcs indicate maximum and
minimum torque limits or ranges.
Vibration monitoring
Unlike piston engines, gas turbines have no reciprocating parts and the
rotating assemblies are finely balanced dynamically. Consequently, they are
much less prone to vibration under normal circumstances and any abnormal
vibration is a clear indication of loss of dynamic balance due to damage. This
may be caused by factors such as erosion, distortion or chipping of turbine
blades, or ingestion damage to fan or compressor blades.
The vibration sensor comprises a permanent magnet suspended on
springs and mounted on the engine and/or fan casing such that it is sensitive
to radial oscillations. A pick-off coil surrounding the permanent magnet is
connected through suitable circuitry to a vibration indicator and a warning
circuit. The principle of operation is illustrated schematically in Figure 7.25.
engine casing
Figure 7.25
Vibration monitor.
210 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
Significant radial vibration or oscillation of the casing will only occur if
there is a loss of dynamic balance in the adjacent rotating assembly. Typical
locations for the vibration sensor are on the HP turbine casing and the fan
casing on turbo-fan engines. As the casing and the attached sensor vibrate
radially, the spring-suspended magnet remains virtually stationary. Thus,
the coil of the sensor unit oscillates rapidly relative to the permanent magnet.
The magnetic field is therefore cut by the coil, inducing a voltage in the coil
proportional to the rate and extent of vibration.
The voltage output of the coil is amplified and integrated by a system of
filters to eliminate known normal vibrations and to isolate vibration frequencies associated with particular components such as turbines, compressors, gearing, etc. The signal is then rectified and supplied to the
vibration indicator on the flight deck, which indicates vibration in units of
amplitude relative to a fixed datum value. The signal also feeds a warning
circuit that illuminates a warning light when vibration exceeds a predetermined value.
Engine vibration is measured in units of relative amplitude.
More recently developed vibration monitoring systems employ piezoelectric sensors in conjunction with CRT displays such as EICAS and ECAM.
Remote signal transmission systems
The flying control surfaces and landing gear systems are not visible from the
cockpit of most aircraft, but knowledge of their positions, particularly secondary control surfaces and landing gear, is important to the flight crew.
In some light aircraft the position of the landing gear and flaps is indicated
to the pilot through a mechanical system of cables, push-pull rods and linkages. In most aircraft, however, transmission of position indicating signals
is electrical.
For simple up, down, locked or unlocked indications the cockpit presentation may be in the form of lights or `doll's eye' captions actuated by
microswitches in the landing gear and flap actuating mechanisms. More
complete indications showing the full range of movement of primary or
secondary flying control surfaces may be transmitted by synchronous
transmission systems of the type illustrated in Figure 7.3 or induction
transmitters as shown in Figure 7.4.
Remote control
Remote control of primary and secondary flight controls and landing gear
may be mechanical, electrical, hydraulic or a combination of all three. In light
Powerplant and System Monitoring Instruments
general aviation aircraft, remote control transmission systems are usually
mechanical, using push-pull rods, chain drives, cable-and-pulley systems
and linkages between the pilot's controls and the systems to be operated.
Larger aircraft typically use hydraulic transmission of movement from the
pilot's controls to hydro-mechanical servo-actuators. In many large modern
transport aircraft the remote transmission of signals from the pilot's controls
to the hydro-mechanical servo-actuators is electrical.
Direct mechanical remote control systems have the advantage of simplicity and ease of maintenance, but are unsuitable for use in larger aircraft
where the control loads are too great for direct manual force. To provide the
power necessary to overcome the aerodynamic loads of large control surfaces and the weight of heavy landing gear, high pressure hydraulic
actuators are required. The principal disadvantage with hydraulic systems is
the requirement for absolutely pressure-tight systems and the fire and corrosion hazards associated with fluid leakages. Electrical transmission of
remote control signals largely overcomes these disadvantages, although
hydraulic actuators are still needed to move the control surfaces and landing
gear. Electrical transmission has the added advantage of direct compatibility
with computerised and automatic control systems.
Sample questions
1. A hydraulic system operates at a pressure of 205 bar. This pressure is
equivalent to:
3015 psi?
2050 psi?
310 kg/cm2?
250 kg/cm2?
2. The pressure sensing element of a manifold pressure gauge typically
A Bourdon tube?
A single aneroid capsule?
A single diaphragm?
Two capsules, one pressure and one evacuated?
3. A synchronous transmission system requires:
A d.c. supply?
An a.c. supply?
No electrical supply since it is self-generating?
A synchronous motor at the transmitting element?
212 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
4. A potentiometer:
Is a solid state electronic device?
Is an electro-magnetic instrument?
Is a rotary variable resistance device?
Is used to measure potential difference?
5. The blue arc on a manifold air pressure gauge indicates:
rich mixture range?
take-off power range?
lean mixture range?
icing hazard range?
6. The resistive type of temperature gauge operates on the principle that:
The resistance of a conductor decreases indirectly with temperature?
The resistance of a conductor increases directly with temperature?
Electrons will travel along a conductor from hot to cold?
Electrons will travel along a conductor from cold to hot?
7. A thermocouple:
Requires an a.c. supply?
Requires a d.c. supply?
Operates on the principle that conductivity varies with temperature?
Requires no external electrical supply?
8. The cylinder head temperature of a piston engine is typically measured
A surface contact thermocouple?
An immersion thermocouple?
A resistive temperature measurement system?
A capacitive temperature measurement system?
9. The yellow arc on a carburettor air intake temperature gauge indicates:
icing hazard range of temperatures?
detonation hazard range of temperatures?
normal operating range of temperatures?
approach power range of temperatures?
10. An electrical tachometer system employs:
a. A two-phase a.c. generator and motor?
b. A three-phase a.c. generator and synchronous motor?
Powerplant and System Monitoring Instruments
c. A d.c. generator and motor?
d. A synchro transmission system and moving coil indicator?
11. When a gas turbine tachometer is indicating 100% rpm, this indicates:
maximum rpm?
normal rpm?
take-off power?
optimum rpm?
12. A tacho-probe rpm measurement system:
Requires an a.c. electrical supply?
Requires a d.c. electrical supply?
Requires no external electrical supply?
Is only suitable for use with piston engines?
13. The impeller of an integrated flow meter:
Is driven by the turbine?
Is driven by an a.c. synchronous motor?
Is driven by the action of the fuel flow?
Drives the turbine at constant speed?
14. A resistive fuel quantity gauging system:
Uses a potentiometer to convert fuel level into an electrical signal?
Uses a capacitor to convert fuel level into an electrical signal?
Typically displays fuel quantity in terms of weight?
Automatically compensates for changes in fuel density due to temperature?
15. The permittivity of aviation kerosene is:
16. A capacitive fuel quantity gauging system is able to measure tank contents by weight because:
The permittivity of fuel varies directly with its density?
There are a number of capacitive probes in each tank?
The permittivity of fuel remains constant regardless of temperature?
The density of fuel remains constant regardless of temperature?
Answers to Sample Questions
Chapter 1: Air Data Instruments
Chapter 2: Gyroscopic Instruments and Compasses
Answers to Sample Questions
Chapter 3: Inertial Navigation Systems
Chapter 4: Electronic Instrumentation
216 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems
Chapter 5: Automatic Flight Control
Chapter 6: In-Flight Protection Systems
Answers to Sample Questions
Chapter 7: Powerplant and System Monitoring Instruments
active control system, 149
air data computer, 26
analogue, 26
digital, 26, 29
air temperature measurement, 194
SAT, 194
TAT, 194
aircraft integrated data system (AIDS), 174
airspeed indicator (ASI), 12
airspeed measurements, 14
ASI errors, 17
square law compensation, 15
tolerance, 18
alternate static source, 4
altimeter errors, 11
altimeter tolerances, 12
altitude, definition of, 9
altitude, rate of change with, 7
aneroid capsule, 6
angle of attack sensing, 168
leading edge vane, 168
plenum chamber, 169
sensor unit, 170
apparent drift, 37
apparent topple, 38
atmospheric pressure, 5
attitude director indicator (ADI), 116, 122
attitude indicator (electrical), 50
acceleration and turning errors, 51
erection speed 52
erection system, 51
warning indications, 52
attitude indicator (pneumatic), 46
acceleration errors, 49
aircraft symbol, 47
construction and principle of operation, 46
erection mechanism, 48
horizon bar, 47
sky plate, 47
turning errors, 50
automatic control, 116
automatic flight control system (AFCS), 129, 134
automatic landing (autoland), 138
alert height, 139
fail operational system, 139
fail passive system, 138
multiplex system, 139
sequence, 139
automatic pitch trim, 151
automatic stabiliser trim unit, 152
automatic thrust control (autothrottle), 116, 141
automatic trim tab pitch trim, 153
autopilot, 129
auto-stabilisation, 129
closed loop, 129
control surface actuation, 132
engagement, 132
inner loop, 129
interlocks, 133
manual inputs, 133
outer loop, 130
single axis, 129
torque limiters, 132
two-axis, 130
autopilot flight director system (AFDS), 134
command modes, 136
control panel, 134
paddle switches, 136
blocked pitot tube, 4, 18, 22
blocked static source, 4, 18, 22, 25
capacitor principle, 205
CAS, 14
cathode ray tube (CRT), 94
cockpit voice recorder, 174
coloured arcs, 187, 194, 196, 209
command bars, 116
compressibility error, 17
control wheel steering (CWS), 133
critical mach number (Mcrit), 19
deep stall, 168
density altitude, 10
deviation (magnetic), 62
dielectric, 204
directional gyro, 41
adjustment procedure, 42
caging mechanism, 42
comparison with magnetic compass, 46
drift calculations, 44
drift compensation, 45
effect of friction, 46
erection system, 43
gimbal error, 44
Dutch roll, 150
dynamic pressure, 1
dynamic vane type VSI, 25
EADI, 95
EAS, 15
EHSI, 97
electrical tachometer, 195
drag cup, 195
electronic flight instrument system (EFIS), 86
ILS modes, 98, 102
MAP mode, 97, 99
PLAN mode, 97, 102
VOR modes, 98, 103
electronic tachometer, 196
engine centralised aircraft monitoring (ECAM), 86, 109
control panel, 109
operating modes, 110
engine indicating and crew alerting system (EICAS), 86,
display select panel, 107
maintenance mode, 105
operational mode, 103
primary display, 103
secondary display, 103
standby engine indicator, 108
status mode, 105
engine pressure ratio (EPR), 145
flight data recorder, 171
types, 172
flight director system, 122
gain programme in approach mode, 128
lateral and vertical beam sensors, 128
modes of operation, 127
warning indications, 126
flight envelope protection, 149
flight hour meter, 199
Hobbs meter, 199
flight management system, 116
alphanumeric keyboard, 120
multipurpose control and display unit (MCDU), 118
fly-by-wire, 149
free gyro, 34
fuel consumption gauge, 199
flow and pressure, 200
flow meter ± integrated, 201
flow meter ± vane type, 200
fuel quantity measurement, 203
capacitive, 204
capacitive probe, 205
compensation, 205
dielectric, 204
permittivity, 205
resistive, 203
full flight regime autothrottle system (FFRATS), 144
ground proximity warning systems, 158
advanced GPWS ± operating modes, 160
basic GPWS ± operating modes, 158
enhanced GPWS, 161
terrain avoidance warning system (TAWS), 161
gyro compass operation, 68
erection system, 68
errors, 68
outputs, 69
precession coil, 68
gyro drives ± pneumatic and electrical, 40
gyro fundamentals, 32
apparent drift, 37
degrees of freedom, 34
drift and topple, 37
earth gyro, 37
rate integrating gyro, 37
real drift, 37
spin axis, 36
tied gyro, 35
transport drift, 38
gyroscopic instruments, 32
attitude indicator, 32, 46
directional gyro, 32, 35
free gyro, 34
gimbal, 34
gyro rotor, 32
gyroscopic properties, 32
Newton's First Law of Motion, 33
precession, 33
rigidity, 33
turn and bank indicator 32, 52
height, definition of, 9
horizontal situation indicator (HSI), 116, 125
IAS, 14
inductive transmission, 185
inertial navigation systems, 74
accelerometers, 75
control and display unit, 86
error corrections, 85
fibre optic gyro, 90
gyro-compassing, 84
gyro-stabilised platform, 77
INS calculations, 81
integrators, 74
levelling, 82
mode selector panel, 84
rate integrating gyroscope, 77
Schuler loop, 86
self-alignment, 82
strap-down system, 90
tuned rotor gyro, 91
wander angle system, 89
inertial reference system (IRS), 75
definition of, 117
instantaneous VSI, 24
instrument error, 11, 17, 22, 25
international standard atmosphere, 5
JAR Ops requirements for autopilots, 133
lag error, 11, 25
latitude nut, 45
limiting airspeeds, 16
LNAV, 136
local speed of sound (LSS), 18
mach meter, 1, 18
display, 22
errors, 22
principle of operation, 20
mach number, 19
mach/TAS calculations, 21
magnetic compass, 56
acceleration errors, 59
angle of dip, 56
aperiodicity, 58
correction card, 62
220 Index
deviation, 62
horizontality, 56
residual deviation, 62
sensitivity, 57
turning errors, 59
variation, 61
manifold air pressure (MAP) gauge, 183
overspeed warning, 167
permittivity, 205
piezo-electric transmitter, 187
pitot pressure, 1, 2
pitot tube/head, 2
pitot-static system, 1, 2
position error, 11
potentiometer, 187
precession, 33
pressure altimeter, 1, 5
pressure error, 12
pressure gauge, 182
pressure operated switches, 188
pressure transducer, 29
radio altimeter, 155
accuracy, 157
frequency range, 155
height ambiguity, 155
principle of operation, 155
system components, 157
ram rise, 28
rate integrating gyro, 77
rate of climb/descent indicator, 1, 22
rate of pressure change with altitude, 7
ratiometer, 186
recovery factor, 28
remote control, 210
remote indicating compass, 65
remote signal transmission systems, 210
rigidity, 33
ring laser gyro, 39
RPM indicator, 195
gas turbine, 197
piston engine, 195
secondary surveillance radar (SSR), 163
sensitive altimeter, 7
servo-assisted altimeter, 10
simple altimeter, 5
slaved gyro compass, 63
collector horns, 65
detector unit, 65
fluxvalve, 64
null-seeking rotor, 66
signal selsyn, 66
stall warning system, 168
standard atmosphere, 1
standby compass, 58
static air temperature (SAT), 28
static pressure, 1
static pressure error, 3
static source, 1, 2
stick pushing, 170
stick shaking, 170
subscale settings, 9
synchronous transmission, 184
tacho-probe, 197
phonic wheel, 198
TAS, 15
temperature error, 12
temperature gauge, 189
temperature measurement probes, 28
terrain avoidance warning system (TAWS), 161
warnings and alerts, 162
terrestrial magnetism, 57
thermocouple, 191
sensors, 193
thrust computation, 144
tied gyro, 35
torque meter, 207
brake horse power, 208
negative torque, 208
total air temperature (TAT), 28
total energy variometer, 26
total pressure, 2
touch control steering (TCS), 133
traffic collision avoidance system (TCAS), 163
crew response, 167
electronic VSI, 164
interrogation equipment, 163
principle of operation, 164
resolution advisory, 164, 166
traffic advisory, 164, 166
visual display, 165
transport drift, 38
true altitude, 10
turn and bank indicator, 52
bank indication, 54
rate gyroscope, 52
rate of turn, 54
turn co-ordinator, 54
vacuum system, 40
variometer 25
vertical speed indicator (VSI), 1, 22
instantaneous, 24
presentation, 24
transonic jump, 25
vibration monitoring, 209
relative amplitude, 210
VNAV, 136
warnings general, 153
altitude alert system, 154
flight warning system, 153
Wheatstone bridge circuit, 190
yaw damper, 150
indicator, 151
linear voltage displacement transmitter (LVDT), 151
rudder power control unit (PCU), 150
yaw rate gyro, 151
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