The operating mode of these displays may be either active or

passive, the definitions of which are as follows:

Light emitting diode - Active - Digital counter displays of engine
performance
Liquid crystal - Passive - monitoring indicators; radio frequency
selector indicators; distance measuring indicators; control
display units of inertial navigation systems.
Electron CRT beam - Active - Weather radar indicators; display
of navigational data; engine performance data: systems status;
check lists.

In a dot matrix display the patterns generated for each

individual
character arc made up of a specific number of illuminated dots
arranged in columns and rows.

When a low~voltage, low~urrent signal is applied to the

segments,
the polarization of the compound is changed together with a
change
in its optical appearance from transparent to reflective

In the case of electronic flight instrument systems, the two CRT

display units (EADI and EHSI) are also used in conjunction with
four
conventional-type indicators to form the basic 'T',

An air data (or manometric) system of an aircraft is one in

which the
total pressure created by the forward motion of an aircraft, and
the
static pressure of the atmosphere surrounding it, are sensed
and
measured in terms of speed, altitude and rate of altitude
change
(vertical speed).
In its basic form the system consists of a pitot-static probe, the
three
primary air data instruments (airspeed indicator, altimeter and
vertical
speed indicator)

Indicated airspeed (IAS) the readings of an airspeed indicator

corrected only for instrument error i.e. the difference between
the
true value and the indicated value.

computed airspeed [Basically, this is IAS with corrections for

position error (PE) applied (see page 34). The term 'computed'
applies specifically to air data computer systems in which PE
corrections are automatically applied to an airspeed sensing
module
via an electrical correction network.

Calibrated airspeed (CAS) This is also associated with air data

computer systems and is the computed airspeed compensated
for the
non-linear, or square-law, response of the airspeed sensing
module
Equivalent airspeed (EAS) this is the airspeed calculated from
the
measured pressure difference p when using the constant sea-
level
value of density p. In air data computer systems, CAS is
automatically compensated for compressibility of air at a pitot
probe
to obtain EAS at varying speeds and altitudes.

True airspeed (T AS) This is EAS compensated for changes in air

temperature and density at various flight levels.
Instantaneous vertical speed indicators (IVS!) These indicators
consist of the same basic elements as conventional VSI’s, but in
addition they employ an accelerometer unit which is designed
to
create a more rapid differential pressure effect, specifically at
the
initiation of a climb or descent.

The probe has an almost negligible time-lag,

and a high recovery factor of approximately 1.00. An axial wire
heating element, supplied with 115 V ac at 400 Hz, is mounted
integral with the probe to prevent the formation of ice.

The purpose of the engine bleed air

injector fitting and tube is to create a negative differential
pressure
within the casing so that outside air is drawn through it at such
a rate
that the heating elements have a negligible effect on the
temperature/resistance characteristics of the sensing element.

The horizontal angle contained between the true and the

magnetic
meridian at any place is known as the magnetic variation or
declination.
'When the direction of the magnetic meridian inclines to the left
of the true
meridian.

Lines are drawn on the charts, and those which join places
having equal variation are called isogonal lines, while those
drawn
through places where the variation is zero are called agonic
lines

The angle the lines of force make with the earth's surface at
any
Compass construction given place is called the angle of dip or
magnetic
inclination, and varies from 0° at the magnetic equator to 90°
at the
magnetic poles.

The bowl is of plastic (Diakon) and so moulded that it has a

magnifying effect on the card and its graduations. It is filled
with a
silicone fluid to make the compass aperiodic
The fluid also provides the system with a certain buoyancy,
thereby
reducing the weight on the pivot and so diminishing the effects
of friction
and wear. Changes in volume of the fluid due to temperature
changes, and
their resulting effects on damping efficiency, are compensated
by a
bellows type of expansion device secured to the rear of the
bowl

Before steps can be taken to minimize the deviations caused by

hardiron
and soft-iron components of aircraft magnetism, their values on
each he.,ding must be obtained and quantitatively analysed
into
coefficients of deviation. There are five coefficients designated
A, B,
C, D and E,

in the case of direct-reading compasses, always relate

only to deviation coefficients B and C. Adjustment for
coefficient A
is effected by repositioning the compass in its mounting by the
requisite number of degrees.

There are three such instruments, namely a gyro horizon

(sometimes
called an artificial horizon), a direction indicator, and a turn-
andbank
indicator.
As a mechanical device a gyroscope may be defined as a
system
containing a heavy metal wheel or rotor, universally mounted
so that
it has three degrees of freedom: (i) spinning freedom, about an
axis
perpendicular through its centre (axis of spin XX,); (ii) tilting
freedom, about a horizontal axis at right angles to the spin axis
(axis
of tilt YY1); and (iii) veering freedom, about a vertical axis
perpendicular to both the other axes

When the rotor is made to spin at high

speed, however, the device then becomes a true gyroscope
possessing
two important fundamental properties: gyroscopic inertia or
rigidity,
and precession. Both these properties depend on the principle
of
conservation of angular momentum

Angular momentum is the product

of the moment of inertia (/) and angular velocity (w) of a body

Rigidity. The property v,hich resists any force tending to

change
the plane of rotor rotation. It is dependent on three· factors: (i)
the
mass of the rotor, (ii) the speed of rotation, and (iii) the
distance at
which the mass acts from the centre, i.e. the radius of gyration.

Precession. The angular change in direction of the plane of

rotation under the influence of an applied force

The rate of' precession

also depends on three factors: (i) the strength and direction of
the
applied force, (ii) the moment of inertia of the rotor, and (iii) the
angular velocity of the rotor. The greater the force, the greater
is
the rate of precession, while the greater the moment of inertia
and
the greater the angular velocity the smaller is the rate of
precession.

The earth rotates about its axis at the rate of 15 °/hour

The instrument employs a horizontal axis

gyroscope and, being non-magnetic, is used in conjunction with
a magnetic compass; it defines the short-term heading changes
during
turns, while the magnetic compass provides a reliable long-
term
heading reference as in sustained straight and level flight

therefore to
minimize these problems the principle is adopted whereby the
pressures are transmitted to a centralized air data computer
(ADC)
unit, which then converts the data into electrical signals and
transmits
these through cables or data busses to the dependent
indicators and
systems. Another advantage of an ADC is that circuits may be
integrated with their principal data modules in such a way that
corrections for pressure error (PE), barometric pressure
changes, and
compressibility effects can be automatically applied; in
addition,
provision can also be made for the calculation of true airspeed
(T AS)
from air temperature data inputs.

An ADC may either be of the analogue type, or of the type

which
processes and transmits data in digital signal format.

A flight director system (FDS) is one in which the display of

pitch
and roll attitudes and heading of an aircraft are integrated with
such
radio navigation systems as automatic direction finding (ADF),
very
high-frequency omnidirectional range (VOR), and instrument
landing
system (ILS) so as to perform a total directive command
function. It
also provides for the transmission of attitude and navigational
data to
an AFCS so that in combination they can operate as an
effective
flight guidance system

The symbol representing the aircraft is fixed and is referenced

against a moving 'sky/ground' background tape on which are
presented an horizon line,

The GS pointer and scale come into view when the FDS is
operating in the GS mode and when there is a valid and reliable
signal from the VHF navigation receiver.

-The command bars, as the name implies, provide the

commands
relating to the changes that are to be made to manoeuvre an
aircraft
info required pitch and/or roll attitudes.

The aircraft symbol is fixed at the centre of the display and it

indicates the position and heading of an aircraft in relation to
the
compass card and the VOR/LOC deviation bar

The servomotor also drives a tach generator that provides rate

feedback #
signals for motor speed control
Battery unit: This unit provides de power for turning the system
on, and is also used as back-up in the event that power from an
aircraft's system is interrupted.

The inertial mode reference panel (IMRP) combines the

functions
of mode selection and control and display of data.

Both ac and de power is required for system operation which

must be
maintained in the event of failures occurring

In a typical gimballed-platform INS, the ac power is supplied

from
an essential busbar, and the de power from a nickel-cadmium
battery
unit which is part of the system installation

The battery supply remains on for a short period (typically 10

seconds).

The inertial navigation unit is provided with a battery charger

circuit which automatically comes into operation when the
battery is
not in use, and whenever its voltage drops below 26.5 V. The
charger is disconnected when the on-charge voltage increases
to
29.5 V.
In aircraft equipped with IR systems, the use of battery units is
eliminated since de power from the busbar of the aircraft's
battery
system is utilized for the starting up of a system. This supply is
also
automatically switched on in the event of a loss of ac power.

Longitude The longitude ofany point is the shortest distance in

the arc along the equator between the prime meridian and the
meridian through the point. It is expressed in degrees and
minutes
and is annotated east or west according to whether the point
lies east
or west of the prime meridian.
Latitude The latitude of any point is the arc of the meridian
between the equator and the point. It is also expressed in
degrees and
minutes, and is annotated north or south according to whether
the
point lies north or south of the equator

Heading (HDG) - The direction in which the nose of an aircraft is

pointing; it
is measured in degrees (000-360) clockwise from true,
magnetic, or
compass north.

Track (TK) - The direction in which an aircraft is moving over

the earth.
Drift Angle (DA) - The angle between HDG and TK due to the
effect of wind.
The direction of drift is always from HDG to TK. Each may be
true or
magnetic but never mixed.

Track Angle Error (TKE) - The angle (left or right) between the
DSR TK. and
the actual TK of an aircraft. It is always measured from DSR TK
to TK.

Because an aircraft can fly in any direction, two acceleration

sensors
('X' and 'Y') are required, and are mounted on a platform in
horizontal planes to sense accelerations 90° apart.

There are two mounting arrangements: the gyro-stabilized

platform,
and the 'strapdown', which will be described later in this
chapter. In
the first arrangement, the accelerometers are mounted on a
platform
which is supported in gimbal rings and stabilized by gyroscopes
and
torque motor systems

The selector switch is provided with two mechanical stops: one

between the 'STBY' and 'ALIGN' mode positions, and the other
between the 'NAV' and 'ATT' mode positions. In order for the
selector knob to move over the stops, it must be pulled out.
The
reason for having the stops is to prevent the 'NA V' mode from
being
inadvertently switched out

ring laser gyroscopes, eliminates the need to allow for 'warm-

up' and gyroscope 'run-up'.

In a typical gyroscope, alignment time periods can vary;

typically, they
would be ten minutes for an IRS and nineteen minutes for an
INS.

A CRT is a thermionic device, i.e. one in which electrons are

liberated as a result of heat energy

Colour CRT displays - These are used in weather radar display

units, and
are the norm for those units designed for the display of data
associated
with the systems installed

Blank screen: Zero or low-level returns.

Green: Low returns (lowest rainfall rate).
Yellow: Moderate returns (moderate rainfall rate).
Red: Strong returns (high-density rainfall rate).

stroke - scanning technique is also used for producing displays

of symbols
and of data in alphanumeric format.

As in the case of conventional flight director systems, ~

An EFIS installation is made up of left (Captain) and -right (First
Officer)
systems. Each system in turn is comprised of two displays
units: an
electronic attitude direction rector indicator (EADI} and
electronic
horizontal situation indicator, (EHSI) a control panel, a symbol
generator (SG), and a remote light sensor unit, A third (centre)
SG is
also incorporated so that its drive signals may be switched to
either
the left or right display units in the, event of failure of the
corresponding, SGs. The signal switching is accomplished
within the
left and right SGs, using electromechanical relays.

In a typical system six colours are assigned for the display of

the
many symbols, failure annunciators, messages and other
alphanumeric
information, and are as follows:
White – Display of present situation information
Green - Display of present situation information where contrast
with
white symbols is required, or for data having lower priority
than white symbols.
Magenta - All 'fly to' information such as flight director
commands,
deviation pointers, active flight path lines.,
Cyan - Sky shading on an EADI and low priority information
such as non-active flight plan map data.
Yellow - Ground shading on an EADI, caution information
display
such as failure warning flap, limit and alert annunciators
and fault messages.
Red – For display of heaviest precipitation levels as detected by
the radar

Symbol generators (SGs)

These provide the analogue, discrete and digital signal
interfaces
between an aircraft's systems, the display units and the control
panel,
and they perform symbol generation, system monitoring, power
control and the main control functions of the EFIS overall

Remote light sensor

This is a photodiode device which responds to the flight deck
ambient
Light conditions and automatically adjusts the brightness of the
CRT
displays to the appropriate level

The autoland status, pitch, roll-armed and engaged modes

are selected on the AFCS control panel, and the decision height
is
selected on the EFIS control panels. Radio altitude is digitally
displayed during an approach, and when the aircraft is between
2500
and 1000 ft above ground level. Below 1000 ft the display
automatically changes to a white circular scale calibrated in
increments of 1000 ft, and the selected decision height is then
displayed as a magenta-coloured marker on the outer scale.
The radio
altitude also appears within the scale as a digital readout. As
the
aircraft descends, segments of the altitude scale are
simultaneously
erased so that the scale continuously diminishes in length in an
anticlockwise
direction.

At the selected decision height plus 50 ft, an aural alert chime

sounds at an increasing rate until the decision height is
reached. At
the decision height, the circular scale changes from white to
amber
and the marker changes from magenta to amber; both the
scale and
marker also flash for several seconds. A reset button is
provided on
the control panel and when pressed it stops the flashing and
causes
the scale and marker to change from amber back to their
normal
colour.
If during the approach the aircraft deviates beyond the normal
ILS
glide slope and/or localizer limits (and when below 500 ft above
ground level), the flight crew are alerted by the respective
deviation
pointers changing colour from white to amber; the pointers also
start
flashing. This alert condition ceases when the deviations return
to
within their normal limits.

Indications of other data such as wind speed and direction,

lateral
and vertical deviations from the selected flight profile, distance
to
waypoint, etc., are also displayed.
The map display also provides two types of predictive
information.
One combines current ground speed and lateral acceleration
into a
prediction of the path over the ground to be followed over the
next
30, 60 and 90 seconds. This is displayed by a curved track
vector,
and since a time cue is included the flight crew are able to
judge
distances in terms of time. The second prediction, which is
displayed
by a range to altitude arc, shows where the aircraft will be
when a
selected target altitude is reached.
In the PLAN mode, a static map background with active route
data
oriented to true north is displayed in the lower part of the HSI
display

The VOR and ILS modes present a compass rose (either

ewanded
or full) with heading orientation display

Failure of data signals from such systems as the ILS and radio
altimeter is displayed on each EADI and EHSI in the form of
yellow flags 'painted' at specific matrix locations on their CRT
screens. In addition, fault messages may also be displayed: for
example, if the associated flight management computer and
weather
radar range disagree with the control panel range data, the
discrepancy message 'WXR/MAP RANGE DISAGREE' appears on
the EHSI.

The display units form part of two principal

systems designated as Engine indicating and crew alerting
system
(EICAS) and electronic centralized aircraft monitoring (ECAM)
systems, which were first introduced in Boeing 757 and 767
aircraft
and the Airbus A310 respectively

the basic system comprises of two display units, a control

panel, and
two computers supplied with analogue and digital signals from
engine
and system sensor.

The upper unit display the primary engine parameters N 1

speed,
EGT and warning and caution messages. In some cases this
unit can
Also display EPR depending on the type of engines installed and
on
the methods of processing data by the thrust management
control
system. The lower unit displays secondary engine parameters,
i.e. N2,
speed, fuel flow, oil quantity, pressure and temperature, and
engine
vibration. In addition, the status of non-engine systems, e.g.
flight
control surface positions, hydraulic system, APU, etc
Read from Page 379 TO 382

Alert messages
The system continuously monitors a large number of inputs
(typica.lly
over 400) from engine and airframe systems' sensors and will
detect
any malfunctioning of systems. If this should occur, then
appropriate
messages are generated and displayed on the upper display
unit in a
sequence corresponding to the !eye! of urgency of action to be
taken.
Up to 11 messages can be displayed, and at the following
levels:
Level A - Warning requiring immediate corrective action. They
are displayed in red. Master warning lights are also illunianted,
and aural warnings (e.g. fire bell) from a central warning
system
are given.
Level B - Cautions requiring immediate crew awareness and
possible action. They are displayed in amber, and also by
message
caution lights. An aural tone is also repeated twice.
Level C - Advisories requiring crew awareness. Also displayed in
amber. No caution lights or aural tones are associated with this
level.

The master warning and caution lights are located adjacent to

the
display units together with a 'Cancel' switch and a 'Recall'
switch.
Pushing the 'Cancel' switch removes only the caution and
advisory
messages from the display; the warning messages cannot be
cancelled. The 'Recall' switch is used to bring back the caution
and
advisory messages into the display. At the same time, the word
'RECALL' appears at the bottom of the display.
Display unit failure
If the lower display unit should fail when secondary information
is
being displayed on it, an amber alert message appears at the
top left,
of the upper display unit, and the information is transferred to it
as
shown in Fig. 16.6. The format of this display is referred to as
'compact', and it may be removed by pressing the 'ENGINE'
switch
on the display select panel. Failure of a display unit causes the
function of the panel 'ST A TUS' switch to be inhibited so that
the
status page format cannot be displayed.

A self-test of the whole system, which can only be activated

when
an aircraft is on the ground and its parking brake set, is
performed
by means of the 'TEST' switch on the maintenance control
panel.

Standby engine indicator

This indicator provides primary engine information in the event
that a
total loss of EICAS displays occurs.

Read Pg.388 – 391

System testing
Each flight warning computer of the system is equipped with a
monitoring module which automatically checks data acquisition
and
processing modules, memories, and the internal power supplies
as
soon as the aircraft's main power supply is applied to the
system. A
power-on test routine is also carried out for correct operation of
the
symbol generator units. During this test the display units
remain
blank.
In the event of failure of the data acquisition and processing
modules, or of the warning light display panel, a 'failure
warning
system' light on the panel is illuminated. Failure of a computer
causes a corresponding annunciator light on the maintenance
panel
captioned 'FWC FAULT', to illuminate. A symbol generator unit
failure causes a 'FAULT' caption on the appropriate push-button
switch on the system control panel to illuminate.

A 'DISPLAYS' push-button switch is provided on the

maintenance
panel and when pressed it initiates a check for correct
operation of
the symbol generator units, and the optical qualities of the
display
units by means of a test pattern display.