ircraft Instruments Integrated Systems

ircraft Instruments Integrated Systems
ircraft Instruments
Integrated Systems
EH J Pallett
IEng, AMRAeS
Foreword by
L FE Coombs
IEng, BSc, MPhil AMRAeS, FRSA
Authorized Licensed Edition of the UK edition, entitled Aircrq(t Instruments and
Integrated Systems, First Edition, by Pallett, published by Pearson Education Limited,
Cop)Tight '.iJ Longman Group UK Ltd 1992
Inclian edition published by Dorling Kindcrslcy India Pvt. Ltd. Copyright <I) 2011
All rights n:scrved. This book is sold subject to the condition that it shall not, by way of
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ISBN 978-81-317-3443-8
f'irst Impression, 2011
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Contents
Foreword vii
Preface ix
.--1
Instrument displays, panels and layouts
2 Air data instruments 25
3 Direct-reading compasses 78
._4
5
6
- 7
8
9
10
U
12
13
14
15
~~6
17
Tables
Gyroscopic flight instruments 97
Synchronous data-transmission systems 135
Digital computers and data transfer 152
Air data computers 161
Magnetic heading reference systems 182
Flight director systems 208
Inertial navigation/reference systems (INS/IRS) 246
Electronic (CRT) displays 283
Electronic flight instrument systems 296
Engine instruments 308
Fuel quantity indicating systems 326
Engine power and control instruments 347
Electronic instruments for engine and airframe systems control 377
Flight management systems 393
1. Standard atmosphere 408
2. Mach number/airspeed relationship 409
3. Temperature/resistance equivalents 409
4. Temperature/millivolt equivalents of typical iron v. constantan
thermocouples 410
5. Temperature/millivolt equivalents of typical copper v.
consta."ltan thermocouples 410
6. Temperature/millivolt equivalents of typical chrome! v.
alumel thermocouples 410
7. Nominal dielectric constants and densities of fuels 411
Principal symbols and abbreviations 412
Appendices
I Conversion factors 413
2 Logic gates and truth tables 418
3 Acronyms and abbreviations 419
V
Exercises 427
Solutions to exercises 439
Index 440
vi
Forewo
by L F E Coombs !Eng, BSc, MPhil, AMRAeS, FRSA
The progress of all types of aviation has depended on providing the
pilot with sufficient information to enable him or her to control the
aircraft safely and to navigate it to its destination.
From 1903 onwards each advance in speed, range, altitude and
versatility has had to be matched by instruments which enable the
crew to maximize an aircraft's potential. In the beginning, i.e. the
Wright 'Flyer' of 1903, the instrumentation was rudimentary,
consisting of only an anemometer for airspeed, a stop watch and an
engine revolution counter. Perhaps the piece of string attached to the
canard structure in front of the pilot, to indicate aircraft attitude
relative to the airflow, can -also be classed as an instrument.
Limited instrumentation was a feature of the aircraft of the f.rst
decade of heavier-than-air powered flight. However, the demands of
wartime flying accelerated the development of instrument, and i:,y
1918 a typical cockpit would have an airspeed indicator, an altimeter,
inclinometer, fuel pressure gauge, oil pressure indicator, rpm
indicator, compass and a clock. Not until the end of the 1920s were
instruments available by which a pilot could maintain attitude and
heading when flying in cloud, or whenever the horizon was obscured.
In the 1930s and '40s, considerable progress was made toward 'blind
flying' instruments. In the 1950s came the 'director'-type attitude
indicators and in the '60s more and m0re ele.ctromechanical
instruments became available. By 1970 solid-state displays were ·
edging their way on to the flight deck. In the past ten years the
electronic tide has swept in to an extent that the modem flight deck is
awash from wall to wall with solid-state displays such as the
electronic instrument systems (EIS) and engine indication and crew
alert systems (EICAS).
With a lifetime's interest in the man-machine interface of the
cockpit, I have depended very m.ich on the knowledge and advice of
many in the aviation industry and 2.Ssociated publishing. Over the
years a 'nose' is acquired which differentiates among the many
authors and separates them into categories. I place E H J Pallett at
the top of the list when it comes to the ability to grasp essentials
from a mass of technological facts, to explain succinctly, and, above
al!, to write with that authority which can only be acquired from long
practical experience.
vii
Any author who writes on the subject of aircraft instrumentation,
and who aims a book, as does Eddie Pallett, at licensed engineers
and flight crew, has a difficult task, a task made more onerous by'the
wide range of instrument types in use. This extends from mechanical,
through electromechanical, to electromechanical with some
electronics, and finally to today's electronic solid-state displays. Both
engineers and flight crews will come across all of these different
technologies. Some aircraft types will include examples of each,
while others will be 90 per cent all-electronic - the so-called 'glass
cockpit'.
Another problem, well tackled in this book, arises from the fact
that modern aircraft have systems rather than individual items of
equipment, such as instrument display units; few instruments,
therefore, can be considered in isolation. Many are just one part of a
system: for example, tne control and display unit (CDU) of a flight
management system is the visual and tactile interface with the pilot.
From this 'tip' depends the rest of the iceberg: the computers and
data links, and the interface units with other systems.
We have -come a long way from the time when the engineer had
only to undo four bolts and two unions, and out came the airspeed
indicator. The title Aircraft instruments and integrated systems is
therefore most apposite.
viii
Preface
The material for this book was initially intended to comprise a third
edition of Aircraft Instruments which for the last twenty years has
served as a reference to this area of the avionics field. Since the last
revision however, many technological changes have taken place, and
in deciding how information on such changes could be included in a
new edition, it soon became evident that the existing 'framework' and
title, would be totally inappropriate. Thus, after considerable
restructuring the compilation of material proceeded on the basis that
it would best serve as a complete replacement of Aircraft Instruments.
As far as the instrumentation requirements for aircraft are
concerned the most significant of the technological changes has been
that of processing data and presenting it in electronic display format.
Instruments and systems utilizing such format are, by virtue of high
levels ·oi digital computer integration and signal distribution via data
'highway' busbar systems, able to project the same quantity of
operational data which would otherwise have to be displayed by a
very large number of conventional 'clockwork' type instruments. The
scene was, therefore, set not only for making drastic reductions in the
utilization of conventional instruments, but also for implementing full
automation of the management of all aspects of in-flight operation of
aircraft.
The development of electronic instrument systems ran parallel with
the launching of the Boeing 757, 767 and Airbus A3 IO aircraft in
1978 as design projects, and these were to become the first of 'new
technology' aircraft to enter commercial service in 1982-3. These
aircraft and several of their descendant types, are now in service
world-wide, together with many types of smaller aircraft, including
helicopters, in which the foregoing technology has also satisfied an
operational need.
Conventional instruments of course still fulfill an important role but
the extent of their utilization now bears a more direct relationship to
types of aircraft. For example, in those already referred to, a
conventional airspeed indicator, altimeter and attitude indicator are
provided to serve as 'standby' references. There are on the other
hand many other types of aircraft in which conventional instruments
still satisfy the full instrumentation requirements appropriate to their
operational needs. The material for this book therefore, covers a
representative selection from the wide range of instruments and
systems that currently come within both areas of technology and the
sequencing of the relevant chapters has been arranged in such a way
as to reflect the transition from one area to the other.
Like its predecesSl'r, the details emphasize fundamental principles
and their applications to civil aircraft instruments and systems, and
overall, are also intended to serve as a basic reference for those
persons who, either incependently or, by way of courses established
by specialist training organizations, are preparing for Aircraft
Maintenance Engineer Licence examinations. It is also hoped that the
details will provide some support to the current technical knowledge
requirements relevant to flight crew examinations. A large number of
'self-test' questions have been compiled and are set out in chapter
sequence at the end of the book.
As with all books of this nature schematic diagrams and
photographs are of great importance in supporting the written details
and so it is hoped that the three hundred or so spread over the
chapters which follow will achieve the desired objective. In reviewing
some current aircraft installations together with the contents of
Aircraft Instruments, and also of another of my books
(Microelectronics in Aircraft Systems) I found that a number of
diagrams and photographs were still appropriate and so it was
expedient to make further use of these. The diagrams relating to 'new
subject' material have, in many cases, been redrawn from my
original 'roughs'. The remaining new diagrams and photographs
(some of which are reproduced in colour) have been supplied to me
from external sources, and in this connection I would particularly like
to express my grateful thanks to Smith's Industries, Aer Lingus, and
Boeing International Corporation for their assistance.
In conclusion, I wish to convey sincere thanks to Leslie Coombs
not only for his help, past and present, in providing material on a
subject of common interest, but in particular for having accepted my
invitation to write the Foreword to this book. It necessitated his
having to read through many pages of draft manuscript, but as this
resulted in comments that required some changes of text, then I am
sure that he too would agree that efforts were not wasted.
Copthorne
W. Sussex
X
E.P.
1 Instrument displays,
panels and layouts
In flight, an aircraft and its operating crew form a 'man-machine'
system loop which, depending on the size and type of aircraft, may
be fairly simple or very complex. The function of the crew within th6
loop is that of controller, and the extent of the control function is
governed by the simplicity or otherwise of the aircraft as an
integrated whole. For example, in manually flying an aircraft, and
manually initiating adjustments to essential systems, the controller's
function is said to be a fully active one. If, on the other hand, the
flight of an aircraft and system's adjustments are automatic in
operation, then the controller's function becomes one of monitoring,
with the possibility of reverting to the active function in the event of
failure of systems.
(Instruments, of course, play an extremely vital role in the control
loop as they are the means of communicating data between systems
and controller. Therefore, in order that a controller may obtain a
maximum of control quality, and also to minimize the mental effort
in interpreting data, it is necessary to pay the utmost regard to the
cootent and format of the data displays. j
/'The most common forms of data display are (a) q~ive, in
~ the variable quantity being measured is presented in terms of a
numerical value and by the relative position between a pointer or
index and a graduated scale, and (b) qualitative, in which the data is
presented in symbolic or pictorial format.
Quantitative displays
There are three principal methods by which data may be displayed:
(i) the circular scale, or more familiarly, the 'clock' type of scale,
(ii) straight scale, and (iii) digital, or counter.
Circular scale
This may be considered as the classical method of displaying data in
quantitative form and is illustrated in Fig. 1.1. The scale base refers
to the graduated line, which may be actual or implied, running from
end to end of the scale and from which the scale marks and line of
travel of the pointer are defined.
1
Figurt /. I Circular scale
q11anti1111ive display.
Implied seal& base
Figure I. 2 Linear and non•
linear scales. (a) Linear;
(b) square-law; (c) logarithmic.
(bl
la)
A.ATE Of
CUMll
0
(C)
Scale or graduation marks are those which constitute the scale of
an instrument. For quantitative displays the number and size of marks
are chosen in order to obtain quick and accurate interpretation of
readings. In general, scales are divided so that the marks represent
units of I, 2 or 5, or decimal multiples thereof, and those marks
which are to be numbered are longer than the remainder.
Spacing of marks is also governed by physical laws related to the
quantity to be measured, but in general they result in spacing that is
either linear or non·linear. Typical examples are illustrated in Fig.
1.2, from which it will also be noted that non·linear displays may be
of the square·law or logarithmic•law type, the physical laws in this
instance being related to airspeed and rate of altitude change
respectively.
The sequence of numbering always increases in a clockwise
direction, thus conforming to what is termed the 'visual expectation•
of the observer. As in the case of marks, numbering is always in
steps of 1, 2 or 5 or decimal multiples thereof. The numbers may be
marked on the dial either inside or outside the scale base.
2
The distance between the centres of the marks indicating the
minimum and maximum values of the chosen range of measurement,
and measured along the scale base, is called the scale length.
Governing factors in the choice of scale length for a particular range
are the size of the instrument, the accuracy with which it needs to be
read, and the conditions under which it is to be observed.
High-range long-scale displays
For the measurement of some quantities - for example, turbine
engine speed, airspeed, and altitude - high measuring ranges are
involved with the result that very long scales are required. This
makes it difficult to display such quantities on single circular scales in
standard-size cases, particularly in connection with the number and
spacing of the marks. If a ,large number of marks are required their
spacing might be too close to permit rapid reading, while, on the
other hand, a reduction in the number of marks in order to 'open up'
the spacing will also give rise to errors when interpreting values at
points between scale marks.
Some of the displays developed as practical solutions are illustrated
in Fig. 1.3. The display shown at (a) is perhaps the simplest way of
accommodating a lengthy scale; by splitting it into two concentric
scales, the inner one is made a continuation of the outer. A single
pointer driven through two revolutions can be used to register against
both scales, but as it can also lead to too frequent misreading, a
Figure I. 3 High-range longscale displays. (a} Concentric
scales: (b) fixed and rotating
scales: (c) common scale and
triple pointers.
I
90
° 10
1
...8 0 ~ ...20-
..70
~o
'gt{
(a)
'8
~...
5P
~
(bl
2...
...7
3 ...
,6tlID4,
I
(c)
3
presentation by two concentrically-mounted pointers of different sizes
is much better. A practical example of this is to be found in some
types of engine speed indicator. In this instance, a large pointer
rotates against an outer scale to indicate hundreds of rev/min, and at
the same time it rotates a smaJler pointer through appropriate ratio
gearing, against an inner scale to indicate thousands of rev/min.
The method shown at (b) is employed in a certain type of
pneumatic airspeed indicator; in its basic concept it is similar to the
one just described. In this case, however, a single pointer rotates
against a circular scale and drives a second scale plate instead of a
pointer. This rotating plate, which records hundreds of knots as the
pointer rotates through complete revolutions, is visible through an
aperture in the main dial of the indicator.
Scale and operating ranges
Instrument scale lengths and ranges usually exceed that actually
required for the operating range of the system with which an
instrument is associated, thus leaving part of the scale unused. This
may appear somewhat wasteful, but an example will show that it
helps in improving the accuracy with which readings may be
observed.
Let us consider a fluid system in which the operating pressure
range is, say, 0-30 lbf/in2 • It would be no problem to design a scale
for the required pressure indicator which would be of a length
equivalent to the system's total operating range, also divided into a
convenient number of parts as shown in Fig. l.4(a). However, under
certain operating conditions of the system concerned, it may be
essential to monitor pressures having such values as 17 or 29 lbf/in2
and to do this accurately in the shortest possible time is not very
easy, as a second look at the diagram will show.
If the scale is now redesigned so that its length and range exceed
the system's operating range and also graduated in the manner noted
Figure 1.4 Reading accuracy.
17
I
(a)
4
(b)
earlier, then as shown at (b) the result makes it much easier to
interpret and to monitor specific operating values.
Straight scale
In addition to the circular scale presentation, a quantitative display
may also be of the straight scale (vertical or horizontal) type. For the
same reason that the sequence of numbering is given in a clockwise
direction on a circular scale, so on a straight scale the sequence is
from bottom to top or from left to right.
1Although such...displays contribute to the saving of panel space and
improved observational accuracy, their application to the more
conventional types of mechanical and/or electro-mechanical
instruments has been limited to those utilizing synchronous datatransmission principle&.) It is pertinent to note at this juncture that in
respect of electronic CRT displays (see Chapter 11) there are no
mechanical restraints, and so straight scales can, therefore, be more
widely applied.
An example of a straight scale presentation of an indicator
operating on the above-mentioned principles is illustrated in Fig.
l .5(a); it is used for indicating the position of an aircraft's landing
flaps. The scales are graduated in degrees, and each pointer is
operated by a synchro (see Chapter 5). The synchros are supplied·
with signals from transmitters actuated respectively by left and right
outboard flap sections.
Another variation of this type of display is shown at (b) of Fig.
1.5. It is known as the moving-tape or thermometer display and was
originally developed for the measurement of parameters essential to
the operation of engines of large transport aircraft. @ach display unit
contains_!!§ervo-drhteILwMe.m~i!L!~l_~ce of a pointer, which moves
in a verticaLpla.ne and_registers .tg<1i11_st a. scal.e Ill a~sirnUaLmanner to
the mercury column of a thermometer._.As will be noted, there is one
display unit for each parameter' the scales being common to each
engine in the particular type of aircraft. When such display.s are
limited to only one or two parameters then, by scanning across the
ends of the tapes, or columns, a much quicker and more accurate
evaluation of changes in engine performance can be obtained as
compared to 'clock' type displays. This fact, and the fact that panel
space can be reduced, are clearly evident from the diagram.
Digital display
A digital, or counter, type of display is one that is generally to be
found operating in conjunction with the circular type of display; two
examples are shown in Fig. 1.6. fin the application to an altimeter
there are two coun~<:_rs: one presents a fixedyressure value which can
5
Figure /.5 Straight scale
displays. (b) gives a
comparison between movingtape and circular scale displays.
(a)
EXHAUST
GAS
ENGINE NO.
2
3
4
EGT 'C
% RPM
% RPM
x100
Engine
No.
1
2
3
4
(bl
EGT
%
RPM
500
470
480
520
89
90
88
90
oc
1 2
X
3 4
1 2
\0
3 4
Figure I. 6 Application of
digital counter displays.
DYNAMIC COUNTER
OISPl.AY
be m~ch,mkaJly, set as a11d wh~.n required, and is known as aJ!S!.!is:_
co.ynt~_.Qi§Qlay; while the other is gearedto th~_~_ltimeter mechanism
and autQma.tically presents changes in _altitude, and is therefore known
as a fl1n'!._mic counte.i: displau It is of interest to note that the
presentation of altitude data by means of a scale and counter is yet
another method of solving the long-scale problem already referred to
on page 3. The counter of the turbine gas temperature (TOT)
indicator is also a dynamic display since, in addition to the main
pointer, it is driven by a servo transmission system (see also page
363).
Dual-indicator displays
r'[these displays are designed principally as a means of conserving
panel space, particularly where the measurement of the various
quantities related to engines is concerne¥ )!'hey are normally of two
basic forms: in one, ~wo separate indicator mechanisms and scales are
contained in one cas~ while in the other, which also has two
mechanisms in one case, the pointers register against a common
i:cale. Typical examples of display combinations are illustrated in Fig.
1.7.
J>perational range markings
These markings take the form of coloured arcs, radial lines and
sectors applied to the scales of instruments, their purpose being to
highlight specific limits of operation of the systems with which the
instruments are associated. The definitions of these marks are as
follows:
RED radial line Maximum and minimum limits
Take-off and precautionary ranges
i - YELLOW arc
Normal operating range
/, GREEN arc
Range in which operation is prohibited
/ _,,RED arc
r
V
7
Figure I. 7 Dual-indicator
displays. The display with three
pointers has a helicopter
application: it shows the speed
of No. I and No. 2 engines,
and of the main rotor.
In the example shown in Fig. 1.8(a), an additional WHITE arc is
provided which serves to indicate the appropriate airspeed range over
which an aircraft's landing flaps may be extended in the take-off,
approach and landing configurations.
The application of sector-type markings is usually confined to those
parts of an operating range in which it is sufficient to know that a
certain condition has been reached rather than knowing actual
quantitative values. For example, it may be necessary for an oxygen
cylinder to be charged when the pressure has dropped to below, say,
500 lbf/in2 • The cylinder pressure gauge would therefore have a red
sector on its dial embracing the marks from O to 500 as at (b) of Fig.
l.8. Thus, if the pointer should register within this sector, this alone
is sufficient indication that recharging is necessary, and it is only of
secondary importance to know what the actual pressure is.
Another method of indicating operating ranges is one that uses
what are termed 'memory bugs'. These take the form of small
pointers which, by means of an adjusting device, can be rotated
around the dial plate of an instrument to pre-set them at appropriate
operating values on the scale. An example of their application to a
Mach/airspeed indicator (see page 47) is shown at (c) of Fig. 1.8.
Qualitative displays
These are of a special type in which the information is presented in a
symbolic or pictorial form to show the condition of a system,
whether the value of an output is increasing or decreasing. or to
show the movement of flight control surfaces as in the example
shown in Fig. 1.9.
Figure I. 8 Operational range
GREEN
markings.
(a)
(b)
Memory 'bugs'
(c)
Figure I. 9 Qualitative display.
WING
UP
I 1-
R
\
'I
(~
\ON \
L
SPOIL
UP
ELEV
+D +
~D~
f Dj
I
I
ON
R
SPOIL
I
RUD
UPR
LWR
Director displays
/_Ihese displays are associatt:d principally with the IJlOnitori~t
' ~ and present it in a manner !_!_!at
indicates to the flight crew what control movements must be made,
either to correct any departure from a desired flight path, or to cause
an aircraft to perform a specific manoeuvr~ It is thus apparent that
in the development of such a display there must be a close
Q"elationship betw~n the direction of control movements and the
instrument pointer, or symbolic-type indicating eleme~ in other
words, movements should be in' the 'natural' sense in order that the
'directives' or 'commands' of the display may be obeyed.
Displays of this nature are specifically applied to the two prim!ll}'
instruments which comprise conventional· flight director systems and
electronic flight instrument systems (see Chapters 9 and 12). One of
the instruments (referred to as an Attitude Director Indicator) has its
display origins in one of the oldest of flight attitude instruments,
namely the gyro horizon (see Chapter 4), and so it serves as a basis
for understanding the concept of director displays. As will be noted
from Fig. 1.10, three elements make up the display of the
instrument: a pointer registering against a bank-angle scale, an
element symbolizing an aircraft, and an element symbolizing the
natural horizon. Both the bank pointer and natural horizon element
are stabilised by a gyroscope. As the instrument is designed for the
display of attitude angles. and as also one of the symbolic elements
can move with respect to the other. then it has two reference axes,
that of the case which is fixed with respect to an aircraft, and that of
the moving element.
Assuming that in level flight an aircraft's pitch attitude changes
such as to bring the nose up, then the movement of the horizon
element relative to the fixed aircraft symbol will be displayed as in
diagram (a). This indicates that the pilot must 'get the nose down'..
Similarly. if an aircraft's bank attitude should change whereby the
left wing, say, goes down, then the display as at (b) would direct the
pilot to 'bank the aircraft to the right'. In both cases the commands
Figure I. JO Director display
(gyro horizon).
(a)
10
(b)
would be satisfied by the pilot moving the appropriate flight controls
in the natural sense.
The display presentation of a typical Attitude Director Indicator is
shown in Fig. l.ll(a), and as will be noted it is fundamentally
similar to that of a gyro horizon. Details of its operation will be
covered in a later chapter, but at this juncture ii suffices to note that
Figure I. /1 Attitude director
display. (a) Aircraft straight
and level; (b) aircraft nose up;
{c) aircraft banked left; (d) 'fly
up' command; (e) 'fly left'
command.
Fixed bank
pointer
t
I
I
(b)
I
I
I
I
I
I
(a)
(C)
~--~
~-----o--"---•
I
(d)
I
I
(e)
11
the horizon symbolic element is driven by servomotors that receive
appropriate attitude displacement signals from a remotely-located
gyroscope unit. Thus, assuming as before a nose-up displacement of
an aircraft, the signals transmitted by the gyroscope unit will cause
the horizon symbolic element to be driven to a position below the
fixed element symbolizing the aircraft, as shown at {b). The pilot is
therefore directed to 'fly down' to the level flight situation as at (a).
If a change in the aircraft's attitude produces, say, a left bank, then
in response to signals from the gyroscope unit the horizon symbolic
element and bank pointer will be driven to the right as shown at (c).
The pilot is therefore directed to 'fly right' to the level flight
situation.
In addition to displaying the foregoing primary attitude changes, an
indicator also includes what is termed a command bar display that
enables a pilot to establish a desired change in aircraft attitude. If, for
example, a climb attitude is to be maintained after take-off, then by
setting a control knob the command bars are motor-driven to a 'fly
up' position as shown at (d) of Fig. 1.11. During the climb the
horizon symbolic element will be driven in the manner explained
earlier, and the command bars will be recentred over the fixed
element so that the display will be as shown in diagram (b).
Roll attitude, or turn commands, are established in a similar
manner, the command bars in this case being rotated in the required
direction; diagram (e) of Fig. 1.11 illustrates a 'fly left' command.
As the aircraft's attitude changes the aircraft symbolic element moves
with the aircraft, while the horizon symbolic element and bank
pointer are driven in the opposite direction. When the command has
been satisfied, the display will then be as shown in diagram (c).
. The scales and pointers shown to the left and bottom of the
indicator also form a director display that is utilized during the
approach and landing sequence under the guidance of an Instrument
Landing System. Details of the operation of this display and of the
second indicator involved in a Flight Director System will be
in
Chapter 9.
Electronic displays
12
With the introduction of digital signal-processing technology into the
field colloquially known as 'avionics', and its application of micro-electronic circuit techniques, it became possible to make drastic
changes to both quantitative and qualitative data display methods. In
fact, the stage has already been reached whereby many of the
conventional 'clock' type instruments which, for so long, have
performed a primary role in data display, can be replaced entirely by
a microprocessing method of 'painting' equivalent data displays on
the screens of cathode ray tube (CRT) display units.
In addition to CRT displays (see Chapter 11), electronic display
Table l. I Applications of electronic displays
Display technology
Operating mode
Light-emitting diode
Liquid crystal
Active
Passive
Electron CRT beam
Active
Typical applications
Digital counter displays of engine performance
monitoring indicators; radio frequency selector
indicators; distance measuring indicators; control
display units of inertial navigation systems.
Weather radar indicators; display of navigational
data; engine performance data: systems status;
check lists.
techniques also include those of light-emitting diode and liquid crystal
elements. Typical examples of their applications are given in Table
1.1. The operating mode of these displays may be either active or
passive, the definitions of which are as follows:
Active:
Passive:
a display using phenomena potentially capable of
producing light when the display elements are
electrically activated.
a display which either transmits light from an auxiliary
light source after modulation by the device, or which
produces a pattern viewed by reflected ambient light.
Display configurations
Displays of the light-emitting diode and liquid crystal type are usually
limited to applications in which a single register of alphanumeric
values is required, and are based on what is termed a seven-segment
matrix configuration or, in some cases, a dot matrix configuration.
Figure l .12(a) illustrates the seven-segment configuration, the
letters which conventionally designate each of the segments, and the
patterns generated for displaying each of the decimal numbers 0-9.
A segmented configuration may also be used for displaying alphabetic
characters as well as numbers, but this requires that the number of
segments be increased, typically from seven up to 13 and/or 16.
Examples of these alphanumeric displays are illustrated at (b) of Fig.
1.12.
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. In the example shown at (c) of Fig.
l .12, the matrix is designated as a 4 x 7 configuration, i.e. it
comprises four columns and seven rows.
Light-emitting diodes (LEDs)
An LED is a solid-state device comprising a forward-biased p-n
junction transistor formed from a slice or chip of gallium arsenide
13
Figure /./2 Electronic
alphanumeric displays.
(a) Seven-segment;
(b) 13- and 16-segment;
(c) a 4 x 7 matrix.
A
F
B
E
C
D
No. of segments
Cl23LIS61B9
6
5
2
5
3
6
5
7
6
(a)
-__ --,
I I-- --11 I__
1--,
I I _(I I__ _I_I I__
1_,1 ,'__ 1v1 " 1
1
I I I \I I__I
L--_ I--_ L_I -1I
I
I_J I I J_ I_J
r-, Ic:i I_~
r-, Ic-1\ r-::i TI
I I I / I I \/ \I I
l,;_J I/ V\I /\ I (._
:_1
l
(b)
J_
1
u1 C_I C
u /7
n o un /1
[I J
• •
•
•
•
•
• •
•
•
•
•
• •
r-4cols.~
(c)
phosphide (GaAsP) moulded into a transparent covering as shown in
Fig. 1.13. When ·current flows through the chip it emits light which
is in direct proportion to the current flow. Light emission in different
colours of the spectrum can, where required, be obtained by varying
the proportions of the elements comprising the chip, and also by a
technique of 'doping' with other elements, e.g. nitrogen.
In a typical seven-segment display format it is usual to employ one
LED per segment and mount it within a reflective cavity with a
14
Figure 1.13 Light-emitting
diode.
CIIYSTAL CHIP
Diffuser plate
AeUectiw cavities
-----r
Eflechve
-
•
segment
-=-~-T
Plastic overlay
Ptashc overlay
CONNECTIONS
plastic overlay and a diffuser plate. The segments are formed as a
sealed integrated circuit pack, the connecting pins of which are
soldered to an associated printed circuit board. Depending on the
application and the number of digits comprising the appropriate
quantitative display, independent digit packs may be used, or
combined in a multiple digit display unit ..
LEDs can also be used in a dot-matrix configuration, and an
example of this as applied to a type of engine speed indicator is
shown in Fig. 1.14. Each dot making up the decimal numbers is an
individual. LED and they are arranged in a 9 x 5 matrix. The counter
is of unique design in that its signal drive circuit causes an apparent
'rolling' of the digits which simulates the action of a mechanical
drum-type counter as it responds to changes in engine speed.
Liquid crystal display (LCD)
The basic structure of a seven-segment LCD is shown in Fig.· 1.15. It
consists of two glass plates coated on their inner surfaces with a thin
film of transparent conducting material (referred to as polarizing film)
such as indium oxide. The material on the front plate is etched to
form the seven segments, each of which forms an electrode. A mirror
image is also etched into the oxide coating of the back glass plate,
but this is not segmented since it constitutes a common return for all
segments. The space between the plates is filled with a liquid crystal
15
Figure I. 14 Engine speed
indica1or wi1h a dot matrix
LED. (Courtesy of Smith's
Industries Ltd.)
Figure I. 15 Structure or an
LCD.
Liquid crystal
[
layer
(typical spacing
= 10 microns)
compound, and the complete assembly is hermetically sealed with a
special thermoplastic material to prevent contamination.
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. The
16
Figure J. 16 Application of
LCD.
magmtude of the optical change is basically a measure of the light
reflected from, or transmitted through, the segment area to the light
reflected from the background area.
unlike an LED, it does not
emit light, but merely acts on light
through it. Depending on
polarizing film orientation, and also on whether the display is
reflective or transmissive, the segments may appear dark on a light
background (as in the case of digital watches and pocket calculators)
or light on a dark background. An example of LCD application is
shown in .Fig. l .16.
neau··uu
displays
A head-up display (HUD) is one in which vital in-flight data are
presented at the same level as a pilot's liµe of sight when he is
viewing external references ahead of the aircraft, i.e. when he is
maintaining a 'head-up' position. This display technique is one that
has been in use for many years in military aviation, and in particular
it has been essential for those airc:-aft designed for carrying out very
high-speed low-level sorties over all kinds of terrain.
As far as civil aviation is concerned, HUD systems have been
designed specifically for use in public transport category aircraft
during the approach and landing phase of flight, but thus far it has
been a matter of choice on the operators' part whether or not to
install systems in their aircraft. This has resulted principally from the
differing views held by operators, pilot representative groups, and
aviation authorities on the benefits to be gained, notably in respect of
a system's contribution to the landing of an aircraft, either
automatically or manually, in low-visibility conditions.
The principle adopted in a HUD system is to display the required
data on the face of a CRT and to project them through a collimating
lens as a symbolic image on to a transparent reflector plate, such that
the image is superimposed on a pilot's normal view, through the
windscreen, of the terrain ahead. The display is a combined
alphanumeric and symbolic one, and since it is focused at infinity it
permits simultaneous scanning of the 'outside world' and the display
without refocusing the eyes. The components of a typical system are
shown in Fig. 1.17.
17
Figure 1.17 Head-up display.
WINDSCREEN
'
"--=-i
I
--1
I
CR TUBE
Data
inputs
PROCESSOR AND
SYMBO~GENERATOR
UNIT
i--~~~~~~~~-~~~~~~....,
(a)
Altitude
Airspeed
150
,:.--,0......- , - - -
•
•
•
Vertical speed
-<
•
Horizon
bars
•
21
22
•
• • i • •
(b)
Heading
The amount of data required for display is governed by the
requirementl': of the various flight phases and operational role of an
aircraft, i.e. military or civil, but the parameters shown in diagram
(b) are common to all. The format and disposition of the displays
corresponding to required parameters can vary between systems; for
example, a heading display may be in the form of a rotating arc at
18
the upper part of the reflector plate, and altitude may be indicated by
the registering of moving dots with a fixed index at one side of the
plate instead of the changing digital counter readout located as shown
in the diagram. Additional data such as decision height, radio altitude
and runway outlines may also be displayed.
Panels and layouts
Instrument grouping
\All instruments essential to the operation of an aircraft are
accommodated on panels~the number and disposition of which vary
in accordance with the number of instruments required for the
appropriate type of aircraft and its cockpit or flight deck layou!) A
main instrument panel positioned in front of pilots is, of course.fa
feature common to all types of aircraft since instruments displaying
primary data must be within the pilot's normal line of vision) The
f panel may_J>e_mounted.in_the verticaLpositioJLor, as is now more
~common-practice, ~opeci. forward at. aboup5° fronLthe_yerticaLto
~-~i~~~arallax error~ l'}'j>}£.alpositions of otherpanels ar.e:
overhead, at the side, and on a contrQL~c!e§.!al locat~<Lcentrally
~tweenyilots, Figure 1.18 illustrates the for~going ~rrangement
appropriate to the Boeing 737-300 series aircraft. Where a flight
engineer is required as a member of an operating crew, then panels
would also be located at the station specifically provided on the flight
deck.
Flight instruments
(Basically there are six_flight instruments whose indications are so coordinated as to create a 'picture' of an aircraft's flight condition and
required control movementsJjthey are the a[r:~Q.e.eJLi.ndicator,
AlJjme1er. gy.r.<>__hg1:iz;on, dires:tioi:i_irn;licator, ver1Lcc3:l~ed-indicator
and turn::aod:banJcindicator. Qt is, therefore, most important for these
instruments to be properly grouped to maintain co-ordination and to
assist a pilot in observing them with the minimum of effort. ?
The first real attempt at establishing a sta.i:idard .method of grouping
was the 'blind flying panel' or ·_~asic six' layout shown in Fig.
l. I 9(a). ,The gyro horizon occupies the top centre position, and since
it provides positive and direct indications of attitude, and attitude
changes in the pitching and rolling planes, it is utilized as the master
instrument. As control of airspeed and altitude are directly related to
attitude, the airspeed indicator, altimeter and vertical speed indicator
flank the gyro horizon and support the interpretation of pitch attitude.
Changes in direction are initiated by banking an aircraft, and the
degree of heading change is obtained from the direction indicator;
this instrument therefore supports the interpretation of roll attitude
19
Figure I. I 8 Flight deck
layout: Boeing 737-300 series
aircraft.
Figure 1.19 Flight instrument
grouping. (a) Basic six:
(b) basic ·r (with flight
director system indicators).
(al
20
(b)
and is positioned directly below the gyro horizon. The tum-and-bank
indicator serves as a secondary reference instrument for heading
changes, so it too supports the interpretation of roll attitude.
(With the development and introduction of new types of aircraft,
and of more comprehensive display presentations afforded by the
indicators of flight director systems, a review of the functions of
certain of the instruments and their relative positions within the group
resulted in the adoption of the 'basic T' arrangement as the current
standa~ As will be noted from diagram (b) of Fig. 1.19, there are
,now four 'key' indicators:~. pitch and roll attitude. an altitude
indicator forming the horizontal bar of the 'T', and a horizonta~
s ~ n (direction) indicator forming the vertical bar. As far as the
positions flanking the latter indicator are concerned, they are taken up
by other less specifically essential flight instruments which, in the
example shown, are the vertical speed indicator and a radiomagnetic
indicator (RMI). In some cases a tum-and-bank indicator, or an
indicator known as a turn co-ordinator, may take the place shown
occupied by the RMI. In many instances involving the use of flight
director system indicators and/or electronic flight instrument system
display units, a tum-and-bank indicator is no longer used.
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', as shown in Fig.
l .20(a). In displays of more recent origin, and now in use in such
aircraft as the Boeing 747-400 (see also Fig. 12.11), the CRT
screens are much larger in size, thus making it possible for the EADI
to display airspeed, altitude and vertical speed data instead of
conventional indicators. The presentation, which also corresponds to
the basic 'T' arrangement is illustrated at (b) of Fig. 1.20.
Power plant instruments
lThe specific grouping of instruments required for the monitoring of
power plant operation is governed primarily by the type of power
plant, the size of aircraft, and therefore the space available for
location of instruments.,)\in a single-engined aircraft this does not
present too much of a problem since the small number of instruments
required may flank the flight instruments 1 thus keeping them within a
small 'scanning range':
·
21
·r
Figure 1.20 Basic
grouping wilh electro1ik flight
instruments.
A,,speed
/,,.
Attitude
Altitude
Vertical
speed
EHSI
'
Heading
(bl
,.~ADI
8
G
Conventional instruments as standby
\Jhe problem is more acute in multi-engined aircraft, the number of
instruments for·all essential parameters doubling up with each engine.
For twin-engined aircraft, and for certain medium-size four-engined
aircraft. the practice is to group the instruments at the centre of the
main instrument panel and between the two groups of flight
instruments.,
In those aircraft having a flight engineer's station, the instruments
are grouped on the control panels at this station. Those instruments
measuring parameters required to be known by the pilot during takeoff, cruising and landing, e.g. rev/min and turbine gas temperature,
are duplicated at the centre of the main instrument panel.
\The positions of the instruments within a group are arranged so
22
that those relating to each power plant correspond to the power plant
positions as seen in plan vieaj. llt will be apparent from the layout of
Fig. 1.21 that by scanning a row of instruments a pilot or engineer
can easily compare the readings of a given parameter, and by
scanning a column of instruments can assess the overall performance ·
Figure I. 21 Power plant
instrument grouping.
I
I
NJ
ENGlNE PRESSURE
RAT!O
EXHAUST GAS
TEMPERATURE
FUEL FLOW
23
Figure 1.22
Power plant
instrument grouping ( solid-state
displays): Primary parameters
pattern of a particular power plant. Another advantage of this
grouping method is that all the instruments for one power plant are
more easily associated with the controls for that power plant.
Figure 1.22 illustrates the grouping arrangement currently adopted
in Boeing 737 -400 series aircraft for the display of the primary
parameters associated with its power plants.
The numeric values corresponding to each parameter are indicated
by LEDs arranged in a dot matrix 'rolling digit' configuration, and
located at the centre of permanently defined scale bases, graduations
and coloured range markings. In addition to the counter displays.
LEDs are also located around the periphery of each scale base, and
in their active state they simulate the rotation of conventional
indicator pointers.
24
.
I
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). The measurement and indication of these three
parameters may be done by connecting the appropriate sensors, either
directly to mechanical-type instruments, or to a remotely-located air
data computer which then transmits the data in electrical signal
format to electro-mechanical or servo-type. instruments.
Since the primary source of air for these measurements is the
earth's atmosphere itself, then it is necessary to have some
understanding of its characteristics before going into the operating
principles of the measuring instrumen,s and systems involved.
The earth's
The earth's atmosphere is the surrounding envelope of air, which is a
mixture of a number of gases, the chief of which are nitrogen and
oxygen. By convention, this gaseous envelope is divided into several
concentric layers extending from the earth's surface, each with its
·own distinctive features; these are shown in Fig. 2.1.
The lowest layer, and the one in which conventional types of
aircraft are. flown, is termed the troposphere, and extends to a
boundary height termed the tropopause.
Above the tropopause, the next layer, termed the stratosphere, also
extends to a boundary height called the stratopause.
At greater heights the remaining atmosphere is divided into further
layers which are termed the chemosphere, ownosphere, ionosphere
and exosphere.
Throughout all these layers the atmosphere undergoes a gradual
transition from its characteristics at sea-level to those at the fringes of
the exosphere where it merges with the completely airless outer
space.
Atmospheric pressure
The atmosphere is held in contact with the earth's surface by the
force of gravity, which produces a pressure within the atmosphere.
Gravitational effects decrease with increasing distances from the
25
Figure 2.1 Earth· s
a1mosphere.
Pressure Temperature 1111111111
Relative density
150
l
I
EXOSPHERE
IONOSPHERE OZONOSPHERE
CHEMOSPHERE
140
130
120
,_
110
Stratopauae 105,000 ft
pressure 8.680 mb
lemperature -44.35°C
i
II
100
ii
·
-
--
I
II
II
.90
Altitude in
'
.ll' '
lhousands of
feet
"
80
1-
111
~
70
StAATOC.Dtt[llt
- -- ---- -65.800 ft
II
60 '-IIII
50
40
30
20
\
,
-i1' "-·-~ ··----- ..
·~
......II
we
10
-----o
Pressure mb
Temp.
•c
Rel. dens.%
'%
, - - -
' ·,
0
----
'
Tropapauae 36,090 ft
pressur1122632mb -· --------·
temper111ure -56.5°C
~
--1-
•·
'·-·
temperature _
decreaaes at
1.98°C per
1000 ft
;;;,.__....._
-·····-'
""""'llli~
1•-
0
100
200
300
400
500
600
700
-60
-30
-40
-30
-20
0
10
20
30
40
800
900
1000
-10
0
10
20
30
40
50
60
70
80
90
100
earth's centre, and thi$ being so, atmospheric pressure decreases
steadily with increases of height above the earth's surface.
The units in which atmospheric pressure is expressed are: pounds
per square inch (psi). inches of mercury (in Hg) and millibars (mb).
Conversion factors for these units are given in Appendix l.
The steady fall in atmospheric pressure has a dominating effect on
the density of the air, which changes in direct proportion to changes
of pressure.
Atmospheric temperature
Another important factor affecting the atmosphere is its temperature.
The air in contact with the earth is heated by conduction and
radiation, and as a result its density decreases as the air starts rising.
In doing so, the pressure drop allows the air to expand, and this in
turn causes a fall in temperature from a known sea-level value. It
falls steadily with increasing height up to the tropopause. and the rate
at which it falls is termed the lapse rate (from the Latin lapsus.
meaning slip). In the stratosphere the temperature at first remains
constant at some reduced value, then increases again to a maximum.
Standard atmosphere
In order to obtain indications of airspeed, altitude and vertical speed,
it is of course necessary to know the relationship between the
pressure, temperature and density variables, and altitude. If such
indications are to be presented with absolute accuracy, direct
measurements of the three variables would have to be taken at all
altitudes and fed into the appropriate instruments as correction
factors. Such measurements, while not impossible, would, however,
demand some rather complicated sensor mechanisms. It has therefore
always been the practice to base all measurements and calculations on
what is termed a standard atmosphere, or one in which the values of
pressure, temperature and density at different altitudes are assumed to
be constant. These assumptions have in turn been based on
established meteorological and physical observations, theories and
measurements, and so the standard atmosphere is accepted
internationally. As far as airspeed indicators, altiIJ1eters and vertical
speed indicators are concerned, the inclusion of the assumed values
of the relevant variables in the calibration laws permits the use of
sensing elements that respond solely to pressure changes.
The assumptions are: (i) the atmospheric pressure at mean sea-level
is equal to 14.7 psi, 1013.25 mb, or 29.921 in Hg; (ii) the
temperature at mean sea-level is l5°C (59°F); (iii) the air
temperature decreases by l.98°C (3.556°F) for every 1000 ft
increase in altitude (the lapse rate already referred to) from 15°C at
mean sea-level to -56.5°C (-69.7°F) at 36 090 ft. Above this
altitude the temperature is assumed to remain constant at - 56.5 °C.
It is from the above mean sea-level values that all other
corresponding values have been calculated and presented in what is
27
termed the International Standard Atmosphere (ISA). Altitudes and
values are given in Table I.
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) and pipelines and drains interconnected as shown
diagrammatically in Fig. 2.2. Sensing of the total, or pitot, pressure
(p,) and of the static pressure (p,) is effected by the probe, which is
Basic air data system
Figure 2.2
Basic air data
VERTICAL SPEED INDICATOR
AJRSP!:ED INOlCATOR
sy:-.rem.
e
DRAINS
STAT!Cl.lNE
PITOT llNE
P!TOT-STATIC PIIOBE
AIRSPEED
ALTITUDE
\
r
-IJ
Stat~ic
pres~sure~~p,
6
J____________ _
!(
tr·'--
IT
J~p,
~
P,
t
~
VERTICAL SPEED
__i
suitably located in the airstream and transmits these pressures to the
sensing elements within the indicators.
The pressure transmission produces small displacements of the
sensing elements in such a manner that displacements corresponding
to (a) airspeed are proportional to the difference between p, and p,,
(b) altitude are directly proportional, to p,. and (c) vertical speed are
proportional to the difference between p, and a 'case' pressure p,
produced by a calibrated metering unit. The displacements are, in
turn, transmitted to an indicating element via an appropriate
magnifying system.
The complexity of an air data system depends primarily upon the
type and size of aircraft, the number of locations at which primary
air data are to be displayed, the types of instrument installed, and the
number of other systems requiring air data inputs. The point about
complexity may be particularly noted from Figs 2.3 and 2.4, which
show in schematic form the systems used in two types of public
transport aircraft currently in service.
Probes
Probes may be either of the combined pitot-static tube type, or of the
single pitot tube type, the latter being used in air data systems that
utilize remotely-located static vents or ports (sec also page 35).
Figure 2. 3 Typical pi tot probe
and static vent system (I).
t.U,:ltt.HIYI
l'l'fCll'll:Cll(
A L T E R N A T ~ -·---,-.:.,;:.___,.-;:u;:------------_c=-_
m:!DPITOT
=STATIC
HOSE AND CONNECTORS
_c:=._
_c::::'.J._
c:ffi::=fil::l
~DRAINS
29
Figure 2.4 Typical pitot probe
and static vent system (2)
STATIC
VENT
R AUX PITOT
, ....,.
too,"""i---·n---1,1---------~
?
;,~ytr''i.
i'w·:1:·:·~,
CABIN PRESS
GAUGES
ISiSiSi'i AUXllllll SIATIC 11!1. I
~mo1
~SUHt
CZZJ Al!IIWH
CCl!3 AUllllllll
STATIC
11101 NO I
a::i:::i) AUlllUll 11101 fill. l
'Tf'
DRAINS
LAUX PITOT
cur,11or
Ftgure 2.5 Basic form of
pitot-statn: probe. I Heating
element. 2 static slots, 3 pilot
tube connection, 4 static tube
connection, 5 heater element
cable, 6 external drain I tile, 7
pitot tut> orain hole
6
r
4
A probe of the combined tube type is shown in basic form in Fig.
2.5. The tubes are mounted concentrically, the pitot tube being inside
the static tube which also forms the casing of the probe. Static
pressure is admitted through small ports around the casing. The
pressures are transmitted from their respective tubes by means of
30
Figure 2.6 Combined pitotstatic probe.
P1 PRESSURE LINE
S1 PRESSURE LINE
HEATER
CONNECTOR
ALIGII PIN CUTOUT
S1 & S2 STATIC PRESSURE
metal pipes which may extend to the rear of the probe, or at right
angles to it, depending on whether the probe is to be mounted at the
leading edge of an aircraft's wing, under a wing, or at the side of a
fuselage. Locations of probes will be covered in more detail i..:nder
the heading of 'Position error'.
A chamber is normally formed between the static hole~ ar.d the
pipe connection to smooth out any turbulent air flowing into i.he holes
which might occur when the probe is yawed, before transmitting it to
the instruments.
Protection against icing is provided by a heating eleme.1t fitted
around the pitot tube, or, as in some designs, around the inner
circumference of the probe casing, and in such a position that the
maximum heating effect is obtained at points where ice build-up is
most likely to occur. The temperature/characteristics of some
elements are such that the current cor.sumption is automatically
regulated according to the temperature conditions to which the probe
is exposed.
Figure 2.6 illustrates a type of combined tube probe; it is supported
on a faired casing which is secured to the side of an aircraft's
fuselage by means of a mounting flange. The assembly incorporates
two sets of static holes (SI and S2) connected to individual pipes
terminating at the mounting flang ~; the use of both sets is shown in
more detail in Fig. 2.4. Pitot pressure is transmitted via an
appropriate connecting union and pipe terminating at the mounting
flange.
An example of a 'pitot tube only' type of probe is shown in Fig.
2. 7, and from this it will be noted that its mounting arrangements are
similar to those adopted for the combined tube type. A typical
application of the probe is shown in more detail in Fig. 2.3.
31
Figure 2. 7 Pitot probe.
PITOT
OPENING
Heating circuit arrangements
The heating elements of some probes require a 28 V de supply for
operation, while others are designed to operate from a 115 V ac
supply, their application to any one type of aircraft being governed
principally by the primary power supply system adopted.
In any heating circuit it is of course necessary to have a control
switch, and it is also usual to provide some form of indication of
whether or not the circuit is functioning correctly. Two typical depowered circuit arrangements are shown in Fig. 2.8.
In the arrangement shown at (a) the control switch, when in the
'ON' position, allows current to flow to the heater via the coil of a
relay which will be energized when there is continuity between the
switch and the grounded side of the heater. If a failure of the heater,
or a break in another section of its circuit occurs, the relay will deenergize and its contacts will then complete the circuit from the
second pole of the switch to illuminate the red light which gives
warning of the failed circuit condition. The broken lines show an
alternative arrangement of the light circuit whereby illumination of an
amber light indicates that the heater circuit is in operation.
In the arrangement shown at (b) an ammeter is connected in series
with the heater element so that not only is circuit continuity
indicated, but also the amount of current being consumed by the
element.
An example of an ac-powered heater circuit is shown in Fig. 2.9;
32
Figure 2.8 Typical probe
heating circuit arrangement
(de). (a) Light and relay;
(b) ammeter.
o.c. llUS8AR
c~rr
-~I I
o.c. tlUSIIAA
-=
R£O
LIGHT
AMIIER
UGHT
--- --'
I
I
~-,
RELAY
HEATER
(a)
Figure 2. 9 Ac-powered
heating circuit.
:f
'r"'
WARHING :~", lNOICATING
(bi
HEATER
DC
AC
Indicator
light
Controll.
1
,
switch
1
i ,
't___
--'VVVV'->.
Probe heater
de power is also used for indicator light circuit operation. With both
power supplies available, and the system control switch in the 'ON'
positio·n, single-phase ac is supplied to the heater element via the
primary winding of a transformer. The de power to the amber
indicator lights passes to ground via a normally closed solid-state
switch, so they will initially remain illuminated.
33
When current is supplied to the heater element, a current is also
induced in the secondary winding of the transformer and is supplied
to a bridge rectifier. The rectified output is, in turn, supplied to the
solid-state switch. When the heating current has reached a sufficient
level, the increased rectified output causes the solid-state switch to
interrupt the indicator light circuit; a 'light out' thus indicates that
'probe heat is on'. Functioning of the indicator lights can be checked
by a press-to-test switch within the body of the light unit.
Position error
The accurate measurement of airspeed and altitude by means of a
combined tube type of probe presents two main difficulties: one, to
design a probe which will not cause any disturbance to the airflow
over it, and the other, to find a suitable location on an aircraft where
the airflow over it will not be affected by attitude changes of the
aircraft. The effects of such disturbances are greatest on the static
pressure sensing section of an air data system, giving rise to what is
termed a position or pressure error (PE). This error may be more
precisely defined as 'the amount by which the local static pressure at
a given point in the flow field differs from the free-stream static
pressure'. As a result of PE, an airspeed indicator and an altimeter
can develop errors in their indications. The indications of a vertical
speed indicator remain unaffected by PE.
As far as airflow over a probe is concerned, we may consider it,
and the aircraft to which it is fitted, as being alike because some of
the factors determining airflow are: shape, size, speed and angle of
attack. The shape and size of a probe are dictated by the speed at
which it is moved through the air; a large-diameter casing, for
example, can present too great a frontal area which at very high
speeds can initiate the build-up of a shock wave which will break
down the airflow over the probe. This shock wave can have an
appreciable effect on the static pressure, extending as it does. for a
distance equal to a given number of diameters from the nose of a
probe. One way of overcoming this is to decrease the casing diameter
and to increase the distance of the static holes from the nose of the
probe. Furthermore, a number of holes may be provided along the
length of the casing of a probe spaced in such a way that some will
always be in the region of undisturbed airflow.
A long, small-diameter probe is an ideal one from an aerodynamic
point of view, but it can present certain practical difficulties; for
example, its 'stiffness' may not be sufficient to prevent vibration at
high speed, and it could also be difficult to accommodate the heating
elements necessary for anti-icing. Thus, in establishing the ultimate
relative dimensions of a probe, a certain amount of compromise must
be accepted.
34
Figure 2. IO Static vent or
port.
FUSELAGE SKIN
When a probe is at some angle of attack to the airflow, it causes
air to flow into the static holes which creates a pressure above that of
the prevailing static pressure, and a corresponding error in static
pressure measurement. The pressures developed at varying angles of
attack depend on such factors as the- axial location of the static holes
along the probe casing, their positions around the circumference and
on their size.
Static vents
From the foregoing, it would appear that, if all these problems are
created by pressure effects at the static holes of a probe, they might
as well be separated from it and positioned elsewhere on an aircraft.
This is, in fact, a solution put into practice on many types of aircraft
by using a single tube type of probe (see Fig. 2. 7) in conjunction
with a static vent located in the side of the fuselage. A typical static
vent is shown in Fig. 2.10.
Location of probes and static vents
The choice of probe locations is largely dependent on the type of
aircraft, speed range and aerodynamic characteristics, and as a result
there is no common standard for all aircraft. Typical locations are:
ahead of a wing tip, under a wing, ahead of a vertical stabilizer tip,
35
at the side of a fuselage nose section, and ahead of a fuselage nose
section.
Independent static vents, when fitted, are always located one on
each side of a fuselage and interconnected so as to balance out
dynamic pressure effects resulting from any yawing or sideslip
motion of an aircraft.
The actual PE due to a chosen location is determined for the
appropriate aircraft type during the initial flight-handling trials of a
prototype, and is finally presented in tabular or graphical format, thus
enabling corrections to be directly applied to the readings of the
relevant air data instruments, and as appropriate to various operating
configurations. In most cases, corrections are performed automatically
and in a variety of ways. One method is to employ aerodynamicallycompensated probes, i.e. probes which are so contoured as to create
a local pressure field which is equal and opposite to that of the
aircraft, so that the resultant PE is close to zero. Other methods more
commonly adopted utilize correction devices either within separate
transducers, or within central air data computers (see page 165).
Alternate pressure sources
If failure of the primary pitot and static pressure sources should
occur, e.g. complete icing-up of a probe due to a heating element
failure, then of course errors will be introduced in instruments'
indications and in other systems dependent on such pressures. As a
safeguard against failure, therefore, a standby system may be
installed in aircraft employing combined tube type probes whereby
static atmospheric pressure and/or pitot pressure from alternate
sources can be selected and connected into the primary system.
The required pressure is selected by means of selector valves
connected between the appropriate pressure sources and the air data
instruments, and located within easy reach of the flight crew. Figure
2.11 illustrates diagrammatically the method adopted in a system
utilizing an alternate static pressure source only. The valves are
shown in the normal operating position, i.e. the probes supply pitot
and static pressures to the instruments on their respective sides of the
aircraft.
In the event of failure of static pressure from one or other probe,
the instruments can be connected to the alternate source by manually
changing over the position of the relevant selector valve.
The layout shown in Fig. 2.12 is one in which an alternate source
of both pitot and static pressures can be selected. Furthermore, it is
an example of a system which utilizes the static holes of a combined
tube type of probe as the alternate static pressure source. The valves
are shown in their normal position, i.e. the probes supply pitot
pressure to the instruments on their respective sides of an aircraft:
Figure 2.1 J Alternate ~tatic
pressure system.
TO CAPTAIN'S
AIR DATA
INSTRUMENTS
TO FIRST OFFICER'S
AIR DATA
INSTRUMENTS
R.H.
PtTOT•STATIC
l'OOIIE
D
ALTERNATE
STATIC \/ENT
=-=
==
PtTOT P'l!fSSURE
HOOMAI.. STATIC
~ ALTERNATE STATIC
and the static pressure is supplied from static vents. In the event of
failure of pitot pressure from one or other probe, the position of the
relevant selector valve must be manually changed over to connect the
air data instruments to the opposite probe. The alternate static source
is selected by means of a valve similar to that employed in the pitot
pressure system, and, as will be seen from Fig. 2.12, it is a
straightforward change-over function.
The probes employed in the system just described are of the type
illustrated in Fig. 2.6, reference to which shows that two sets of
static holes (front and rear) are connected to separate pipes at the
mounting base. In addition to being connected to their respective
selector valves, the probes are also coupled to each other by a crossconnection of the static holes and pipes; thus, the front set of holes
are connected to the rear set on opposite probes. This balances out
any pressure differences which might be caused by the location of the
static holes along the fore-and-aft axis of the probes.
Pipelines and drains
Pitot and static pressures are transmitted through seamless and
corrosion-resistant metal (light alloy and/or tungum) pipelines, and,
37
Figure 2.12 Alternate pilot
pressure and static pressure
system.
TO CAPTAIN'S
AIR DATA
INSTRUMENTS
.----.,
SELECTOR
VALVES
TO FIRST OFFICER'S
AIR DATA
INSTRUMENTS
,-.._
L.H.
R.H.
PITOT·STATIC
PROBE
PITOT• STATIC
PROBE
STATIC VENTS
PITOT
PRESSURE
{
::JIIIC LH .
:IIII: R.H
===
STATIC {
NORMAL
PRESSURE ~ ALTERNATE
where connections to components mounted on anti-vibration
mountings are required, flexible pipelines are used. The diameter of
pipelines is related to the distance from the pressure sources to the
instruments in order to eliminate pressure drop and time-lag factors.
In order for an air data system to operate effectively under all
flight conditions, provision must also be made for the elimination of
water that may enter the system as a result of condensation, rain,
snow, etc., thus reducing the probability of 'slugs' of water blocking
the lines. Such provision takes the form of drain holes in probes,
drain traps and valves in the system's pipelines. Drain holes provided
in probes are of such a diameter that they do not introduce errors in
instrument indications.
The method of draining the pipelines varies between aircraft types,
and some examples are shown in Fig. 2.13. Drain traps are designed
to have a capacity sufficient to allow for the accumulation of the
maximum amount of water that could enter a system between
servicing periods. Valves are of the self-closing type so that they
cannot be left in the open position after drainage of accumulated
water.
38
Figure 2. I3 Air data system
drains. (a) Typical valve
construction; (b) Transparent
water trap and drain valve; (c)
combined sump and drain
valve; (d) separate water traps
and drain valves; (e) combined
sump and drain valve with sight
gauge.
(b)
SIGHT GAUGE
OIIAIIG£ FLOAT
/
BAYONET FITT1116 CAP
(Cl
(d)
(e)
Air data instruments
The three primary air data instruments may be either of the pure
'pneumatic' type, or the servo-operated type. Pneumatic-type
instruments are those which are connected to probes and/or static
vents, and therefore respond to the pressures transmitted directly to
them. They are commonly used in the more basic air data systems
installed in many types o( small aircraft, while in the more complex
systems adopted in large public transport aircraft, they are used only
in a standby role.
Servo-operated instruments are, on the other hand, of the indirect
type in that they respond to electrical signals generated by pressure
transducers within central air data computers (CADCs) to which
probes and static vents are connected. The fundamental principles of
these instruments will be described-in a later chapter.
39
Figure 2.14 Pitot pressure.
---------- -
~
----::..Y..:iii: __-,_--:_-_- --------
-------
------------
STAGNATION POINT
~Airspeed indicators (pneumatic)
These indicators measure speed in terms of the difference between
the pitot and static pressures detected by either a com4iped pitot-static
probe, or a pitot probe and static vent, as appropriate:_J
L
Pitot pressure This may be defined as the additional pressure
produced on a surface when a flowing fluid is brought to rest, or
stagnation, at the surface.
Let us consider a pitot probe placed in a fluid with its open end
shown in Fig. 2.14. When the fluid flows at a
facing upstream
certain velocity V over the probe it will be brought to rest at the
nose; this point is known as the stagnation point. If the fluid is an
ideal one, i.e. is not viscous, then the total energy is equal to the
sum of the potential energy, the kinetic energy and pressure energy,
and remains constant. In connection with this probe, however, the
potential energy is neglected, thus leaving the sum of the remaining
two terms as the constant.
In coming to rest at the stagnation point, kinetic energy of the fluid
is converted into pressure energy. This means that work must be
done by the mass of fluid and this raises an equal volume of fluid
above the level of the fluid stream. The work done in raising the
fluid is equal to the product of its mass, the height through which it
is raised, -and acceleration due to gravity. It is also equal to the
product of the ratio of the mass (m) to density (p) and pressure (p );
thus,
as
/ Work done
I
l...
= !!!. p
p
TheRinetic energy ofa mass m before being brought to rest is
equ"afto ml72";"wFiere V is the speed, and since this is
---·--·-·'·'·-·····
pressure energy,
t
,---·
m
I
-p=y
p
------- -- -~- --- -
Therefore
lp =
t
t
~ t i t y p V2 is additional to the static pressure in the region
of the fluid flow, and is usually referred to as the dynamic pressure,
denoted by the letter Q.
The factor assumes that the fluid is an ideal one and so does not
take into account the fact that the shape of a body subject to fluid
flow may not bring the fluid to rest at the stagnation point. This
coefficient is, however, determined by experiment and for pitot
pressure probes it has been found that its value corresponds almost
exactly to the theoretical one.
The pV 2 law, as it is usually called in connection with airspeed
measurement, does not allow for the effects of compressibility of air
as speed increases. In order therefore to minimize 'compressibility
errors' in indication, the calibration law is modified as follows:
t
t
p
= f pV2
(1 + t :: )
= pressure difference (mmH20)
= density of air at sea-level
= speed of aircraft (mph or knots)
a 0 = speed of sound at sea-level (mph)
where p
p
V
Airspeed terminology
.Jndicated airspeed (IAS) fj:he readings of an airspeed indicator
corrected only for instrument erroJJ i.e. the difference between the
true value and the indicated value. Errors and appropriate corrections
to be applied are determined by comparison against calibration
equipment having high standards of accuracy.
_computed airspeed [Basically, this is IAS with corrections for
position error (PEDapplied (see page 34). The term 'computed'
appiies 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 modulei
,._ Equivalent airspeed (EAS) {fhis 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.
41
Figure 2.15 Airspeed
terminology.
Pitot pressure ( p , ) - - - - - - - - - - - o . 1
Static pressure (P,)-----1-----11,,1
AIRSPEED
SENSING
MODULE
P.E. correction
PNEUMATIC
AIRSPEED
INDICATOR
Indicated
airspeed (IAS)
1-------0------- Computed
airspeed
Square-law
compensation
Mach speed
Altitude
Total air
temperature
Static
pressure (p,)
COMPRESSIBILITY
COMPENSATION
Equivalent
airspeed (EAS)
AIR DENSITY
COMPENSATION
True
airspeed (TAS)
True airspeed (T AS) This is EAS compensated for changes in air
temperature and density at various flight levels. This is also done
automatically in air data computer systems.
The foregoing airspeeds are summarized pictorially in Fig. 2.15.
limiting speeds
Vmo
Mmo
Maximum operating speed in knots.
Maximum operating speed in terms of Mach number.
_ Typical indicator {The mechanism of a typical pneumatic-type
airspeed indicator is illustrated in Fig. 2.16. The pressure-sensing
element is a metal capsule, the interior of which is connected to the
pitot pr~ssure connector via a short length of capillary tul-?Je which
damps out pressure surges. Static pressure is exerted on the exterior
of the capsule and is fed 1nto the· instrument case via the second
connector. Except for this connector the case is sealed.
Displacements of the capsule in accordance with what is called the
'square-law' are transmitted via a magnifying lever system, gearfog,
and a square-law compensating device to the pointer, which moves
over a linear scale calibrated in knots. eperature compensation is
achieved by a bimetallic strip arranged to vary the magnification of
42
Figure 2. 16 Typical pneumatic
airspeed indicator.
t
Pitot
pressure
t
Static
pressure
Figure 2. I 7 Square-law
characteristics. (a) Effect of
linear deflection/pressure
response; (b) effect of direct
magnification.
I
'150
200
250-
CAPSULE DfFUCTION (II)
(b)
the lever system in opposition to the effects of temperature on system
and capsule sensitivity~
Square-law compensation Since airspeed indicators measure a
differential pressure which varies with the square of the airspeed, it
follows that, if the deflections of the capsules responded linearly to
the pressure, the response characteristic in relation to speed would be
similar to that shown in Fig. 2.17(a). If also the capsule were
coupled to the pointer mechanism so that its deflections were directly
magnified, the instrument scale would be of the type indicated at (b).
The non-linearity of such a scale makes it difficult to read
accurately, particularly at the low end of the speed range;
furthermore, the scale length for a wide speed range would be too
great to accommodate conveniently in the standard dial sizes.
Therefore, to obtain the desired linearity a method of controlling
either the capsule characteristic, or the dimensioning of the coupling
43
Figure 2. I 8
compensator.
spring length
spring makes
screws.
Tuning spring
OX = effective
diminishing as
contact with
RANGING
SCREWS
RANGING SPRING
element conveying capsule deflections to the pointer, is necessary. Of
the two methods the latter is the more practical because means of
adjustment can be incorporated to overcome the effects of capsule
'drift' plus other mechanical irregularities as determined during
calibration ..
The principle of a commonly used version of the foregoing method
is one in which the length of a lever is altered as progressive
deflections of the capsule take place, causing the mechanism, and
pointer movement, to be increased for small deflections and
decreased for large deflections. In other words, it is a principle of
variable magnification.
Another type of square-law compensating device is shown in Fig.
2.18. It consists of a special ranging or 'tuning' spring which bears
against the capsule and applies a controlled retarding force to capsule
expansion. The retarding force is governed by sets of ranging screws
which are pre-adjusted to contact the spring at appropriate points as it
is lifted by the expanding capsule. As speed and differential pressure
increase, the spring rate increases and its effective length is
shortened; thus linearity is obtained directly at the capsule and
eliminates the need for a variable magnifying lever system. In some
types of servo-operated indicators, a specially profiled cam provides
square-law compensation (see also page 171).
/Machmeters and Mach/airspeed indicators
In order for aircraft to operate at speeds approaching and exceeding
that of sound, their aerodynamic profiles and structural design must
be such that they minimize, or ideally overcome, the limiting effects
that high-velocity airflow and its associated forces could otherwise
have on in-flight behaviour of aircraft. Since the speed of sound
44
Figure 2. /9 Machmeler.
I Airspeed capsule. 2 allitude
capsule. 3 altilude rocking
shaft, 4 sliding rocking shaft.
5 calibra1ion spring (square-law
compensation}, 6 calibralion
screws (square-law
compensation}.
6
---7'"-....._
I
~
\
~
\
-
PITOT
PRESSUIIE
...---
MOVEMENTS OUE TO AIRSPEED CHANGES
- -
MOVEMENTS DUE TO ALTITUDE CHANGES
depends on atmospheric pressure and density, it will vary with
altitude, and this suggests that for an aircraft to operate within speed
limits commensurate with structural safety, a different speed would
have to be maintained for each altitude. This obviously is not a
practical solution, and so it is therefore necessary to have a means
whereby the ratio of an aircraft's speed, V, and the speed of sound,
a, can be computed from pressure measurement and indicated in a
. conventional manner; This ratio, Via, is termed the Mach number
{M), and the instrument which measures it is termed a Machmeter.
,-·A Machmeter is a compound air data instrument which, as may be
seen from Fig. 2.19, accepts two variables and uses them to compute
tte re~~i:_atioJfl'he first variable is air.speed and therefore a
mechanism based on that of a conventional airspeed indicator is
adopted to measure this in terms of .!h.e. pressure differenc.5: p, - p.,
where p, is the -~_tal_<:>t:._[l_itQ_Lpressure, and Ps is th~§~essure."J
[!he second variable is q[titude' and this is also measured in the
conlentional manner, i.e. by means of an aneroid capsule sensitive to
p ..)eeflections of the capsules of both mechanisms are transmitted to
the mdicator pointer by rocking shafts and lever_§] the dividing
function of the altitude unit being accomplished' by an intermediate
sliding rocking shaft.
Let us assume that the aircraft is flying under standard sea-level
conditions at a speed V of 500 mph. The speed of sound at sea-level
45
is approximately 760 mph, therefore the Mach number is 500/760 =
0.65. Now, the speed measured by the airspeed mechanism is, as we
have already seen, equal to the pressure difference p,-Ps, and so the
sliding rocking shaft and levers A, B, C and D will be set to angular
positions determined by this difference. The speed of sound cannot be
measured by the instrument, but since it is governed by static
pressure conditions, the altimeter mechanism can do the next best
thing and that is to measure p, and feed this into the indicating
system, thereby setting a datum position for the point of contact
between the levers C and D. Thus a Machmeter indicates the Mach
number Via in terms of the pressure ratio (p,-p,)lp5 , and for the
speed and altitude conditions assumed the pointer will indicate 0.65.
What happens at altitudes above sea-level? As already pointed out,
the speed of sound decreases with altitude, and if an aircraft is flown
at the same speed at all altitudes, it gets closer to and can exceed the
speed of sound. For example, the speed of sound at 10 000 ft
decreases to approximately 650 mph, and if an aircraft is flown at
500 mph at this altitude, the Mach number will be 500/650 = 0.75,
a IO per cent increase over its sea-level value. It is for this reason
that critical Mach numbers (Mer;,) are established for the various
types of high-speed aircraft, and being constant with respect to
altitude it is convenient to express any speed limitations in terms of
such numbers.
We may now consider how the altitude mechanism of the
Machmeter fiinctions in order to achieve this, by taking the case of
an aircraft having an Mer;, of, say, 0.65. At sea-level and as based on
our earlier assumption, the measured airspeed would be 500 mph to
maintain M 1, = 0.65. Now, if the aircraft is to climb to and level
off at a flight altitude of IO 000 ft, during the climb the decrease of
p, causes a change in the pressure ratio. It affects the pressure
difference p, - p, in the same manner as a conventional indicator is
affected, i.e. the measured airspeed is decreasl;'d. The airspeed
mechanism therefore tends to make the pointer indicate a lower Mach
number. However, the altitude mechanism simultaneously responds to
the decrease in p,, its capsule expanding and causing the sliding
rocking shaft to carry lever C_ towards the pivot point of lever D.
The magnification ratio between the two levers is therefore altered
as the altitude mechanism divides p, - Ps by p,, lever D being forced
down so as to make the pointer maintain a constant Mach number of
0.65.
The critical Mach number for a particular type of aircraft is
indicated by a pre-adjusted lubber mark located over the dial of the
Machmetc,
0
,
Mach/airspeed indicator This indicator is one which combines•the
functions of both a conventional airspeed indicator and a Mach meter,
MACH NO. SCALE
Figure 2.20 Mach/airspeed
AIRSPEED POINTER
indicator.
LIMIT SPEED
tV,.o) POINTER
EXTERNAL INDEX~~
POINTER
COMMAND BUG
SETIING KNOB
EXTERNAL INDEX
POINTER
COMMAND BUG
V,.o SETIING
KNOB
and presents the requisite information in the manner shown in Fig.
2.20. The mechanism consists of two measuring elements which drive
their own indicating elements, i.e. a pointer and a fixed scale t9
indicate airspeed, and a rotating dial and scale calibrated to indicate
Mach number. A second pointer, known as the velocity maximum
operating (V,,w) pointer, is also provided for the purpose of indicating
the maximum safe speed of an aircraft over its operating altitude
range; in other words, it is an indicator of critical Mach number. The
pointer is striped red and white and can be pre-adjusted to the desired
limiting speed value, by pulling out and rotating the setting knob in
the bottom right-hand corner of the indicator bezel. The adjustment is
made on the ground against charted information appropriate to the
operational requirements of the particular type of aircraft. The
purpose of the setting knob in the bottom left-hand corner of the
bezel is to enable the pilot to position a command 'bug' with respect
to the airspeed scale, thereby setting an airspeed value which may be
used as a datum for an autothrottle control system, or as a fast/slow
speed indicator. Two external index pointers around the bezel may be
manually set to any desired reference speed, e.g. the take-off speeds
V 1 and VR.
In operation, the airspeed measuring and indicating elements
respond to the difference between pitot and static pressures in the
conventional manner, and changes in static pressure with changes in
altitude cause the Mach number scale to rotate (anti-clockwise with
increasing altitude) relative to the Vmo pointer. When the limiting
speed is reached, and the corresponding Mach number graduation
coincides with the Vmo pointer setting, mechanical contact is made
between the scale and pointer actuating assemblies so that continued
rotation of the scale will also cause the pointer to rotate in unison.
47
The pointer rotates against the tension of a hairspring which returns
the pointer to its originally selected position when the Mach speed
decreases to below the limiting speed. It will be noted from Fig. 2.20
that at the high end of the speed range, the airspeed pointer can also
register against the Mach scale, thereby giving a readout of speed in
equivalent units. The necessary computation is effected by calibrating
the scales to logarithmic functions of pitot and static pressures.
In addition to their basic indicating function, Mach/airspeed
indicators can also be designed to acutate switch units coupled to
visual or audio devices which give warning when such speeds as
Mach limiting, or landing gear extension are reached. In aircraft
having an autothrottie system, certain types of Mach/airspeed
indicator are de!>igned to provide a speed errnr output which is
proportional to the difference between the reading indicated by the
airspeed pointer and 1;1e setting of the command 'bug'. This is
accomplished by means of a CT/CX synchro (see page 140)
combination which senses the positions of the airspeed pointer and
the command bug, and produces an output error signal which, after
amplification, is then supplied to the autothrottle system.
Indicated/computed airspeed indicator
An example of this type of indicator is show~1 in Fig. 2.21. It is very
similar in construction and presentation to the Mach/airspeed
indicator in that it employs pitot and static pressure-sensing elements
which position the appropriate pointers. It has, however. the
additional feature of indicating the airspeed computed by a central air
data computer (see Chapter 7). The indicating element for this
purpose is a servomotor-driven digital counter, the motor being
supplied with signals from a synchronous transmission system. In the
MAXIMl'M OPERATING
SPEtO POINTER
Figure 2. 21 Indicated/
computed airspeed indicator.
COMPUTED AIRSPEED COUNTER
COMMAND AIRSPEED
COUNTER
INDICATED AIRS?EEO
i>OINTE1l
I
COMMAND AIRSPEED
SET :;NOB
48
COMPUTED AIRSPEED
ON/OFF SWITCH
event of failure of such signals a. yellow warning flag obscures the
counter drums. A check on the operation of the failure monitoring
and flag circuits. can be made by moving the calibrated airspeed
(CAS) switch from its normally 'ON' position to 'OFF'.
As in the case of certain types of Mach/airspeed indicators,
provision is made for setting in a command airspeed signal and for
transmitting it to an autothrottle system which will adjust the engine
power to attain a commanded speed. In the example illustrated, the
command set knob mechanically adjusts a synchrotel (see page 149)
which also senses indicated airspeed. Thus, the synchrotel establishes
the airspeed error signal output required by the autothrottle computer.
A readout of the command speed is given on a digital counter which
is also mechanically set by the command speed knob.
Figure 2.22
altimeter.
Pneumatic-type
,t!ltimeters (pneumatic)
CThesei~struments operate on the aneroid barometer princ1pl~ in
other words they respond to changes in atmospheric pressure, and in
accordance with appropriate calibration laws they indicate these
changes in terms of equivalent altitude values.
The dial presentation and mechanical features of a typical
pneumatic type of altimeter are shown in Fig. 2.22. iThe prcssui esensing element consists Of !".Y-~Sules, which transmit their
MECHANISM
BODY
SPIRAL
GEAR
DIAL
l!JMETAL
l~f',,.n"'J';-,,,.,C.--1000 FT
ORUM
COMPENSATOR
BARO
SET KNOB
49
Figure 2.23 Conversion of
pressure/height relationship to a
linear scale.
INSTRUMENT INDICATION
WJTt\D~i~i'1.~~ v
HEIGHT
1 REVOLUTION - 1000 FT
(FT)
PRESSURE - HEIGHT LAW
50.000
4
""!""
"" ""
\I\.
20.000
.,v
/
V --
,;',,. '
10.000
""
/
/
\
30.000
I
0
0
0
3
\
40.000
+-
t
--
0.12
0.16
,
I
I/
1000
.,v
0
0.04
0.08
'"-
200 400: 600 800
ATMOSPHERIC PRESSURE
ON CAPSULES
(MILLIBARS)
/
+- 7'
/
V
0.20
VARIABLE MAGNIFICATION CAPSULE
PRESSURE-DEFLECTION
LEVER AND GEAR
DEFLECTION. CHARACTERISTICS OF
SYSTEM
(INCHES)
CAPSULES
2
EXAMPLE = 20 000 ft
1 PRESSURE = 466 rnbar
2 CAPSULE DEFLECTION= 0.087 in
3 POINTER ROTATION= 20 revs
4 INDICATED HEIGHT= 20 000 ft
deflections in response to pressure changes. to a single pointer and
altitude drum via sector gears and pinior§{fhe direction of the solid
arrows shown in the diagram corresponds to the movements obtained
under increasing altitude conditions} The complete mechanism is
contained within a casing which, with the exception of the static
pressure connection, forms a sealed unit.
In order to derive a linear altitude scale from the non-linear
pressure/altitude relationship (sec Fig. 2.1 again) it is necessary to
incomorate some form of conversion within the altimeter mechanism.
This conversion is represented by the graphical example shown in
Fig. 2.23. Typically, linearity is obtained by a suitable choice of
material for the capsules and their corresponding deflections (curve 2)
and also choice of deflection characteristics of the variable
magnification and lever system for transmitting the reJevant
deflections to the pointer (curve 3). The resultant of both curves
produces the linear scale as at curve 4. To cater for variations
between deflection characteristics of individ·ual capsules, and so allow
for calibration, adjustments are always provided whereby the lever
and gear system magnification may be matched to suit the capsule
characteristics.
(The pressure-sensing element of the altimeter is compensated for
changes in temperature of the air supplied to it by a bi-metal
compensator device connected in the magnification lever system\ The
50
--
temperature coefficient of the instrument is chiefly due to the change
of elasticity of the capsule material with change of temperature; this,
in tum, varies the degree of deflectiQn of the capsules in relation 'to
the pressure acting external to them.1.,for example, if at sea-level the
temperature should decrease, the elasticity of the capsules would
increase; in other words, and from the definition of elasticity, the
capsules have a greater tendency 'to return to their original size' and
so would expand and cause the altimeter to over-read. At higher
altitudes the same effects on elasticity will take place, but since the
pressures acting on the capsules will also have decreased, then, by
comparison, the expansion of the capsules becomes progressively
greater: The bi-metal compensator is simultaneously affected by the
decrease in ambient temperature, but by virtue of its charac;:teristics it
exerts forces through the lever system to oppose the error-producing
deflection of the capsules)
Barometric pressure setting As pointed out earlier in this chapter,
the basis for the calibration of air data instruments is the ISA and its
assumed values. As far as altimeters are concerned, they will, under
ISA conditions, indicate what is termed pressure altitude. In practice,
however, atmospheric pressure and temperature are continually
changing, and so under these 'non-standard' conditions altimeters
would be in error and would then display what is termed indicated
altitude.
We may consider these errors by taking the case of a simple
altimeter situated at various levels. In standard conditions, and at a
sea-level airfield, an altimeter would respond to a pressure of
1013.25 mb (29.92 in Hg) and indicate the pressure altitude of zero
feet. Similarly, at an airfield level of 1000 ft, it would respond to a
standard pressure of 977.4 mb (28.86 in Hg) and indicate a pressure
altitude of 1000 ft. Assuming that at the sea-level airfield the pressure
falls to 1012.2 mb (29.89 in Hg), the altimeter will indicate that the
airfield is approximately 30 ft above sea-level; in other words, it will
be in error by +30 ft. Again, if the pressure increases to 1014.2 mb
(29.95 in Hg), the altimeter in responding to the pressure change will
indicate that the airfield is approximately 30 ft below sea-ltwel, an
error of - 30 ft.
In a similar manner, errors would be introduced in the readings of
such an altimeter in flight and whenever the atmospheric pressure at
any particular altitude departed from the assumed standard value. For
example, when an aeroplane flying at 5000 ft enters a region in
which the pressure has fallen from the standard value of 842.98 mb
to, say, 837 mb, the altimeter will indicate an altitude of
approximately 5190 ft.
Since the ISA also assumes certain temperature values at all
altitudes, then consequently non-standard values can also cause errors
51
Figure 2.24 Effect of
atmospheric temperature on an
altimeter.
I
I
C
.I
8 1-
C'4: u
H 5000FT'--e.L
I~
___
A - - - - - ' - - A 1-·-•----'-·-A
2
in altimeter readings. Variations in temperature cause differences of
air density and therefore differences in weight and pressure of the
air. This may be seen from the three columns shown in Fig. 2.24. At
point A the altimeter measures the pressure of the column AC. At
point H, which is, say, at an altitude of 5000 ft above A, the
pressure on the altimeter is less by that of the part AB below it. If
the temperature of the air in part AB increases, the column of air
wifl expand to A 1Bi, and so the pressure on the altimeter at H wiil
now be less by that of A 1H. The pressure of A 1B1 is, however, still
the same as that of AB, and so the pressure of A 1H must be less than
that of AH. Thus the altimeter, in rising from A 1 to 5000 ft, will
register a smaller reduction of pressure than when it rose from point
A to 5000 ft. In other words, it will read less than 5000 ft. Similarly,
when the temperature of the air between points A and H decreases,
the part AB of the column reduces to A 2B2 and the change of
pressure on the altimeter in rising from A 2 to 5000 ft will be not only
the pressure of A2B2 (which equals AB) but also the pressure of B2H.
The altimeter will thus read a greater pressure Mop and will indicate
an altitude greater than 5000 ft. The relationship between the various
altitudes associated with flight operations is presented graphically in
Fig. 2.25.
It will be apparent from the foregoing that, although the simple
form of altimeter performs its basic function of measuring changes in
atmospheric pressure accurately enough, the corresponding altitude
indications are of little value unless they are corrected to standard
pressure data. In order, therefore, to compensate for altitude errors
52
Figure 2.25 Relation between
various altitudes.
PMSSUl!E AT THIS LEVEL.
1<>00 IN HG
PRESSURE ALTITU0£ IN
,·, DENSITY ALTITUOE lN
STANDARD ATh!OSPHERE. 16250 FT
STANllARt> A~RE,•13400 FT
j__
ABSOLUTE ALTITUO!:
13000 FT
Alftf1£L[) El.£VATION
2
FT
due to atmospheric pressure changes, altimeters are provided with a
manually-operated setting device which allows prevailing ground
pressure values to be preset.
In the altimeter shown in Fig. 2.22, the adjustment device consists
of two drum counters (one calibrated in in. Hg and the other in mb)
interconnected through gearing to a setting knob. When the knob is
rotated then, as shown by the dotted arrows, both counters can be set
to indicate the prevailing barometric pressure; i.e. the static pressure,
in the equivalent units of measurement. Likewise it will be noted that
the setting knob is also geared to the sensing element mechanism
body, so that this mechanism can also be rotated. The deflected
position of the capsules appropriate to whatever pressure is acting on
them at the time will not be disturbed by rotation of the mechanism.
However, in order to maintain the correct pressure/altitude
relationship, rotation of the setting knob will cause the altitude
pointer and drum to rotate and so indicate the altitude corresponding
to the pressures set on the counters. The underlying principle of this
may be understood by considering the setting device to be a millibar
scale having a simple geared connection to the altitude pointer as
shown in Fig. 2.26.
At (a) the altimeter is assumed to be subjected to standard
conditions; thus the millibar scale, in this case, when set to
1013.25 mb, positions the pointer at du; 0 ft graduation. If the setting
is then changed to, say, 1003 mu as at (b), the scale will be rotated
clockwise, causing the altitude pointer to rotate anti-clockwise and to
indicate - 270 ft. If now the altimeter is raised through 270 ft as at
53
Figure 2.26 Principle of
baromelric pressure selling.
(8)
lb)
(c)
(c) a pressure decrease of 10 mb will be sensed by the capsule and its
corresponding deflection will cause the altitude pointer to return to
the zero graduation. Thus, whatever pressure is set on the millibar
scale, the altimeter will indicate zero when subjected to that pressure.
Similarly, any setting of the altitude pointer automatically adjusts the
millibar scale reading to indicate the pressure at which the altitude
indicated will be zero.
'Q' code for altimeter setting The setting of altimeters to the
barometric pressures prevailing at various flight levels and airports is
part of flight operating techniques, and is essential for maintaining
adequate separation between aircraft, and also terrain clearance during
take-off and landing. In order to make the settings flight crew are
dependent on observed meteorological data which are requested and
transmitted from air traffic control. The requests and transmissions
are adopted universally and form part of the ICAO 'Q' code of
communication.
There are two code letter groups commonly used in connection
with altimeter setting procedures, and they are defined as follows:
QFE Setting the barometric pressure prevailing at an airport to
make the altimeter read zero on landing at, and taking off from, that
airport. The zero reading is regardless of the airport's elevation
above sea-level.
QNH Setting the barometric pressure to make the altimeter read
airport elevation above sea-level on· landing and take-off. The
pressure set is a value reduced to mean sea-level in accordance with
ISA. When used for landing and take-off, the setting is generally
known as 'airport QNH'. Any value is only valid in the immediate
vicinity of the airport concerned.
Since an altimeter with a QNH setting reads altitude above sealevel, the setting is also useful in determining terrain clearance when
an aircraft is en route. For this purpose, the UK and surrounding seas
are divided into fourteen Altimeter Setting Regions, each transmitting
an hourly 'regional QNH' forecast.
There is also a third setting and this is referred to as the Standard
Altimeter Setting (SAS), in which the barometric pressure counters
54
Figure 2.27 Altitude. devation
and height.
Flight levels - SAS
Transition altitude
Height - QFE
t
Altitude - QNH
Elevation
Sea-level
are set to the ISA values of 1013.2 mb or 29.92 in Hg. It is used for
flights above a prescribed transition altitude and has the advantage
that with all aircraft using the same airspace and flying on the same
altimeter setting. the requisite limits of separation between aircraft
can more readily be maintained. The transition altitude within UK
airspace is usually 3000 ft to 6000 ft, and from these data altitudes
are quoted as flighr levels: e.g. 4000 ft is FL 40 and 15 000 ft is FL
150.
The following definitions, together with Fig. 2 .27, show how the
terms 'altitude'. 'elevation· and 'height' are used in relation to
altimeter setting procedures.
Altitude is the vertical distance of a level, point or object
considered as a point above mean sea-level. Thus, an altimeter
indicates an altitude when a QNH is set.
Elevation is the vertical distance of a fixed point above or below
mean sea-level. For altimeter settings the QFE datum used is the
airport elevation which is the highest usable point on the landing
area. Where a runway is below the airport elevation, the QFE datum
used is the elevation of the touchdown point, referred to as
touchdown elevation.
Height is the vertical distance of a level, point or object considered
as a point measured from a specified datum. Thus, an altimeter
indicates a height above airport elevation (the specified datum) when
a QFE is set.
~ertical speed indicators
These indicators (also known as rate-of-climb indicators) are the third
of the primary group of air data instruments, and are very sensitive
55
Figure 2.28 Vertical speed
indicator presentation.
differential pressure gauges, designed to indicate the rate of altitude
change from variations in static pressure alone.
Since the rate at which the static pressure changes is involved in
determining vertical speed, a time factor has to be introduced as a
pressure function. This is accomplished by incorporating a special air
metering unit in the sensing system, its purpose being to create a lag
in static pressure across the system and so establish the required
pressure differences.
A pneumatic type of indicator consists basically of three principal
components: a capsule-type sensing element, an indicating element
and a metering unit, all of which are housed in a sealed case
connected to the static pressure source. The dial presentation is such
that zero is at the 9 o'clock position; thus the pointer is horizontal in
the straight and level flight attitude, and can move from this position
to indicate climb and descent in the correct sense. Certain types of
indicator employ a linear scale, but in the majority of applications,
indicators having a scale calibrated to indicate the logarithm of the
rate of pressure change are preferred. The reason for this is that a
logarithmic scale is more open near the zero graduation, and so
provides for better readability and for more accurate observation of
variations from· level flight conditions. A typical example of this
presentation is shown in Fig. 2.28.
An indicator mechanism is shown in schematic form in Fig. 2.29,
from which it will be noted that the metering unit forms part of the
static pressure connection and is connected to the interior of the
capsule by a length of capillary tube. This tube serves the same
purpose as the one employed in a pneumatic type of airspeed
indicator, i.e. it prevents pressure surges reaching the capsule. It is,
however, of greater length due to the fact that the capsule is much
more flexible and sensitive to pressure changes. The other end of the
metering unit is open to the interior of the case to apply static
pressure to the exterior of the capsule. Let us now see how the
instrument operates under the three flight conditions shown in the
diagram.
56
figure 2. 29 Principle of
venical speed indicator.
STATIC
PRESSURE-
i={ \
\
CAl'!U.ARY
IEl==::::lll•
(j~t
DESCENT
Level !light
!!,,'CREASING
STATIC PRESSURE
Descent
i
DECREASING
STATIC PR!:SSURE
DESCENT
Climb
In level flight, air at the prevailing static pressure is admitted to the
interior of the capsule, and also to the instrument case via the
metering unit. Thus, there is no difference of pressure across the
capsule and the pointer indicates zero.
At the instant of commencing a descent, there will still be no
57
Figure 2. 30 Cons1ruction of a
typical vertical speed indicator.
I Rocking shaft assembly.
2 sector. 3 hand-staff pinion.
4 gearwheel, 5 eccen1ric shaft
assembly, 6 capsule plate
assembly, 7 calibration springs,
8 capsule, 9 capillary tube,
10 caiibration bracket. 11 static
connection, 12 metering unit.
13 mechanism body.
14 hairspring, 15 link.
16 balance weight.
3
pressure difference, but as the aircraft is now descending into
conditions of higher static pressure then such pressure will be directly
sensed inside the capsule. The pressure inside the case, however, will
not build up at the same rate as the capsule pressure, because in
having to pass through the metering unit the airflow into the case is
restricted. Thus, a differential pressure is created across the capsule,
causing it to distend and so make the pointer indicate a descent.
During a climb, an aircraft will of course pass through conditions
of decreasing static pressure, but as the metering unit will then
restrict the airflow out of the case, a differential pressure is created,
as a result of case pressure now being greater than that inside the
capsule, causing it to collapse and so make the pointer indicate a
climb.
Apart from the changes of static pressure with changes of altitude;
air temperature, density and viscosity changes are other very
important variables which must be taken into account, particularly as
the instrument depends on rates of airflow. In addition, \he
volumetric capacities of cases and capsules must also be considered in
order to obtain the constant differential pressures necessary for the
indication of specific rates of climb and descent. Metering units are
designed to compensate for the effects of the variables over the
ranges normally encountered, and so from the theoretical point of
view a vertical speed indicator is a somewhat sophisticated
instrument.
The construction of a typical indicator is shown in Fig. 2.30. It
consists of a cast aluminium-alloy body which forms the support for
58
all the principal components with the exception of the metering unit,
which is secured to the rear of the instrument case. Displacements of
the capsule in response to differential pressure changes are
transmitted to the pointer via a balanced link and rocking-shaft
assembly, and a quadrant and pinion. The flange of the metering unit
connects with the static pressure connection of the indicator case. and
with the capsule via the capillary tube.
In order to achieve the correct relationship between the capsule's
pressure/deflection characteristics and the pointer position at all points
of the scale, forces are exerted on the capsule by two pre-adjusted
calibration springs. The upper spring and its adjusting screws control
the rate of descent calibration, while the lower spring and screws
control that of rate of climb.
An adjustment device is provided at the front of the indicator for
settiog the pointer to zero, and when operated it moves the capsule
assembly up or d~wn to position the pointer via the magnifying lever
and- gearing system. The range of adjustment around zero depends on
tne scale range of any one' type of indicator, but ± 200 and ± 400
ft/min are typical values.
Instantaneous vertical speed indicators (IVS!) These indicators
consist of the same basic elements as conventional VSis, 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 basic principle is illustrated in
Fig. 2.31.
The accelerometer comprises a small cylinder, or dashpot,
containing a piston held in balance by a spring and by its own mass.
The cylinder is connected in a capillary tube leading to the capsule,
and is thus open directly to the static pressure source. When a change
in vertical speed is initiated, the piston is immediately displayed
under the influence of a vertical acceleration force. and this creates
an immediate pressure change inside the capsule. For example, at
initiation of a descent. the piston moves up and thereby decreases the
volume of chamber 'A· to produce an immediate increase of pressure
inside the capsule. The capsule displacement in turn produces
instantaneous deflection of the indicator pointer over the descent
portion of the scaie. At initiation of an ascent, the converse of the
foregoing responses would apply. The accelerometer response decays
in each case after a few seconds, but by this time the change in
actual static pressure becomes effective, rn that a pressure differential
is produced by the metering unit in the conventional manner. The
purpose of the restrictor in the bypass line is to prevent any loss of
pressure change effects created by displacements at the acceleration
pump.
59
Figure 2.31 Instantaneous
venical speed indicalor.
BYPASS
RESTRICTION
CAPSULE
VERTICAL
ACCELERATION
PUMP
-SECTOR GEAR
Descent
t
i
Ascent
"'------------
/
___ ___
.._
.., ._
...._,ps
Air temperature sensing
Air temperature is another of the basic parameters used to establish
data vital to the performance monitoring of aircraft and engines, e.g.
true airspeed measurement, temperature control, thrust settings,
fuel/air ratio settings, etc. of turbine engines, and it is therefore
necessary to provide a means of in-flight measurement.
The temperature which would overall be the most ideal is that of
air under pure static conditions at the flight levels compatible with the
operating range of any particular type of aircraft concerned. The
measurement of static air temperature (SAT) by direct means is.
however, not possible for all types of aircraft for the reason that
measurements can be affected by the adiabatic compression of air
resulting from increases in airspeed. At speeds below 0.2 Mach, the
air temperature is very close to static conditions, but at higher
speeds, and as a result of changes in boundary· layer behaviour and
the effects of friction, the temperature is raised to a value appreciably
higher than SAT; this increase is referred to as ram rise.
Figure 2.32 Direct-reading air
temperature indicator.
LIGHT SHIELD
For use in aircraft capable of high Mach speeds, and for efficient
control and management of the overall performance of their engines,
it is customary to sense and measure the maximum temperature rise
possible. This parameter is referred to as total air temperature (TAT)
and is derived when the air is brought to rest (or nearly so) without
further addition or removal of heat. If the corresponding SAT value
is to be determined and indicated, it is necessary to calculate the
value of tam rise and then subtract it from that of TAT. Details of
the method by which this is normally accomplished will be given in
Chapter 7.
Various types of sensor may be adopted for the sensing of air
temperature. The simplest type, and one which is used in some types
of small low-speed aircraft for the indication of SAT, is a directreading indicator which operates on the principle of expansion and
contraction of a bi-metallic element when subjected to temperature
changes. The element is arranged in the form of a helix anchored at
one end of a metal sheath or probe; the opposite, or free end of the
helix, is attached to the spindle of a pointer. As the helix expands or
contracts, it winds or unwinds causing the pointer to rotate over the
scale of a dial fixed to the probe. The thermometer is secured
through a fixing hole in the side window of a cockpit, or in the
wrap-around portion of a windscreen, so that the probe protrudes into
the airsiream. An example of this thermometer and its installation in
one type of helicopter is shown in Fig. 2.32.
The measurement of TAT requires a more sophisticated measuring
technique, and because the proportion of ram rise due to adiabatic
61
Figure 2.33 Total air
temperature probe (I).
AIRCRAFT FUSELAGE SKIN
1
SENSING ELEMENT,.,-
1' Iii
r1
CENTRE BODY -·
__.. 5-POLE CONNECTOR
compression is dependent on the ability of a sensor to sense and
recover this rise, then a TAT sensor must itself be of a mort:
sophisticated design. In this context, the sensitivity of a sensor is
normally expressed as a percentage termed the recovery factor. Thus,
a sensor having a factor of 0.80 would measure SAT plus 80 per cent
of the ram rise.
TAT sensors are of the probe type, and one example is shown in
Fig. 2 .33. The probe is in the form of a small" strut and air intake
made of nickel-plated berylfa:m copper which provides good thermal
conductivity and strength. It is secured at a pre-determined location in
the front fuselage section of an aircraft (typically at the side, or upper
surface of the nose) and outside of any boundary layer which may
exist. In flight, the air flows tluough the probe in the manner
indicated; separation of any water particles from the air is effected by
the airflow being caused to turn through a right angle before passing
round the sensing element. The bled holes in the intake casing
62
Figure 2.34 Total air
temperature probe (2).
AIR fl.Oll----1
AIR EXIT TO S10Ef'ORTS
11---} """' '"' "' ""
EJECTOR FITTING
ELECTRICAL COlll4!:CTOR~~-,_~()--'
ENGINE BLEED AIR IN
permit boundary layer air to be drawn off under the influence of the
higher pressure that is created within the intake and casing of the
probe.
A pure platinum wire resistance-type sensing element is used and is
hermetically sealed within two concentric platinum tubes. The
element is wound on the inner tube, and since they are both of the
same metal, a close match of thermal expansion and minimizing of
thermal strain is ensured. The probe has an almost negligible timelag, 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, and is of the
self-compensating type in that as the temperature rises so does the
element resistance rise, thereby reducing the heater current. The
heater dissipates a nominal 260 W under in-flight icing conditions,
and can have an effect on indicated air temperature readings. The
errors involved, however, are small, some typical values obtained
experimentally bei".lg 0.9°C at 0.1 Mach decreasing to O. l5°C at
Mach 1.0.
A secord type of TAT probe is shown in Fig. 2.34. The principal
differences between it and the one just described relate to the air
intake configuration and the manner in which airflow is directed
through it and the probe casing. 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
63
that the heating elements have a negligible effect on the
temperature/resistance characteristics of the sensing element.
In some cases, an auxiliary sensing element is provided in a probe.
The purpose of this element is to transmit a signal to other systems
requiring air temperature information. An example of this would be
the airspeed measuring circuit of an ADC for the computing of true
airspeed (see Chapter 7).
Air temperature indicators As in the case of other instruments, TAT
indicators can, as a result of the instrumentation arrangements
adopted for each particular type of aircraft, vary in the manner in
which they display the relevant data. Some of the variations are
illustrated in Fig. 2.35.
The circuit of a probe and a basic conventional pointer and scale
type of indicator is shown in Fig. 2.36. The system is supplied with
115 V ac which is then stepped down and rectified by a power supply
module within the indicator. The probe element forms one part of a
resistance bridge circuit, and as the element's resistance changes with
temperature, the bridge is unbalanced, causing current to flow
through the moving coil of the indicator.
Figure 2.37 illustrates the circuit arrangement of a servo-operated
indicator employing a mechanical drum-type digital counter display.
The generation of the appropriate temperature signals is also
accomplished by means of a de bridge circuit, but in this case
unbalanced conditions are monitored by a solid-state chopper circuit
which produces an error signal to drive an ac servomotor via an
operational amplifier. The motor then drives the counter drums, and
at the same time positions the wiper contact of a potentiometer to
start rebalancing of the circuit, until at some constant temperature
condition the circuit is 'nulled'.
In order to indicate whether temperatures are either positive or
negative, the rebalancing/feedback system also activates a 'sign
changer' and an indicator drum, and a switch which reverses the
polarity of the bridge circuit when the temperature indications pass
through zero.
Detection of failure of the 26 V ac power to the indicator, and
sensing of an excessive null voltage in the rebalancing/feedback
system, is provided by a failure monitor circuit module. This controls
an 'OFF' flag which under normal conditions is held out of view by
an energized solenoid.
The internal arrangement of an LCD (see page 15) type of
indicator is schematically shown in Fig. 2.38. The temperature data
signals are transmitted from a digital type of ADC (see Chapter 7)
via a data bus and receiver to a microcomputer. The power supply to
the computer is connected via supply, low voltage and failure monitor
modules. In addition to TAT, the indicator can also display SAT and
64
Figure 2.35 TAT displays.
(a) Conventional pointer and
scale; (b) servo/digital counter;
(c) LCD; (d) electronic (CRT).
Fwcr:a«
l"J$1l
SELECTOII
ecnoo
(c}
(b)
Cyan
(d)
65
115V ac
Figure 2.36 TAT incicator
system.
,_.. ..
-· . ----- ....
L.,_
I,
I
_j
·-
I'
I
I
I
__ .i
~
'
:
i
lndic~----·
I
_i __ _
Probe
Heater
supply
Figure 2.37 Servo-operated
TAT indicator.
·---·---·---·--~
Rebalancing
potentiometer
SIGN
CHANGER
TRIGGER
Sign indicator
-------
4
:_ _ d ~ - - ~ - ~
I
Input from
TAT probe
I
I
DC BRIDGE
I
CHOPPER
J
A-M-P-'L-IFI_E_R:::::::_--·:::::::__,·
POWER
SUPPLY
DC
MOOULE
I
26V ac
supply-+~~==:::.-.-::.-::_-::_-::_-::_:::_-.-_----.::::::-.:::::_-::_-=_-_-__·:._-_-_-_-_-.====-=-=-~- ___ .---- .I
---
Mechanical drive
11•
L __
I
I
I " ' POWER
SUPPLY
.. 1
1r·-H
FAILURE
MONITOR
i..--t
-----·--·---·
4 MHz
OSCILLATOR
ARING 429
RECEIVER
LOW-VOLTAGE
MONITOR
Microc~mputer
Clock
Data input
Function
Control select
Test
Reset
Monitor
----------------
MEMORY
/1----J FUNCTION
'-.......---1 SELECT
LCD
DRIVER
..:I.::_
II•
-----· -----.
4
·-----·---F-un-ction :elect
switch
Figure 2.39 Function mode
selection sequence.
TAT
,-,.c
-,-,,--, ,-,
- -·-
.!.
I
Function selector
push-button
=,,-, ,-,.c
'-f '_,l :
lCT
-"-'•'-'
SAT
TAS
0
TAS, each of which can be selected in sequence by a push-button
function select switch. When power is first applied, the indicator
displays TAT, as in Fig. 2.39; to select SAT the switch is pushed in,
and then pushed in again to display TAS. Pushing the switch in for a
third time returns the display to TAT. A test input facility is
provided, and when activated it causes the display to alternate
between all seven segments (of each of the three digits), 'ON' for
two seconds, and blank for one second.
As noted earlier, indications of SAT can be derived by subtracting
the ram rise from the measured values of TAT. Since this is normally
done b:},' also supplying TAT signals to the speed-measurir.g module
of an ADC, the operating principles of SAT indication will be
covered in Chapter 7.
Details of the coloured display shown at (d) of Fig. 2.35 will be
given in Chapter 16.
Air data alerting and
warning systems
68
In connection with the in-flight operation of aircraft, it is necessary to
impose limitations in respect of certain operating parameters
compatible with the airworthiness standards to which each type of
aircraft is certificated. It is also necessary for systems to be provided
which will, both visually and aurally, alert and warn a flight crew
whenever the imposed operational limitations are being exceeded.
The number of parameters to be monitored in this way varies in
relation to the type of aircraft and the number of systems required for
its operation overall. As far as air data measuring systems are
concerned, the principal parameters are airspeed and altitude, so let
us now consider the operating principles of associated alerting and
warning systems typical of those used in some of the larger types of
public transport aircraft.
Figure 2.40 Mach warning
system.
ii - - - - - - - - - - - - - - - - - - - ,
2
PITOT PRESSURE
==::::===-;:::/
aJ
'CLACKER'
TEST SWITC~
L _ ___________ - - SW1TCH UNIT
=
Mach warning system
This system provides an aural warning when an aircraft's speed
reaches the maximum operating value in terms of Mach number, i.e.
Mmo (a typical value is 0.84M). The warning is in addition to any
limiting speed reference pointers or 'bugs' that are provided in
Mach/airspeed indicators (see Fig. 2.20 again).
The system consists of a switch unit which, as can be seen from
Fig. 2.40, comprises airspeed and altitude sensing units connected to
an aircraft's pitot probe and static vent system in a manner similar to
that of a Machmeter. It will also be noted that in lieu of a pointer
actuating system, the sensing units actuate the contacts of a switch
which is connected to a 28 V de power source ..
At speeds below the limiting value, the switch contacts remain
closed and the de passing through them energizes a control relay. The
contacts of this relay interrupt the ground connection to an aural
warning device generally referred to as a 'clacker' because of the
sound it emits when in operation. When the limiting Mach speed at
any given altitude is reached, the airspeed sensing unit causes the
switch contacts to open, thereby de-energizing the control relay so
that its contacts now complete a connection from the 'clacker' to
ground. Since the 'clacker' i:. directly supplied with de, then it will
be activated to provide the apprupriate warning, which is emitted at a
specific frequency (typically 7 Hz).
A toggle switch that is spring-loaded to 'OFF' is provided for the
69
Figure 2. 4/
Combined
indicator and switch unit
system.
28V de
l
CAPT INDICATOR
Speed
signals
From
CADC
{
Limiting
•
parameter
AURAL
WARNING UNIT
F/0 INDICATOR
purpose of functional checking the system. When placed in the
'TEST' position, it allows de to flow to the ground side of the switch
unit control relay, thereby providing a bias sufficient to de-energize
the relay and so cause the 'clacker' to be activated.
In some aircraft systems, Mach/airspeed indicators W'ith 'built-in'
warning switch units may be used and so arranged that they operate
two independent 'clackers'. In the exam:,le shown in Fig. 2.41, the
indicator in the captain's group of flight instruments is servo-operated
by signals from an ADC. The other indicator, which is in the first
officer's group of flight instruments, is also of the servo-operated
70
type, but contains a switch unit that is connected directly to the pitot
probe and static vent system. The 'ciacker' units associated with the
indicators are respectively designated as 'aural warning I' and 'aural
warning 2'.
The captain's indicator contains an overspeed circuit module that is
supplied by the ADC with prevailing speed data and also the limiting
V,,w and Mmo values appropriate to the type of aircraft. The circuit
module is, in turn, connected to a solid-state switch (S 1) that is
powered 'open' at speeds below Vm and Mmo· If, however, these
speeds are reached, then S 1 is relaxed to provide a ground connection
for the de supply to 'aural warning I' clacker unit which thus gives
the necessary warning.
The contacts of the switch unit in the first officer's indicator are
connected to a relay, and since these contacts remain closed at speeds
below maximum values, the relay is de-energized. When the
maximum speed is reached, the relay coil circuit is interrupted and its
contacts then change over to provide a ground connection for the de
supply which activates 'aural warning 2' clacker unit.
Test switches are provided for checking the operation of each
clacker by simulation of overspeed conditions. When switch 1 is
operated de is applied to the overs peed circuit module in the captain's
inr!icator, and causes the switch S 1 to relax. The operation of switch
2 applies de to the relay coil such that it is shorted out against the
standing supply from the closed airspeed switch; the relay 'is therefore
de-energized to provide a ground connection for 'aural warning 2'
clacker unit.
The indicators themselves provide visual indications of overspeed
and these are discernible when the airspeed pointers become
positioned coincident with pre-set maximum limit pointers (see Figs
2.20 and 2.21).
0
Altitud*: alerting system
This system is designed to alert a flight crew, by aural and visual
means, of an aircraft's approach to, or deviation from, a pre-selected
altitude. The components of a typical system are shown in Fig.
2.42(a).
An aircraft's pressure altitude is provided as a signal input to the
alert controller unit from an altimeter via a coarse/fine synchro
system. The selected altitude is set by means of a knob on the
controller, and is indicated by a digital counter which is geared to the
rotors of control and resolver synchros, so that they produce a
corresponding signal. The signal is compared with the pressure
altitude signal. and the resulting difference is supplied to level
detection circuits within the controller. At predetermined values of
rotor voltages of both synchros, two signals are produced and
n
Figure 2.42 Altitude alerting
system.
Pressure---altitude
Audio
AL;,~~oe
I2 I7 I5 IO IO I
CONTROLLER UNIT
Annunciator light
(a)
AUOIC) WARNING
ON FOR TWO SECONDS. AND
WARNING LIGHTS ILLUMINATED
H11900 F T ) - - - -
H2·{300FTJ - - - - - - - - - - - - - - -
SELECTED
ALTITUDE
(bl
supplied as inputs to a logic circuit and timing network which
controls the aural and visual alerting devices.
The sequence of alerting is shown at (b) of Fig. 2.42. As an
aircraft descends or climbs to the preselected altitude the difference
signal is reduced, and the logic circuit so processes the input signals
that, at a pre-set outer limit H 1 (typically 900 ft) above or below
preselected altitude, one signal activates the aural alerting device
which remains on for two seconds; the annunciator light is also
illuminated. The light remains on until at a further pre-set inner limit
H2 (typically 300 ft) above or below preselected altitude, the second
J
signal causes the annunciator light to be extinguished. As an aircraft
approaches the preselected altitude, the synchro system approaches
the 'null' position, and no further alerting takes place. If an aircraft
should subsequently depart from the preselected altitude, the
controller logic circuit changes the alerting sequence such that the
indications correspond to those given during the approach through
outer limit Hi, i.e. aural alert on for two seconds, and annunciator
light illuminated.
Angle of attack
sensing
The angle of attack (AoA), or alpha (a) angle, is the angle between
the chord line of the wing of an aircraft and the direction of the
relative airflow, and is a major factor in determining the magnitude
of lift generated by a wing. Lift increases as a increases up to some
critical value at which it begins to decrease due to separation of the
slow-moving air (the boundary layer) from the upper surface of the
wing, which, in turn, results in separation and turbulence of the main
airflow. The wing, therefore, assumes a stalled condition, and since it
occurs at a particular angle rather than a particular speed, the critical
AoA is also referred to as the stalling angle. The angle relates to the
design of aerofoil section adopted for the wings of any one particular
type of aircraft, and so, of course, its value varies accordingly;
typically it is between 12 ° and 18 °.
The manner in which an aircraft responds as it approaches and
reaches a stalled condition depends on many other factors, such as
wing configuration, i.e. high, low, swept-back, and also on whether
the horizontal stabilizer is in the 'T' -tail configuration. Other factors
relate to the prevailing speed of an aircraft, which largely depends on
engine power settings, flap angles, bank angles and rates of change
of pitch. The appropriate responses are pre-determined for each type
of aircraft in order to derive specificaliy relevant procedures for
recovering from what is, after all, an undesirable situation.
An aircraft will, in its own characteristic manner, provide warning
of a stalled condition, e.g. by buffeting, gentle or severe pitch-down
attitude change, and/or 'wing drop', and although recoverable, in a
situation such as an approach when an aircraft is running out of
airspace beneath itself, these inherent warnings could come too late!
It is, therefore, necessary to provide a means whereby a can be
sensed directly, and at some value just below that at which a stalled
condition can occur it can provide an early warning of its onset.
Stall warning systems
The simplest form of system, and one which is adopted in several
types of small aircraft, consists of a hinged-vane-type senso.r mounted
..<~::;;::::;::;':''::::':'.:•.'.::--=~-·
·-- ~....
73
Figure 2.43 Alpha sensor.
in the leading edge of a wing so that the vane protrudes into the
airstream. In normal level flight conditions, the airstream maintains
the vane in a parallel position. If the aircraft's attitude changes such
that a increases, then, by definition, the airflow will meet the leading
edge at an increasing angle, and so cause the vane to be deflected.
When a reaches that at which the warning unit has been preset, the
vane activates a switch to complete a circuit to an aural warning unit
in the cockpit.
In larger tyges of aircraft, stall warning systems are designed to
perfon_n <!Jl!QI~ai;:J!Y,.e funchon .• ~m.ihai:~ii£e..eTt__~er of the -.-sticksfiaker' or 'stick push or nudger' type; for some aircraft
configurations 'tney ~ ...conii)InatiOU:-Figure 2.43 iflus.trates the type of sensor normally used for these
systems. It consists of a precision counter-balanced aerod namic .ne
which positions t e rotonr· sync ro. The vane is protected against
ice formation by an internal heater element. The complete unit is
accurately aligned by means of index pins at the side of the front
fuselage section of an aircraft. Since the pitch attitude of an aircraft
is also changed by the extension of its flaps, the sensor synchro is
also interconnected with a synchro within the transmitter of the flap
position indicating system, in order to modify the a signal output as a
function of flap position.
.,.
Stick-shaking is accomplished by a motor which is secured to a
control column and drives a weighted ring that is deliberately
unbalanced to set up vibrations of the column, to simulate the natural
buffeting associated with a stalled condition.
Sensor signals, and signals for the testing of a system, are
processed through a circuit moduie unit located on a flight deck
panel. Control switches for normal operation and for testing are also
provided in this unit. Sensing relays and shock strut microswitches on
the nose landing gear are included in the circuit of a system to permit
operational change-over from ground to air.
The circuit of a typical system is shown in basic form in Fig. 2.44.
When the aircraft is on the ground and electrical power is on, the
74
Figure 2.44 Stick-shaker
system.
115V
ac
28V de
r·
LG
microsw1tch
AIRLi.
Bias of!'
II
Demodulator
,, signals modified
by flap position
Flap pos,hon
1ransm11!er
---
Heater suµpli-y- - ' - - ~
@:::::s:i
Synchro supply
Sensor
Stick-shaker
contacts of the landing gear microswitches complete a de circuit to a
sensing relay K 1 which, on being energized, supplies an ac voltage
(in this case 11.8.V) to the circuit module amplifier. The output is
then supplied to a demodulator whose circuit is designed to 'bias off
the ac voltage from the contacts of K 1, so that the solid-state switch
SS 1 remains open to isolate the stick-shaker motor from its de supply.
The vane heater element circuit is also isolated from its ac supply by
the opening of the second set of contacts of K 1• The sensor synchro
is supplied directly from the ac power source.
During take-off, and when the nose gear 'lifts off, the
microswitches operate to de-energize relay K1, and with the system
control switch at 'NORMAL', the system is fully activated. The only
signal now supplied to the amplifier and demodulator is the modified
a signal.
In normal flight, the signal produced and supplied as input to the
amplifier is less than a nominal value of 20 mV, and in phase with
the ac vo!tage supplied as a reference to the demodulator. ff the
aircraft's attitude should approach that of a stalled condition, the a
signal will exceed_ 20 mV and become out-of-phase. The demodulator
then produces a resultant voltage which triggers the switch SS 1 to
connect a 28 V de supply direct to the stick-shaker motor, which.then
starts vibrating the conirol column.
75
A confidence check on system operation may be carried out by
placing the circuit module control switch in the 'TEST' position. This
energizes a relay which switches the sensor signal to the motor of an
indicator, the dial of which will be rotated by the motor if there is
circuit continuity. Since the switch isolates the sensor circuit from the
amplifier, the reference voltage to the demodulator triggers the switch
SS 1 to operate the stick-shaker motor. The control switch also has a
'HEATER OFF' position which isolates the vane heater circuit from
its power supply, thus enabling the vane to be manipulated manually
without inflicting burns.
In most cases, two systems are installed in an aircraft, so that a
sensor is located on each side of the front fuselage section, and a
stick-shaker motor on each pilot's control column.
In certain types of aircraft the sensor signals are transmitted to an
air data computer, which then supplies an output, corresponding to
actual a angle, to a comparator circuit within an electronic module of
the stall warning system. The comparator is also supplied with signals
from a central processor unit (also within the module) which
processes a programme to determine maximum a angles based on the
relationship between flap position and three positions of the leading
edge slats. The positions are: retracted, partially extended and fully
extended, and so signals corresponding to three different computed
angles are processed for comparison with an actual a angle signal. If
the latter is higher than a computed maximum, the circuit to the
stick-shaker motor is completed.
,vS;ick-pushers
In some types of ai:-craft, particularly those with rear-mounted
engines and a 'T' -tail configuration. it is possible for what is termed
a 'deep' or 'super' stall situation to develop. When such aircraft first
get into a stalled condition then, as in all cases, the air flowing from
the wings is of a turbulent nature. and if the a angle is such that the
engines are subjected to this airflow. loss of power will occur as a
result of surging and possible 'flame-out'. If. then, the stall develops
still further, the horizontal stabilizer will also be subjected to the
turbulent airflow with a resultant loss of pitch control. The aircraft
then sinks rapidly in the deep Slalled attitude, from which recovery is
difficult, if not impossible. lhis was a lesson that was learned, with
tragic results, during the flight testing of two of the earliest types of
commercial aircraft configured as mentioned, namely, the BAC 1-11
and HS 'Trident'.
In order to prevent the development of a deep stall situation,
1
warning systems are installed which, in acfciit'rn11 to stick-shaking,
utilize the a sensor signals to cause a forward push on the control
columns and downward deflection of the elevators. The manner in
76
which this is accomplished varies; in some aircraft, the signals are
transmitted to a linear actuator which is mechanically connected to
the feel and centring unit of the elevator control system. In aircraft
having computerized flight control systems, a sensor signals are
transmitted to the elevator control channel of the flight control
computer. Whenever stick-push is activated, the elevator control
channels of automatic flight control systems are automatically
disengaged via an interlock system.
Indicators
There is no standard requirement for angle of attack indicators to be
installed in aircraft, with the result that the adoption of any one
available type is left as an option on the part of an aircraft
manufacturer and/or operator. When selected for installation,
however, they must not be used as the only means of providing stall
warning, but as a supplement to an appropriate type of stick-shake
and stick-push system.
Indicators are connected to the alpha sensors of a stall warning
system, and display the relevant data in a variety of ways, depending
on their design. In some cases a conventional pointer and scale type
of display is used, while in aircraft having electronic flight instrument
display systems, the data can be programmed into computers such
that it is displayed against a ,vertical scale, usually located adjacent to
that indicating vertical speed, on the attitude director indicator.
Another type of indicator currently in use has a pointer which is
referenced against horizontal yellow, green and red bands; a dividing
line between the yellow and green bands signifies the angle at which
the stick-shaker operates.
77
3 Direct-reading compasses
Compasses of this type were the first of the many airborne flight and
navigational aids ever to be introduced in aircraft, their primary
function being to show the direction in which an aircraft is heading
with respect to the earth's magnetic meridian.
As far as present-day aircraft are concerned, the use of directreading compasses as a primary directional reference source is
confined to small types of air9raft whose design and operating
requirements are at a fairly basic level. In the more sophisticated
types of aircraft, directional references are derived from flight
instrument systems and navigational aids based on advanced
technology, and although airworthiness requirements still necessitate
the installation of direct-reading compasses, they are relegated to a
secondary role.
The operating principle of a direct-reading compass is based on
established fundamentals of magnetism, and on the reaction between
the magnetic field of a suitably suspended magnetic element, and that
of terrestrial magnetism.
Terrestrial magnetism
78
The surface of the earth is surrounded by a weak magnetic field
which culminates in two internal magnetic poles situated near the
North and South true or geographic poles. That this is so is obvious
from the fact that a magne.t freely suspended at various parts of the
earth's surface will be found to settle in a definite direction, which
varies with locality. A plane passing through the magnet and the
centre of the earth would trace out on the earth's surface an
imaginary line called the magnetic meridian as shown in Fig. 3.1.
It would thus appear that the earth's magnetic field is similar to
that which would be expected at the surface if a short but strongly
magnetized bar magnet were located at the centre. This partly
explains the fact that the magnetic poles are relatively large areas,
due to the spreading out of the lines of force, and it also gives a
reason for the direction of the field being horizontal in the vicinity of
the equator. The origin of the earth's field is still not precisely
known, but, for purposes of explanation, -the supposition of a bar
magnet at its centre is useful in visualizing the general form of the
magnetic field as it is known to be.
The field differs from that of an ordinary magnet in several
Figure 3. I Terrestrial
magnetism. i_ines AA, BB and
CC are isoclinals.
ANGLE Of 0d'
(INCREASING FROM EOOATOfl)
TRUE
NORTH
VARIATION
(EASTERLY}
VARIATION
{WESTERLY)
respects. Its points of maximum intensity, or strength,. are not at the
magnetic poles (theoretically they should be) but occur four other
positi'ons, two near each pole, known as magnetic foci. Moreover, the
poles themselves are continually changing their positions, and at any
point on the earth's surface the field is not symme:rical and is subject
to changes both periodic and irregular.
a,
Magnet'ic variation
As meridians and parallels are constructed with reference to the true
or geographic North and South pcles, so can magnetic meridians be
constructed with reference to the magnetic poles. If a map were
prepared to show both true and magnetic meridians, it would be
observed that these intersect each other at angles varying from 0° to
180° at different parts of the earth, diverging from each other
sometimes in one direction and sometimes in the other. 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, the variation is said to be westerly, while an
inclination to the right produces easterly variation. It varies in amount
from 0° along those lines where the magnetic and true meridians run
together, to 180° in places between the true and magnetic poles. At
some places on the earth, where the ferrous nature of the rock
disturbs the main magnetic field, local attraction exists and abnormal
variation occurs which may cause large changes in its value over very
short distances. While the variation differs all over the world, it does
not maintain a constant value in any one place, and the following
changes, themselves not constant, may be experienced: (i) Secular
change, which takes place over long periods due to the changing
positions of the magnetic poles relative to the true poles; (ii) Annual
change, which is a small seasonal fluctuation superimposed on the
secular change; (iii) Diurnal or daily change.
Information regarding variation and its changes are given on special
charts. 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
Magnetic dip
As stated earlier, a freely suspended magnet will settle in a definite
direction at any point on the earth's surface and will lie parallel to
the magnetic meridian at that point. It will not, however, lie parallel
to the earth's surface at all points for the reason that the lines of
force themselves are not horizontal, as may be seen from Fig. 3.2.
These lines emerge vertically from the North magnetic pole, bend
over and descend vertically into the South magnetic pole; it is only at
what is known as the magnetic equator that they pass horizontally
along the earth's surface. If, therefore, a suspended magnet is carried
along a meridian from north to south, it will be on end, red end
down, at the start, horizontal near the equator, and finish up again on
end but with the blue end down.
The angle the lines of force make with the earth's surface at any
Figure 3.2 Relationship
between magnetic components
and dip.
a - c = Vertical component Z
c - b = Horizontal component H
a - b = total force T
Given angle of dip IJ and H,
z
-
H
= tan IJ and Z ·= H tan IJ
H
,:=cos8andT=Hcos/J
r2 = H2
80
+
z2
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 angle of dip at all places undergoes changes similar to those
described for variation and is also shown on charts of the world.
Places on these charts having the same dip angle are joined by lines
known .as isoclinals, while those at which the angle is zero are joined
by a line known as the aclinic line or magnetic equator, of which
mention has already been made.
Earth's total force
When a magnet freely suspended in the earth's field comes to rest, it
does so under the influence of the total force of the field. This total
force is resolved into its horizontal and vertical components, termed
H and Z respectively. The relationship between these components and
dip is shown in Fig. 3.2.
As in the case of variation and dip, charts of the world are
published showing the values of the components for all places on the
earth's surface, together with the mean annual change. Lines of equal
H and Z forces are referred to as isodynamic lines.
The earth's magnetic force may be stated either as a relative value
or an absolute value. If stated as a relative value, and in the case of
compasses this is the case, it is given relative to the H force at
Greenwich.
Compass
construction
Direct-reading compasses have the following common principal
features: a magnet system housed in a bowl; liquid damping; liquid
expansion compensation; and deviation compensation. The majority of
compasses currently in use are of the card type, and the construction
of two examples is illustrated in Fig. 3.3. '
The magnet system of the example shown at (a) comprises an
annular cobalt-steel magnet to which is attached a light-alloy card.
graduated in increments of 10°, and referenced against a lubber line
fixed to the interior of the bowl. The system is pendulously
suspended by an iridium-tipped pivot resting in a sapphire cup
supported in a holder or stem. The pivot point is above the centre of
gravity of the magnet system which is balanced in such a way as to
minimize the effects of angle of dip over as wide a range of latitudes
North and South as possible.
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, i.e. to ensure that after
the magnet system has been deflected, it returns to equilibrium
directly without oscillating or overshooting. The fluid also provides
81
·e· AND ·c- CORRECTOR
Figure 3.3 Typical card
compasses.
INDICATORS
HORIZONTAL ('B· AND
CORRECTORS
MOUNTING PLATE
I
·c)
""''° ""'
BELLOWS
STEM ANO BRACKET
ASSEMBLY
(a)
MAGNET
SYSTEM
BOWL
BOWL
LIQUID
EXPANSION
CAPSULES
(b)
82
ELECTRICAL
CONNECTOi1
MAGNET (1 of 2)
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.
Compensation of the effects of deviation due to longitudinal and
lateral components of aircraft magnetism (see page 87) is provided
by permanent magnet coefficient 'B' and 'C' corrector assemblies
secured to the compass mounting plate.
The compass shown at (b) of Fig. 3.3 is designed for direct
mounting on a panel. Its magnet system is similar to the one
described earlier except that needle-type magnets are used. The bowl
is in the form of a brass case which is sealed by a front bezel plate.
Changes in liquid volume are compensated by a capsule type of
expansion device. A permanent-magnet deviation compensator is
located at the underside of the bowl, the coefficient 'B' and 'C'
spindles being accessible from the front of the compass. A small
lamp is provided for illuminating the card of the magnet system.
Compass location
The location of a compass in any one type of aircraft is of
importance, and is pre-determined during the design stage by taking
into account the effects which mechanical and electrical equipment in
cockpit or flight deck areas may have on indications. In this
connection it is usual to apply the compass safe distance rule which,
precisely defined, is 'the minimum distance at which equipment may
be safely positioned from a compass without specified design values
of maximum deviation being exceeded under all operating
conditions'. The distance is measured from the centre of a compass
magnet system to the nearest point on the surface of equipment.
Values are quoted by manufacturers as part of the operating data
appropriate to their equipment.
Errors in indication
The pendulous suspension of a magnet system, although satisfactory
from the point of view of counteracting dip, is unfortunately a
potential source of errors under in-flight operating conditions in
which certain force components are imposed on the system. There are
two main errors that result from such components, namely
acceleration error and turning error.
Acceleration error
This may be broadly defined as the error, caused by the effect of the
earth's field component Z, in the directional properties of the magnet
83
Figure 3.4 Acceleration
errors. (a) Acceleration on
nonherly heading in northern
hemisphere: (b) deceleration on
nonherly heading in nonhem
hemisphere: (c) acceleration on
easterly heading in northern
hemisphere; (d) deceleration on
easterly heading in nonhern
hemisphere.
-N--r:::::::::=~
p
----.:--S
'
R---\ C.G.
I
(b)
N
---
EASTl:RLY
W!:STERLY
DEVIATION
DEV!ATION
;~
~
N
-\
I
p
R
(c)
-e
-e
R
C.G.
I
s
(d)
s
system when its centre of gravity is displaced from its normal
position, such errors being governed by the heading on which
acceleration or deceleration takes place.
When accelerating or decelerating on any fixed heading, a force is
applied to the magnet system at the point of suspension P, this being
its only connection. The reaction to this force will be equal and
opposite and must act through the centre of gravity, which is below
and offset from P due to the slight dip of the magnet system. The
two forces constitute a couple which, dependent on the heading being
flown, causes the magnet system merely to change its dip offset
angle, or to rotate in azimuth.
Consider now an acceleration on a northerly heading in the
northern hemisphere. The forces brought into play will be as shown
in Fig. 3.4(a). Since both the point P and centre of gravity are in the
plane of the magnetic meridian, the reaction R causes the 'N' end of
the magnet system to go down, thus increasing the dip offset angle
without any azimuth rotation. Conversely, when decelerating, the
reaction R tilts the 'S' end of the magnet system as shown at (b).
In either the northern or southern hemispheres, acceleration or
deceleration on headings other than the N-S meridian will produce
azimuth rotation of the magnet system and consequent errors.
When an acceleration occurs on an easterly heading in the northern
hemisphere, as at (c) of Fig. 3.4, a force will again act through point
P, and the reaction R through the centre of gravity. In this case,
however, they are acting away from each other and the couple
produced tends to rotate the magnet system in a clockwise direction,
thus indicating an apparent turn to the north, or what is termed
easterly deviation. The reverse effects occur during a deceleration,
producing an apparent turn to the south or westerly deviation.
Hence, in the northern hemisphere, acceleration causes easterly
deviation on easterly headings, and westerly deviation on westerly
headings, whilst deceleration has the reverse effect. In the southern
hemisphere the results will be reversed in each case.
As northerly or southerly hea<lir,gs are approached, the magnitude
of the apparent deviation decreases, the acceleration error varying as
the sine of the compass heading.
One further point may be mentioned in connection with these
errors, and that is the effect of aircraft attitude changes. If an aircraft
flying level is put into a climb at the sa:ne speed, the effect on its
compass magnet system will be the same as if the aircraft had
decelerated. If the change in attitude is also accompanied by a change
in speed, the apparent deviation may be q:iite considerable.
Turning errors
During a turn, the point P of a compass magnet system is carried
with the aircraft along the curved path of the turn. The system's
centre of gravity, being offset, is subjected to the centrifugal
acceleration force produced by the tum, causing the system to swing
outwards and to rotate so that apparent deviations, or turning errors,
will be observed. In addition, the magnet system tends to mabtain a
position parallel to the transverse plane of the aircraft, thus giving it
a lateral tilt the angle of which is governed by the aircraft's bank
angle. For a correctly banked tum, the tilt angle would be maintained
equal to the bank angle, because the resultant of centrifugal force and
gravity ·lies normal to the aircraft's transverse plane, and also to the
plane through the point P and centre of gravity of the magnet system.
In this case, centrifugal force itself would have no effect other than
to exert a pull on the centre of gravity and so decrease the offset dip
angle of the magnet system.
As soon as the system is tilted, however, and regardless of whether
or not the aircraft is correctly banked, the system is free to move
under the influence of the earth's component Z which will then have
a component in the lateral plane of the system, causing it to rotate
and further increase the turning error.
The extent and direction of the error is dependent upon the
aircraft's heading, the magnet system tilt angle, and the dip. In order
to form a clearer understanding of its effects, we may consider a few
examples of heading changes from the magnetic meridian, and in
both the northern and southern hemispheres.
Turning from a nonherly heading towards east or west
Figure 3.5(a) represents the magnet system of a compass in an
aircraft flying on a northerly heading in the northern hemisphere. Let
us assume that a change in heading to the eastward is required. As
85
Figure 3.5 Turning errors.
c,@r.,@• '1J ,@,
'.c)
(11)
t
(e
C.F.
s
N
SOUTHERN
c.F
(b)
N,
(g)
s
t
f
.... c.F~c.<1
(vc.G.C.G.
[E
p
E
l p.... E ~
I --,
''
C.F
(d)
t --
~
~
(f)
~:--
(hi
s
'
....... _
_,,./ \.._-;.F.C.~.J--_
soon as the turn commences, the centrifugal acceleration acts on the
centre of gravity causing the system to rotate in the same direction as
the turn, and since the system is tilted, the earth's component Z
exerts a pull on the N end causing further rotation of the system.
Now, the magnitude of system rotation is dependent on the rate at
which turning and banking of the aircraft is carried out, and resulting
from this three possible indications may be registered: (i) a turn of
the correct sense, but smaller than that actually carried out when th~
magnet system turns at a slower rate than the aircraft; (ii) no tum at
all when the system and aircraft are turning at the same rate; (iii) a
turn in the opposite sense when the system turns at a rate faster than
the aircraft. The same effects will occur if the heading changes from
N to W whilst flying in the northern hemisphere.
In the southern hemisphere (diagram (b)) the effects are somewhat
different. The south magnetic pole is now the doll)inant pole and so
the offset dip angle of the magnet system changes to displace the
centre of gravity to the north of point P. We may again consider the
case of an aircraft turning eastward from a northerly heading. Since
the centre of gravity is now north of point P, the centrifugal
acceleration acting on it causes the magnet system to rotate more
rapidly in the opposite direction to the turn, i.e. indicating a turn in
the correct sense but of greater magnitude than is actually carried
out.
Turning from a southerly heading towards east or west
If the turns are executed in the northern hemisphere (Fig. 3.5(c)) then
because the magnet system's centre of gravity is still south of point
86
P, the rotation of the system and the indications registered will be the
same as turning from a northerly heading in the northern hemisphere.
In turning from a southerly heading in the southern hemisphere
(Fig. 3.5(d)) the magnet system's centre of gravity is north of the
point P and produces the same effects as turning from a northerly
heading in the southern hemisphere.
In all the above cases, the greatest effect on compass indications
will be found when turns commence near to northerly or southerly
headings, being most pronounced when turning through north. For
this reason the term nonherly turning error is often used when
describing the effects of centrifugal acceleration on compass magnet
systems.
Turning through east or west
When turning from an easterly or westerly heading in either the
northern or southern hemispheres (diagrams (e)-(h)) no errors will
result because the centrifugal acceleration acts in a vertical plane
through the magnet system's centre of gravity and point P. The
centre of gravity is merely deflected to the north or south of point P.
thus increasing or decreasing the magnet system's pendulous
resistance to dip.
A point which may be noted in connection with turns from E or W
is that when the N or S end of the magnet system is tilted up, the
line of the system is nearer to the direction where the directive force
is zero, i.e. at right angles to the line of dip. Thus, if a compass has
not been accurately adjusted during a 'swing', any uncorrected
deviating force will become dominant and so cause indications of
apparent turns.
Aircraft magnetism
and Its effects on
compasses
Magnetism is unavoidably present in aircraft in varying amounts, anJ
can therefore also produce errors in the indications of compasses.
However, by analysis it is divided into two main types and also
resolved into components acting in definite directions, so that steps
can be taken to minimize the errors, or deviations as they are called,
resulting from such components.
The two types of magnetism can be further divided in the same
W'iJ.Y that magnetic materials are classified according to their ability to
be magnetized, namely hard-iron and soft-iron.
Hard-iron magnetism is of a peramenent nature and is caused, for
example, by the presence of magnetically 'hard' materials in an
aircraft's structure, in power plants and.other equipment, the earth's
field building itself into such materials during the many varied
manufacturing and assembly processes involved in the overall
construction of an aircraft.
87
Soft-iron magnetism is of a temporary nature and is caused by the
metallic materials of an aircraft which are magnetically 'soft'
becoming magnetized due to induction by the earth's field. The effect
of this type of magnetism. is dependent on an aircraft's heading,
attitude and its geographical position.
There is also a third type of magnetism, due to the sub-permanent
magnetism of what is called 'intermediate' iron, which can be
retained for varying periods. Such magnetism depends, not only on
heading, attitude and geographical position of an aircraft, but also on
the nature of its previous motion, vibrations, lightning strikes and
other external effects
The various magnetic components which cause deviations are
designated by letters, those for permanent hard-iron magnetism being
capitals, and those for soft-iron magnetism being small letters. The
resulting deviations are termed easterly when positive, and westerly
when negative.
Components of hard-iron magnetism
The total effect of this type of magnetism at a compass position may
be considered as having originated from equivalent bar magnets lying
longitudinally, laterally and vertically, as shown in Fig. 3.6. The
components are respectively denoted as P, Q and R, and are either
positive or negative depending on the locations of the blue poles of
the equivalent magnets. The strength of these components does not
vary with heading or change of latitude, but may do so with time due
to a weakening of aircraft magnetism. The deviations caused by each
of the components are set out in Table 3.1.
Figure 3.6 Components of
hard-iron magnetism.
~ u,
~I~
!!;j
-R
COMPASS
POS!TlON
---
-P
l+R
I
Table 3.1
Component
Axis
Component
polarity
Aircraft heading
Nonh
East
South
West
Deviations
p
Horizontal
+ve
-ve
0
0
max. +ve
max. -ve
0
0
max. -ve
max. +ve
Q
Lateral
+ve
-ve
max. +ve
max. -ve
0
0
max. -ve
max. +ve
0
0
Aircraft nose up
+ve
-ve
0
0
max. +ve
max. -ve
0
0
max. -ve
max. +ve
Aircraft nose down
R
Vertical
+ve
-ve
0
0
max. -ve
max. +ve
0
0
max. +ve
max. -ve
Aircraft banked to port
+ve
-ve
max. +ve
max. -ve
0
0
max. -ve
max. +ve
0
0
Aircraft banked to stb'd
+ve
-ve
max. -ve
max. +ve
0
0
max. +ve
max. ·-ve
0
0
Notes: I. +ve and -ve deviations are termed easterly and westerly respectively.
2. Component R effective only in the aircraft altitudes indicated.
Components of soft-iron magnetism
The effect of this type of magnetism may be considered as originating
from a piece of soft-iron in which magnetism has been induced by
the earth's field. This field, as we already know, has two components
designated H and Z, but in the analysis of soft-iron magnetism H is
resolved into two additional components X and Y. These, together
with component Z, are also related to the three principal axes of an
aircraft, namely X - longitudinal, Y - lateral and Z - vertical.
The polarities and strengths of components X and Y vary with
changes in aircraft heading relative to the fixed direction of the
earth's component H. Components X, Y and Z also change with
geographical location because this results in changes in the earth's
field strength and direction. A change in the polarity of component Z
will only occur with a change in magnetic hemisphere.
00
Table 3.2
Component
Axis
Component
Aircraft heading
polarity - - - - - - - - - - - - - - - - - - - - - - - - - - - - SW
West
South
NW
North
NE
East
SE
Deviations
aX
bY
Longitudinal
+ve
-ve
0
0
max. +ve
max. -ve
0
0
max. -ve
max. +ve
0
0
max. +ve
max. -vc
0
0
max. -ve
max. +ve
+ve
-ve
0
0
-ve
+ve
max. -ve
max. +ve
-ve
+ve
0
0
-ve
+ve
max. -ve
max. +ve
-ve
+ve
0
0
-ve
cl
+ve
-ve
dX
+ve
-ve
max. +ve
+ve
max. -ve
-ve
+ve
-ve
Same as - ve component aX.
Same as + ve component aX.
JZ
+ve
-ve
• Same as corresponding polarities of component Q.
gX
+ve
-ve
eY
hY
Lateral
Vertical
+vc
-ve
• Same as corresponding polarities of component P.
0
0
+ve
max. +ve
+ve
-ve
ma>:. -ve
-ve
+ve
• Same as corresponding polarities of component R.
+ve
-ve
kZ
* See Table 3.1.
Note: I. +ve and -ve deviations are termed easterly and westerly respectively.
2. The polarities and direction of components cZ and jZ depend on whether an aircraft is in the northern or southern
hemisphere.
Each of the three components produce three soft-iron components
that are designatej aX, b Y, cZ; dX, e Y, JZ; and gX, h Y, kZ. The
deviations caused by such components are set out in Table 3.2.
Total magnetic effect
The total effect of the magnetic fields that produce deviating forces
relative to each of the three axes of an aircraft is determined by
algebraically summing the quantities appropriate to each of the related
components. Thus:
TOTAL X 1 (longitudinal) = X+P+aX+bY+cZ
TOTAL Y' (lateral)
= Y+Q+dX+eY+JZ
TOTAL Z 1 (vertical)
= Z+R+gX+hY+kZ
Figure 3. 7 Relationship
between aircraft magnetism and
deviation coefficients.
l'.IORIZONTAl
COMPONENT 0
·~-,~~r ~(1.
_:~~,ITT:·~,i-· _
AAIIOl./HT Of . _ . .
DEVIATION
!~ N T C
COEFFlCIENT 8
·-·T·-·
+
rs
MAX.OEVL,\TlON
MAX DEVIATION
ON
ON EiW
JLIAA"'
COEfl'lCl£NTA\
--------.-CONSTANT
DEVltATION
I
CORRECTOR MAGNETS
ELECTROMAGNETIC OR
MECHANICAL/MAGNETIC
REALIGN
COMPASS
Deviation coefficients
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, termed positive or negative as the case may be, and
expressed in degrees. The relationship between them and the
components of aircraft magnetism is shown in Fig. 3.7.
Coefficient A
This represents a constant deviation and may be termed as either real
A, which is caused by the soft-iron components bY and dX, or
apparent A, which is a deviation produced by non-magnetic causes
such as misalignment of a direct-reading compass or of a flux
detector unit where appropriate (see page 182) with respect to an
aircraft's longitudinal axis. In practice it is not necessary to
distinguish between them, since they are both understood to be
included in the term coefficient A.
The coefficient is calculated by taking the average of the algebraic
differences between deviations measured on a number of equidistant
91
compass headings; this also applies to the other four coefficients. In
the case of A, the average may be determined from deviations ·on
either the four cardinal headings or, for greater accuracy, on these
headings plus the four quadrantal headings. Thus:
A
=
Deviation on N+E+S+W
4
or
A
=
Deviation on N+NE+E+SE+S+SW+W+NW
8
The coefficient is positive or negative, depending on whether the
constant deviation which it represents is easterly or westerly.
· Coefficient B
This represents the resultant deviation due to the presence, either
together or separately, of hard-iron component P and soft-iron
component cZ. When these components are of like signs, they cause
deviation in the same direction, but when of unlike signs they tend to
counteract each other. The coefficient is calculated from the formula:
B
= Deviation on E -
Deviation on W
2
Since components P and cZ cause deviation which varies as the
sine of an aircraft's heading 8, then deviation due to coefficient B
may also be expressed as Bx sin 0.
Coefficient C
This represents the resultant deviation due to the presence, either
together or separately, of hard-iron component Q and soft-iron
component jZ. When of like and unlike signs these components cause
deviations whose directions are the same as those caused by
components P and cZ. The coefficient is calculated from the formula:
C
=
Deviation on N - Deviation on S
2
Since components Q and jZ cause deviation which varies as the
cosine of an aircraft's headin~. then deviation due to coefficient C
may also be expressed as C x cos ().
Coefficient D
This represents the deviation due to the presence, either together or
separately, of components aX and eY which cause deviations of the
same direction when they are of unlike signs and counteract each
other when of like signs. When a +aX or a -eY predominates, o.r
when they are present together, the coefficient is said to be positive,
whilst a -aX or a +eY predominating or together cause a negative
coefficient D. It is calculated from the formula:
D
=
(Dev. on NE+Dev. on SW)-(Dev. on SE+Dev. on NW)
4
The deviations caused by components aX and e Y vary as the sine of
twice an aircraft's heading; therefore deviations may also be
expressed as D x sin 2fJ.
Coefficient E
This coefficient represents the deviation due to the presence of
components bY and dX of like signs. When a +bY and a +dX are
combined, coefficient Eis said to be positive, whilst a combination of
a -by and a -dX gives a negative coefficient; the two components
must in each case be equal in magnitude. The coefficient is calculated
from the formula:
E
=
(Dev. on N + Oev. on S) - (Dev'. on E +Dev. on W)
4
The deviations caused by the components b Y and dX vary as the
cosine of twice an aircraft's heading; therefore deviations may also be
expressed as E x cos 20.
The total deviation on an uncorrected compass for any given
direction of an aircraft's heading by compass may be expressed by
the equation:
Total deviation
=A
+ BsinO + CcosO + Dsin20 + Ecos2fJ
Deviation compensation
In order to determine by what amount compass readings are affected
by hard- and soft-iron magnetism, a special calibration procedure
known as 'swinging' is carried out so that adjustments can be made
to compensate for the deviations.
These adjustments are effected by compensator or corrector magnet
devices which, 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.
A compensator forms an integral part of a compass (see Fig. 3.3)
and in common with the majority of those in current use it contains
two pairs of permanent magnets which can be rotated through gearing
as shown in Fig. 3.8.
One pair of magnets is positioned laterally to provide a variable
longitudinal, field, thereby permitting adjustment for coefficient B,
93
Figure 3.8 Deviation
compensator device.
B
E/W ADJUSTER
while the other pair is positioned longitudinally to provide a variable
lateral field and so permit adjustment for coefficient C. Thus, the
fields are effective in neutralizing the deviations on only the two
cardinal headings appropriate to each of the coefficients.
The manner in which compensation is carried out may be
understood by considering the case of an adjustment having to be
made for coefficient B, as indicated in Fig. 3.9. When the
appropriate compensator magnets are in the neutral position, as
shown at (a), the fields produced are equal and opposite, and if as
also shown the aircraft is heading north, then the total field is aligned
with the earth's component Hand the compass magnet system.
Variation of the field strength by rotating the magnets will, therefore,
have no effect. This would also be the case if the aircraft was
heading south.
At (b) of Fig. 3.9, the aircraft is represented as heading east and,
as before, the compensator magnets are in the neutral position. The
total field of the magnets, however, is now at right angles to the
earth's component and the compass magnet system, and so if the
magnets are now rotated from the neutral position to the positions
shown at (c), the distance between poles N2 and S 1 is smaller. Since
tl1e intensity of a field varies in inverse proportion to the square of
the distance from its source, then in this case, a strong field will exist
between the poles of the magnets. The north-seeking pole of the
compass magnet system will, therefore, experience a greater repulsive
force, resulting in deflection of the system through 'an appropriate
number of degrees. If the magnets are rotated so as to strengthen the
field between poles N 1 and S 2 , the compass magnet system will then,
of course, be deflected in rhe opposite direction. Deflection of the
compass magnet system would be obtained in a similar manner with
the aircraft heading west.
The coefficient C compens;;ting magnets also produce similar
effects but, as a study of diagrams (d) to (f) will show, deflections of
the compass magnet system are only obtainable when an aircraft is
heading either north or south.
It will be apparent from the foregoing operating sequences that
maximum compensation of deviation on either side of cardinal
headings is obtained when the magnets are in complete alignment.
94
t
t
Figure 3. 9 Operation of a
compensator. (a) Aircraft
heading north; (b) and (c)
aircraft heading east; (d) and
(e) aircraft heading north;
(f) aircraft heading east.
t
The gears on which magnets are mounted are connected to
operating heads which, depending on the type of compass, are
operated either by a key or by means of a screwdriver. Indication of
the neutral position of magnets is given by aligning datum marks,
located typically on the ends of magnet-operating spindles and on the
compensator casing.
96
4 Gyroscopic flight
instruments
In addition to the airspeed indicator, the altimeter and the vertical
speed indicator, a basic group of flight instruments also comprises
instruments which provide direct indication of an aircraft's attitude.
There are three such instruments, namely a gyro horizon (sometimes
called an artificial horizon), a direction indicator, and a turn-andbank indicator. The complete group constitutes what is tenned the
'basic six' arrangement, details of which were given in Chapter l
(see page 20).
The three additional instruments utilize a gyroscopic type of sensing
element, the properties of which need to be understood before going
into the construction and operating details of each instrument.
The gyroscope and
Its properties
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 YY 1); and (iii) veering freedom, about a vertical axis
perpendicular to both the other axes (axis of veer ZZ 1).
The three degrees of freedom are obtained by mounting the rotor in
two concentrically pivoted rings, called inner
outer rings. The
whole assembly is known as the gimbal system of a free or space
gyroscope. The gimbal system is mounted in a frame as shown in
Fig. 4.1, so that in its normal operating position, all the axes are
mutually at right angles to one another and intersect at the centre of
gravity of the rotor.
The system will not exhibit gyroscopic properties unless the rotor is
spinning; for example, if a weight is suspended on the inner ring, it
will merely displace the ring about its axis YY I because there is no
resistance to the weight. 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, whlch means that the angular
momentum of a body about a given point remains constant unless
and
97
Figure 4. I Elements of a
gyroscope.
FRAME
ROTOR
XX, SPIN AXIS
YY 1 TILT AXIS
ZZ 1 VEER AXIS
some force is applied to change it. Angular momentum is the product
of the moment of inertia (/) and angular velocity (w) of a body
referred to a given point - the centre of gravity in the case of a
gyroscope.
If a weight is now suspended from the inner gimbal ring with the
rotor spinning it will be found that the ring will support the weight,
thus demonstrating the first fundamental property of rigidity. It will
also be found, however, that the complete gimbal'system will start
rotating about the axis ZZi, such rotation demonstrating the second
property of precession.
These rather intriguing properties can be exhibited by any system
in which a rotating mass is involved, and a simple example of such a
system is the bicycle. If we lift the front wheel off the ground, spin it
at high speed, and then turn the handlebars, we will feel rigidity
resisting us, and precession trying to twist the handlebars from our
grasp. Other familiar mechanical systems possessing gyroscopic
properties are aircraft propellers, and jet engine compressor and
turbine assemblies.
The two gyroscopic properties may be more closely defined as
follows:
98
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 change in
direction takes place, not in line with the force, but ahvays at a
point 90° away in the direction of rotation. The rate ot' 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.
Precession of a rotor will continue, while the force is applied,
until the plane of rotation becomes coincident with that of the
force. At this point there will be no further resistance to the force
and so precession will cease.
The axis about which a force is applied is termetl the input axis,
and the one about which precession takes place is termed the output
axis.
Determining the direction of precession
The direction in which a gyroscope will precess under the influence
of an applied force may be determined by means of vectors and by
solving certain gyrodynamic problems, but for illustration and
practical demonstration purposes, there is an easy way of determining
the direction in which precession will take place, and also of finding
out where a force must be applied for a required direction. It is done
by representing all forces as acting directly on the rotor itself.
At (a) in Fig. 4.2, the rotor is shown spinning in a clockwise
direction and with a force F applied upwards on the inner ring. In
transmitting this force to the rim of the rotor, as will be noted from
(b), it will act in a horizontal direction. Let us assume for a moment
that the rotor is broken into segments and concern ourselves with two
of them at opposite sides of the rim as shown at (c). Each segment
has motion m in the direction of rotor rotation, so that when force F
is applied there is a tendency for each segment to move in the
direction of the force. This motion is resisted by rigidity, but the
segments will turn about the axis ZZ 1 so that their direction of
motion is along the resultant of motion m and force F. The other
segments will be affected in the same way; the.-efore, in being
combined to form the solid m..ss of a rotor it will precess at an
angular velocity proportional to me applied force (see diagrams (d)
and (e)).
99
Figure 4.2 Gyroscopic
precession (I). (a) Gyro resists
force; (b) transmission of
force; (c) effect on rotor
segments; (d) generation of
precession; (e) effect of
precession.
FOftCE
I
Z1
z,
z
(b)
I
(8)
-X1
z
I
(C)
_..,
I
I
X
(d)
(el
In the example illustrated in Fig. 4.3(a), a force Fis applied to the
outer ring; this is the same as transmitting the force to the rotor rim
at the point shown at (b). As in the previous case this results in the
direction of motion changing to the resultant of motion m and force
F. This time, however, the rotor precesses about the axis YY I as
indicated at (d) and (e).
100
z
Figure 4. 3 Gyroscopic
precession (2). (a) Gyro resists
force; (b) transmission of
force; (d effect on rotor
segments; (d) generation of
precession; (e) effect of
precession.
I
z
Pl.AN£ Of
II
SPIN
x-
Z;
Z1
(a)
(b)
M
\.
F
--·
,,,......--- ,,,. ...
x--'1, /
PRECESSION
,. ,.
F
\
,/,.
\
(C)
M
z
z
PLANE OF
PRECESSION
I
')
..
·X1
-
y .....
• -X1
--"'
>
-"-v1
I
z,
I
z,
(di
References
established by
gyroscopes
(e)
For use in aircraft, gyroscopes must establish two essential reference
datums: one for the detection of pitch and roll attitude changes, and
the other for the detection of changes about the vertical axis, i.e. a
directional reference. These datums are established by using vertical
and horizontal spin-axis gyroscopes respectively as shown in Fig.
4.4. Both types utilize their fundamental properties in the following
101
Figure 4.4 References
established by gyroscopes.
X
z
i
VERTICAi. AXIS
GYROSCOPE
Dll!ECTIONAL
r-Y,
~x,
:..-x-,
z,
HOOIZQ!flAI. AXIS
GYROSCOPE
manner: rigidity provides a stabilized reference unaffected by
movement of the supporting body, and precession controls the effects
of apparent and real drift thus maintaining stabilized datums.
It will also be noted from Fig. 4.4 that the pitch, roll, and
directional attitudes of an aircraft are determined by its displacement
with respect to each appropriate gyroscope. For this reason,
therefore, the gyroscopes are referred to as displacement-type
gyroscopes. Each one has three degrees of freedom and,
consequently, three mutual axes, but for the purpose of attitude
sensing, the spin axis is discounted since no useful attitude reference
is provided when displacements take place about the spin axis alone.
Thus, in the practical case, the two types of gyroscope are further
classified as two-axis displacement gyroscopes.
Limitations of a free
gyroscope
In flight, the attitudes of an aircraft must be referenced with respect
to the earth's surface, and this being so requires that a free or space
gyroscope, thus far considered, be corrected for drift with respect to
the earth's rotation, called apparent drift, and for wander as a result
of carrying a gyro:,cope over the earth's surface, called transport
wander.
Apparent drift
The earth rotates about its axis at the rate of 15 °/hour, and in
association with gyrodynamics, this is termed earth rate (w,). When a
102
Figure 4.5 Drift and transport
wander.
A
B
C
LOCAL NORTH
ALIGNED WlTH EARTH'S AXIS
LOCAL VERTICAL
w, EARTH'S RAT£
>,
LATITUDE
(11)
X --·
HOURS
AFTER 6
z
(c)
free gyroscope is positioned at any point on the earth's surface, it
will sense, depending on the latitude at which it is positioned, and on
the orientation of its spin and input axes, various components of w~ as
an angular input. Thus, to an observer on the earth having no sense
of the earth's rotation, the gyroscope would appear to veer or drift.
This may be seen from Fig. 4.S(a) which illustrates a horizontal-axis
gyroscope at a latitude A: At 'A', the input axis is aligned with the
local N-S component of w,; therefore, to an observer at latitude A
103
the gimbal system would appear to drift clockwise (opposite to the
earth's rotation) in a horizontal plane relative to the frame, and at a
rate equal to 15° cos A. When the input axis is aligned with that of
the earth ('B'), drift would also be apparent, but at a rate equal to
w,, i.e. 15° /hour. If the input axis is now aligned with the local
vertical component of w, ('C' in the diagram) the apparent drift
would be equal to 15° sin A.
In order to further illustrate drift, we may consider diagram (b) of
Fig. 4.5, which is a plan view of a free horizontal-axis gyroscope
positioned at the North pole with its input axis (ZZ 1) aligned with
that of the earth. After three hours the earth will have rotated through
45 °, and the gyroscope will appear to have drifted through the same
amount but in the opposite direction. After six hours the earth's
rotation and apparent drift will be 90°, and so on through a complete
24-hour period.
If the same gyroscope were to be positioned so that its input axis
ZZ 1 was aligned with the E-W component of w, at any point, its
spin axis would then be vertical; in other words, it becomes a
vertical-axis gyroscope. Since the plane of rotation is coincident with
that of the earth, there will be no apparent drift.
Real drift
Real drift results from imperfections in a gyroscope such as bearing
friction and gimbal system unbalance. Such imperfections cause
unwanted precession which can only be minimized by applying
precision engineering techniques to the design and construction.
Transport wander
Let us again consider a horizontal-axis gyroscope which is S!!t up
initially at the North pole, with its input axis aligned with that of the
earth. In this position it will exhibit an apparent drift equal to w,.
Assume now that it is carried to a lower latitude, ,and with its input
axis aligned with the local vertical component of w,. During the
period of transport it will have appeared to an observer on the earth
that the spin axis has tilted in a vertical plane, until at the new
latitude it appears to be in the position shown at (c) of Fig. 4.5. This
apparent tilt, or transpon wander, would also be observed if, during
transport, the input axis were aligned with either a local N -S
component, or a local E-W component of w,.
Transport wander will, of course, appear simultaneously with drift,
and so for a complete rotation of the earth, the gyroscope as a whole
would appear to make a conical movement. The angular velocity or
transpon rate of this movement will be decreased or increased
depending on whether the E-W component of an aircraft's speed is
towards east or west. The N -S component of the speed will increase
the maximum divergence of the gyroscope axis from the vertical, the
amount of divergence depending on whether the aircraft's speed has a
N or S component and also on whether the gyroscope is situated in
the northern or southern hemisphere.
The relationship between w., transport wander, and input axis
alignment are summarized in the following table:
Input axis alignment
Earth rate
Transport wander
Local north
Local east
Local vertical
w, cos 11.
nil
w, sin 11.
u
V
R
i!_ tan X
R
R
= earth's angular velocity; 11. = latitude; R = earth's radius; V
N-S component of transport velocity; U = E-W component of
transport velocity.
w,
=
If the input axis of a gyroscope were to be positioned such that its
spin axis was vertical, then during transport it would only exhibit
transport wander.
Control of drift and transport wander
Before a free gyroscope can be of practical use, drift and transport
wander must be controlled so that the plane of spin of the rotor is
maintained relative to the earth; in other words, it requires
conversion to what is termed an earth gyroscope.
· Drift, as already pointed out, relates only to horizontal-axis
gyroscopes, and it can be controlled either by (i) calculating
corrections using the earth rate formulae given in the preceding table
and applying them as appropriate, e.g. to the readings of a direction
indicator; (ii) applying fixed torques which unbalance the gyroscope
and cause it to precess at a rate equal and opposite to w,; (iii)
applying torques having a similar effect to that stated in (ii) but
which can be varied according to the latitude in which the gyroscope
is being used.
The control of transport wander is normally achieved by using
gravity-sensing devices which automatically detect tilting of the
gyroscope's spin axis, and applying the appropriate corrective
torques.
The operation of some typical control methods will be described
later under the headings of the appropriate flight instruments.
Displacement
gyroscope limitations
Depending on the orientation of its gimbal system, a displacement
gyroscope can be subject. to certain operating limitations; one is
referred to as gimbal lock and the other as gimbal e"or.
Gimbal lock
This occurs when the gimbal orientation is such that the spin axis
becomes coincident with one or other of the axes of freedom which
serve as attitude displacement references. Let us consider, for
example, the case of the spin axis of a vertical-axis gyroscope shown
in Fig. 4.4 becoming coincident with the ZZ 1 axis. This means, in
effect, that the gyroscope would 'lose' its spin axis, and since the
plane of spin would be at 90° to the ZZ 1 axis but in the same plane
as displacements in roll, then the stable roll attitude reference would
also be lost. If, in this 'locked' condition of the gimbal system, the
gyroscope as a whole were to be mrned, then the forces acting on the
gimbal system would cause the system to precess or topple.
Gimbal error
This is an error which is also related to gimbal system orientation,
and it occurs whenever the gyroscope as a whole is displaced with its
gimbal rings not mutually at right angles to each other. The error is
particularly relevant to horizontal-axis gyroscopes when used in
direction indicating instruments (see page 127).
Methods of operating
gyroscopic flight
instruments
There are two principal methods of driving the rotors of gyroscopic
flight instruments: pneumatic and electric.
Pneumatic
The pneumatic method is adopted in a number of small types of
aircraft, and may be either vacuum or pressure. A typical vacuum
system is shown schematically in Fig. 4.6; it consists of an enginedriven pump that is connected through pipelines to the aopropriate
flight instruments. A vacuum indicator, a relief valve, and a central
air filter are also provided. In operation the pump creates a vacuum
that is regulated by the valve at a value between 3.5 and 4.5 in Hg.
Some types of tum-and-bank indicator may operate at a lower value
106
Figure 4.() Vacuum-operated
system.
Relief valve
Pump
Vacuum gauge
Vacuum connection
fl';:::=::=:= to relief valve
II
and instruments
Engine -
bleed air
Pressure
regulator
Ejector/venturi
and this is obtained by including an additional relief valve in the
main supply line.
Each instrument case has two connections: one is made to the
pump and the other is made internally to a spinning jet system that is
open to the surrounding atmosphere via the central air filter. When
vacuum is applied, the pressure within the cases of the instruments is
re,duced to allow surrounding air to enter' and emerge through the
spinning jets. The jets are positioned adjacent to a series of recesses
(commonly called 'buckets') formed in the periphery of each
gyroscope rotor, so that as the airstreams impinge on the 'buckets',
the rotors are rotated at high speed.
An example of a relief valve is shown in Fig. 4.7. During system
operation the valve remains closed by compression of the spring, the
tension of which is pre-adjusted to obtain the required vacuum so that
air pressure acting on the outside of the valve is balanced against
spring tension. If for some reason the adjusted value should be
exceeded, the outside air pressure would overcome spring tension,
thus opening the valve to allow outside air to flow into the system
until the balanced condition was once again restored.
In' some small types of turbine··engine aircraft that have
pneumatically-operated instruments installed, the vacuum is created
107
Figure 4. 7 Relief valve.
COMPRESSION
SPRING
OUTSIDE AIR PRESSURE
by bleeding air from the engine compressor and passing it through an
ejector/venturi (see Fig. 4.6).
A pressure-operated system is, as far as principal components are
concerned, not unlike a vacuum system but, as will be noted by
comparing Figs 4.6 and 4.8, a changeover of system inlet and outlet
connections is necessary.
Electric
In electrically-operated instruments, the gyroscopes are special
adaptations of ac or de motors that are designed to be driven from
the appropriate power supply systems of an aircraft. In current
applications, ac motors are adopted in gyro horizons, while de motors
are more common to tum-and-bank indicators. Gyroscopes used for
the purpose of direction indicating can also be motor-driven, but they
normally form part of a magnetic heading reference system, or of the
more widely adopted flight director systems. These systems will be
covered in later chapters.
Gyro horizon
principle
100
A gyro horizon indicates the pitch and roll attitude of an aircraft
relative to its vertical axis, and so for this porpose it employs a
Figure 4. 8 Pressure-operated
system.
Pressure
In-line fitter
regulator
Pressure
gauge
Pump
Gyroscopic
instruments
Air discharge
displacement gyroscope whose spin axis is vertical. Indications of
attitude are presented by the relative positions of two elements, one
sym~lizing the aircraft itself, the other in the fonn of a bar
stabilized by the gyroscope and symbolizing the natural horizon.
Supplementary indications of roll are presented by the position of a
stabilized pointer and a fixed roll angle ,scale. Two methods of
presentation are shown in Fig. 4.9.
The gimbal system (see Fig. 4.10) is arranged so that the inner
ring forms the rotor casing and is pivoted parallel to an aircraft's
lateral axis YY 1; the outer ring is pivoted at the front and rear ends
of the instrument case, parallel to the longitudinal axis ZZ 1• The
element symbolizing the aircraft may be either rigidily fixed to the
case, or it may be externally adjustable for setting a particular pitch
trim reference.
In operation the gimbal system is stabilized so that in level flight
the three axes are mutually at right angles. When there is a change in
an aircraft's attitude, it goes into a climb, say, the instrument case
and outer ring wm move about the axis YY I of the stabilized inner
ring. The horizon bar is pivoted at the side and to the rear of the
outer ring, and engages an actuating pin fixed to the inner ring, thus
forming a magnifying lever system. The pin passes through a curved
slot in the outer ring. In a climb attitude the bar pivot carries the rear
109
MltllATUflf AIRCRl>.FT
Figure 4. 9 Gyro horizon
presentations. (a) Bottom bank
scale; (b) 1op bank scale.
/
(a)
(b)
end of the bar upwaids so that it pivots about the stabilized actuating
pin. The front end of the bar is therefore moved downwards through
a
angle than that of the outer ring, and since the movement is
relative to the symbolic aircraft element, the bar wili indicate a climb
attitude.
Changes in the lateral attitude of an
i.e. roHing, ais1p1a<:e
the instrument case about the axis ZZ1, and the whole stabilized
gimbal system. Hence, lateral attitude
are indicated by
movement of the symbolic aircraft element relative to the horizon
bar, and also by relative movement between the mil angle scale and
pointer.
Freedom of gimbal system movement about the roll and
axes
is 360° and 85 ° respectively, the latter being restricted by means of a
'resilient stop'. The reason for this restriction is to
lock (see page 106).
A typical instrument of the vacuum-driven type is shown in Fig.
4. H. The rotor is pivoted in ball-bearings within the inner
ring/casing which is, in rum,
in outer ring bearings. The
upper bearing of the rotor is
to
for the
effects of differential expansion between the rotor shaft and casing
under varying temperature conditions. A
plate which
Figure 4.10 Principle of gyro
horizon. l Symbolic aircraft, 2
rotor. ;; outer ring, 4 inner
ring. 5 balance weight, 6 pivot
point, 7 actuating pin, 8
horizon bar, 9 roll pointer and
scale.
X
X
I
-z,
r
Y--
I
x,
CLIMB ATTITUOE
x,
BANK TO l'OllT
Figure 4. I I Pneumatic type of
gyro horizon. I Sky plate,
2 inner gimbal ring, 3 resilient
stop, 4 balance nut,
5 temperature compensator,
6 rotor, 7 actuating pin, 8 outer
gimbal ring, 9 ~c!uator ann.
!O pendulous vane unit,
11 buffer stops, l2 bank
pointer. 13 horirnntal bar.
13
12
.
11
GIMBAl ANO !\OTOll ASSEMELY
111
symbolizes the sky is fixed to the front end of the outer ring and
carries the roll pointer which registers against the roll angle scale.
A vacuum supply connection is provided at the rear of the
instrument case, together with a filtered air inlet. The latter is
positioned over the outer ring rear-bearing support and pjvot which
are drilled to communicate with a channel in the outer ring. This
channel terminates in diametrically-opposed spinning jets within the
rotor casing, the underside of which has a number of outlet holes in
it.
When the vacuu,n system is in operation, the air pressure within
the instrument case becomes lower than that of the surrounding air,
which is then able to pass through the filtered inlet and to the
spinning jets. The air issuing from the jets impinges on the rotor
buckets, thus imparting even driving forces to spin the rotor at
approximately 15 000 rev/min in an anti-clockwise direction as
viewed from above. After spinning the rotor, the air passes through a
pendulous vane unit (see page 114) attached to the underside of the
rotor casing, and is finally drawn off by the vacuum source.
Electric gyro horizon
This instrument is made up of the same basic elements as a
pneumatic type, with the exception that the gyroscope is an ac
squirrel-cage induction motor which operates from a 115 V, 400 Hz,
three-phase supply source.
One of the essential requirements of any gyroscope is to have the
mass of the rotor concentrated as near to the periphery as possible,
thus ensuring maximum inertia. This presents no difficulty where
solid. metal rotors are concerned, but when adopting electric motors
as gyroscopes some rearrangement of their basic design is necessary
in order to achieve the desired effect. An induction motor normally
has its rotor revolving inside its stator, but to make one small enough
to be accommodated within the confines of an instrument would mean
too small a rotor mass and inertia. This is overcome by designing the
rotor and its bearings so that it rotates on the outside of the stator;
thus, for the same required size of motor the rotor mass is
concentrated further from the centre, thereby increasing the radius of
gyration and inertia.
The motor assembly· is carried in a housing which forms the inner
ring, this in tum being supported in the outer ring bearings. The
horizon bar assembly is pivoted and actuated in a manner similar to
that already described on page 109. The ac power supply is fed to
the motor stator via slip rings, wire brushes and finger contact
assemblies, thereby allowing for all gimbal ring movements.
When power is applied, a rotating magnetic field is set up in the
stator; the field, in turn, inducing a current in the squirrel-cage rotor.
112
The effect of this current is to produce magnetic fields which interact
with the stator's rotating field causing the rotor to turn at a speed of
approximately 20 000-23 000 rev/min. A solenoid-operated 'power
off warning flag is also provided.
Standby attitude indicators
Many aircraft currently in service employ flight director systems, or
more sophisticated electronic flight instrument systems, all of which
comprise indicators having the capability of displaying not only
attitude data, but also the data from other navigational systems. In
such cases, therefore, the role of a conventional gyro horizon is
relegated to that of secondary or standby, for use as a reference in
the· event of any failure that might occur in the attitude display
sections of the aforementioned primary systems.
An example of one type of gyro horizon designed for use as a
standby attitude indicator is shown in Fig. 4.12. Its gyroscope is
powered by l 15 V, three-phase ac supplied by a static inverter
which, in tum, is powered by 28 V de from the battery busbar of an
aircraft. Power from such a source is always available, thereby
ensuring continuity of indicator operation. In place of the mor:!
conventional stabilized horizon bar method of displaying attitude, a
stabilized spherical element is adopted as the reference. The ~1pper
half of the element is coloured blue to display climb attitudes, and is
divided, by an horizon line, from the lower half which is in black
and displays descending attitudes. Each half is graduated in 10°
increments, the upper one up to 80°, and the lower up to 60°. Roll
or bank angles are indicated in the conventional manner.
A pitch trim adjustment and a fast-erection facility are provided,
both being controlled by a knob in the lower right-hand .comer of the
indicator bezel. When the knob is rotated in its normally 'in'
Figure 4.12 Standby attitude
indicator.
POWER OFP FLAG
PlTCH ERECTION/TRIM KNOii
113
position, the aircraft symbol may be positioned through :1: 5 °, thereby
establishing a variable pitch trim reference. Pulling the knob out and
holding it actuates a fast-erection circuit (see also page 119).
Erection systems for gy:ro horizons
These systems are designed for the purpose of erecting the gyroscope
to, and maintaining it in, its vertical spin-axis position during
operation. The systems adopted depend on the particular design of
gyro horizon, but they are all of the gravity-sensing type and in
general fall into two main categories: mechanical and electrical.
Figure 4. I 3 Pendulous vane
unit. (a) Construction; (b)
precession due to air reaction;
(c) gyro in vertical position;
(d) gyro tilled.
Pendulous vane unit
This is a mechanical system adopted for the gyro horizon described
on page 110. It is fastened to the underside of the rotor housing and
as indicated in Fig. 4.13(a) it consists of four knife-edged,
pendulously-suspended vanes clamped in pairs on two intersecting
shafts and passing through the unit body. One shaft is parallel to the
axis YY 1 and the other to the axis ZZ 1• In the sides of the body there
are four elongated ports (A, B, C and D), one under each vane.
After having spun the rotor, air is exhausted through the poris,
emerging as four streams and in the directions indicated. The reaction
of the air as it flows through the ports applies a force to the unit
AIR FLOW FROM
ROTOR HOUSING
~f:.i.
(5)
·:h· f
v,,-n
(b)
(d)
114
body. The vanes, under the influence of gravity, always hang in the
vertical position and it is this feature that is utilized to govern the
airflow from the ports and to control the reaction forces applied to
the gyroscope.
When the gyroscope is in its normal vertical position as at (b) the
knife edges of the vanes bisect each of the ports, making all four
openings equal. All four air reactions are therefore equal and the
resultant forces about each axis are in balance.
If now the gyroscope is displaced, so that, for example, its top is
tilted towards the front of the instrument as at (c), the pair of vanes
on the axis YY 1 remain vertical, thus opening the port D and dosing
the port B. The increased reaction of the air from D results in a
force being applied to the body in the direction of the arrow, about
axis XX 1• This force is equivalent to one applied on the underside of
the rotor and to the left, or at the top of the rotor at point F as
shown at (d). Precession back to the vertical position will therefore
take place at point P, and the vanes will again bisect the ports to
equalize the air reactions.
Figure 4. I 4 Ball-type erection
unit. (a) Gyro vertical.
(b) tilted away from front of
instrument; (c) precession to
vertical.
Ball-type erection unit
This mechanical system is applied to some designs of electric gyro
horizon; it utilizes the precessional forces resulting from gravity on a
number of steel balls displaced within a rotating holder suspend.;d
from the gyroscope housing as shown in Fig. 4.14. The balls are free
fl'!ONT Of
INSTRUMENT
SPIN AXIS
--
X
'
'
I
z-~~~T~~-~1
(cl
X',
y
z
0
Y11.£FT SIDE
(a)
Z1
z,
z
(bl
115
to roll across a radiused erecting disc and into and out of a number
of specially profiled hooks in a plate fixed around the inner edge of
the holder. The 5.pacing of the hooks is chosen so as to regulate the
release of the balls when the gyroscope tilts, and to shift their mass
to the proper point on the erecting disc to apply the force required
for precession. Rotation of the holder takes place through reduction
gearing from the gyro,cope's rotor shaft; the speed of the holder is
approximately 25 rev/min.
When the gyroscope is in its normal operating position, as shown
at (a), the balls change position as the holder rotates but their mass
remains concentrated at the centre of the disc. Under this condition,
gravity exerts its greatest pull at the centre of the mass, and therefore
all forces about the principal axes of the gyroscope are in balance.
At (b) the gyroscope's vertical axis is shown displaced about pitch
axis YY I away from the front of the instrument. The displacement of
the ball holder causes the balls to roll towards the hooks, which at
that instant are on the low side; therefore the force due to gravity is
now shifted to this side. Since the hooked plate is rotating (clockwise
viewed from above), the balls and the point at which the force is
acting will be carried round to the left-hand side of the holder. In this
position the balls remain hooked and their mass remains concentrated
to allow a force to be exerted at the left-hand side of the holder as
indicated at (c). This force may also be considered as acting directly
on the left-hand bearing of the gyroscope housing and outer ring.
Transferring this point of applied force to the rotor rim, precession
will then take place about axis YY I to counteract the displacement.
As the erector mechanism continues to rotate, the balls will be
carried round to the high side of the holder, but one by one they will
roll into the hooks at. the lower side. Thus, their mass is once again
concentrated at this side, allowing the force and precession to be
maintained as they are carried around to the left-hand side. This
action continues with diminishing movement of the balls as the
gyroscope erects to its normal vertical position, at which the balls are
at the centre of the disc and the force due to gravity is again
concentrated at the centre of the mass.
Displacement of the gyroscope in other directions about its lateral
or longitudinal axes will result in similar actions to those described.
Torque motor and levelling switch system
This system is used in a number of electrically-operated gyro
horizons and the remote vertical gyroscope units associated with
flight director systems. It consists of two torque control motors
operated independently by liquid levelling switches, which are
mounted, one parallel to the lateral axis, and the other parallel to the
longitudinal axis. The disposition of motors and switches is illustrated
diagrammatically in Fig. 4.15.
116
Figure 4.15 Torque motor and
levelling switch erection
system.
PITCH TORQUE MOTOR
SPIN AXIS
1x
ROLL TORQUE MOTOR
x1
The laterally-mounted switch detects roll displacement and is
connected to its torque motor so that a corrective force is applied
around the pitch axis. Pitch displacements are detected by the
longitudinally-mounted switch, which is connected to its torque motor
so that corrective forces are applied around the roll axis. Each switch
is in the form of a sealed glass tube containing three electrodes and a
small quantity of either mercury or an electrolytic solution.
Each torque motor consists of a stator and a squirrel-cage rotor.
The roll torque motor has its stator fixed to the outer ring of the
gyroscope, and its rotor fixed to the inner ring. The stator of the
pitch motor is fixed to the instrument frame, and its rotor fixed to the
outer ring.
The electrical interconnection of the components that comprise each
system is indicated in Fig. 4.16(a). In the case of mercury levelling
switches, the mercury will lie at the centre of the tubes and, being in
contact with the centre electrode, will supply a voltage to the
reference windings of their respective torque motors, only when the
gyro is running and in its normal operating position. The two outer
electrodes are connected one to each section (designated 'A' and 'B')
of their respective torque motor control windings; thus, in the normal
operating position of the gyro the control winding circuit is open.
When the gyro is displaced about one of its axes, the appropriate
levelling switch will also be displaced so that the mercury bridges the
gap between the centre electrode and one or other outer electrode.
This completes a circuit to either the 'A' or 'B' section of the
117
Figure 4. /6 Circuit diagram of
erection system.
---....,
------<>115V
I
I
I
____ ,,.
'
_,. AIRCRAFT SUPPLY VOLTAGE
V ANO I
(a)
TO CONTROL WINDING
~
'A'
'B'
1
(b)
(c)
respective torque motor control winding, depending on the direction
of gyro displacement.
In order for the torque motor to apply the necessary corrective
torque to the gimbal system, the magnetic field of the motor stator
must be made to rotate. As will be noted from Fig. 4.16, the voltage
to the reference winding is applied via a capacitor, and so, as in any
ac circuit containing capacitance, the phase of the current is shifted
so as to lead the voltage by 90°.
The control winding circuit has no capaeitance, and so the voltage
and current flowing through it are in phase; therefore, and because
the control and reference windings are both supplied from the same
power source, reference winding current must also lead control
winding current by 90°. This out-of-phase arrangement (called phase
quadrature} applies also to the magnetic field set up by each
winding.
118
Thus, with current and a magnetic field flowing through the
appropriate half of the control winding resulting from a displacement
of the gyro, a resultant field is produced which rotates within the
torque motor stator in either a clockwise or anti-clockwise direction.
As the field rotates, it cuts the conductors of the rotor and induces a
current in them; this in turn produces a magnetic field that interacts
with the stator field and creates a tendency for the rotor to rotate
with the stator field. This tendency is opposed because of the rigidity
of the gyro, and consequently a reactive torque is set up in the
torque motor and is exerted on the associated gimbal ring to precess
the gyro and levelling switch to their normal operating position.
In the case of levelling systems utilizing electrolytic solution-type
switches, the supply of current to the control winding sections of a
torque motor is controlled in a different manner to that of mercury
levelling switch systems. The reason for this is, as will be noted from
Fig. 4. l6(b), that the electrodes are always immersed in the
electrolytic solution, and the circuits to the control winding of a
torque motor are always closed. In the normal stabilized vertical
position of the gyro, the switch electrodes are in equal amounts of
electrolyte and so the currents flowing in each section of a torque
motor control winding are equally opposed. Since the electromagnetic
effects on the rotors are also equally opposed then no torques will be
applied to the gimbal system.
When a switch is displaced it causes a change in the amount of
surface area of electrolyte in contact with the electrodes and, in turn,
an imbalance in the electrical resistance of the control winding
circuit. This may be noted from diagram (c): at the 'low end'
electrode there is a greater amount of electrolyte and, in accordance
with basic electrical principles, this means low resistance and so more
current wili flow in that half of the control winding connected to that
electrode. A corresponding rotating field is therefore produced to set
up a reactive torque in the torque motor for precessing the gyro and
levelling switch to the normal stabilized position and in the same
manner as that described earlier.
Fast-erection systems
These systems are used in some types of electrically-operated gyro
horizons for the purpose of bringing their gyros to the vertical
position as quickly as possible from large angles of tilt, particularly
during starting.
A control device is therefore provided which, in a typical system,
activates a set of contacts to introduce a higher voltage and current
flow through the control phase windings of the erection torque
control motors. The resulting higher torque applied thereby increases
the precession rate. To prevent overheating of the stator coils a time
limit (typically 15 seconds) is imposed on system operation. Certain
119
types of gyro horizon utilize a system whereby the gimbal system is
mechanically caged when the operating knob is operated.
Erection rate
This is the term used to define the time taken, in degrees per minute,
for a vertical gyroscope to take up its normal operating position
under the control of its particular type of erection system. Ideally, the
rate should be as fast as possible under all conditions, but in practice
such factors as speed, turning and acceleration of an aircraft, and the
earth's rotation, all have their effect and must be taken into account.
Thus, erection systems are designed so that, for small angular
displacements of a gyroscope from the vertical, the erecting couple is
proportional to the displacement, while for larger displacements it is
made constant. It is also arranged that the couple gives equal erection
rates for any rotor axis displacement in any direction in order to
reduce the possibility of a slow cumulative error during manoeuvring
of an aircraft. Normal rates provided by some typical erection
systems are 8°/min for pneumatic-type gyro horizons, and 3-5°/min
for those that are electrically operated.
Errors due to acceleration and turning
Since gyro horizon erection devices are of the pendulous type, it is
possible for them to be displaced by the forces acting during the
acceleration and turning of an aircraft, and unless provision is made
to counteract this the gyroscope spin axis would be precessed to a
false vertical position, thereby presenting a false attitude indication.
For example, let us consider the effects of a rapid acceleration in the
flight direction, firstly on the vane type of erection device, and
secondly on the levelling switch and torque motor type.
As shown in Fig. 4.17(a) the acceleration force will deflect the two
athwartships-mounted vanes to the rear, thus opening the right-hand
port. The greater reaction of the air flowing through this port applies
a force to the underside of the rotor causing it to precess forward
about the axis YY 1• The horizon bar is thus displaced downwards,
presenting a false climb indication.
In the case of levelling switches (Fig. 4 .17 (b)), an acceleration
force will deflect the liquid in the one related to pitch to the rear of
its tube. A circuit will thus be completed to the control winding of
the pitch torque motor which, in the manner described on page 117,
will precess the gyroscupe forward and will therefore also produce a
false climb indication.
In both cases the precession is due to a natural response of the
gyroscope, and the pendulous vanes and/or liquid always return to
their neutral positions, but for as long as the disturbing forces
remain, such positions apply only to a false vertical. When the fotces
120
Figure 4. J7 Acceleration
error. (a) Vane-type erection
sys1en;i; (b) levelling switch and
torque motor erection system.
(ON UNOERS1D£ OF
ROTOR)
FALSE INDICATION
OF ASCENT
FALSE
VERTICAL
\
PITCH TORQUE
MOTOR
X
\d)
I
I
..---,
(bl
X
are removed the false climb indication will remain initially and then
gradually diminish under the influence of precession, restoring the
gyroscope to its normally true vertical position.
· It should be apparent from the foregoing that, during periods of
deceleration, a gyro horizon will present a false indication of descent.
When an aircraft turns, false indications about both the pitch and
roll axes can occur, due to what are termed 'gimbailing effects'
brought about by forces acting on both sets of pendulous vanes or
both levelling switches, as appropriate. There are, in fact, two errors
•foe to turning: erection errors and pendulosity errors.
121
Erection errors As an aircraft enters a turn, the gyroscope's spin
axis will initially remain in the vertical position and so an accurate
indication of the roll or bank angle will be presented. In this position,
however, the longitudinally-mounted pendulous vanes, or the roll
levelling switch, are acted upon by centrifugal force. This force will
be applied to the gyroscope in such a direction that it will tend to
precess towards the perpendicular along which the resultant of
centrifugal and gravity forces are acting. Thus, the gyroscope erects
to a false vertical and introduces an error in roll indication. Such
errors may be compensated by one of the following methods: (i)
inclination of the gyroscope's spin axis; (ii) erection cut-out; and (iii)
pitch-bank compensation.
Inclined spin axis This method is based on the idea that, if the top
of the axis can describe a circle about itself during a turn, then only
a single constant error will result. In its application, the method is
mechanical in form and varies with the type of gyro horizon, but in
all cases. the result is to impart a constant tilt of the axis from the
vertical (typically 1.6° or 2.5°). In pneumatic types of gyro horizon,
the athwartships-mounted pendulous vanes are balanced so that the
gyroscope is precessed to the tilted position; in certain electric gyro
horizons, the pitch levelling switch is fixed in a tilted position so tl;at
the gyroscope is precessed away from the true vertical in order to
overcome what it detects as a pitch error. The linkages between
gyroscope and horizon bar are so arranged that during level flight the
horizon bar will indicate this condition.
Erection cut-nut An example of this method as applicable to
electrically-operated gyro horizons incorporates additional liquid-level
switches positioned in pairs on the pitch and roll axes. The switches
are connected to the torque motors in such a way that under the
influence of forces they isolate the control winding circuits from the
main erection switches. Operation of the system may be understood
from Fig. 4.18 which shows the arrangement applicable to pitch
erection cut-out. The pairs of switches are set at an angle to each
other in order to differentiate between acceleration and deceleration
forces.
Assuming that an acceleration occurs, the electrolyte in the pitch
erection switch will be displaced in the opposite direction. and as we
have already learned, the change in electrical resistance of the
electrolyte will produce an unbalanced condition in the torque motor
control winding circuit, and a torque tending to precess the gyro; in
this case, to a false vertical position. At the same time, however, the
acceleration force displaces the electrolyte in cut-out switch 'A' to
complete a circuit to a solid-state switch which then operates to open
the ground connection of the control winding; torque motor operation
Figure 4.18 Erection cut-out.
-Acceleration
Pitch
erection
115V_ _.._ _ _ ___, switch
ac
-
-
---Acceleration
L
----ac
26V
is thereby prevented. Switch 'B' performs a similar function but
under the influence of a decleration force.
Roll erection cut-out is accomplished by another identicallyarranged pair of switches connected to the roll torque motor, except
of course that they are angled to differentiate between the centrifugal
forces corresponding to left· and right turns.
Pitch-roll (bank) erection This method is a combined one
{incorporated in some gyro horizons utilizing mercury-type levelling
switches) in which the roll levelling switch is disconnected during a
turn by the pitch levelling switch. It is intended to correct the varying
pitch and roll errors and operates only when the rate of tum causes a
centrifugal acceleration exceeding 0.18 g, which is equivalent to a
10° tilt of the roll erection switch. As shown schematically in Fig.
4.19, two additional switches, connected in a double-pole changeover
123
Figure 4.19 Pilch-roll (bank)
erection.
MERCURY DISPLACED
TO MAKE CIRCUIT TO
ROLL TORQUE MOTOR
I
I
MERCURY DISPLACED
TO INTERRUPT CIRCUIT
TO ROLL TORQUE
MOTOR
ROLL MERCURY
SWITCH
t
• i
=
SUPPLY
PITCH
MERCURY SWITCH
---
t:::=(>
B
I
I
t
SUPPLY WHEN CENTRll'UGAL
c::::,p ACCELERATION LESS THAN 0-18g
.,.._ SUPPlY WHEN CENTfflfUGAL
ACCELERATION MORE THAN 0·18g
- - - SUPPLY IN CHANGEOVER,fUNCTION
arrangement, are provided and are interconnected with the normal
erection systems.
Let us consider first a turn to the left and one creating a centrifugal
acceleration less than 0.18 g. In such a turn, the mercury in the roll
levelling switch will be displaced to the right and wiil bridge the gap
between the supply and right-hand electrodes, thus completing a
circµit to the roll torque motor. This is the same as if the gyroscope
axis had been tilted to the right at the commencement of the turn; the
roll torque motor will therefore precess the gyroscope back to a false
vertical, and left of the true one. At the same time, the axis tilts
forward due to gimballing effect, and the mercury in the ptich
switch, being unaffected by centrifugal acceleration, moves forward
and completes a circuit to the pitch torque motor, which precesses the
gyroscope rearwards. The two additional switches, which are also
mounted about the roll axis, do not come into operation since the
mercury in them is not displaced sufficiently to contact the right-hand
electrodes. Thus, with acceleration less than 0.18 g there is no
compensation.
When acceleration is in excess of 0.18 g, the mercury in the roll
levelling switch is displaced to the end of the tube and so disconnects
the normal supply to its torque motor, i.e. it acts as an erection.cutout. The pitch switch, however, still responds to a forward tilt and
remains connected to its torque motor, and, as will be noted from the
diagram, it also connects a supply to the lower of the two additional
switches. Since the mercury in these switches is also displaced by the
acceleration, a circuit is completed from the lower switch to the roll
124
torque motor, which precesses the gyroscope axis to the right to
reduce the roll error. At the same time, the pitch switch completes a
circuit to its torque motor, which then precesses the gyroscope axis
rearward, so reducing the pitch error. Thus, during turns a constant
control is applied about both the pitch and roll axes by the pitch
levelling switch.
The changeover function of the additional switches depends on the
direction of gyroscope tilt in pitch. This is indicated by the broken
arrows in Fig. 4.19; the gyroscope and the pitch levelling switch now
being tilted rearward, the latter connects a supply to the upper
additional switch so that the direction of the supply to the roll torque
motor is changed, and the gyro5cope is precessed to the left. The
change in direction of the supply to the roll torque motor is also
dependent on the direction of turn, as a study of Fig. 4.19 will show.
Since the forces produced depend on an aircraft's speed and rate of
tum, then all erection errors will vary accordingly, thus making it
difficult to compensate for them under all conditions. It is usual,
therefore, particularly for instruments employing the inclined axis
method of compensation, to base compensation on standard values,
e.g. a rate l turn of 180°/min at a speed of 200 mph.
Direction indicator
This indicator was the first gyroscopic instrument to be introduced as
a heading indicator, and although for most aircraft currently in
service it has been superseded by remote-indicating compass systems
and flight diiector systems, there are still applications of it in its
pneumatically-operated form. The instrument employs a horizontalaxis 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. In addition,
of course, the direction indicator overcomes the effects of magnetic
dip, and of turning and acceleration error inherent in the magnetic
compass (see Chapter 3).
In its basic form, the outer ring of the gyroscope carries a circular
card, graduated in degrees, and referenced against a lubber line fixed
to the gyroscope frame. When the rotor is spinning, the gimbal
system and card arc stabilized so that, by turning the frame, the
number of degrees through which it is turning may be read on the
card.
The manner in which this simple principle is applied to practical
indicators varies between types, but we may consider the vacuumdriven version illustrated in Fig. 4.20(a) and (b), which is used in the
basic instrumentation of some types of small aircraft.
The rotor is enclosed in a case, or shroud, and supported in an
125
Figure 4.20 Direction
indicator.
0
0
0
0
SYNCHRONISER RING
0
(a)
(b)
(C)
inner ring which is mounted in an outer ring, the bearings of which
are located at the top and bottom of the indicator case. The front of
the case contains a cut-out through which the card is visible, and also
the lubber line reference.
When the vacuum system is in operation, the reduced pressure
created within the case allows surrounding air to enter through a
filtered inlet and to pass through channels in the gimbal rings to
emerge finally through jets. The air issuing from the jets impinges on
the rotor 'buckets', causing the rotor to rotate at speeds between
12 000 and 18 000 rev/min.
A caging and setting knob is provided at the front of the case to set
the indicator on the same heading as that of the magnetic compass.
When this knob is pushed in, a!l arm is lifted thereby locking the
inner ring at right angles to the outer ring, and at the same time
meshing gearing between the knob and the outer ring. Thus, a
heading can be set by rotating the knob and the whole gimbal system.
The reason for caging the inner ring is to prevent it from precessing
126
when the outer ring is rotated, and to ensure that, on uncaging, their
axes are mutually at right angles.
Control
or drift
Drift, as we have already learned (see page 102), is a fundamental
characteristic of a horizontal-axis type of gyroscope, and so for
practical direction indicating purposes, earth rate error, transport
wander, and real drift must be controlled. This is generally effected
by gimbal ring balancing and by erection devices.
Gimbal ring balancing
The method of controlling earth rate error is deliberately to unbalance
the inner ring so that a constant force and precession are applied to
the gimbal system. The imbalance is effected by a nut fastened to the
inner ring, and adjusted during initial calibration to apply sufficient
outer ring precession to cancel out the drift at the latitutde in which it
is calibrated. For all practical purposes, this adjustment is quite
effective up to 60° of latitude on the earth's surface.
Erection devices
These form part of the rotor air-drive system and are so arranged that
they sense misalignment of the rotor axis in terms of an unequal air
reaction. In the indicator already described, this is accomplished by
exhausting air over a wedge-shaped plate secured to the outer ring as
shown in Fig. 4.20(c).
In the normal horizontal position of the rotor axis, the air flowing
from the outlet of the casing is equally divided, and the reaction of
the air applies equal and opposite forces to the faces of the wedge.
When the rotor is tilted, the air outlet is no longer bisected by the
wedge; thus, the reaction forces are unbalanced, and if the greater
force is visualized as being applied to the rotor rim, then it and the
inner ring will be precessed until the forces are again equal and
opposite.
Gimbal errors
A definition of gimbal error has already been given (see page 106).
In the case of a direction indicator, errors are dependent upon: (i) the
angle of climb, descent, or roll; (ii) the angle between the rotor axis
and longitudinal axis of an aircraft. Fig, 4.21 illustrates the gimbal
system geometry When an aircraft is in particular attitudes.
At (a) an aircraft is represented as flying straight and level on an
easterly heading, and as the gimbal system geometry is such that the
127
Figure 4.21 Gimbal errors.
.FRAME
(cl
{el
rotor axis lies N-.S, the three axes of the system are mutually at
right angles, and the heading will be indicated without error. The
same would also be true if an aircraft were flying on a westerly
heading.
If an aircraft rolls to the left or right on either an easterly or
westerly heading, or executes a left or right turn, the outer gimbal
ring will be carried about the axis of the stabilized inner rin3
(diagram (b)). In this condition the cardinal headings, or changes of
heading during turns, would also be indicated without error.
At {c) an aircraft is assumed to be descending so that, in addition
to the outer gimbal ring being tilted forward about the rotor axis, the
inner ring also rotates, both rings maintaining the same relationship
to each other. Again, there is no gimbal error; this would also apply
in the case of a climbing attitude.
When an aircraft carries out a manoeuvre which combines changes
128
in roll and pitch attitudes, e.g. the banked descent shown at (d), the
outer ring is made to rotate about its own axis, thus introducing a
gimbal error causing the indicator to show a change of heading.
If an aircraft is flying on an intercardinal heading. the rotor axis
will be at some angle to the aircraft's longitudinal axis. as at (e), and
gimhalling errors will occur during turns. rolling in straight and level
flight, pitch attitude changes or combinations of these.
When the heading is such that the aircraft's longitudinal axis is
aligned with that of the gyroscope rotor. rolling of the aircraft on a
constant heading will not produce gimballing error because the
gimbal system also rotates about the rotor axis. If, however, rolling
is combined with a pitch attitude change. the effect is the same as the
combined manoeuvre noted earlier {diagram (d)).
Whenever the angular relationship between the gimbal rings is
disturbed during a manoeuvre. an indicator's erection device will be
attempting to re-erect the rotor into a new plane of rotation and will
cause false erection, ,he magnitude of which depends on how long
the erecting force is allowed to operate. i.e. on the duration of the
manoeuvre. The magnitude of the force itself will depend on the
angle of the rotor to the device. Thus. on completion of a manoeuvre
it is possible to have an error due to false erection. and during a
manoeuvre an error can be caused which is a combination or' toth
gimballing effect and false erection.
Turn-and-bank
indicator
This indicator contains two independent mechanisms: a
gyroscopically-controlled pointer mechanism for the detection and
indication of the rate at which an aircraft turns. and a mechanism for
the detection and indication of hank and/or slip. The dial presentation
of a typical indicator is shown in Fig. 4.22(a).
Rate gyroscope
For the detection of rates of turn. a rate gyroscope is used and is
arranged in the manner shown at (b) in Fig. 4.22. It differs in two
respects from the displacement gyrcscopes thus far described: it has
only one gimbal ring. and it has a calibrated spring connected
between the gimbal ring and cas;:1g to restrain movement about the
longitudinal axis YY 1, i.e. it is a single-axis gyroscope.
When the indicator is in its normal operating position the rotor spin
axis. due to the spring restraint. will always be horizontal and the
turn pointer will he at the zero datum mark. With the rotor spinning.
its rigidity will further ensure that the zero position is maintained.
Let us assume that the indicator is turned to the left about a
vertical input axis. The rigidity of the rotor will resist the turning
129
Figure 4. 22 Turn-and-bank/
slip indicator.
(a)
STRETCHING Of SPRING
fOR LEFT AND RIGHT TURNS
INPUT
AXIS
F
PRECESSION AXIS
y_...--
_.,
'"--.
x,
(b)
movement, which it detects as an equivalent force being applied to its
rim at point F. The gimbal ring and rotor will therefore be tilted
about the longitudinal axis as a result of precession at point P.
As the gimbal ring tilts, it stretches the calibrated spring until the
force it exerts prevents further deflection of the gimbal ring. Since
precession of a rate gyroscope is equal to its angular momentum and
the rate of turn, then the spring force is a measure of the rate of
turn. The actual movement of the gimbal ring from the zero position
can, therefore, be taken as the required measure of mm rate.
In practice, the gimbal ring deflection is generally not more than
6°, the reason for this being to reduce the error due to the rate of
turn component not being at right angles to the spin axis during
gimbal ring deflection
130
The rate of turn pointer is actuated by the gimbal ring and a
magnifying system which moves the pointer in the correct sense over
a scale calibrated in what are termed 'standard rates'. Although they
are not always marked on a scale, they are classified by the numbers
1 to 4 and i::orrespond to turn rates of 180°, 360°, 540° and 720°
per minute respectively. The marks at either side of zero of the
indicator scale shown in Fig. 4.22 correspond to a Rate 1 turn.
A system for damping out oscillations of the gyroscope is also
incorporated and is adjusted so that the turn pointer will respond to
fast rate of turn changes and at the same time respond to a definite
tum rate instantly.
It should be noted that a rate gyroscope requires no erecting device
or correction for random precession, for the simple reason that it is
always centred by the control spring. For this reason also, it is
unnecessary for the rotor to rotate at high speed, a typical speed
range being 4000-4500 rev/min. The most important factor in
connection with speed is that it must be maintained constant, since
precession of the rotor is directly proportional to its speed.
Bank indication
In addition to the primary indication of turn rate, it is also necessary
to have an indication that an aircraft is correctly banked for the
particular turn. A secondary indicating mechanism is therefore
provided which depends for its operation on the effect of gravitational
and centrifugal forces. A method commonly used for bank indication
is one utilizing a ball in a curved liquid-filled glass tube as illustrated
in Fig. 4.23.
Figure 4.23 Ball-type bank
indicating element. (a) Level
flight; (b) correctly banked;
(c) underbanked (skidding out
,,f tum); (d) overbanked
(slipping into the turn).
W
w
{a}
(bl
y
y
.
H---,c.F.
~-~
W ---- R
(cl
R
>
(d)
w
131
In normal level flight (diagram (a)) the ball is held at the centre of
the tube by the force of gravity. Let us assume now that the aircraft
turns to the left at a certain airspeed and bank angle as at diagram
(b). The indicator case and the tube move with the aircraft, of
course, and because of the turn, centrifugal force in addition to that
of gravity acts upon the ball and tends to displace it outwards from
the centre of the tube. However, when the turn is executed at the
correct bank angle and :natched with airspeed, then there is a
balanced condition between the two forces and so the resultant force
holds the ball at the centre of the tube as shown. If the airspeed were
to be increased during the turn, then the bank angle and centrifugal
force would also be increased, but so long as the bank angle is
correct for the appropriate conditions, the new resultant force will
still hold the ball at the centre of the tube.
If the bank angle for a particular rate of turn is not correct, say
underbanked as in diagram (c), then the aircraft will tend to skid out
of the turn. Centrifugal force will predominate under such conditions
and will displace the ball from its central position. When the turn is
overbanked, as at (d), the aircraft will tend to slip into the turn and
so the force due to gravity will n..>w have the predominant effect on
the ball. It will thus be displace(. from centre in the opposite
direction to that of an underbanked turn.
Typical indicator
The mechanism of a typical pneumatic type of indicator is shown in
Fig. 4.24. Air enters through a filtered inlet situated at the rear of
the case and passes through a jet from which it is directed onto the
rotor buckets. The direction of spin is in the direction of flight.
Adjustment of gyroscope sensitivity is provided by a screw attached
Figure 4. 24 Mechanism of a
pneumatic-type turn-and-bank
indicator.
9
10
132
to one end of the rate control spring. A stop is provided to limit
gimbal ring movement to an angle which causes slightly more than
full-scale deflection (left or right) of the rate of turn pointer.
A feature common to all indicators is damping of gimbal ring
movement to provide 'dead beat' indications. In this particular type,
the damping device is in the form of a piston, linked to the gimbal
ring, and moving in a cylinder or dashpot. As the piston moves in
the cylinder, air passes through a small bleed hole, the size of which
can be adjusted to provide the required degree of damping.
The slip indicator is of the ball and liquid-filled tube type, the
operation of which has already been described.
Turn coordinator
This instrument is a development of the turn-and-bank indicator, and
is adopted in lieu of this in a number of small types of aircraft. The
primary difference, other than the display presentation, is in the
setting of the precession or output axis of the rate gyroscope. This is
set at about 30° with respect to an aircraft's longitudinal axis, thus
making the gyroscope sensitive to rolling or banking as well as
turning. Since a turn is initiated by banking, then the gyroscope will
precess, and thereby move the aircraft symbol to indicate the
direction of the bank, enabling a pilot to anticipate the resulting turn.
The turn is then controlled to the required rate as indicated by the
alignment of the symbol with the graduations on the outer scale. In
the example illustrated in Fig. 4.25 the graduations correspond to a
Figure 4.25 Turn coordinator.
133
rate 2 (2-min) turn. Coordination of the turn is indicated by the balltype indicator remaining centred in the normal way (see page 13 l ).
In some indicators, a pendulous type of indicator may be adopted for
this purpose.
The gyroscope is a de motor operating at approximately 6000
rev/min; in some cases, a constant-frequency ac motor may be used.
The annotation 'no pitch information' on the indicator scale is given
to avoid confusion in pitch control which might result from the
similarity of the presentation to that of a gyro horizon.
Damping of the gyroscope may be effected by using a silicone fluid
or, as in the indicator illustrated, by a graphite plunger sliding in a
glass tube.
134
5 Synchronous
data-transmission
systems
The instruments that have been described thus far are very basic in
concept, in that the display of appropriate data is achieved by
mechanical-type elements which require either a direct connection to
remotely-located sensing elements, as in the case of pneumatic
airspeed indicators, altimeters and vertical speed indicators, or which
provide indications directly from their sensing elements, as in
compasses and gyroscopic instruments. While such instruments can
still satisfy a primary role requirement in the instrumentation of small
types of aircraft, their application to large aircraft is restricted by
increase in distances between the locations of sensing elements and
flight deck panels. In order, therefore, to overcome some of the
problems that can arise, e.g. having to run lengthy pipelines between
sensors and instruments, an electrical method of transmitting changes
in measured quantities is adopted. This is not new, of course, having
originated in the days when it became necessary to improve the
measuring accuracy of those instruments associated with the operating
parameters of engines and, in particular, their use in multi-installation
arrangements.
As part of the technique of 'remote indicating', synchronous datatransmission systems, or synchro systems as they are generically
known, were introduced. They consist of transmitting and receiving
elements which,
varying circuit configurations, are now utilized
not only in certain engine instruments, but also in analogue-type air
data computers, remote-indicating compasses, and in flight director
systems for heading and aircraft attitude sensing.
in
Categories of synchro
systems
Synchro systems are divided into four main categories as follows:
I. Torque This is the simplest form of synchro, in which torque is
derived solely from the input to its transmitting element; no
amplification of this torque takes place. Moderate torque only is
developed at the output shaft of the receiving element, and for this
reason the system is used for data-indicating purposes, e.g. oil or
135
fuel pressure, and for the indication of the position of mechanical
controlling devices, e.g. airflow control valves.
2. Control This type of synchro normally forms part of a
servomechanism to provide the requisite signals which, after
amplification, are used for the control of a drive motor.
3. Resolver This· is used where precise angular measurements are
required. It converts voltages, which represent the cartesian
coordinates of a point, into a shaft position and a voltage which
together represent the polar coordinates of the point. They can
also be used for conversion from polar to cartesian coordinates.
Typical applications are in analogue computers, remote-indicating
compasses, and flight director systems.
4. Differential This type of synchro is used where it is necessary to
detect and transmit error signals representative of two angular
positions, and in such a manner that the difference or the sum of
the angles can be indicated. They can be utilized in conjunction
with either torque, control or resolver synchro systems.
The foregoing synchros are designated by standardized
abbreviations and symbols as in Table 5.1.
Torque synchro system
This consists of a TX element and a TR element interconnected as
shown in Fig. 5.1. Both elements are electrically similar, each
consisting of a rotor carrying a single winding around a laminated
iron core, mounted co-axially within a stator core carrying windings
that are spaced 120° apart. The principal physical differences
between the two elements are that the TR is usually fitted with a
mechanical damper to reduce oscillation, and that the TX rotor is
mechanically rotated whereas the TR rotor is rotated by a magnetic
field produced by the TX.
Both rotors are supplied from a source of single-phase ac power via
sliprings and fine wire brushes, and the co~responding stator
connections are ~oined by transmission-lines to form a closed circuit.
When the axis of a rotor is aligned with that of the S 1 winding cf a
stator, a synchro is in what is termed the electrical zero position.
This serves as a datum when testing a synchro for accuracy.
When power is applied to the system, current flows in both rotor
windings and sets up an alternating magnetic flux. Since the rotor
windings and stator windings of the TX and TR correspond, in effect,
to the primary and secondary windings of a transformer, then the flux
induces an alternating voltage in the stator winding coils. The induced
voltages are at maximum, and in phase with the rotor voltages, in the
'electrical zero' position, and are zero when the rotors are at 90° to
this position. At 180°, the induced voltages are again maximum but
136
Table 5.1
Synchro type
Abbreviation
Torque:
Transmitter
Receiver
TX
TR
Comrol:
Transmitter
ex
Receiver
CT
Circuit
Symbol
~
..
S2
k
S2
""""'
m
Resolver
Differential (in a
torque system):
Transmitter
Receiver
RS
~
STATOR
S1
G
. r.irt
TDX
TDR
.
m~Ot
Ill
..
Differential (in a
control system):
Transmitter
Receiver
.,
CDX
CDR
out of phase with the rotor voltages. At 270°, the induced voltages
will again be zero. Due to the 120° spacing of the stator coils, the
voltages appearing between the connection points S 1, S2 and S3 will
be the sums or differences between the voltages induced in the stator
coils, depending on whether such voltages are in phase or 180° out
of phase.
When the TX and TR rotors are in the same angular positions, e.g.
the electrical zero position, the voltages induced in the stators are
equal and opposite, and so no current will flow in the stator coils. If,
however, the rotors occupy different angular positions, e.g. the TX
motor is moved through an angle of 30° while the TR rotor is at
'electrical zero', then an unbalanced condition between voltages
induced in the stator coils arises.
This imbalance causes current to flow through the closed circuit
between TX and TR stators. The magnitude and phase of the currents
are in proportion to that of the induced voltages. The currents are
greatest in the coil sections of stator windings when the voltage
137
Figun 5.1
Torque synchro
ElECTRICAl
ZERO
system.
I
S1
-
ELECTRICAL
ZERO
I
INPUT SHAFT
-
CURRENT
ROTOR FIELDS
==!>
STATOR f)ELDS
AC SUPPLY
imbalance is greatest; thus, with the TX rotor at 30°, the imbalance
is greatest in the section comprising coils S 1 and S3 • The greatest
current therefore flows through these coils to the corresponding ones
in the TR stator.
The current flow through the TX stator produces a resultant
magnetic field, and as with normal transformer action, this field and
the one produced by current flowing in the rotor winding must
always be in balance; the directions of the fields are, therefore, in
opposition.
The CUITent flowing through the TR stator also produces a resultant
magnetic field, but as the direction of current flow is opposite to that
through the TX stator, then the direction of the field will also be
opposite. The interaction of this field with that of the rotor will
develop a torque and thereby turn the rotor from 'electrical zero' to
the same position as that of the TX. As the rotor turns, the imbalance
of induced voltages decreases, and in turn the currents produced by
them also decrease. When the TR rotor synchronizes with the 30°
position of the TX rotor, its field will be in alignment with the
resultant field, its voltages will be balanced, and current no longer
flows between the stators. The system is then said to be at the 'null'
position.
The foregoing action takes place as a result of positioning the TX
138
rotor at any other angle; it is apparent, therefore, that if this rotor
were to be continuously rotated, then rotating magnetic fields would
be set up to provide synchronized and continuous rotation of the TR
rotor.
During operation, and because of the current flow in the stator
coils of the TX, a torque is also set up tending to tum its rotor out of
alignment. This torque, however, is overcome by the loads exerted
by the prime mover that actuates the rx.
In the event that the rotor and stator connections of a TX and TR
system are changed over from the normal symmetrical arrangement
indicated in Fig. 5.1, then different operating results will be
produced. The TR rotor can still move synchronously with the rotor,
but it can do so from a different reference position, or in the reverse
direction. The possible rearrangements are illustrated in Fig. 5.2.
Figure 5. 2 Interchange of TX
synchro connections.
(a) Symmetrical connections data coincide; (b) rotor
connections reversed - output
datum advanced 180°;
(c) cyclic shift of stator
connections - output datum
advanced 240°; (d) two stator
leads interchanged - output
rotor reverses direction of
rotation.
139
Control synchro system
A control synchro system differs from the one just described in that
its function is to produce an error voltage signal in the receiver
element, as opposed to the production of a rotor torque. Typical uses
of the system are in servomechanisms such as: altimeters and
Mach/airspeed indicators that operate in conjunction with air data
computers, and in the indicators of flight director systems.
The interconnection of the two elements of the system are shown in
Fig. 5.3. The transmitter is designated as CX, and the receiver as
CT, signifiying control transfonner. The CX is similar to a TX, and
from the diagram it will be noted that the single-phase ac power is
connected to its rotor only. The CT rotor is not energized since it
acts merely as an inductive winding for detecting the magnitude and
phase of error voltage signals which are supplied to an amplifier.
Other differences in a CT element are as follows:
(a) the rotor winding is on a cylindrical core, to ensure that the rotor
is not subjected to any torque when the magnetic field of the CT
stator is displaced.
(b) it operates as a single-phase transformer with the stator windings
acting as the primary, and the rotor winding as the secondary.
(c) the stator winding coils are of high impedance to limit the
alternating currents through them.
(d) it is at electrical zero when the rotor is at 90° as shown.
When the CX rotor is energized, then, as in the case of a TX,
voltages are induced in the stator windings to produce resultant
magnetic fields when the rotor is at electrical zero (see diagram (a))
The induced voltages are applied to the CT stator coils and the
alternating flux produced induces a voltage in the rotor. The
magnitude of this voltage depends on the relative position of the
rotor; with the rotor at its 'electrical zero' position of 90° as shown,
the induced voltage is zero.
If the CX rotor is now rotated clockwise from its electrical zero
position as in diagram (b), the resultant flux in the CT stator will be
displaced from its datum point by the same angle, and relative to the
rotor at that instant. An error voltage is therefore induced in the
rotor, and with the connections as shown, the magnitude of the
voltage increases from zero, .md is also in phase with the voltage
applied to the CX rotor. For an anti-clockwise rotation of the CX
rotor from electrical zero (as in diagram (c)), the error
induced in the CT rotor again increases in magnitude, but this time it
is in anti-phase with the voltage applied to the rotor.
As commonly used in servomechanisms, the error voltage signal
from a CT rotor is supplied to an anplifier which, in turn, supplies
its output to the control phase of an ac servomotor. The other phase
Figure 5.3 Control synchro
system.
hn ('\ n
Applied
voltage ~ T i m e
ac
0° error
supply
I
I
No output
voltage
S2
Time
AMPI.IFIED ERROil
VOLTAGE TD CONTROL
PHASE
I
Clockwise
error signal I
Time
In phase
with
I applied
I voltage
t
I
1Anticlockwise 1
error signal!
ac
I
supply
Time
In antiphase
with
applied
voltage
0
Anti-clockwise
rotation
......,.
Rotor fields
• • ..,._ Stator fields
(the reference phase) of this motor is continuously supplied with ac.
Since the control phase of a two-phase motor can either lead or lag
the reference phase voltage, then the phase of the error voltage will
determine the direction in which the motor will rotate, and its
magnitude will determine its speed of rotation.
141
The servomotor drives the mechanism being controlled, and in
addition it turns the CT rotor in the direction appropriate to that in
which the CX rotor has been turned, thereby reducing the error
voltage. At zero error voltage, the synchros are at 'null' and the
mechanism being controlled is at the new datum established by the
positioning of the CX rotor.
In some control synchro servomechanisms, the servomotor also
drives a tacho-generator which produces a feedback signal that is
supplied to the amplifier for the purpose of controlling the rate at
which the servomotor rotates.
Differential synchros
In some applications it is necessary to detect and transmit error
signals representative of two angular positions, and in such a manner
that the receiver element of a synchro system will indicate the
algebraic difference or the sum of the two angles. This is achieved by
introducing a differential synchro into either a torque or control
synchro system, and then using it as a transmitter. Unlike TX or CX
synchros, the rotor of a differential synchro also has three starconnected windings; the rotor core is of cylindrical shape. When
utilized in torque or control synchro systems, differential synchros
are designated as TDX and CDX respectively.
Figure 5.4 shows the three synchros comprising a TDX system,
their interconnection in this case being set up for detecting the
algebraic difference between two inputs. One input shaft controls the
angular position of the TX rotor, and the second input shaft controls
the angular position of the TDX rotor. Clockwise rotations of the
rotors are taken as positive, and anti-clockwise rotations as negative.
The TDX rotor windings are connected to the TR stator windings.
The TX and TDX rotors are at their electrical zero positions, and so
the resultant magnetic fields are as shown. The effects of the voltages
• induced in the TR stator will, therefore, produce resultant fields such
that its rotor will also be at electrical zero.
If now the TX rotor is rotated clockwise through 60° while the
TDX rotor remains at electrical zero as in diagram (a), the signals
generated in the stator of TX will be transmitted, unmodified, to the
TR stator windings and so the resultant fields will rotate its rotor
clockwise through 60°.
In diagram (b) the TX rotor is shown at the electrical zero position,
while the TDX rotor is rotated clockwise through 15°. The fields of
both synchros remain at electrical zero because their position is
determined solely by the orientation of the TX rotor. However, a 15°
clockwise rofation of the TDX rotor without a change in the position
of its field is equivalent to moving the rotor field 15° anti-clockwise
whilst leaving the rotor at electrical zero. This relative angular
142
TX
Figure 5.4 TDX synchro
TDX
TR
system.
V
INPUT SHAFT
OUTPUT SHAFT
INPUT SHAFT
I
SYNCHROS AT ELECTRICAL ZERO
AC SUPPLY
(al
~
\
•
ELECTRICAL ZERO
(bl
(c)
R1~9l
R2
52
R3 •
-
ROTOR FIELDS
=!> STATOR FIELDS
S3
CIRCUIT SYMBOL OF
DIFFERENTIAL SYNCHIIO
143
Figure
5.5 Algebraic addition.
AC SUPPLY
change is duplicated in the TR stator and so its rotor will aligrl itself
V(i~ the field, i.e. for a 15° clockwise rotation of the TDX rotor, the
TX rotor will rotate 15° anti-clockwise.
When the algebraic addition of two inputs is required, the TX and
TOX stator connections Si and S3 , and also the TDX rotor
connections R2 and R3 to the stator connections S2 and S3 of the TR,
are interchanged as shown in Fig. 5.5.
Figure 5.6 illustrates the effects produced by interchanging the
connections of a TDX system.
An alternative method of integrating two inputs into a single output
is to utilize a differential synchro as a receiver (TOR) in conjunction
with two TXs (one for each input) as shown in Fig. 5.7. The
interconnections are arranged to obtain the algebraic difference
between the two inputs.
As in the case of the basic TR synchro, the rotor of a TOR always
aligns itself so that its magnetic field coincides with that of the
receiver stator. Assume now that the rotor of TX (A) is displaced
60° clockwise and the rotor of TX (B) is displaced 15° from their
electrical zero positions. This displacement of the rotor of TR (A)
displaces the stator magnetic field of TDR 60° clockwise from the
electrical zero position. The magnetic field of the TDR rotor is
derived from TX (B), thus a 15° displacement of its stator magnetic
field also causes a 15° displacement of the field of the TDR rotor.
The consequent rotation of the rotor to align its magnetic field with
that of its stator is thus 45° clockwise, thereby indicating the
difference (60° - 15°) in the angular displacements of the two TX
rotors.
If the connections R2 and
of the TDR rotor and the connections
S2 and S3 are interchanged, the angular rotation of the rotor of TX
(B) from the electrical zero position gives rise to an equal and
opposite displacement of the TDR rotor field. The angular movement
of this rotor to align its magnetic field with that of its stator is now
144
Figure 5.6 Interchange of
TDX system connections.
TX
TDX
TFl
4$°CL0CKW1SE
O»TWO INPUT & OUTPUT LEADS INTERCHANGED, TR ANGLE= TX ANG!.!:+ TDX ANGLE
'
15"CI.OCKWISE
{,:)TWO INPUT I.EADS INTERCHANG!!D,l'R ANGU: •-(TX ANGLE +TDX ANGLE)
"
15°CLOCKW1SE
75° clockwise, thus indicating the algebraic sum of the two
to
the TDR.
In the same way that differential synchros can be used· in
synchro systems, so they can be used in systems utilizing control
synchros; the basic
is shown in Fig. 5.8.
As was
out at the
, resolver
are employed to convert
the cartesian
coordinates of a
into a shaft position and a voltage which
145
Figure 5. 7 TOR arrangement.
(TRANSMITTER TX)
TRANSMITTER (TX)
R,
R,
System at electrical zero position
TX rotor displacements from electrical zero position
-
ROTOR FIELDS
.....c, STATOR FIELDS
together represent the polar coordinates of a point. Let us now see
what is meant by these terms and how they are related to a voltage.
If a vector representing an alternating voltage is drawn, then, as
indicated in Fig. 5.9, it can be defined in terms of its length
(designated r) and also of the angle (} it makes with a horizontal axis
X; these are referred to as the polar coordinates. This same vector
can .also be defined in terms of x and y, where:
X
= r COS(}
and
y
=
r sin(}
These two expressions are the cartesian coordinates.
Unlike torque or control synchros, a resolver has four stator
windings and four rotor windings arranged as shown in Fig. 5. lO(a).
Stator windings S I and Si are in series and have a common axis
which is at right angles to that formed by S3 and S4 in series. A
similar arrangement applies to the rotor windings. For the purpose of
simpiification, arrangements are usually shown as at the right of the
diagram.
146
Figure 5.8 CDX synchro
system.
cox
ex
CT
51
ANGLE Of FIELD
PROPORTIONAL TO
e, - e,
..... ROTOR FIELDS
==Q> STATOR FIELDS
Figure 5. 9 Polar and cartesian
coordinates.
y
Figure 5. lO(b) illustrates a common application of a resolver
whereby polar coordinates are converted to cartesian. An alternating
voltage is applied to one of the .rotor windings and represents the
length r of the vector shown in Fig. 5.9. The other winding is
unused and is normally shorted-out to improve accuracy and to limit
spurious response.
The flux produced by the rotor current links with the stator
windings, and voltages are induced in them depending on the relative
147
Figure 5.10 Resolver synchro.
ROTOR
A1
k
51
C
JJlCl
R:)
w
INPUT SHAFT
(a)
0
---·----RI
FLUX
It ~
LA~SUP?LY
R2
1
I
I
I
I
I
I
I
l
I
I
~---~RCOS8
Anr':~
i
I
I
,-~'"'"'"'
1
I
I
I
I
_J
_____ 0
tJLj
M
I
51
v~~1M&1&°1~~~il R
fu0
I
[
,
STATOR
R SIN B
S3
I
I
8
(b)
position of the rotor. In the position shown in Fig. 5.10, maximum
voltage is induced in the stator coils aligned with the rotor windings
in use, i.e. S 1 and S2 aligned with R1 and R2 • No voltage is induced
in S3 and S4 which are at right angles to the rotor flux.
Movement of the rotor at a constant speed will therefore induce
sinusoidal voltages across the two stator coils; these variations will be
equal to r cos 8 in S 1 and S2 and to r sin 8 in S3 and S4 • The sum of
these two defines, in cartesian coordinates, the input voltage and
rotor shaft rotation, which together are the polar coordinates (r and i1
of Fig. 5.9).
Figure 5 .11 shows the arrangement in which cartesian coordinates
are converted to polar. An alternating voltage V, = r cos 8 is applied
to the cosine winding S 1 and Sz, and a voltage Vy = r sin 8 is
applied to the sine winding S3 and S4 • An alternating flux of
amplitude and direction dependent on these voltages and representing
the cartesian coordinates is, therefore, produced inside the stator.
One of the rotor windings R 1 and R2 is connected to an amplifier
and a servomotor which drives the output load, and also the rotor in
such a direction as to return the rotor to a 'null' position, i.e. at 90°
148
Figure 5. IJ Conversion of
cartesian lo polar coordinates.
f\1
,.-
..., ~
~
e
a
lClO( MECHANICAL DBIVE
to the stator flux; the motor then stops. Rotor winding R3 and R4
must now lie parallel to the stator field and has induced in it a
voltage proportional to the amplitude of the alternating flux; in other
words, proportional to the length of vector r (Fig. 5.9) and equal to
j(V; + V~). The shaft position then represents the angle fJ. Thus, the
input defined in cartesian coordinates has been converted to an output
in terms of polar coordinates.
Synchrotel
A synchrotel is generally used as a low-torque CT and is
interconnected with a CX synchro, but unlike the CT synchro applied
to the more conventional type of control system, it serves as a signal
transmitting element. The construction of a synchrotel, and its
interconnections, are shown schematically in Fig. 5.12. It employs a
stator core carrying three windings at 120° apart, but, as will be
noted, the rotor section differs from the conventional form of
construction in that: (i) the rotor, which is made of aluminium, is
hollow and of oblique section; (ii) the rotor rotates in an air gap
fonned between a fixed cylindrical core, the stator core, and the
single-phase rotor winding which is also fixed. The rotor shaft is
supported in jewelled bearings and concentric with the cylindrical
core, and is mechanically connected to the element whose position or
displacement is to be measured.
In a typical application, e.g. measurement of a fluid pressure, the
synchrotel rotor is connected to the appropriate pressure-sensing
element. The CT is located within a panel-mounted indicator,
together with a servo-amplifier and a servomotor. The CT rotor is
energized by a 26 Volts, 400 Hz single-phase ac supply which
induces voltages in its stator; since the stator is connected to the
149
'
Figure 5.12 Synchrotel.
150
v'___, ,.,. .
synchrotel stator then a resultant radial alternating flux is established
across it.
For any particular pressure applied to the sensing element there
will be a corresponding position of the synchrotel rotor and, due to
its oblique shape, sections of it will be cut by the radial stator flux.
Currents are thus produced in the rotor and, since it is pivoted
around the cylindrical core, an axial component of flux will be
created in the core. The core flux will also induce an alternating
voltage in the fixed rotor winding, and the amplitude of this voltage
will be a sinusoidal function of the relative positions of the rotor and
stator flux. This voltage is fed, via the servo-amplifier, to the control
phase of the two-phase servomotor which drives the CT rotor round
in its stator, thereby causing a change in synchrotel stator flux, to the
point where no voltage is induced in the rotor winding, i.e. the CT is
driven to the 'null' position. This position corresponds to the pressure
measured by the sensing unit at that instant, and since the servomotor
also drives the indicator pointer, then this will also be positioned to
register the pressure on the indicator scale.
Synchros and
electronic display
systems
The circuits of synchro systems, irrespective of their type and data
measurement application, are of the linear or analogue type, i.e. their
output signals vary continuously as a given function of the input. This
does not, however, preclude their use in conjunction with electronic
display systems whose circuits process all incoming data in a binary
digit and coded signal format. Operating details of the devices used
for converting analogue output signals from synchros into digital
format will be given in later chapters.
151
6
Instruments and certain associated integrated systems utilizing signals
from digital computers have been in use for a very long time, and
they have operated alongside, and been integrated with, those systems
dependent upon outputs from the earlier types of analog computer.
For example, the digital computer of an inertial navigation system
(INS) can be supplied with altitude and airspeed data signals from an
analog type of air data computer (ADC). In such applications. signal
interfacing devices referred to as analog-to-digital (AID) and digitalto-analog (DIA) converters are also used. The AID conversion is also
necessary in applications whereby analog sensors are retained for the
supply of signal outputs to digital computers.
In the mid- to late- '70s, the operational requirements for aircraft
became more demanding and, in consequence, design concepts
underwent drastic changes which were to pave the way for greater
automation of in-flight management of aircraft and their systems. The
integration of systems which, of necessity, were to assume higher
levels of importance demanded greater capacity for processing of
their data outputs, and a faster means of data transfer; the application
of mixed computer technology was, therefore, no longer acceptable.
Thus 'new technology' types of aircraft have come into commercial
service, each having a large number of digital computers operating
within what is still an analog environment, and distributing their data
in binary-coded format via a 'data highway' bus system.
computer
fundamentals
152
A digital computer is essentially a device which uses circuits that
respond to, and produces signals of. two values: namely. logic high
or binary l, and logic low or binary 0. It is capable of performing
operations on data represented as a series of discrete impulses
arranged in the binary-coded or 'bit" format. Figure 6.1 illustrates
what is generally termed the organization (sometimes architecture) of
the principal hardware elements of a computer. The central processor
unit (CPU) executes the individual machine instructions which make
up the computer program.
The binary-coded format of the program consists of an operation
code which tells the computer what operation it is to start with next,
Figure 6. I Computer
organization.
Central
processor
unit
Address bus
Data bus
Control bus
Fig11r,' 6.2 Typical CPU
Data bus
\)r_ganization.
Control
signals
lnslruction
decoding &
CPU control
Instruction
registers
Arithmetic
loglc unit
CPU
control
Address
bus
and an operand which is the data to be operated on. The program,
together with procedures and associated documentation, form what is
termed the software.
The CPU contains a number of registers or temporary storage units
which can each store a single byte or word, an arithmetic logic unit
(ALU) which performs the binary arithmetic and logic functions
associated with data manipulation, and a timing and control section
for co-ordinating CPU internal operation so that fetching and
execution of the instructions specified by a program is performed.
The typical organization of the CPU is shown in Fig. 6.2.
Communication between the CPU and memory, and the
input/output ports, is by means of a computer highway consisting of
three separate busses: the data bus, address bus and control bus. The
term 'bus' signifies a group of conductors carrying one 'bit' per
conductor, and is represented on diagrams by a broad arrow
identified by function.
The data bus carries the data associated with a memory or
input/output transfer, and the number of lines constituting the bus is
the same as the number of bits (e.g. eight) in the CPU's word length.
153
The bus is usually bi-directional, i.e. the CPU can write data to be
read by a memory, or it can read data from the bus presented by the
memory. Thus, data transfer between the two can be effected over a
single set of data lines. AH information transferred under program
control travels on the data bus via the CPU.
The address bus specifies the memory locations or input/output
ports involved in a transfer. The number of bits constituting this bus
has no direct relationship to the data bus word-length and depends on
the operation being performed; for example, at the beginning of an
instruction cycle the CPU must supply the address of the next
instruction in sequence to be fetched from the memory. During the
execution of the instruction, data may then be required to move
between the CPU and either the memory or an input/output port. If
this is the case, then the data memory address or input/output address
must be placed on the address bus by the CPU. Typically, the bus
contains 16 lines, and so gives the CPU the capability of addressing
up to 2 16 or 65 536 individual locations.
The control bus is also bi-directional since, being made up of
individual control lines for CPU memory and input/output control, it
synchronizes the transfer of 'read-out' and/or 'write-in' data along the
data bus.
Memory
This consists of a number of storage locations for instruction words
whose bit patterns define specific functions to be performed, and for
data words to be used for carrying out the operations specified by the
instruction words. Each memory word is given a numbered location
or address which is itself a binary word. There are two types of
memory:
Random-Access Memory (RAM) in which stored data at any
location can be changed by 'writing in' new data at that location.
It can therefore also be called a read/write memory.
Read-Only Memory (ROM) in which the binary information it
contains is permanently stored in it. The data, which can be
accessed in random fashion, are written in at the time of
manufacture, and so the specific program cannot usually be
changed afterwards.
In the organization of some digital computers, the two types of
memory are used together.
Memories are also classified as: volatile, i.e. one which loses its
stored data when the power supply is switched off, or non-volatile,
i.e. one that retains stored data even though the power supply is off.
154
Capacity and addressable locations
The capacity of a memory relates to 'bit storage' and is quoted in
kilobits (K); the prefix 'kilo'· does not stand for 1000 in the usual
sense, b;,it for 2 10 or !024. Thus, 8K = 8 x I 024 = 8192 bits
capacity.
The number of addressable locations in a memory is dependent on
its number of input/output data lines, and is derived from the uit
storage capacity divided by the number of data lines. This is because
each address location generally contains as many bits as it can pass
through the data bus. If, for example, a l K memory has only one
data line, it will have 1024 separate addressable locations, but with
four data lines it can only be addressed at 256 locations. The number
of lines are decided by design and specified in the appropriate
manufacturer's data sheets.
Input/Output (I/0) ports
These form the interface between a computer and the sources of input
data and subsequent output data, and are generally under the control
of the CPU. Special I/0 instructions are used to transfer data into
and out of the computer.
More sophisticated I/0 units can recognize signals from extra
peripheral devices called interrupts that can change the operating
sequence of the program. In some cases. units permit direct
communications between the memory and an external peripheral
device without interference from the CPU: such a function is called
direct memory access (DMA).
Computer languages
In the same way that we humans communicate with each other
through language, so a digital computer must use a language of one
sort or another to carry out its functions. There is, however, a big
difference between us and the computer in that when we are, say.
given an instruction to do something, the understanding of our own
language enables us to understand the instruction directly, and apart
from acting upon it, no other conversion is required. This is not true
for a computer, because when we want to give it an instruction a
conversion from our language into the binary-coded language must
first be carried out. The digital code is called the machine code, and
if instructions for the computer can be programmed directly in this
code, the program is written in machine language. and overall it is
called a machine language program.
The task of converting to machine code is usually delegated to the
computer, which follows what is called an assembler program telling
the computer what to do. The choice of instruction can be made by
155
selecting a mnemonic which is an abbreviation of what the instruction
does. This programming with mnemonic instructions is called
assembly language programming because, after the sequence is
written, it is fed into the assembler program which makes the
conversion to machine code and assembles it into the memory in the
proper order.
Such mnemonic codes are still unlike the human language, and so a
higher-level language-programming concept can be adopted in which
instructions are written in a problem-oriented or procedure-oriented
notation with each statement corresponding to several machine code
instructions. The conversion of the statements tq machine code is
done by a more involved computer program called a compiler. The
easier the programming is made by bringing the machine language
closer to the human language, the more complex is the computer
program needed to convert the machine language statements in
machine code. Once the conversion is available, however, it can be
used over and over again as necessary.
Data conversion
The majority of data to be processed by digital computers are, in the
first instance, in analog form, and so in order for the computers to
carry out their interpreting function, the data must be converted to a
binary-coded format. In many cases it is also necessary for data to be
converted from digital to analog format. The convel'6ion devices used
for such purposes are of the integrated logic circuit type, and
respectively they perform encoding and decoding functions as shown
graphically in Fig. 6.3.
The ideal AID converter ))as a 'staircase' transfer characteristic,
with the analog input quantized into a number of levels corresponding
to the number of 'bits' resolution. The true analog value
corresponding to a given output code is centred between two decision
levels. For the ideal DIA converter, there is a one-to-one
correspondence between input and output.
Data transfer
The transfer of data between the individual computer systems of an
aircraft is a necessary feature of in-flight operation. For example,
under automatically-controlled flight conditions, an automatic flight
control system operates in conjunction with an INS, ADC, FDS
(flight director system) and radio navigation systems, and all these
involve an exchange of data to provide appropriate commands to the
control system.
When conventional techniques are used to interconnect all the units
comprising individual systems, the extent of the cabling required is
156
Figure 6.3 Data conversion.
(a) Analog-digital;
(b) digital-analog.
111
,,
110
101
""'0
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100
Code centre
point
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3
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(b)
Binary input code
considerable, particularly as individual wires must transfer signals
dedicated to each of the parameters being monitored. In maximizing
the utilization of digital computer-based systems, therefore, it became
necessary to adopt an alternative method by which the exchange of
information could be effected by a network of single data busses,
known as a data highway, within an aircraft. In other words, this is
an adaptation of the data highway concept that is utilized within
digital computers themselves.
Each data bus consists of shielded and twisted pairs of wires, and
the voltage difference between them encodes a binary O or bin~ry I.
157
All outgoing encoded data from the computers are identified by an
additional binary-coded word called a label. The label takes up the
first eight bits of each word and is octal-coded, i.e. coded to the
base 8.
The designation of labels to particular functions is arranged by an
aircraft manufacturer in relation to each of the specific systems
installed in the type of aircraft concerned, and in accordance with
standard specifications. A specification accepted as an air transport
industry standards reference for the transfer of digital data is known
as ARINC 429 (ARINC is the abbreviated name of a US organization
'Aeronautical Radio Incorporated'). As separate bus systems are
predictable for the different classes of aircraft systems, ARINC 429
includes some duplication of labels where it is known that the use of
a common label on the same bus for two different purposes will
occur. For example, label 315 8 defines 'wind shear' for navigational
purposes, but for flight control systems the same label defines
'stabilizer position' .
Systems providing data outputs (referred to as transmitters) each
have their own data bus connecting them to the 'receiver' systems in
need of the data, as shown in Fig. 6.4. The shielding of the wires
comprising each data bus is connected to ground and, in particular, at
each branch to receivers. The maximum number of receivers that can
be connected to the same bus line is 20.
The digital computers of the different aircraft systems process data
in the form of specific messages or parallel binary words. The
messages are converted and transmitted in serial form, the reason for
this being that weight of transmission lines is reduced, and also
reliability is improved. The serial messages are then adapted into
high- and low-voltage levels, and transmitted along the data bus lines
in the form of strings of pulses. These comprise the word strings of a
message and correspond to those appropriate to all of the systems
detailed in the ARINC 429 specification. Each word is formed of 32
bits, each bit being either a binary l or a binary 0. As noted earlier,
the first eight bits comprise the label which identifies the source of
the message; the remaining bits are assigned to data, parity, sign and
status or validity. Two examples of a message are also shown in Fig.
6.4; one is labelled 'DME distance' and the other 'Radio Altitude'.
In each case, bits 9 and W are assigned to what is termed a Series
Destination Identifier (SDI); fois applies when specific words need to
be directed to a specific system of a multi-system installation, or
when the source system needs to be recognized from the word
content. In the examples indicar~d the systems are, of course, the
DME and Radio Altimeter respectively.
The bits 11 to 29 are those assigned to the actual data being
transmitted which, in the case of our examples, are distance in
nautical miles, and radio altitude in foet. The groups of binary ls and
TO OTHER
RECEIVERS
Figure 6.4 Data transfer.
Para!lel/
serial
conversion
(shift register)
t
Level
adapting &
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line
transmitter
converting
serial to
parallel
Serial messsga
RECEIVER
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Data = 2450.5 ft
'
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OME distan~e
Label
radio altitude
I
Os (in bits 11 to 29) correspond to equivalent decimal numbers,
which indicate that the DME system computer .is transmitting encoded
data corresponding to a distance of 257 .86 nautical miles, while the
Radio Altimeter computer is transmitting data. corresponding to an
altitude of 2450.5 feet.
Bits 30 and 31 are assigned to what is tef'l'led the Sign/Status
159
Figure 6.5 Refreshment rate.
4
50 mtsec
"'I
25 m/sec
25 m/sec
I
i-0-~- ~-
-0-·
:25 m/sec
25 m/sec
Matrix (SSM), which refers to plus, minus, north, south, left, right,
etc. of binary-coded decimal numeric data. They also refer to the
validity of data, and failure warning.
The detection of errors in codes and their correction is a very
important aspect in the transmission of digital data, and for this
purpose a parity check method is provided whereby a computer can
test whether bits. in a binary word have been accidentally changed
during transmission. The test is done by automatic summation of the
bits comprising a word to determine whether the total number is odd
or even, and by calculating what is termed a parity bit: this forms the
last bit of a word, i.e. bit 32. If, for example, there is an odd
number of binary ls among the first 31 bits, the parity bit is set to
'1' to make the word of 'even parity'. 'Odd parity' can also be used
where the parity bit is set to binary O to make the total number of
binary 1s odd. This latter form of parity is adopted in the ARINC
429 specification.
Data are transmitted in batches at a specified repetition or
refreshment rate along the appropriate busses, and at either high
speed (100 kilobits/sec) or low speed (12-14.5 kilobits/sec)
according to the frequency at which interfacing systems require an
update of information. This is shown in Fig. 6.5.
A standards specification of comparatively recent origin is the
ARINC 629. It relates to a data bus system called Digital
Autonomous Terminal Access Communications DATAC, conceived
by Boeing for use in the B777. Unlike ARINC 429 it is a two-way
bus requiring fewer wires and having a very much faster data transfer
rate.
7 Air data computers
The term 'air data', as we learned from Chapter 2, relates to the
sensing and transmission of pitot and static pressures to indicators
which, on the basis of physical laws, are specifically designed to
measure such pressures in terms of airspeed, altitude and rate of
altitude change. In addition to these three indicators, however, there
are many other systems whose operation depends on an air data
input. The utilization of such systems in an aircraft does, in turn,
depend on its size and operational category.
Although it would not be impossible to connect these systems to
pressure probes and/or vents by pipelines, then, as may be imagined,
the amount of 'plumbing' required would have to be considerably
increased. Apart -from causing additional weight problems, there
could be others associated with maintenance. In order 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 tc-msmits
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 c·hanges, and
compressibility effects can be automatically applied; in addition,
provision can also be made for the calculation of true airspeed (TAS)
from air temperature data inputs. The modulator arrangement of an
ADC, its associated indicators, and details of systems that utilize air
data inputs are shown in Fig. 7. 1
An ADC may either be of the analogue type, or of the type which
processes and transmits data in digital signal format. The latter type
is now more widely used, but as analog computers are still adopted
in some types of aircraft we can, at this stage, and by way of
introduction to ADC operating principles overall, consider a typical
analog arrangement.
Analog ADC
The arrangement of the basic modules of this type of computer and
their interfacing is shown in Fig. 7.2. Each module constitutes what
is termed a servomechanism, and is comprised of certain mechanical
elements, and synchros which perform the various functions already
described in Chapter 5. The output signals from each module are
161
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Gust response
Electronic engine contr
Cabin pressurization
Electronic units of fligh
control systems, e.g. fl
slats, stabilizer trim. ya
damper
SYSTEMS UTILIZING AIR DATA INPUTS
--- - --- - --
Flight diredor
Automatic flight control
h~ertial navigatfon
Altitude reporting
Ground proximity warning
Flight management
Flight recorder
Stall warning
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transmitted to their relevant indicators which, as we shall see later in
this chapter, are of the servo-operated type.
Transducers
The pipelines from an aircraft's pitot pressure probe and static vent
system are connected vi1 a manifold in the computer mounting to
pressure transducers in f1e computed airspeed and altitude modules.
The transducers are of the electro-mechanical type, the constructional
features of which vary dependent on those adopted by any one
manufacturer. One example we may consider is known as a forcebalance transducer which, as can be seen from the schematic diagram
in Fig. 7.3, consists of a capsule-type pressure sensor that actuates an
'E' and 'I' bar pick-off element.
The 'E' bar has an ac-powered primary input winding on its centre
limb, and a secondary output winding on each of its outer limbs;
these windings are connected to an amplifier. The 'I' bar is
mechanically connected to the capsule, the displacements of which
pivot the bar such that the gaps between its ends and the outer limbs
of the 'E' bar are increased or decreased. The 'I' bar is also
interconnected with a servomotor via a torsion bar, gear train and a
cam follower; the servomotor also forms part of the synchro system
appropriate to the computed airspeed and altitude modules.
In the static condition, i.e. a capsule is not subjected to a pressure
change, the gaps between the ends of the T bar and outer limbs of
Figure 7. 3
transducer.
Force-balance
AMPLIFIER
PRESSURE FUNCTION
CAM
PRESSURE CAM
FOLLOWER
164
TORSl(>N
BAR
CAPSULE
the 'E' bar are equal. When ac is applied to the primary winding
then magnetic fields will be produced which, in the static condition,
will be equal and opposite; thus, no signals will be induced in the
secondary coils.
When a pressure change occurs, the capsule responds accordingly
and the force it produces displaces the 'I' bar so that one air gap
increases and the other decreases. The resulting changes in the
magnetic fields cause out-of-balance signals to be induced in the
secondary coils. After amplification, the signals are applied to the
control phase of the servomotor, which then drives the cam follower,
and torsion bar, to produce opposing torsional effects which start
balancing the force exerted by the capsule, to 'back-off the signals
induced in the secondary coils. Whe'fl a constant pressure condition is
attained, equilibrium between capsule force and torsion is established,
and no further amplified signals are supplied to the servomotor.
In some types of ADC, the pressure transducers take the form of a
solid-state circuit device which utilizes what is termed the
piezoelectric* effect, i.e. the generation of electrical signals by
certain crystalline materials when subjected to pressure. The device
consists of quartz disks with a metallic pattern deposited on them,
and arranged in a thin stack such that they serve as a flexible
diaphragm. Thus, when subjected to pressure changes, the resultant
flexing sets up an electrical polarization in the disks so that electrical
charges are produced. The polarity of the charges depends on the
direction of flexing, in other words, on whether the pressure applied
is increasing or decreasing. All ouptut signals are supplied to the
appropriate type of transmission link adopted for the airspeed and
altitude modules.
Let us now refer once again to Fig. 7.2, in order to see how
pressure transducer output signals are processed and transmitted by
the various· modules for the purpose of operating their associated
indicators.
Module operation
In the case of the computed airspeed module, the servomotor, in
response to the amplified output signals from the transducer, drives
the rotor of a CX synchro whose stator is connected to a CT synchro
within the indicator.
The servomotor also drives, via differential gearing, the rotor of an
RS that forms part of a static source error correction (SSEC) network
which, as shown in Fig. 7 .2, originates in the Mach module of the
computer. The circuit of this network is pre-adjusted so that the
signal input to the RS is a correction factor signal corresponding to
the position error (PE) of the aircraft (see also page 34) in which
* From the Greek piezein,
meaning 'to press'.
165
the ADC is installed. The output signals from the RS are supplied to
the pressure transducer circuit so that its output, which is a measure
of the pressure difference p, -p,, is in turn also corrected. Thus, the
servomotor and CX synchro rotor position are controlled to produce
an output compensated for PE as a function of Mach number.
In operation, the servomotor also drives a tachogenerator which
supplies rate feedback signals to the control amplifier to reduce the
input error voltage signals. and thereby prevent the motor from
·overshooting' its controlled positions.
The altitude module is comprised of a servomechanism
arrangement, whose only difference from the one just described is
that it operates in response to signals which are a measure of the
pressure p,. In addition to supplying signals to a servo-operated
altimeter, the module also determines rates of altitude change. i.e.
vertical speed (V /S). and produces the corresponding signals. Since
the rate of change involves a time factor, the measurement of V/S is
accomplished by supplying the rate signals produced by the
servomotor-driven tachogenerator to an integrating amplifier. This is
a device that performs the mathematical operation of integration so
that its output is substantially the integral with respect to time of the
input to the device. After integration, the signals are amplified and
supplied to a servo-operated VSI and/or to V/S mode select modules
which form part of the pitch channels of automatic flighfcontrol and
tlight director systems.
An indication of speed in terms of Mach number can. as we
learned from Chapter 2 (see page 46). be derived by measuring it in
terms of the pressure ratio p, ~ p/p, In the case of basic
pneumatically-operated indicators [his, as we also learned.
necessitates that altitude and speed measuring elements be used in
combination with a mechanism that will perform the required dividing
function. Fundamentally. this arrangement also applies to the Mach
module of an ADC. but in adopting synchronous transmission and
servomechanism methods of accurately measuring the three
parameters involved, it is incumbent to use an equally accurate
method of performing the dividing function. In the example of ADC
shown in Fig. 7.2, the dividing is done by means of a differential
synchro in combination with a torque synchro system.
The differential synchro (TDX) is part of the computed airspeed
module servomechanism. the TX synchro is in that of the altitude
module, while the TR synchro is part of the Mach module
servomechanism.
When the altitude and computed airspeecl modules are in operation,
the TX synchro rotor will be driven to some angular position within
its stator corresponding to the pressure sensed by the altitude module
transducer. The signals induced in the stator will be of a related
value. and these are transmitted to the TDX. In response to the
166
signals produced by the transducer of the computed airspeed module,
the TDX rotor will also be at some corresponding angular position
within its stator. Since the angular positions of the TX and TDX
rotors are different, then, by virtue of the connection arrangements
between the two synchros, the output signals from the TDX are the
difference between those produced by its rotor and the TX synchro,
and in terms of the required pressure ratio.
The signals are transmitted to the control amplifier in the Mach
module via the TR synchro. The servomechanism arrangement of this
module is the same as that of the computed airspeed and altitude
modules. The CX synchro transmits signals in terms of Mach number
to a digital counter which may be individually mounted on a panel or
combined with a Mach/airspeed indicator. The 'nulling Ollt' of signals
under constant speed and altitude conditions is obtained by driving
the TR synchro rotor from the servomotor.
For the measurement of true airspeed (T AS) it is necessary to
utilize signals that are a measure of total air temperature (TAT).
These signals are generated by externaliy mounted sensing probes
(see page 62) and, in addition to an independent indicator, they are
also transmitted to the TAS module of the ADC via a potentiometric
network in the Mach module as shown in Fig. 7.2. This network
serves as a function generator in that it produces TAS output signals
that correspond to the values of a specified function of independent
variable inputs, in this case TAT and Mach speed. The output signals
are supplied to drive and control amplifiers for the operation of a
servomechanism consisting of a motor and CX synchro, the output
from which is supplied to an independent T AS indicator. The
servomotor also drives a 'follow-up' device which provides a signal
to the drive amplifier for the purpose of balancing out incoming T AS
signals.
Failure warning
Each module of the ADC incorporates a warning logic circuit
network which activates a warning flag in the associated indicators in
the event of loss of the respective data signals. Annunciator lights
corresponding to each module are provided on the end panel of the
computer, and are also illuminated in the event of failures. Once a
warning circuit has been triggered it remains latched.
Indicators
The indicators that are used in conjunction with an ADC of the
analog type just described also contain servomechanisms, and when
connected to the computer they each form a complete servo loop with
the respective modules of the computer. These indicators may, in
167
some applications, be of the combined pneumatic and servo type, as
for example the airspeed indicator shown in F_ig. 2.16 of Chapter 2,
or they may be entirely servo-operated.
Airspeed indicators
In the case of the indicator referred to above, its indicated (IAS) and
maximum operating speed pointers are operated by pressure-sensing
capsules within the indicator, while a servomechanism is used for
driving a digital counter for the display of computed airspeed. The
servomechanism, which is illustrated in Fig. 7.4, operates in response
to the signals supplied to its CT synchro by the relevant module of
the ADC (see also Fig. 7 .2).
A failure monitor circuit is also incorporated in the indicator and
comes into operation in the event of loss of power, or data signal
input from the ADC, and also if excessive 'nulling' occurs in the
digital counter servo loop. The circuit controls a solenoid-operated
flag such that it obscures the digital counter display. A check on flag
operation can be carried out by moving a computed airspeed switch
(see also Fig. 2.21) to its 'off position, thereby isolating the
excitation circuit of the CT synchro.
This indicator is also used in conjunction with an autothrottle
system, the purpose of which is to adjust the power settings of
engines in order to acquire, and then maintain, a commanded
airspeed. The system is also integrated with an aircraft's automatic
flight control system (AFCS). Airspeed commands may be selected
either from the AFCS mode select panel, or by a command set knob
in the airspeed indicator.
Figure 7.4 Servo-operated
airspeed indicator.
'
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-
-
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MONITOR
FAILURE MONITOR
.:....;.v_oc""------------------'
211
FAILURE WARNING FLA0_...,
SIGNAL IFAOM CENTA.Al
AIR OATA COMPUTEAJ
168
FiguT'I! 7.5 Command airspeed
circuit arrangement.
AIRSPEED
ERROR
IOUTl'UT TO
AUTO THROTTLE
COMPUTER!
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oc
The command airspeed circuit arrangement within the indicator is
shown in Fig. 7 .5, and from this it will be noted that it consists of a
CT synchro system, and a synchrotel mechanically connected with the
command speed set knob, a reference marker and a command speed
indicator.
Under normal operating conditions of the autothrottle system and
the AFCS, command airspeeds are set on a digital counter display on
the AFCS mode select panel. This setting also positions a CX
synchro rotor so that it can transmit equivalent signals to the indicator
servomechanism for the purpose of positioning the speed reference
marker and command speed indicator. In order that it may do so,
however, the clutch in the drive must, of course, be disengaged by
pulling out the command set knob; at the same time, a switch in the
CT synchro excitation circuit is held in the closed position.
The servomotor is also mechanically coupled to the synchrotel
transmitter, which differs from that described in Chapter 5 in that its
stator can also be rotated. The rotor is mechanically positioned within
the stator by the indicated airspeed pointer mechanism. The relative
positions of the two therefore produce an error signal ouptut
representing the difference between indicated airspeed and
commanded airspeed at any one instant. This output is then supplied
to the autothrottle system computer which then causes the power
output of the engines to be automatically adjusted to attain the
commanded speed.
Fifptre 7. 6
Mach speed
indicator.
MACH
Figure 7. 7 Servo-operated
Mach/airspeed indicator.
MANUAL MOOE FLAG - MAX AIRSPHO POINTER
MACH READOUT
SPEED REFERENCE
MARKER
------- EXTERNAL INDEX
MARKER (TYPICAL)
AIRSPEED FAILURE FLAG
AIRSPEED POINTER _..AIRSPEED READOUT
SPHO REFERENCE KNOB /
If it is required to set a command airspeed on the indicator itself,
the set knob must be pushed in to engage the servomechanism drive
dutch. This action also opens the switch in the excitation circuit of
the CT synchro, thereby isolating it from the servomotor. Rotation of
the set knob now provides for manual positioning of the reference
marker and synchrotel stator and, therefore, manual control of the
output signals to the autothrottle system computer. The command
speed digital counter is also rotated, but its display is obscured by a
yeltow 'MAN' flag, the solenoid control circuit of which is also
isolated when the set knob is pushed in. In the event that a command
airspeed exceeding a certain value (in this example 250 knots) is set,
a black flag is triggered to obscure the counter display.
Figure 7 .6 illustrates a display presentation of a Mach speed
indicator that is used in conjunction with the indicated/computed
airspeed indicator just described. The digital counter is servo-operated
by a CT synchro supplied with input signals from the Mach module
of the ADC.
The display presentation of a pure servo-operated indicator
(referred to as a Mach/airspeed indicator) is shown in Fig. 7.7; it
may be used in conjunction with an ADC of either the analog or
170
digital type. Computed airspeed is displayed in knots by a distinctlyshaped pointer and by a digital counter. The indication of speed in
terms of Mach number is shown by a digital counter display. A
striped pointer, which is also servo-driven, provides an indication of
V'"" and Mm 0 (see page 42).
The speed reference knob and marker perform the same function as
that of the indicated/computed airspeed indicator described earlier.
The other markers are 'memory bugs' that are pre-set to indicate
certain operating speeds appropriate to the type of aircraft, e.g. takeoff speeds, flap extension speed.
Five warning and indicating flags are provided as follows:
I. airspeed flag to indicate a failure in the airspeed circuit within the
indicator or ADC; 2. Mach flag to indicate failure in the Mach
circuits; 3. v,,IO flag to indicate failure in the Vmv and Mmo circuits; 4.
'INOP' flag that comes into view to indicate that the speed reference
marker is inoperative; and 5. 'M' flag that operates in conjunction
with the 'INOP' flag to indicate manual setting of the speed reference
marker.
The internal circuit arrangement of the indicator is shown in Fig.
7 .8. Power requirements are 26 V ac for synchro operation, and this
is distributed within the indicator via a power supply module. The
module also supplies 28 V de for the operation of servomotors,
amplifiers, flag monitor circuits, etc.
Signals corresponding to computed airspeed are supplied from the
relevant module in the ADC to a CT synchro, the error signal output
from which is amplified to drive the servomotor connected to the
airspeed pointer and digital counter. At the same time, it drives the
synchro rotor to 'null out' the error signal. The servomotor also
drives, through 2: l gearing, a potentiometer which supplies a de
signal to an 'anti-ambiguity' circuit connected to the servo amplifier.
The purpose of the circuit is to ensure that the airspeed pointer is not
driven to a position 180° out with respect to 'null'.
The airspeed servomotor also drives a synchro transmission loop,
the purpose of which is to transmit computed airspeed to an
autothrottle system.
The airspeed pointer and counter drive mechanism also incorporates
a specially calibrated cam and follower to provide square-law
compensation (see page 43). As the cam rotates it varies the
magnification rate of the pointer movement so as to maintain linearity
as speed increases.
The maximum airspeed pointer is driven by a servomotor which
receives its signals from a V'"" and overspeed processor circuit
module via an amplifier. The· pointer is always driven to, and
'nulled' out at, a scale reading higher than that of the airspeed
pointer, by signals from a synchro whose rotor is also driven by the
servomotor. If airspeed is increased to the maximum value, the
171
Figure 7.8 Mach/airspeed
indica1or circuit.
--------lJ!l§.P~Q9Q.U~~~Q.R_!YE
__.,.. _________
.,.
M
I
AMBIGUITY
CIRCUIT
__:._ _ _ _ _ _ _ ___,
I
~
w
~-{:0
OUTPUT {_....;...I_ __,__ _ _ _ _ _ _--'
AIRSPEED
I
I
I
SQUARE-LAW
:
~PENSATION,.1,.,
~ A I R r E E D PDlNTER DRIVE
ANTI·
AUTOTHRO_TTL.......;E.;....._
I
I
·e·
-1-j
VALID--;.-------------r-..+-.::..:.---4
I
I
rv
@:TI
r
DC
COMMANOl ~
- - - - - - - -1-----I M
I
I
AIRSPEED
I
INPUT
~oj
"'-
vu!D
OUTPtJT
I ;
}
---+J~;~QA~~D~~~;U;UTO)
_____
,,_:.., ___________ _
1!. Jl,cf
4!!1'j_ll!,L
FLAG
-. - - - - - i - - - - - - = - - - - - - - - - & " " " ' ) - ' . ' f R S P E E D CURSOR
~
. .iNffl!' FLAG
-
DC
I
~-,
I
~1
I
~~CH
~~EEO
_iC_ti_':_LAG
~CHWAA,~~----------l
T E S T - - + - - - - - - - - - - - - - - - - . . . . . i VMO AND
ADC VALID
OVERSPEED
V
PROCESSOR
CURVE SELECT
COARSE
ALTITUDE
Lt
J
""l
FINE-{
t
ALTITUDE
SYNCHRO
TO DIGITAL
CONVERTER
SYNCHRO
MULTIPLEXER
M
VwofMoo POINTER DRIVE
-,---------
1
1-------...;'~..,f"'"1,:.V~-~LAG
----=---{=0
I
~-'!?!!-E'!""
'-------{=0
airspeed pointer will be driven to coincide with the maximum
airspeed pointer position, and the higher airspeed signal will be
detected by the V""' and overspced processor. This produces an output
signal that triggers a solid-state overspeed switch, causing it to
activate a Mach/airspeed aural warning system (see page 70).
In order to drive the Mach number counter of the indicator,
synchro output signals corresponding to computed airspeed and
altitude are supplied from the respective modules of the ADC. As can
be seen from Fig. 7.8, the signals are supplied to a synchro
multiplexer, and then after conversion from synchro to digital they
are fed to the Vmn and overspeed processor; after amplification they
drive the Mach counter via its servomotor. The motor also drives a
synchro whose output is fed back to the multiplexer to 'null' the
signals corresponding to Mach number, when constant speed values
are obtained.
The setting of command airspeeds and associated signals for
autothrottle system operation is done in a similar manner to that
described earlier. For automatic operation, i.e. settings made on an
AFCS mode select panel, or, in some cases, on a display unit of a
Performance Data Computer (PDC) system, the speed reference knob
of the indicator remains in its normal pushed-in position. In this
position a clutch is disengaged, and a switch in the servomotor circuit
is closed to provide a path to ground as shown in Fig. 7.8. When the
commanded airspeed is set a command signal is supplied via an
amplifier to the speed reference marker servomotor so that it now
rotates the marker to the commanded speed. The servomotor also
drives the rotor of a synchro (indicated 'A') which then becomes desynchronized with respect to a second synchro (indicated 'B'). Thus,
an error voltage signal corresponding to the difference between
computed and commanded airspeeds is transmitted to the autothrottle
system. As the airspeed changes in response to the commanded
engine power change, the airspeed pointer and counter are driven so
as to indicate the speed change. At the same time, the rotor of
synchro 'B' is rotated in order to reduce the error signal voltage
produced by synchro 'A'. When the null position is reached, no
further output is supplied to the autothrottle system and the airspeed
pointer and counter are then at the commanded airspeed. A de
potentiometer is also driven by the servomotor to provide position
feedback.
For manual operation, the speed reference knob is pulled out to
engage a gear type of clutch, and, as can be seen from Fig. 7 .8, the
switch in the marker drive motor circuit now rotates the marker;
through a 2: I ratio gear it also rotates the rotor of synchro 'A' to
establish an error signal for transmission to the autothrottle system in
the same manner as that resulting from automatic operation.
A logic circuit is provided in the speed reference system, its
purpose being to monitor the system (while the speed reference knob
is pushed in) for loss of power, nulling of the synchro/servo system,
and validity of the input signals from an AFCS mode select panel or
PDC display unit. If an invalid reference display should occur, the
'INOP' flag appears as shown in Fig. 7.8. This flag also appears
when the reference knob is pulled out for manual operation (the
monitor circuit is disabled in this case) together with the 'MANUAL'
flag.
The remaining flags, i.e. Vmo• 'MACH' and 'A/S', appear under
the conditions referred to earlier.
173
The indication of true airspeed (T AS) is provided by a digital
counter type of indicator, the servomotor of which is supplied with
signals from the TAS module of the ADC. A failure monitor circuit
is incorporated in the ·indicator for the operation of a yellow 'OFF'
flag.
Altimeters
Figure 7. 9 Pneumatic/servooperaled altimeter.
The display presentation of one example of pneumatic/servo-operated
altimeter, and the basic arrangement of its mechanism, are shown in
Fig. 7 .9. The pneumatic section consists of two capsules which in
responding to changes in static pressure admitted to the indicator case
drive the pointer and digital counter in a manner similar to that of a
conventional pneumatic altimeter. The pointer and counter are also
driven by signals supplied to a CT synchro from the altitude module
of the ADC, and since these signals are of higher resolution and
accuracy, pointer and counter operation is predominantly controlled
through the servo drive. The pneumatic section, therefore, performs a
standby role so that it can provide altitude indications in the event of
failures in the synchronous transmission loop. A control knob,
located at the front of the instrument, is provided for use in such
cases, and when moved from the 'CADC' position to 'STBY', it
PRESSURE
ALTITUDE
SIGNALS
TO AFCS
--------1
--------,I
I
I
I
I
ALTITUDE INPUT
SIGNALS
FROM ADC
--1-----.....i
.,._-+------
BP SET
SIGNALS
TO ADC
\
'-
--- --- ---,
I
I
I
I
STATICPRESSURE
174
PRESSURE~===~
SENSOR
Figure 7.10 Servo-operated
altimeter.
Coarse
altilude
input
isolates the CT synchro signal circuit and also electrically activates a
red 'STBY' flag. The flag is also automatically operated by a failure
monitor circuit similar to that incorporated in Mach/airspeed
indicators.
As in the case of pure pneumatically-operated altimeters, indicated
altitudes :,.re corrected to standard pressure data by means of a
barometric pressure setting knob and counter mechanism. In addition,
however, the mechanism also positions the CT synchro stator with
respect to its rotor, thereby modifying the input signal from the ADC
altitude module. The resulting error signal voltage induced in the
rotor therefore drives the servomotor and, via the differential gearing,
the altitude poi,1ter and counter are driven to the required pressure
altitude value. The servomotor also drives the CT synchro stator for
'nulling out' the error voltage signal. The barometric pressure setting
knob also positions the stator of a second CT synchro provided for
the purpose of supplying equivalent pressure altitude signals to the
altitude selection facility and pitch control computer of an AFCS.
The internal arrangement and display presentation of a servooperated altimeter is shown in Fig. 7.10. Although used principally
with ADCs of the digital type, it may in some cases be interfaced
with some analog types of ADC.
Altitude signals designated as coarse and fine are transmi:td from
2-speed
switch
G
,... ___ a
T
t-----+-- ·----,"'""'=-~
I
I
I
Fine
altitude
inpul
1
I
I
,~:
J,..-.~'
l
I
I
I
----·-----------~-J
y
l
I
1,. __ _.____
I
I
I
I
I
I
I
l----1-------- _ I
I
I
- - - - - - - - LOGIC
' - - - - - - - - - - CIRCUIT
I
I
I
----IE}------,--- --___ :
i::z::@ill
I
I
I
I
I
Non-linear
mechanism
---------------------~
175
the ADC altitude module to the stator windings of corresponding
resolver synchros. The rotors are mechanically interconnected by a
I :27 ratio gear train, and their sine windings are connected to a
solid-state switch referred to as a speed switch. The purpose of the
switch is to control the servomotor operation so that fine altitude
signals are supplied to its amplifier at altitudes below 1000 ft, while
at altitudes above this value, coarse altitude signals are supplied. The
servomotor drives the altitude pointer and counter through a clutch
and gear train, and directly drives a tachogenerator which provides
rate feedback signals to the amplifier. The reduction and 'nulling out'
of altitude error signals is effected by driving the rotors of the
resolvers through a differential gear.
The cosine windings of the resolvers are connected to a logic
circuit that monitors the presence of the following: I. coarse and fine
servo 'nulls' (sine windings); 2. coarse and fine excitation (cosine
windings); 3. indicator power supplies; and 4. valid altitude data. If
either of these is unreliable, a solenoid-operated 'OFF' warning flag
is activated to obscure the digital counter display of altitude.
Barometric pressure setting is done in a manner similar to that of
the altimeter described earlier, except that the setting knob rotates the
stators of two resolver synchros for establishing the error voltage
signals necessary to obtain the required pressure altitude indications.
The purpose of the 'non-linear' mechanism shown in the diagram is
to compensate automatically for the non-linear relationship between
barometric pressure and altitude, so that for any setting of the
pressure counters, the corresponding altitude will be indicated. The
pressure setting knob also changes the rotor position of a third ·
resolver, the purpose of which is to supply pressure-corrected signals
to such systems as AFCS and altitude alerting.
The second .knob in the bottom right-hand corner of the altimeter
permits the setting of a reference marker to align with an altitude
indication corresponding to a specific operating condition. The
purpose of the servo-driven 'NEG' flag is to obscure the digital
counter display at altitudes below sea-level.
Static air temperature indication
The most basic method of obtaining an indication of SAT is to use
charts of pre-calculated values of ram rise related to sensing probe
recovery factors (see page 62) and Mach number and then subtract
the values from the readings of the TAT indicator. Such conversion
charts are provided by manufacturers and normally form part of an
aircraft's operations or flight manual. It is, however, more
advantageous to provide an automatic method of conversion so that
corrections, in the form of electrical signals, can be applied to the
TAT signal output to derive SAT, and then utilize the corrected
signals to operate a separate indicator.
In the case of the analog ADC described in this chapter, the
conversion and correction is effected by a circuit network whose
electrical characteristics are matched to those of the TAT sensing
probe. As shown in Fig. 7 .2, the network is incorporated within the
Mach module of the computer to accept TAT probe output signals, as
well as a drive input from the servo loop in order to vary the SAT
signals as a function of Mach number.
The circuit arrangement of one example of SAT indicator is shown
in Fig. 7 .11. It utilizes a drum type of counter, the left-hand and
right-hand sections of which display temperatures in the plus and
minus parts, respectively, of the range. The centre drum displays the
sign of the temperature being indicated; the drums not in use are
automatically masked. The computed SAT is supplied as a de analog
voltage to a chopper circuit, and is compared with a voltage supplied
as a reference via a re-balanced potentiometer. The chopper circuit
produces a 400 Hz ac error signal representative of the difference
between the two inputs, which is then amplified to drive the
servomotor and counters. The motor also repositions the re-balancing
potentiometer to 'null out' the error signal. In the event of loss of de
or ac power, or an excessive 'null' voltage in the re-balance/feedback
system, an 'OFF' flag is triggered by a failure monitor circuit to
obscure the counter display.
Digital ADC
The modular arrangement and data signal flow of a typical computer
are shown in schematic form in Fig. 7 .12. It processes the same
basic parameters as the one already described, but with the major
difference that all the signals corresponding to the variables measured
are converted and transmitted in digital format. The pilot and static
pressure sensors are of the piezoeiectric crystal type (see also page
165) and their frequency-modulated signals are supplied to the
altitude, computed airspeed, and Mach calculation circuit modules via
a frequency-to-digital converter. The analog inputs from the synchros
of angle of attack (alpha) sensors, and altimeter barometric pressure
setting controls, are converted by means of synchro-to-digital
converters. Outputs from all modules of the computer are supplied to
an ARINC 429 transmitter connected to four data busses from which
all interfacing systems requiring air data are then supplied. The
purpose of the discretes coder module is to monitor signals relating to
the status and integrity of particular circuits, e.g. the heater circuits
of TAT protes, pitot probes, and angle of attack sensors, and to
initiate appropriate warnings. In order for the computer automatically
to take into account the pressure error of the air data system of a
11'1
I
I
1
I
'I
UT{-
OG
!
20_ V
UNIT
I
h
'
·-·-·
L__.
~ONITOR
~of~
.
~
""""°" ! ~
,1~1~=ceuI
MONITOR
POWER
.
l
,
Ii
·-·
FAILUAE_M~
-
,,,,,~
_:::-1I
~---·--- ---·---·
-{>-
'
I
L
G;J
I
I
I
I
I
I
~!
·
- - - - - - ----------- - I- - CHOPPER
POWER
SUPPLY
I
COM···MON INPUT
ENCE
{
DC
I
-.
an
To
COMPUTATION
TAT
CORRECTED
AOA
TEMPERATURE
COMPENSATION
ANO
Clll.16RATION
SAT
COMPUTATION
SSEC
L.---·-·-·-·
ES•--1-------,..i
COOER
~
MOO\JI.E
SUPP\.Y
R
CONVERTER
TOOIGITAI.
SYNCHRO
OIGITIII.
CONVERTER
PRESSURE
SENSOR
.AJCTYPE
FREQUENCY
TO
NC T_YPE
EEC
IAU
GPWS
FCC
FM¢
EICAS
EFIS
·-·==>Oala-
CIII.C.OV.TION
MACH
CALCULATION
AIRSPEED
COMPUTEO
CIII.CULATION
ALTITUOE
L----------·---------------------
FREQVENC\'
I.IOOUI.ATEO
CONVERTER
TO DIGIT Al
R--i------4>!
--
T I MSYNCHRO
l!ttRh'
TltfGS
,-·
I
.d
AAINC
429
TRANSMITTER
=;l
Ground p«»umity wa,rvno syttem
Inertial reference UNI
Electrorne eogtne control
E)edronic ffigh1 11\SttUmatftt system
Engmo indtc.ating and cotrtrol l'f51om
Ffighl_COffll)IJ...
FlighlcooUol"""1jlUW
TAT'
SAT
TAS
I
IRU
y,,,.
FMC
GPWS
EICAS
TC&t
Fligh
F/O
EFIS
EEC
IRU
°""
Allltudo
Yaw
$lall warnin
Slab. -
FCC
Tl.IC
Figure 7.13 Digi1aliservooperated altimeter.
nsv~
v o!1sc
ac
I sPOWER
V DC
.,.
SUPPL
-15 V DC
DECODE
INPUT
FLAG
DRIVER
LAOC _ _ _.......,
OFF
NEG
H!Jll
RAOC _ _ __
I
particular type of aircraft, and also its stall characteristics, it is
'matched' by programming the SSEC and angle of attack modules
with the relevant data.
The indicators associated with a digital ADC are of the pure servooperated type, and as an example of their operation generally, we
may consider the altimeter circuit shown in Fig. 7.B. Data may be
supplied from either of two ADCs as selected by the triggering of a
solid-state switch; the right ADC is shown. Under changing altitude
conditions the corresponding signals pass to a microprocessor, and
from this unit they are transmitted to a D/ A converter which then
provides the drive signals for operating the servomotor, and the
altitude pointer and counter mechanisms. At the same time the minor
drives two CX synchros which supply coarse and fine analog inputs
to an input multiplexer. The output signals from this unit are then
compared with those from the DI A converter, and the difference
between them (as a result of altitude change) is fed back into the
microprocessor via an AID converter. The signals will 'nun out'
when a constant altitude condition has been attained. The setting of
barometric pressures is done in the usual manner, i.e. by means of a
set knob and digital c~unters, and, as will be noted, the stators of two
resolver synchros are also repositioned. These produce sine- and
cosine-related signals which are fed back to the corresponding
synchro-to-digital converter in the computer (see Fig. 7.12). The
change in the converter output signal is supplied to the altimeter, via
the relevant data bus, so that its servomotor will drive the pointer and
counter mechanism to indicate the attitude change corresponding to
the barometric pressure setting. If input signal failure or a negative
altitude condition should occur, the microprocessor activates decoder
and flag driver circuits which then cause the appropriate flag to
appear across the altitude counter display.
181
8
agnetic heading
reference systems
A magnetic heading reference system (MHRS), sometimes called a
remote-indicating compass, is basically one in which an inductive
type of element detects an aircraft's heading with respect to the
horizontal component of the earth's magnetic field in terms of flux
and induced voltage changes, and then transmits these changes via a
synchronous/servo and stabilized reference system to a heading
indicator. Thus, in concept an MHRS is a combination of the
functions of a direct-reading magnetic compass and a direction
indicator, but one in which the individual errors associated with these
two instruments are considerably reduced.
In practice, there are two types of MHRS: (i) that in which the
detector element monitors a directional gyroscope unit linked with a
heading indicator, and (ii) one in which the detector element operates
in conjuncti9n with the platform of an inertial navigation system
(INS) to supply magnetic heading and stabilized heading data
respectively to a compass coupler unit linked with a heading
indicator.
Detector elements
182
These elements (variously known as flux detectors, or flux valve<;),
unlike those of direct-reading compasses, are of the fixed type
which detect the effect of the earth· s magnetic field as an
electromagnetically induced voltage and control a heading indicator
by means of a variable secondary output voltage signal. In general,
the construction of an element is as shown in Fig. 8.1. It takes the
form of a three-spoked wheel, slit through the rim between the
spokes so that they, and their section of rim, act as three individual
flux collectors.
Around the hub of the wheel is an exciter coil which has an
electrical function corresponding to that of the primary winding of a
transformer. Coils are also wound around the spokes, and these
correspond in function to that of a transformer's secondary windings.
The reason for adopting a triple spoke and coil arrangement will be
made clear later in this chapter, but at this stage the operating
principle can be understood by considering a single-turn coil placed
in a magnetic field. The magnetic flux passing through the coil is a
Figure 8.1 Detector element
construction.
SECONOARV
PICK-OFF COIi.$
1.AMINATEO
COLLECTOR HORNS
SPOKES
EXCITER (PRIMARY) COIL
maximum when it is aligned with the direction of the field, zero
when it lies at right angles to the field, and maximum but of opposite
sense when the coil is turned 180° from its original position. Fig.
8.2(a) shows that for a coil placed at an angle fJ to a field of strength
H, the field can be resolved into two components, one along the coil
equal to H cos 8, and the other at right angles to the coil equal to H
sin fJ. This latter component produces no effective flux through the
coil so that the total flux passing through it is proportional to the
cosine of the angle between the coil axis and the direction of the
field. In graphical form this total flux may be represented as at (b).
If the coil were to be positioned in an aircraft so that it lay in the
horizontal plane with its axis fixed on, or parallel w, the aircraft's
longitudinal axis, then it would be affected by the earth's horizontal
component and the flux passing through the coil would be
proportional to the magnetic heading of the aircraft. It is therefore
apparent that in this arrangement we have the basis of an MHR
system able to detect the earth's magnetic field without the use of a
permanent magnet. Unfortunately, this simple system would be of
little practical use because, in order to determine the magnetic
heading, it would be necessary to measure the magnetic flux, and
there is no simple and direct means of doing this. If, however, a flux
can be produced which changes with the earth's field component
linked with the coil, then we can measure the voltage induced by the
changing flux, and interpret the voltage changes so obtained in terms
of heading changes. This is achieved by adopting the construction
method shown at (c) of Fig. 8.2.
Each spoke consists of a top and bottom leg suitably insulated from
each other and shaped so as to enclose the hub core around which the
primary coil is wound. The material from which the spokes are made
is an alloy especially chosen for its characteristic property of being
easily magnetized but losing almost all of its magnetism once the
183
Figure 8, l
coil.
Detector element
FIELD H
I
t
t
I
lt
I
I
I
I
I
I
(a)
(b)
I.AMINATEO SPIDERS
OF SPOKE
LAMINATED
COLLECTOR HORNS
EXcrrtR COIL SUPPLIED
AT 23·6V, 400 HZ
(C)
external magnetizing force is removed (a typical material is one
known as Permalloy). With this arrangement there are two sources of
flux to be considered: (i) the alternating flux in the legs due to the
current flowing in the primary coil; this flux is at the same frequency
as the current and is proportional to its amplitude; (ii) the static flux
due to the earth's component H, the maximum value of which
depends upon the magnitude of H and the cosine of the angle
between H and the axis of the detecting element.
If we consider first that the axis of the element lies at right angles
to H, the static flux linked with the coils will be zero. Thus, with an
184
Figure 8. 3 Total flux-detector
spoke al right angles to the
eanh·s field.
EARTH'S COMPONENT H
1+-
FLUX IN BOTTOM LEG
+
/ \
l-'7....--j,flh"'11,iit-+-I: :
II
:
:
:
:
:
I
I
I! I I I
EXCITER coo. VOLTAGE : : : : :
I
I
I
I
I
I
1
J /
:
I
l
I :
I;IJ:
/
"'"""'"'
I
I
I
I
I
I
J
11 I I I I I 11
t
l
1 I
I
l
I I
FLUX IN TOP LEG
+
TOTAL FLUX IS ZERO
Figure 8. 4 BIH curve and
hysteresis loops.
IRON
I
I
I
-H
I
+H
I
I
I
/
/
-8
alternating voltage applied to the primary coil, the total flux linked
with the secondary will be the sum of only the alternating fluxes in
the top and bottom legs and must therefore also be zero as shown
graphically, in Fig. 8.3.
The transition from primary coil flux to flux in the legs of the
detector element is governed by the magnetic characteristics of the
material, such characteristics being determined from the
magnetization or BIH curve. In Fig. 8.4, the curve for Permalloy is
compared with that for iron to illustrate how easily it may be
magnetized. There are several other points abo1..t Fig. 8.4 which
185
should also be noted because they illustrate the definitions of certain
terms used in connection with the magnetization of materials, and at
the same time show other advantages of permalloy. These are:
1. Permeability: which is the ratio of magnetic flux density B to
field strength or magnetizing force H; the steepness of the curve
shows that Permalloy has a high permeability.
2. Saturation point: the point at which the magnetization curve
starts levelling off, indicating that the material is completely
magnetized. Permalloy is more susceptible to magnetic induction than
iron, as shown by its higher saturation point.
3. Hysteresis* curve and loop: these are plotted to indicate the
lagging behind of the induced magnetism when, after reaching
saturation, the magnetizing force is reduced to zero from both the
positive and negative directions, and also to determine the ability of a
material to retain magnetism. The magnetism remaining is known as
remanence or remanent flux density, and it will be noted that
Permalloy has an extremely low remanence, thus making it admirably
suitable for use in detector elements.
4. Coercivity: this refers to the amount of negative magnetizing
force (coercive force) necessary to comple~ely demagnetize a .
material, and is represented by the distances OC; and OCP. Coercivity
and not remanence determines the power of retaining magnetism.
In order to show the characteristics of the flux waves produced in
the legs of a detector, a graphical representation in the form of that
illustrated in Fig. 8.5 is adopted.
The waveshapes of the alternating primary fields ·are drawn across
the axis B of the BIH curve, and those of the corresponding flux
densities in the legs are then deduced from them by projection along
the H axis. The total flux density produced in the legs is the sum of
the individual curves, and with the detector element at right angles to
the earth's horizontal component H then, like the static flux linked
with the secondary coil, it will be zero. Since the total flux density
does not change, the output voltage in the secondary coil must also
be zero.
Let us now consider the effects of saturation when the detector
element lies at any angle other than a right angle to the horizontal
component Has indicated in Fig. 8.6(a). The alternating flux due to
the primary coil changes the reluctance, i.e. the magnetic resistance,
of the material, thus allowing the static flux due to component H to
flow into and out of the spoke in proportion to the reluctance
changes. This effect is analogous to that produced by the opening and
closing of a valve, hence the name flux valve being applied to a
detector element.
During those stages of the primary flux cycle when the reluctance
* From the Greek hysteros, meaning 'later'.
188
Figure 8.5
Flux wave
B
Lharactcristics
FLUX IN BOTTOM LEG
:rf-
1
I
H
I
I
EXCITER COIL
VOLTAGE AND
MAGNETIZING
FORCE IN TOP
LEG
-
SATURATION
-POINT
.
\
QI-----+-,----l-T-IM-E
_j ___ J_
I
-
\
I
\
/
~- - - -POINT
SATURATION
FLUX IN TOP LEG
B
---~
, I 'excrrER COIL0
VOLTAGE AND
MAGNETIZING
FORCE IN
BOTTOM LEG
TOTAL FLUX AND SECONDARY
PICK·OFF COIL OUTPUT IS ZERO
is greatest, the static flux links with the secondary coil, and the effect
of this is to displace the axis, or datum, about which the magnetizing
force alternates. The amount of this displacement depends upon the
angle between the earth's field component Hand the flux detector
axis. This is shown graphically at (b) of Fig. 8.6.
If we now apply a graphical representation similar to Fig. 8.5, and
include the static flux of component H, the result will be as shown at
(c) of Fig. 8.6. It should be particularly noted that a flattening of the
peaks of the flux waves in each leg of a spoke has been produced.
The reason for this is that the amplitude of the primary coil excitation
current is so adjusted that, whenever the datum for the magnetizing
forces is displaced, the flux material is driven into saturation. Thus a
positive shift of the datum drives the material into saturation in the
direction shown, and produces a flattening of the positive peaks of
the fluxes in a spoke. Similarly, the negative peaks will be flattened
as a result of a negative shift driving the material into saturation at
the other end of the BIH curve. The total flux linked with the
secondary coil is, as before, the sum of the fluxes in each leg. and is
of the waveshape as also indicated in Fig. 8.6.
When the detector element is turned into other positions relative to
the earth's field, then dependent on its heading the depressions of the
total flux value become deeper and shallower. Thus the desired
changes of flux are obtained and a voltage is induced in the
secondary coil. The magnitude of this induced voltage depends upon
the change of flux due to the static flux linked \\ ith the secondary
187
Fi11ure 8. 6 Effect of eanh 's
component H. (a) Deteclor at
an angle to component H.
1b) displacement of axis due to
,latic flux; (c) lolal flux and
emf.
+
l/''\
IrtT
EXCITER COil VOLTAGE
(a)
DISPLACEMENT OF AXIS
DVE TO EARTH'S COMPONENT
___ L
o~----.~---'------.-TIME
--a,;..
(b)
t
B
TOTAL FLUX
LINKED
WITH
~~~~....::C#--1-~~-l-~~~.i..:.....;.,1....1:..1....i..;.....;.,l....JIU-.l.!.......~l..l..-l-l.~~S!:COl!IDARV
H--
-µ'-~'~
I I
. I
MAGNETIZING
FORCE IN
TOP LEG
(c)
188
I I
E.M.F.
INOUCEO IN
S!:COl!IOARY
PICK·Off
COIL
coil which, in turn, depends upon the value of the effective static
flux. As pointed out earlier, the value of the static flux for any
position of the detector element is a function of the cosine of the
magnetic heading; thus the magnitude of the induced voltage must
also be a measure of the heading.
One final point to be considered concerns the frequency of the
output voltage and current from the secondary pick-off coil and its
relationship to that in the primary excitation coil. During each halfcycle of the primary voltage, the reluctance of the material goes from
minimum to maximum and back to minimum, and in flowing through
the material the static flux cuts the pick-off coil twice. Therefore, in
each half-cycle of primary voltage, two surges of current are induced
in the pick-off coil, or for every complete cycle of the primary, two
complete cycles are induced in the secondary pick-off coil.
The ac supply for primary excitation has a frequency of 400 Hz;
therefore, the resultant emf induced in the secondary pick-off coil has
a frequency of 800 Hz, as shown in Fig. 8.6(c), and an amplitude
directly proportional to the earth's magnetic component in line with
the particular spoke of the detector element.
Having thus far studied the operation of a single spoke of an
element. the reasons for havi1,1g three may now be examined a little
more closely. If we again refer to Fig. 8.3, and also bear in mind the
fact that the flux density is proportional to the cosine of the magnetic
heading, it will be apparent that for one detector spoke there will be
two headings corresponding to zero flux, and two corresponding to a
maximum. Assuming for a moment that we were to connect an ac
voltmeter to the detector, the same voltage reading would be obtained
for both maximum values because the voltmeter cannot take into
account the direction of the voltage. For any other value of flux there
will be four headings corresponding to a single reading of the
voltmeter. However, by employing a triple spoke and coil
arrangement at a spacing of 120°, the paths taken by the earth's field
through the spokes, and for 360° rotation, will be as shown in Fig.
8.7. Thus, varying magnitudes of flux and induced voltage can be
obtained and related to all headings of the detector element without
ambiguity of directional reference. The resultant of the voltages
induced in the three spokes at any one time can be represented by a
single vector which is parallel to the earth's component H.
Figure 8.8 is a sectional view of a practical detector element. The
spokes and ~oil assemblies are pendulbusly suspended from a
universal joint which allows a limited amount of freedom in pitch and
roll, to enable the element to sense the maximum effect of the earth's
component H. It has no freedom in azimuJh. The case in which the
assymblies· are mounted is hermetically sealed and partially filled with
fluid to damp out excessive oscillations of the assemblies. The
complete unit is secured to an aircraft's structure at locations which
189
Fi11ure 8. 7 Path of earth ·s
field through a detector.
EARTH'S COMPONENT f,
000°
Figure 8. 8 Practical detector
clement. l Mounting nange
(ring seal assembly). 2 contact
assembly, 3 terminal. 4 cover,
5 pivot, 6 bowl. 7 pendulous
weight, 8 prim~ry (excitation)
mil. 9 spider leg. lO secondary
coil. 11 collector horns.
12 pivot.
· ow•
1so•
270°
2
afford maximum protection against the deviating effects of aircraft
magnetism. Typical locations are in wing tips and vertical stabilizers.
One of the slots in the mounting flange is calibrated a limited number
of degrees on each side of a zero position corresponding to an
aircraft installation datum, for adjustment of deviation coefficient A
(see also page 93). In the example illustrated, provision is made at
the top" cover of the casing for electrical connections and attachment
of a deviation compensating device.
In many MHRS installations, the detector elements are of the preindexed type, i.e. they can be removed and replaced without
subsequent adjustments having to be made for coefficient A
compensation. The element (see Fig. 8.9) is supplied by its
manufacturer, together with a mounting plate on which it has been
accurately <!ligned; its fixing, or indexing, screws are sealed and must
not thereafter be removed. The complete unit is then secured to an
alignment reference bracket wbich, having already been accurately
aligned parallel to the longitudinal axis of an aircraft, ensures that the
detector element is similarly aligned. The forward indexing screw of
the bracket is also sealed and must not be removed.
From the foregoing description of a detector element, the similarity
Figure 8. 9 Pre-indexed
detector elemenl.
ALIGNMENT REFERENCE
BRACKET
I
DETECTOR
INDEXING SCREW (3)
(00 NOT REMOVE)
SUPPORT
BRACKET
betwen its operation and that of a basic type of synchro transmitter
will no doubt have been observed. Such an observation is not
incorrect, of course, and in fact if the detector were directly
connected to an electrically-matched receiver synchro such as a TR,
then in combination they would form a simple MHR or remoteindicating compass system. It would not, however, be very accurate
in its indications since, by vinue of the limited pendulous suspension
arrangement of a detector element, errors can occur as a result of its
tilting under the influence of acceleration forces, e.g. during speed
changes on a constant heading, or turning of an aircraft.
In order therefore to compensate for these effects and so reduce
errors, it is necessary to incorporate within the system a means by
which the long-term azimuth or magnetic reference established by the
detector element is continuously stabilized and monitored. A
stabilizing technique of early origin, but nonetheless still widely used,
is that in which a horizontal-axis gyroscope unit is referenced initially
to the magnetic meridian, and then, in order to maintain this
relationship, precessional forces created by a slaving synchro/torque
motor system are applied to the gyroscope. The degree of control of
the detector element over the gyroscope, i.e. the monitoring rate, is
of considerable importance. For example, during a turn the detector
monitoring rate
element heading is likely to be in error, and so
must be such that ,the induced heading is that of the gyroscope. At
the same time, there must be sufficient control to correct for drift of
191
the gyroscope. The gyroscope, therefore, provides short-term azimuth
references.
Monitored gyroscope
system
As will be seen from Fig. 8.10, this type of MHRS is comprised of
five individual units: (i) detector element; (ii) slaving/servo amplifier;
(iii) directional gyroscope unit (DGU); (iv) radiomagnetic indicator
(RMI); and (v) a deviation compensator.
The units (i) to (iv) are interconnected through a transmission loop
consisting of control synchros (see page 140) which produce the
required slaving and monitoring signals appropriate to the heading
reference signals transmitted by the detector element. This element
can, therefore, be considered as a special form of CX synchro,
whereby the transmitter rotor field is represented by the resultant of
the earth's field component H, as shown in Fig. 8.11. The secondary
pick-off coils are connected to the corresponding windings of the
stator of a CT receiver synchro whose function is to produce error
signals which, after amplification, precess the DGU, thereby slaving
it to the detected magnetic heading reference.
When the detector element is positioned as shown at (a) of Fig.
8.11, the path of the earth's field component H through the spokes
will cause a maximum voltage signal to be induced in the pick-off
coil A, while in coils B and C, signals of half the amplitude and of
opposing phase will be induced. These signals produce fluxes in the
CT synchro stator to establish a resultant field which is in alignment
Figure 8.10 Monitored
gyroscope system.
r
-c==:}-~
Radio n a v . - - - - - ,
data
DEVIATION
COMPENSATOR
Directional reference and heading error
@
.-----1
DETECTOR
ELEMENT
~
RMI
Slaving synchro
SLAVING
Power
Excitation
...__
_ _ _ supply
module
SERVO
AMPLIFIER
----"'!
==~
Heading servo
synchro
DIRECTIONAL GYRO
Servo synchro
Compass card motor
Heading signals
to flight director
HSI
Slaving.
(a) Heading= 000°; (b)
heading = 090°.
Figure 8. I I
~
COIL A
~~
~
(a)
RESULT ANT OF EARTH"S
FIELD COMPONENT
THROUGH DETECTOR
"'
:,\
I I\
R
RESULT ANT OF FIELD DUE
TO INDUCED VOLT AGE SIGNA~S
ICOILA
le
Po
C
COIL C
-
- - - - EARTH"S FIELD
-
INDUCED VOLTAGE SIGNALS
(b)
with that passing through the detector element. If, at that instant, the
CT rotor is at right angles to the resultant field, no voltage will be
induced in its winding. The DGU alignment will corre!>pond to that
of the resultant earth's field, and the RMI will indicate the heading
0000.
Let us now assume that the detector element is turned through 90°
say; the disposition of the pick-off coils will then be as shown at (b}
of Fig. -8.11. No signal voltage will be induced in coil A, but coils B
and C have increased voltage signals induced in them, the signal in
coil C being opposite to what it was on the previous heading. The
resultant flux across the CT stator will also have rotated through 90°,
and assuming for a moment that the rotor is still at the original
position, then the flux will induce maximum voltage in the rotor.
This heading error voltage signai is supplied to the slaving module of
the amplifier in which its phase is detected, and after amplification it
is supplied to a slaving torque motor which then precesses the
gyroscope to the new magnetic heading reference. At the same time,
the DGU operates a servo control synchro system, the function of
which is to rotate the slaving synchro rotor so as to start 'nulling' out
the heading error signal, and also to rotate the compass card of the
RMI as the heading change takes place.
193
Figure 8. i2 Operation of
monitored gyroscope system.
-------1
RADIOMAGNETIC INDICATOR
-
GNAL FLOW
:--
::..t, DIRECTIONAL REFERENCE
I'L
-i> HEADING ERROR
.._,,: MONITORING ANO
~ PRECESSION
-SERVO
-LOOP ERROR
'
I
!
'--+,-+----+---bl
· _J
SERVO AMPLIFIE~
--C,, SERVO DRIVE
* sigrlals
He!l(ling data
to other
sys,tems, e.g.
automatic flight
control
*
:
SELECTED HEADING
DATA ex
~DIFFERENTiAL GEAR
SET HEADING
KNOB
SYNCKRON!ZlNG
KNOB
In practice, the rotation of the field in the slaving synchro, and
slaving of the DGU, occur simultaneously with the turning of the
detector element so that synchronism between the element and the
DGU is continuously maintained.
The synchronous transmission link between the four principal units
of the MHRS is shown in more detail in Fig. 8.12. The rotor of the
servo CX synchro is rotated whenever the DGU is precessed, or
slaved, to a magnetic heading reference, and the signals thereby
induced in its stator are applied to the CT synchro in the RMI.
During slaving, the rotors of both synchros wili be misaiigned. and a
servo loop error voltage is therefore induced ir. the CT rotor and then
applied to the servo module of the amplifier. After amplification, !he
voltage signal is applied to a servomotor which is mechanically
. Thus,
coupled to the CT rotor and to the he.:ding dial of the
both the rotor and dial are rotated, the latter Indicating the direction
of the heading change taking place. On cessation of the heading
change, the rotor reaches 1 'null' position, and as there wiH no
longer be an input to the servo module of !he amplifier. the
servomotor ceases to rotate and the RMt indi.cates the new heading.
The servomotor also drives a tachogenerator which supplies feedback
194
signals to the amplifier to damp out any oscillations of the servo
system.
:OGU
This unit is locatt".d at a remote point in an aircraft (typically in an
electronics equpment compartment), and although it normally forms
part of an MHR system, it is also designed to serve as a centralized
source of heading data for use in the operation of flight director and
automatic flight control systems. For example, and as shown in Fig.
8.12, an additional CX synchro is provided in the RMI for the
purpose of supplying magnetic heading data to the horizontal situation
indicator (HSI) of a flight director system (see page 216) and the
DGU is provided with a similar synchro for supplying this data to an
automatic flight control system (AFCS).
in a typical unit (see Fig. 8.13) the gyroscope is based on a twophase induction motor operating from a l 15 V ac power source. The
inner gimbal ring is fitted with hemispherical covers to totally enclose
the motor, and this assembly is filled with helium and hermetically
sealed. The rotors of the slaving/heading CX synchro, and the
additional heading data CX synchro, are mounted on the top of the
Figure 8. 13 Example of a
DGU.
195
outer gimbal ring; the stators are fixed to the frame supporting the
gimbal ring. The slaving torque motor is positioned on one side of
the outer gimbal r:ng, and on the inner gimbal ring axis, to provide
precession of the ou:er gimbal ring appropriate to magnetic heading
reference changes; the precession rate is I 0 -2°/min.
In order to control drift of the gyroscope (see page 102) its spin
axis is maintained in th! horizontal position by a liquid levelling
switch and torque motor system which operates in a similar manner
to the erection system used in electrically-operated gyro horizons (see
page 116). The levelling switch is mounted on the inner gimbal ring
and is connected to the control winding of the torque motor mounted
on tlie lower part of the outer gimbal ring.
A speed-monitoring circuit system (sometimes called ·spin-down
braking') is also incorporated to prevent the oscillating effect, or
nutation, of the gyroscope which can occur when, at low rotational
speeds, its gimbal ring axes are not mutually perpendicular. The
system holds the gimbal system steady for short periods during the
start-up and run~down stages of gyroscope operaton. Circuits are also
provided for the monitoring of the slaving/heading synchro output
signals, and the input signals from the servo module of the system's
amplifier. Should the signals from any one of these monitoring
circuits become invalid, a relay is energized to complete circuits
controlling warning flags in the RMI, and the HSI of a flight director
system.
The complete gyroscope assembly is mounted on anti-vibration
mountings contained in a base which provides for attachment at the
appropriate location in an aircraft. and also for the connection of the
relevant electrical circuits.
Radio magnetlc indicator (RMI)
This is a triple display type of indicator which derives its name from
the fact that, in addition to magnetic heading data. it also displays the
magnetic bearing of an aircraft with respect to ground-based
transmitting stations of radio navigation systems. The systems
concerned are: ADF (Automatic Direction Finding) and VOR (Very
high~frequency Omnidirectional Range).
The display presentations of two examples of RMI are shown in
Fig. 8.14. Magnetic heading data is displaced by the heading card
which is rotated relative to a fixed lubber line in the manner already
described. Magnetic bearing indications are provided by two
concentrically-mounted pointers, one called a 'double-b1.!f.'.-pointer and
the other a 'sin~te:-ba( poii;it('.r. Both are referencecf against the
and are positioned by synchros that are supplied with
the appropriate bearing signals from the ADF and VOR navigation
receivers.
lieaa'ing'carcf,
Figure 8. 14 RMI displays.
LUBBER LINE
SYNCHRONIZING
ANNUNCIATOR
VOA 2/AOF 2 POINTER
VOA 1/AOF 1 POINTER
VOA 2/AOF 2 SELECTOR
VOA 1/AOF 1
SELECTOR
COMPASS
FAILURE FLAG
(a)
LEFT DISTANCE
DISPLAY
LEFT BEARING
POINTER
RIGHT BEARING
POINTER
(b)
ai~~=~,~~~:~~:a:st:~.~f~1;~~~:~~i[;h:n:nf&!~n::
indi;idually s~l~teii'so thatiheir outptii signals can ·operate the
':§ITiI~fiJii·i:~2rr~sPQ~di~gtnanner: Fo'r exampie; when the
'VOR-1/ADF-l' and 'VOR-2/ADF-2' selector knobs are each at the
'VOR' position, the single-bar pointer and double-bar pointer synchro
systems will, respectively, respond to bearing signals received from
the No. l and No. 2 VOR system receivers. A similar response will
be obtained with both selector knubs at the •ADF' position. Bearing
information can also be displayed with one selector knob set at
'ADF' and the other at 'VOR'.
197
Figure 8. 15 Magnetic bearing
and heading display.
N
ADI'
The signals transmitted to the synchros are such that the pointers
always point to the stations from which the signals are received. This
may be seen from Fig. 8.15, which is a representation of how the air
position of an aircraft may be determined from the display of
magnetic bearings and magnetic heading.
The function of the synchronizing knob and annunciator
incorporated in the indicator illustrated at (a) of Fig. 8.14 will be
described later.
MHRS integration
with an INS
198
When an MHRS is to be integrated with an INS (see Fig. 8.16) there
is no longer a requirement for it to have its own individual
directional gyroscope unit. The reason for this is that short-term
stabilizing of heading references is· readily available from the inertial
platform of the INS. The system also differs from the one described
earlier in that the slaving/heading and servo signal transmission and
control circuits are contained within a unit referred to as a compass
coupler. The interconnection between all relevant units of a typical
integrated system is shown in more detail in Fig. 8.17.
The stabilized heading reference, or platform heading signal, is
derived from an azimuth CX synchro the rotor of which is positioned
by the gimbal system of the inertial platform during changes in
aircraft heading. The signals so produced are supplied to the rotor
Figure 8.16 MHRS integration
with an INS.
COMPASS
COUPLER
AMI
CONTROLLER
RADIO
....------. TRUE
INS
PLATFORM
.-----------,...i
INS
1-H_D_G_-,+--<>
PLATFORM HOG.
COMPUTER
INS
HSI
RADIO/IN$ RELAY
windings of a CDX synchro and a resultant field vector is produced
in the conventional manner (see page 142). During a heading change,
the detector element will also induce voltage signals corresponding to
magnetic heading, and these are supplied in the normal manner to a
slaving CT synchro. In this case, however, the amplified heading
error signals are supplied to a stepper motor via logic circuit control
modules which perform the functions of frequency comparison and
voltage/frequency conversion. In addition a polarity detector circuit is
provided to determine the direction of stepper motor rotation. The
motor is mechanically coupled to the CDX synchro rotor and, as its
name implies, it is one whose shaft rotates a step at a time as it
responds to signals supplied to its windings. In the motor control
circuit module shown in Fig. 8.17, the output pulses from the
voltage/frequency converter are combined with the polarity detector
output to energize the motor windings in pairs. The sequence, and
frequency, of energizing determines the direction and rate
respectively of step rotation; each step corresponds to 1.3 min of arc.
As may be seen from Fig. 8.17, the stator of the CDX synchro is
connected to a CT synchro which utilizes the platform heading
signals for driving a heading servomotor. The motor is mechanically
coupled to the rotor of this synchro, the rotor of a CX synchro which
supplies heading output signals to the RMI, and also to the magnetic
heading CT synchro. Thus, during a heading change, the servomotor
199
~
El
oz ...
a:o
l: a:
01zz
iii8
I
'-----...,;..._____,
Figure 8.17 MHRS/INS signal
lransmission.
200
is driven at a rate to maintain synchronism between the platform
heading and magnetic heading signals, and so no heading error
signals are supplied to the stepper motor. The servomotor drives a
tachogenerator which supplies rate feedback signals for speed control
of the motor, and it also repositions the rotor of the servo CT
synchro to ensure 'nulling out' of the heading reference signals on
completion of a heading change.
In the event of de-synchronizing occurring, e.g. as a result of
inertial platform drift, the signal produced by the CDX synchro
would cause the heading servomotor to drive the slaving input CT
synchro rotor out of 'null', thereby producing a heading error signal.
This signal, after amplification, is supplied to the stepper motor
which then repositions the CDX synchro rotor in a direction opposite
to that of the field vector input produced by the drift. This vector is
therefore 'physically' repositioned so as to produce a corresponding
directional change in the stator of the CDX. The resulting reversal of
the stator output signal to the heading servo system then causes the
servomotor to drive the synchros until the heading error signal caused
by drift has b.een completely 'nulled'.
Synchronizing
Whenever heading error signals are produced as a result of, say,
selection of a new magnetic heading to be flown, or drift of a DGU
or inertial platform, it must be ensured that in the 'nuiling out' of
such signals, synchronization of the slaving circuit system and RMI
indications with the apP,ropriate heading references is maintained. In
order to accomplish this it is necessary, therefore, to provide
additional circuitry and devices that will control and annunciate
synchronized conditions and any departure therefrom.
The rate at which synchronization is carried out depends in the first
instance on the magnitude of the heading error, e.g. if it is less than
2 °, synchronization takes place at a slow rate of 1° -2 °/min, this
being the normal automatic slaving rate of typical systems. In the
event that errors are greater than 2 °, synchronization must take place
at a faster rate, and to achieve this, the circuits also include manual
and automatic control facilities.
A typical annunciator consists of a de micro-ammeter, the centrezero position of which indicates the synchronized state of the slaving
system. Depending on the type of MHRS, the instrument may be
incorporated within the RMI (see Fig. 8.14(a)) or within a control
panel as shown in Fig. 8.18. If the system becomes desynchronized,
the annunicator pointer will be deflected to one or other side of zero.
For example, if the servo CT synchro output signal to the RMI is
such that it produces an indicated heading that is less than the sensed
magnetic heading, the annunciator pointer is deflected to the left of
201
Figu" 8. /8 Control panel
annunciator.
ANNUNCIATOR
MODE SELECTOR
SWITCH
SYNCHRONIZING
KNOB
zero. Conversely, a deflection to the right signifies a greater indicated
heading. Annunciator pointers are also deflected during turning of an
aircraft as a result of the heading error signals produced by
displacement of the detector element. Under synchronized conditions,
the annunciator should, ideally, remain steady at its centre-zero
position. During flight, however, it oscillates slowly due to
pendulosity effects on the detector element, and this can in fact serve
as a useful indication that slaving is taking place.
In the monitored gyroscope.system described earlier (see page 192)
fast synchronizing is initiated by manually operating a synchronizing
knob which is also incorporated in the RMI. Referring to Fig.
8.14(a) again, it will be noted that the knob is marked with arrows
and signs which correspond to the deflected positions of the
annunciator pointer. Thus, if, as in the example already noted, desynchronization produces a deflection to the left of zero, the plus sign
signifies that the heading· indicated by the RMI must be increased to
regain synchronism. The synchronizing knob is therefore rotated in
the direction of its plus sign and, in so doing, it rotate&.the stator of
the RMI heading CT synchro (see Fig. 8.12) to induce a large error
signal in its rotor. This signal, after amplification by the servo
amplifier, drives the servomotor and compass card at a much faster
synchronizing rate, which typically is 300° /min. At the same time,
the servomotor drives the rotor to start 'nu!Hng out' the error signal,
and it also rotates the slaving CT synchro rotor to produce a slaving
signal for precessing the DGU into synchronism with the magnetic
heading reference established by the detector element. When the error
signal is reduced to 2 °. synchronizing takes place at the normal slow
rate of 1° -2 °/min. In dual MHR systems, the foregoing
synchronizing process is also activated when switching from one
system to another.
In the case of an MHRS/INS integrated system, the slaving circuit
has to be maintained in synchronism with both magnetic and inertial
platform heading references, and so the appropriate control circuit is
connected to that of the stepper motor (see Fig. 8.17). In response to
error signals produced by a de-synchronized condition, the stepper
motor controls the relative positions between the stator and rotor of
202
the CDX synchro, the resultant output of which drives the
servomotor to 'null out' the error signals in the manner already
described. Since the direction of motor rotation to attain synchronism
is automatically determined by passing error signals through a
polarity detector, it is not necessary to utilize a synchronizing knob
as in the case of a monitored gyroscope system. A slow
synchronizing rate of I 0 -2°/min up to 2° heading error is also
utilized, but for larger errors the stepper motor is driven at an
increased rate of 600° -800° /min.
The fast synchronizing rate is also activated when: power is
initially applied to the system; the source~ of heading references are
changed over as is possible in dual systems; a system is switched to
the slaving mode from the DG mode, when there is a valid inertial
platform reference signal, and the magnetic heading is greater
than 2°
MHFIS operating
modes
MHR systems provide for the selection of two modes of operation,
namely slaved and DG. The slaved mode is the one normally
selected, and provides for operation in the manner already described.
When operating in this mode, however, the accuracy of a system is
affected by the range of latitudes over which an aircraft is flown. The
reason for this is that the magnetic intensity of the earth's field
component H varies with latitude such that beyond 70° north or south
of the equator, it becomes an unreliable primary heading reference.
Such a reference would also be obtained if, regardless of latitude, a
malfunctioning of heading reference signal circuits were to occur.
Thus. as in the case of direct-reading compasses and direction
indicators (see page 125). an MHRS can also be selected to operate
in the DG mode to obtain a short-term stability reference irrespective
of magnetic field variations. Once selected, the heading information
displayed must be frequently updated in order to maintain its
integrity.
In a monitored gyroscope type of system, the selection of the DG
mode disables the slaving control circuit in the DGU so that its
gyroscope then functions as a basic direction indicator for controlling
the heading servo loop coupled to the compass cards of the RMI. and
the HSI of the flight director system. In an MHRS/INS integrated
system, the slaving control and stepper motor circuit in the compass
coupler is similarly disabled, and the heading setvo loop coupling the
RMI and HSI compass cards is controlled solely by inertial platform
heading references established by the azimuth gyroscope and synchro.
Heading selection
The method of selecting a magnetic heading- to be flown varies
between types of MHRS and their integration with other associated
203
Fi!iurc' 8. /9 Heading selection
- MHRS/INS.
'SET HOG' SWITCH 800-1200°/min (C·W rotation
of AMI compass card)
2
2
} C-W rotation of AMI
0
~c_-_40_0_
,_mi_n_ _ __,,_ compass card
1
From synchronizing
logic circuit
Fast or slow
L-----'r-..., synchronizing signal
To stopper motor
control, servo and
output synchro loop
200-400°/min
AC.W rotation of AMI
..__ _ _ _ _ _ _ _ _ _ _ _a_oo_-_12=0..c.0°"'"'/m""i.c..n_ _ _ _
., } compass card
systems. In a basic monitored gyroscope type (see Fig. 8.12) the RMI
is provided with a 'SET HDG' knob which, on being rotated,
positions a heading 'bug' relative to the compass card; it also
positions the rotor. of a heading data CX synchro whose stator is
connected to that of a CT synchro in the roll control module of an
AFCS computer. The resulting error signal corresponds in magnitude
to the selected heading, and after processing by the roll control
module the aircraft is automatically turned onto this heading. As the
DGU and RMI respond normally to the changing position of the
detector element, the heading data synchro rotor is repositioned to
'null out' the error signal supplied to the roll control module, and the
compass card is rotated to indicate the new heading with reference to
the heading 'bug'.
When a system is integrated with a flight director system, the HSI
provides the facility for selecting heading changes, and its operation
will be described in Chapter 9.
In the MHR/IN system thus far used as an example, heading
selection is accomplished· by a 'SET HDG' switch (see Fig. 8.19) on
the system control panel. The switch is supplied with direct current
and is of the limited-travel rotary type which is spring-loaded back to
its centre position. It has two positions left and right of centre, and
the contacts corresponding to these positions are connected to the
stepper motor control circuit in the compass coupler. When the
switch is selected to the first right-hand position, the de supply passes
through a resistor and causes the stepper motor to rotate at a rate of
200° -400° /min, and through the servo and output synchro loop, the
RMI compass card is rotated ir; a clockwise direction. In the second
right-hand position there is no resistance in the control circuit, and so
when selected the de supply rotates the stepper motor at a faster rate
of 800°-1200°/min. A similar operation results when the first and
second left-hand positions are selected, except that the compass card
of the RMI is rotated in an anti-clockwise direction. When the switch
is released, the system reverts to the normal slaved mode and slow
synchronizing rate.
Heading selections can also be made when the MHRS is operating
in the DG mode, except that on releasing the switch to its centre
position, the heading will remain at the 'set' position until the
aircraft's heading changes, at which moment the servo loop positions
the synchros to the relative heading change.
Deviation
compensation
Deviation is, as pointed out in Chapter 3, an error in heading
indication that results from the effects of hard- and soft-iron
components of aircraft magnetism (see page 87) on the detector
element of a compass. Although the detector elements of MHR
systems have the advantage over those of direct-reading compasses,
in that they are fixed in azimuth, and can be located at specifically
chosen remote points in an aircraft, they are not entirely immune
from extraneous fields that may be present in their vicinity. The
principal reason for this is that the element material has a high
permeability and so is very receptive to magnetic flux (see page 186).
Thus, any flux additional to that of the desired earth's field
component will displace the H axis of the material's BIH curve to a
false datum, and thereby induce heading error signals. It is therefore
necessary to incorporate deviation compensation devices in an MHR
system.
As far as compensation for the deviation coefficient A is concerned,
it is normally the practice to utilize detector elements of the preindexed type as referred to on page 190. In some early designs of
detector element, compensation was effected by rotating the element
in its mounting by the requisite amount, and referencing it against a
scale and fixed datum mark.
An electromagnetic method is normally adopted for the
compensation of deviation coefficients B and C, and this i:, illustrated
in the basic circuit diagram of Fig. 8.20. The potentiometers, which
are incorporated in a remotely-located compensator unit, are
connected to the pick-off coils of the detector element. When rotated
with respect to calibrated scales, they inject very small de signals into
the coils, so that the fields they produce are sufficient to oppose those
causing deviations. The output of the detector element is thereby
modified to correct the readings of the RMI via the synchronous
transmission loop. In some types of electromagnetic compensator,
provision is also made for coefficient A adjustment. This is achieved
by the inclusion of a differential synchro between the detector
element and the servo synchro within the RMI. When the position of
the differential synchro rotor is adjusted in the appropriate direction,
205
Figure 8.20 Electromagnetic
compensation.
TO COMPASS{
COUPLER
ANO/OR RMI
----,j~----------------------------'
$ DROPPING RESISTORS
Figure 8.21 Dual monitored
gyroscope system.
CAPT. HSI
{
COMPASS CARO
-~CO~M:P~A~S~C:A~R=O~---------,
PRE-SET HDG & COURSE
PRE-SET HDG } F/0 HSI
& COURSE
VHF NAV. 1
OMEGANAV.1
=~
OGU
No. 1
_ _ _ _..,,. VHF NAV. 2
FLIGHT RECORDER
OMEGA NAV. 2
,---..;1--+-o...._l
LI
I
DGU
No.2
__J
No. 2
TRANSFER RELAYS
1-----ll>i
AMPLlFIEf\
AMPLIFIER
MAG.HOG
F/0 RMI
it offsets the deviation by changing the magnitude of the heading
signals transmitted by the detector element.
Dual systems
The interconnection of dual systems of a typical monitored gyroscope
type is shown in Fig. 8.21. The transfer relays are controlled by a
Figure 8.22 Dual MHRS/INS.
F/0
HSI
CAPT
HSI
TRANSFER
RELAY
I
PLATFORM HOG
REFERENCE
{INS!)
COMPASS
COUPLER
COMPASS
COUPLER
No 1
No. 2
PLATFORM HOG
REFERENCE
(INS2)
DETECTOR
DETECTOR
AMI
RMI
panel-mounted selector switch, and in the 'normal' position :tie
systems operate independently of each other. In the eveni. of failure
of magnetic heading data input to one or other system, the operating
system can be selected to take over by energizing ,he t!ppropriate
transfer relay. For example, if the input to the capt.tin's or No. l
system should fail, the selector switch is moved to a position
placarded 'BOTH ON 2'., and so, as may be seen from the diagram,
the contacts of the No. 1 transfer relay change over to connect an
alternate data input from the first officer's or No. 2 system to the
captain's system.
The arrangement of a typical dual MHR/IN system is illustrated in
Fig. 8.22; this is also drawn to represent normal independent
operation. The method of transferilng magnetic heading data from
one MHR system to the other is similar to that described above. The
system also incorporates selector switches placarded 'RADIO' and
'INS', and as will be noted the~· are connected independently between
each INS and each flight director system's HSI. When the switches
are in the 'RADIO' positions each HSI is supplied with magnetic
heading data, while in the 'INS' positions these indicators are
supplied with true heading data from the respective inertial platforms.
Further details of these aspects of data transfer and switching will be
given in Chapter IO.
207
9 Flight director systems
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 components comprising a typical FDS, their connections and
signal interfacing with the systems providing essential navigational
data are shown in Fig. 9.1.
Vertical gyroscope
unit (VGU)
This unit performs the same function as a gy;:-o horizon, i.e. it
establishes a stabilized reference about the pitch and roll axes of an
aircraft. Instead, however, of providing attitude displays by direct
means, it is designed to operate a synchro system which produces,
and transmits, attitude-related signals to a computer (sometimes
referred to as a 'steering' computer) and to an amplifier unit. After
processing and amplification, the signals are then transmitted to
servo-operated indicating elements within a separate attitude director
indicator (ADI). The synchro system also supplies attitude-related
signals to the appropriate control channels of an AFCS. The
gyroscope and its levelling switch and torque motor system is
basically the same as that adopted in electrically-operated gyro
horizons (see pages 116 and 119).
The synchro system referred to earlier senses changes in pitch and
roil attitudes by means of a CX synchro positioned on each
corresponding axis of the gyroscope's gimbal 3ystem. The stator of
the roll synchro is secured to the frame of the unit, while its rotor is
secured to the outer gimbal ring. The pitch synchro has its stator
secured to the outer gimbal ring, and its rotor secured to the inner
gimbal ring. The stators supply attitude error signals to corresponding
CT synchros in the ADI, and also to pitch and roll circuit modules
of the computer.
Computer
This unit contains all the solid-state circuit module boards, or cards,
necessary for the processing of attitude reference and command
Figure 9.1 Typical FDS and
signal interfacing.
28V DC
Annunciator
Panel
115 V AC
VERTICAL
GYRO
Ptcch & rofl attitude reference
UNIT
Attitude
change
commands
AIR
DATA
COMPUTER
Altitude hold
Inst.
Ampl1her
Computer
~~~IGATION VOR:ILS beam signal deviations
RECEIVER
Signals
from
r-
{
Pre-sefec1 J
haading I
ground-based
transmitters
MARKER
BEACON
RECEIVER
Signals for s1arung ghde slope gain programme
Pre-select:
course
1
I
I
II
I
I
lI
I
f
I
I
Magnetic headtng reference
I
I
lI
I
8-------
I
I
I
True heading
-----------------------------J
D
FDS COMPONENTS
signals. Logic circuit boards are also provided for the purpose of
adjusting the scaling and gain values of signals appropriate to the
type of FDS and aircraft in which it is installed. In many cases it is
usual for a status code number to be quoted on the front panei of a
computer; this number relates to the required pre-adjusted scaling and
gain values which are listed in the form of charts in maintenance and
overhaul manuals.
Instrument amplifier
The primary function of this unit is to convert the attitude reference
and command signals supplied to it by the computer into servoactuating power signals for driving the display elements of the FDS
indicators. Like the computer, all circuits are of the solid-state type
contained on plug-in type module boards or cards.
Attitude director
indicator (ADI)
Figure 9.2 Altitude director
indicator.
This indicator, like the gyro horizon it basically resembles. provides
information on an aircraft's pitch and roll attitude. In addition,
however, and as may be seen froJ'!l Fig. 9.2, it .provides attitude
commands and information related to an aircraft's position with
respect to the glide slope (GS) and localizer (LOC) beams transmitted
by an ILS. A series of warning flags are also provided, and a ball-intube indicator provides indication of slip during turns.
The symbol representing the aircraft is fixed and is referenced
against a moving 'sky/ground' background tape on which are
presented an horizon line, and markings spaced at a specified number
of degrees to indicate pitch-up and pitch-down attitudes. The tape is
positioned around two rollers which, on being driven by a servomotor
and gear train, move the tape up or down as appropriate to the pitch
attitude change (see Fig. 9.3). The servomotor is activated by
amplified signals from a CT synchro connected to the pitch CX
synchro of the VGU, and its direction of rotation is determined by
Figure 9. 3 Pitch and roll
attitude indication.
Ring gear
0----
Tape
and rollers
Pttch motor
--0
Roll motor
•
the phase relationship between the CX synchro error signal voltages
and the servomotor excitation voltage.
Roll attitude is indicated on the roll attitude scale by the relative
position of a bank pointer. The pointer is fixed to the ring gear which
supports the pitch attitude tape rollers, and is also coupled to a
servomotor via a gear train. This servomotor is activated by
amplified signals from a CT synchro connected to the roll synchro of
the VGU, and as in the case of the pitch servomotor its direction of
rotation is determined by the phase relationship between error signal
voltages and servomotor excitation voltage. A differential gear is
provided in the drive system so that whenever there is a change in
roll attitude, the pitch attitude tape also rotates together with the roll
attitude pointer.
The GS and LOC indicating elements are respectively located at the
left and bottom of the ADI's attitude display. Each element consists
of a scale and a pointer that is deflected by miniature-type microammeter movements that respond to the appropriate beam deviation
signals supplied from an aircraft's VHF radio navigation receivers via
the FDS computer. The scales of the indicators are shown in a little
more detail in Fig. 9.4. The pointers represent the position of the
beams relative to an aircraft, and the pointer positions in relation to
the commands should be particularly noted. The dots on the GS scale
represent 75"µ.A (one dot) and 150µ.A (two dots) of aircraft deviation
above or below the beam (75µ.A ::: 0.35°; 150,uA = 0. 7°). The
LOC scale has a single dot to the left and right of the centre position,
each dot also representing 75µ.A. but in this case = 1" of aircraft
deviation. The LOC pointer is distinctly shaped to represent
211
Figure 9.4 GS and LOC
indicating elements.
Aircraft below
the beam;
'fly up'
command
Aircraft above
the beam; 'fly down' r - >
command
i..-
75 µA 1°
150 µA
0.7°
e
75 µA
0.35°
Aircraft lo left
of beam; 'fly right'
command
Aircraft to right
of beam; 'fly left'
command
e
symbolically the converged shape of the runway as it appears during
an approach. In some ADis the pointer is also deflected upwards by
signals from a radio altimeter to simulate the runway coming up to
the aircraft. At touchdown this 'rising runway' symbol, as it is
called, just touches the fixed aircraft symbol.
The QS pointer and scale come into view when the FDS is
op&ating in the GS mode and when there is a valid and reliable
signal from the VHF navigation receiver. At all other times it is
covered by a red warning flag controlled through the signal circuit.
The LOC pointer and scale are normally covered by a shutter which
is retracted when valid and reliable signals are supplied during the
VOR/LOC mode of operation of the FDS.
Two further red warning flags are provided, one placarded
'GYRO' and the other 'COMPUTER'. The 'GYRO' flag is
controlled by a monitoring circuit within the VGU and warns cf loss
of ~wer to this unit and any failures occurring in the attitude
r~f~r,~rt_Ce£ircuits. The 'COMPYTER' flag. monit.grs __ilE~ warns of
failures in inputs to the computer)tself, the instrument amplifier, and
ilieAt'>r····.
.
·-
-·-!ne·command bars, as the name implies, provide the ~()_I11Inands
relating io th~ changesJhl!Ll!!'..e. to~e. m~<:lt': to 111Jm~J!YJe an aircraft
info reguired pitch and/or roll attitudes. They are driven by_
servomotors that receive their input signals from the pitch and roll
channels of the computer via the instrument amplifier. Although the
two bars are not physically connected to each other, they move
together up or down to represent pitch commands, and tilt to the left
or right to represent roll commands (see Fig. 9.5). When the
commands being generated by the computer are satisfied, the bars
take up a position coincident with the top of the fixed aircraft
symbol.
212
Figure 9.5 Command bars
denection.
Fly up
Fly down
Fly left
Fly right
Commands satisfied
Operation
As far as aircraft attitudes are concerned, the ADI displays the
information in two ways: 1. as a primary attitude reference, and 2. as
command attitude changes.
Primary attitude reference
The interconnection between the VGU and the ADI are shown in
simplified form in Fig. 9.6.
When the gyroscope is operating and is stabilized with its axis
vertical, both CX synchros are at their 'null' positions and so there is
no output from their stator windings. The ADI will therefore indicate
a level flight attitude. If a displacement of the aircraft occurs about,
say, its pitch axis, the outer gimbal ring is also displaced and carries
the pitch synchro stator around its stabilized rotor. This produces a
displacement signal voltage which is then supplied to the stator of the
CT synchro in the ADI. Because at that moment an unbalanced
condition exists between rotors and stators of both synchros, then the
output from the CT synchro rotor is an induced error signal voltage
of a phase and magnitude appropriate to the direction and degree of
displacement.
As may be seen from the diagram, this error signal voltage is
supplied to the pitch channel of the instrument amplifier, and after
amplificatiQn it drives the corresponding servomotor in the ADI to
position the attitude tape upwards or downwards appropriate to the
aircraft's displacement about the pitch axis. At the same time, the
213
Figure 9. 6
Primary attitude
INST. AMPLIFIER
references.
ERROR $1GNALS
PITCH
ROI.L
SYNCHRO
PITCH
AXIS
ATTITUDE
DISPLACEMENT
SIGNALS
ROLL
AXIS
PITCH
SYNCHRO.
VGU
ADI
servomotor repositions the CT synchro rotor until displacement of the
aircraft ceases, at which point no further error signal is induced. The
servomotor also drives a tachogenerator which produces an output
proportional to the motor speed and out of phase with the error signal
voltage. This output is fed back to the pitch channel of the instrument
amplifier to provide damping of motor rotation and attitude tape
movement.
A similar operating sequence takes place when an aircraft is
displaced about its roll axis, except, of course, that the ADI' s attitude
tape is rotated left or right in response to signals from the roll CX
synchro.
If an aircraft is displaced simultaneously about the pitch and roll
axes, as for example in a climbi.ng tum, both channels will operate
and; by means of the differential gearing in the drive system, the
ADI's attitude tape will be positioned to indicate such a tum.
Command attitude changes
As mentioned earlier, command attitude changes are indicated by the
movements of the command bars in the required directions. The
method of driving the bars is shown in Fig. 9.7, and as will be noted
it is a little more complex than that adopted for primary attitude
references. The principal reason for this is that command signals can
214
Figure 9. 7 Command attitude
changes.
HOG.
SEL.
ROLL•
ATTITUDE
VOR/LOC
DEV"N
COMMAND FU
ROLL COMMAND
SIGNAL
ROLL
CHANNEL
FEEDBACK
COURSE
ERROR
MAN. PITCH
ATTITUDE
ALT.
HOLD
GS (AUTO)
DEV'N
FEEDBACK
PITCH COMMAND
SIGNAL
PITCH
CHANNEL
COMMAND FU
MAN.
GS
FD COMPUTE.A
• FOi' rec&ntring command bars aher
required attitude established
INST. AMPLIFIER
ADI
®
~
Summ,ng pc»nts
D1tf. gearing
be supplied from several external sources as governed by selected
modes of FDS operation. Details of these modes and their command
signals will be covered at a later stage, but for the moment we may
consider one of them, namely the 'manual pitch attitude mode', in
order to see how the signals are generated, and also how the
command bars are operated.
The manual pitch attitude mode is one which enables a pilot to
select a pitch attitude reference at any time there is no other mode
selected for controlling about the pitch axis. An example of this
would be the selection of a desired climb attitude that is to be
maintained after take-off. The selection is made before take-off by
rotating a pitch command knob on the FDS mode controller. The
control knob also rotates the rotor of a synchro which then supplies
an error signal to the pitch channels of the computer and instrument
amplifier as shown in Fig. 9.7. The amplified signal then drives a
motor in the ADI to position the command bars above the fixed
aircraft symbol. At the same time, the motor drives a tachogenerator
that provides a rate feedback signal for motor speed control purposes,
and it also positions the rotor of a CT synchro. The output signal
from this synchro serves as a follow-up to the command signal, and
to null it out so that the motor and the command bars are stopped at
the selected pitch up command position. When the aircraft takes off
and rotates to the desired climb attitude, the pitch attitude tape of the
ADI will be positioned to indicate the climb in the manner explained
earlier, and the command bars will be recentred over the fixed
aircraft symbol.
215
Figure 9.8 Manual pitch
attitude mode.
6. After aircraft gets to
desired altitude and
levels, bar deflects up.
Recentre by turning pitct
command knob until bari
centred on fixed airplane
reference.
1. Aircraft on the
ground
2. F/0 mode selector to
HOG mode (typical
outbound heading)
3. Pilch command knob
to desired climb
attitude
4. Command bar movas
up
5. When aircraft rotates
to desired climb attitude
attitude command bar
centres
After the aircraft reaches its required altitude and levels off, the
command bars deflect upwards, and are then recentred by turning the
pitch command knob back to its zero position. The foregoing
sequence is shown pictorially in Fig. 9.8.
Horizontal situation
Indicator (HSI)
This indicator derives its name from the fact that its display, as can
be seen froni Fig. 9.9, presents a pictorial plan view of an aircraft's
situation in the horizontal plane in the form of its heading,
VOR/LOC deviation, and data relating to flight to and from a VOR
station. In addition, it displays deviations from the GS beam and
distance from a distance measuring equipment (DME) station.
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. This bar is also
sometimes called a lateral deviation bar. Selector knobs at the botrom
corners of the indicator permit the setting of a desired magnetic
heading and a VOR/LOC course.
Heading display
The primary display element of the indicator is that related to an
aircraft's magnetic heading, and so it is integrated with a magnetic
216
Figure 9. 9 Horizontal si•uation
indicator.
heading reference system (MHRS). In aircraft equipped with an
inertial navigation system (INS) the indicator is also integrated with
the computer of that system so that it can be selected to displa)' either
true heading or magnetic heading.
Figure 9.10 shows the arrangement of the heading display section
in simplified form. It consists of an azimuth or compass card which
is mounted on a ring gear driven through a gear train by a
servomotor. Headings are indicated by the position of the card with
.E~~~o. a fixed lubber line. The servomotor also drives a
tachogenerator that provides rate feedback signals for motor speed
control. In order to select a magnetic heading, a heading marker is
provided, and can be positioned relative to the compass card by
rotating the heading selector knob. The differential gear shown is to
permit relative movement between the marker and card, and also to
allow the marker to rotate with the card when a change in heading
takes place.
Heading signals are supplied from the MHRS via its RMI (or
217
Lubber line
Figur,- 9.10 Heading selection
and display.
I
Compass or
azimuth
card
Heading
change
Signals
I 3'
I
Feedback
I
I
I
I
$r------,g····-Differential
gear
~!command
@
signals
Heading error
synchro
~
Selector knob
compass coupler unit in the case of integration with an INS) and as
will be noted they a;e fed to a CT azimuth synchro in the HSI. On a
constant heading, the synchro is at 'null' with that of the MHRS, and
so the compass cards of both the RMI and HSI indicate the same
heading.
When it is required to change an aircraft's heading, the FDS is
operated in the heading (HDG) mode to provide roll commands. To
select a heading change the heading select knob is rotated to position
the heading marker against the corresponding graduation mark on the
compass card. At the same time, the rotor of a heading error CT
synchro is rotated to produce an error signal in its stator. This signal
is supplied to the roll channel of the FDS computer and then to the
ADI as a roll command. The command bars are therefore deflected in
the manner already described (see Fig. 9.5), to indicate the direction
in which the aircraft is to be turned onto the new heading.
As the aircraft turns, a heading change signal is produced by the
MHRS to rotate the RMI compass card. The signal, as shown in Fig.
9.10, is also supplied to the azimuth CT synchro in the HSI, but as
the rotor of this synchro will, at that moment, be desynchronized in
relation to the MHRS synchro, it produces an error signal. This
signal is then amplified and supplied to the servomotor that is
coupled through a gear train to the compass card. Thus, the card and
also the heading marker are rotated in the appropriate direction. It
218
will also be noted from Fig. 9.10 that the motor drives the azimuth
CT synchro rotor; this is done to reduce the error signal
progressively as the aircraft turns onto the selected heading. In other
words, the motor brings the synchro into synchronism with the
MHRS so that the HSI compass card 'repeats' the heading indication
of the RMI. When the aircraft has levelled off on the new heading,
synchronism is attained and the heading is indicated by alignment of
the heading marker with the lubber line.
Aircraft position with respect to VOR/LOC beams
The secondary section of an HSI display relates to an aircraft's lateral
position with respect to a VOR station and to the localizer beam of
an ILS; the basic arrangement of this section is shown in Fig. 9. I 1.
consists of a deviation bar that can be deflected left or right over a
scale plate, and in relation to a course pointer or markii} The
deflections perform a directive command function, i.e. a deflection to
the left is a 'fly left' command to capture a beam, and a deflection to
the right is a 'fly right' comman{·
.
Each dot shown on the scale corresponds to a I O deflection~CThe
bar is deflected by a de meter movement supplied with signals fron,
the radio navigation receiver. When there are no deviation signals
Fix11re 9.1 I Course selection
and display.
Course marker
Selected course
counter
II II
O
O
O
I
Meter
Righi
Phase
comparison
Left
Phase
shift
Course
datum synchro
I
J
~
~
I
~
Selector knob
219
present, the bar is aligned with the course marker as shown. In
addition to deflection, the bar, its scale and course marker can be
rotated: (i) relative to the compass card whenever the course to a
VOR station or ILS localizer is selected, and (ii) rotated with the
compass card when the aircraft turns onto the selected course.
Selection of a desired VOR radial. or localizer course is carried out
by rotating the course selector knob until the course marker coincides
with the desired value on the compass card. The deviation bar, its
meter movement and scale also rotate with the marker. At the same
time, the control knob drives a digital counter to the corresponding
course indication. The gear train comprising the drive from the
selector knob is coupled to the rotor of a course resolver (RS)
synchro associated with the VOR/LOC navigation receiver, and to the
rotor of a course datum CX synchro. Both rotors are, therefore, set
to some angular position with respect to their stators when the course
selector knob is rotated.
In the case of selecting a course to fly onto a desired VOR radial,
changing the position of the course RS synchro rotor causes it to shift
the phase of the low-frequency (30 Hz) reference modulating signal
received from the station by the aircraft's radio navigation receiver.
The signal is then compared with the station's variable modulating
signal (also 30 Hz) in a phase comparator circuit, the output of which
is supplied to the meter movement to deflect the deviation bar left or
right as appropriate.
The course datum CX synchro performs the same function as that
of the heading error synchro, i.e. it produces a roll command signal
which is supplied to the ADI command bars to deflect them in the
direction in which the aircraft is to be turned to intercept the VOR
beam radial. As the aircraft turns in response to the roll command,
the compass card rotates in response to the signals sensed by the
MHRS. and through the differential gearing the card also rotates the
deviation scale, the bar and the course marker. When the beam is
being approached, the signals from the RS synchro, and the phase
comparator, to the deviation bar meter movement are being reduced
and so the bar deflection is towards the fixed aircraft symbol. The
output from the course datum CT synchro also changes to deflect the
ADI command bars in the opposite direction, thereby commanding
that the aircraft be rolled out to a w'ings-Ievel attitude and on course
to the VOR station.
Flight along the beam is indicated by alignment of the deviation bar
with the course arrow. The effects of any cross-winds during the
flight along the beam are automatically corrected by a compensating
'wash-out' circuit in the FDS computer, to establish the 'crab angle'
necessary for the aircraft to stay on course. This angle is also
indicated on the HSI by the positio:i of the deviation bar relative to
the fixed lubber line.
220
To-from indicators
Indication of whether an aircraft is flying to or from a station is
provided by an arrow-shaped marker that is positioned by a de meter
movement. The meter is supplied with signals from a phase
comparator circuit when the VOR station frequency is tuned in.
Referring to Fig. 9.10 once again, the marker is positioned in the
direction of the course marker, thus indicating that the selected
course is to the station selected. When an aircraft flies from the
station, the meter movement deflects the arrow through 180°. In
some types of FDS, the HSI has two separate meter movements and
arrow-shaped markers.
LOC mode
When the FDS is operating in the LOC mode, the HSI functions in
the same ma.nner as that just described for the VOR mode, but with
two exceptions. Firstly, the output to the meter movement controlling
the deviation bar results from amplitude comparison of signals either
side of localizer beam centre, and secondly, the to-from arrow
remains out of view since no to-from signals are transmitted in the
LOC mode.
W~ning flags
As in the case of the ADI, red warning flags are provided in an HSI
to indicate that GS, LOC and VOR signals are unreliable or have
completely failed. In addition a flag is provided to give similar
indications in the case of the MHRS.
DME indicator
This indicator receives signals from the interrogator of the distance
measuring equipment (DME) carried in an aircraft, and displays the
distance in nautical miles to be flown to a selected DME ground
station. If the system is not valid the indicator display is obscured by
an electrically-operated shutter.
Radio altitude
In some HSis an indicator light is provided and is connected to a
radio altimeter system such that it illuminates when an aircraft
reaches a specified minimum altitude, referred to as a 'decision
height', during the final stages of an automatically-controlled
approach.
221
Figure 9.12 Mode controller.
VOR AUTO
LOC APP
ON
&
0
OFF
ALT HOLD
MODE SEL
Mode controller
PITCH CMD
A typical controller is shown in Fig. 9.12; it provides three principal
control functions: (i) mode selection, (ii) altitude hold, and (iii)
manual pitch command.
Mode selection is done by means of a rotary selector switch which
can be manually placed in any of the indicated positions. In all
positions except GA (go-around) a pin drops into a detent to maintain
the switch in position. If the switch is in either the 'AUTO APPR' or
the 'MAN' position, and the GA manoeuvre has to be carried out
(see page 227), the pins are automatically retracted from their
detents, and a spring returns the switch to the GA position.
The six positions of the switch and the system's response in each
case are as follows:
·
GA
OFF
HDG
VOR/LOC
Causes the ADI command bars to display a pitch-up
command and a wings-level attitude.
Disables the pitch and roll outputs from the computer
causing the ADI command bars to be deflected out of
view.
Allows selection of a magnetic heading which, as
indicated earlier, is done by means of the selector
knob on the HSI. While in this mode, the computer
pitch channel can be operated in the 'MANUAL
PITCH' mode by means of the pitch command
selector, or the 'ALT HOLD' mode by attitude
command signals generated and supplied by an ADC.
Allows selection of a VOR radial or a localizer beam
for lateral guidance, so that input command signals
can be produced in the manner already described. The
pitch channel of the computer can be operated in
either the 'MANUAL PITCH' mode or the 'ALT
HOLD' mode.
AUTO/ APPR Allows the use of the ILS beams (LOC and GS) for
lateral and vertical guidance during an automatic
approach. In the roll channel, the input commands are
LOC deviation, course error and roll attitude. The
pitch channel is also armed for capture of the GS
beam. Before capture of this beam, however, the pitch
channel can be operated in the 'MANUAL PITCH' or
'ALT HOLD' modes. At GS capture, both these
modes are automatically cancelled, and pitch command
signals are generated within the computer by a
combination of GS deviation and pitch attitude signals.
MAN/GS
This mode is selected for use under three different
conditions: (i) to establish a fixed· intercept angle of
either a VOR or LOC beam; (ii) to force a GS beam
capture condition when making an approach which is
above the beam; and (iii) to force a LOC or GS beam
capture condition if it is known that the beam-sensing
circuits of the computer are inoperative.
The function of the 'ALT HOLD' switch is to establish a command
signal that positions the ADI command bars in pitch so that they are
referenced to the altitude sensed by an ADC. The switch ic he'.d in
the 'ON' position by a solenoid provided valid signal conditions
exist. If the switch is on at GS capture, it will automatically c-e
returned to the 'OFF' position by de-energizing of the solenoid.
Further details of the altitude hold function will be given later.
As described earlier, the 'PITCH COMMAND' knob permits
selection of a desired pitch attitude of an aircraft when tile FDS
computer is operating in the 'MAN PITCH' mode. In a typical
system, a minimum of 15° of pitch-up command or 10° pitch-down
command can be generated for the input to the pitch channel.
fDS/AFCS mode
controllers
The basic attitude and navigational data displayed by the indicators of
an FDS, together with the selection of operating modes, are features
that are also common to the operating requirements of an AFCS.
They can, therefore, be developed as a natural complement to each
other, and by matching their characteristics with those relating to the
aerodynamics and flight control of any one type of aircraft, they can
be fully integrated to serve as an overall flight guidance system. 1\s
far as the operation of such a system is concerned, one of the items
of FDS 'hardware' having a common purpose is the mode controller.
In some systems this may be provided as a separate unit (as, for
example, the one shown in Fig. 9 .12) for use in combination with the
control panel of an AFCS. It is, however, a ,more logical approach to
223
Figure 9.13 FDS/AFCS mode
c-ont roller.
PRE SET COURSE
CONTROLS & INDICATORS
CAPTAIN'S
FLIGHT DIRECTOR
ON-OFF SWITCH
F/O'S
FLIGHT DIRECTOR
ON -OFF SWITCH
OOURSE
SELECT
SWITCH
F/O'S PTW
_n~ fi}
-
AUTOPll.01' IINOAe!
PRE SET HEADING
CAPTAIN'S
FLIGHT DIRECTOR
PITCH TRIM WHEEL C'1'Wl
SELECT KNOB &
INDICATOR
NAV MOOE
SELECT
SWITCH
ALT HOLD/ALT SELECT
SWITCH & GREEN
LIGHTS
ALTITUDE SELECT
KNOB & INDICATOR
eliminate the panel by combining its selective functions and switching
circuits within the AFCS control panel itself, and this is a method
now adopted in the majority of present-day systems. Figure 9.13
illustrates an example of an FDS/AFCS mode controller, or mode
select panel, based on that used in some series of Boeing 747. The
controls annotated perform functions associated with the FDS as
follows:
Flight director switches
Switch on an output to the command bars
of their respective ADis, and also control
power to the annunciator panel. They do
not control power to the computers.
Pitch control wheels
Control pitch command bars of their
respective ADls when no other pitch
mode has been selected. They are
disabled electrically when a pitch mode is
captured.
Heading selector
This replaces the selector that is normally
provided in an HSI, and performs the
same function, i.e. it establishes errcr
signals for the purpose of roll control,
and also drives the heading marker in the
HSI.
Nav. mode selector switch Apart from the 'INS' and 'LAND'
positions, which are specific to the
aircraft concerned, the functions of the
other positions are the same as those
described earlier, namely, 'HDG' selects
the signals from the MHRS, 'VOR/LOC'
224
Course selectors
Course transfer switch
flight mode
annunciators
provides appropriate deviation signals
from the VHF navigation receivers, and
'ILS' provides localizer and GS deviation
signals from the receivers.
These are connected, one to each of two
MHR systems provided in the aircraft
concerned, and perform the same function
as the selector normally provided in an
HSI, i.e. they establish course error
signals for roll control, and also drive the
HSI course pointer.
This is a three-position switch for
determining which MHRS and radio
navigation receiver are to be used for
supplying signals to the roll channels of
the FDS/AFCS computers. In either the
No. l or No. 2 positions, the switch is
held in the selected position by a solenoid
whenever the nav. mode selector switch is
at either the 'INS', 'HDG' or
'VOR/LOC' position. If the 'ILS' or
'LAND' positions are selected, the switch
is spring-returned to the 'MULTI'
position, in which the distribution of
signals between computers is varied, and
outputs from a third VHF navigation
receiver are introduced.
The purpose of these units is to annunciate, by coloured lights, the
conditions appropriate to the flight modes selected on a mode
controller. The lights are grouped so that those at the left of the panel
relate to an FDS, while those at the right relate to an AFCS. Two
examples are shown in Fig. 9.14.
The annunciator illustrated at (a) is referred to as an approach
progress display and, as will be noted, the FDS group consists of
VOR/LOC, glide slope and go-around (GA) lights. If a VOR or LOC
mode has been selected, the VOR/LOC light illuminates amber, and
it remains on untii the appropriate beam has been captured, when it
then changes to green. The glide slope light illuminates amber when
the approach mode is selected, then changes to green when the GS
beam has been captured. The GA annunciator light illuminates green
whenever the GA mode is selected as a result of an automatic
approach having to be aborted. In addition to VOR/LOC and glide
slope, the AFCS group of the display has annunciator lights for
altitude select and altitude hold modes.
Figure 9. I 4 Flight mode
annunciator?t.
GA
ALT SEl
ALT HOLD ;::•:
(a)
F/D
A/?·B
ALT SELECT
ANNUNCIATOR
A/P·C
P/RST
A/P
I
TRIPLE
A/T
P/RST
TEST
TEST 1 1 - - - AMBEII
GREEN 11---Hl'IE
TEST
NAV
RED
BLUE
G/S
NA/ MODE
ANNUNCIATOR
A/P
PHOTO-CELL
DIMMING
GLIDE SLOPE
ANNUNCIATOR
GO-AROUND
ANNUNCIATOR
FLARE
ANNUNCIATOR
(b)
The lights may be checked for functioning by individual press-totest facilities, and provision is also made for bulb replacements.
The annunciator shown in (b) of Fig. 9.14 is of the type used in
conjunction with the FDS/ AFCS combination adopted in some series
of Boeing 747 (see also page 224). The lights are so arranged that
captions appropriate to each mode are illuminated white whenever a
mode is selected, while the lower half of ea..:h light is illuminated
green when the desired funccion has been satisfied. The arrow-shaped
symbol of the go-around annunciator light is illuminated green when
the associated manoeuvre has been initiated by activation of goaround switches on the engine thrust levers. Testing of the white and
green lights is done by depressing the cap of the FDS 'ALT/S'
annunciator; this also checks amber lights in the AFCS and
autothrottle annunciators in the top row of the panel. This group of
annunciators also contains red lights which, together with a blue light
in the 'TEST' annunciator, may be checked by depressing the cap of
the AFCS 'ALT/S' annunciator. Dimming of the lights is controlled
226
by two photocells which operate in conjunction with a separate
master switch when selected to a 'dim' position.
Go-around switches
When an automatic approach has to be aborted, engine thrust has to
be increased in order to climb the aircraft into what is termed a 'goaround' (GA) manoeuvre. The FDS also has to go into the GA mode
of operation, and this is done by switches (one for each pilot's FDS)
located on the engine power levers as shown in Fig. 9.15.
The switches are of the 'momentary press' type so that they can be
conveniently operated by the palm of the hand as the power levers
are pushed forward. Closing of the switches causes the mode selector
switches on the mode control panels to move to the 'GA' position
from either the 'AUTO APPR' position or 'MAN GS' position,
whichever mode was in operation at the time. This action places the
computer in the GA mode, and it then provides pre-set bias signals to
each pilot's ADI, causing the command bars to deflect to a 'pitch-up'
and 'wings-level' command position. The amount of command bias
required varies between different types of aircraft and their flight
characteristics; in one particular type the command is 14 ° pitch up.
Figure 9.15 Go-around
switches.
227
Figure 9.16 Heading select
commands.
~ @+t8
4-
@++·8 +"
r,:;-,.
COMMAND BARS
3
~ \::!,/CENTRED
AS NC
'lllllllillllilD~©---({)---~\
°'"
t.\ NC
HOG 090°
INDEX 090°
FD IN HOG MODE
COMMAND BARS
CENTRED
\V SET HOG
OO<
®
SET HOG INDEX 180°
COMMAND BARS
INDICATE 'BANK RIGHT'
~
,rnnme CH'"G'°
-~ll';.
BYCORRECTAMOUNT
~1 ©
I 4
~
I
I
-t8 CD
1
II
081
,..
.,-
Operating sequences
AS A/C APPROACHES
NEW HEADING
COMMAND BARS
DEFLEC.T IN OPPOSITE
DIRECTION.
8 NC
~g~~X'~~GR~~~L
TS IN
ROLLING OUT
WINGS LEVEL ON NEW
HEADING.
Heading changes
Figure 9.16 illustrates in pictorial form the sequence of ADI and HSI
display presentations (as would be seen in the direction of flight)
when, with the FDS operating in the 'HDG' mode, it is required to
change an aircraft's magnetic heading. For purposes of explanation it
is assumed that the change is· to be from 090° to 180°.
Position I The HSI heading marker is aligned with the lubber line to
show the present heading of 090°. There are no command
signals to the ADI and so it displays a level flight attitude.
Position 2 The heading select knob is now rotated to position the
heading marker at 180° on the compass card. Since the
aircraft must be turned to the right to attain this heading,
the heading error CT synchro in the HSI sends a
command signal to the ADI, via the computer roll
channel, and in response the command bars are deflected
to the right.
Position 3 The aircraft is headed into the turn and, at the correct
bank attitude, the command bars are centred and the ADI
then indicates this attitude. The MHRS detects the heading
change, and its signals to the HSI produce rotation of the
compass card and marker towards the lubber line in the
manner described on page 218. The compass card rotates
in the opposite direction to the turn because the MHRS is
continuously being slaved to magnetic North. As the
228
aircraft approaches the new heading, and since it must be
rolled out to a wings-level attitude, the ADI command
bars now receive a signal that deflects them lo a 'fly left'
command position.
Position 4 The aircraft has rolled out to wings level as shown by the
ADI display, and is now flying on the selected heading as
indicated by alignment of the heading marker and lubber
line of the HSI.
Flying to a VOR station
Assuming that an aircraft is flying on a magnetic heading of 180°,
and that it is then required to fly to a VOR station on a transmitted
beam radial of 090°, the display sequence would be as shown in Fig.
9.17.
Position 1 The aircraft is being flown with the FDS operating in the
'HDG' and 'ALT HOLD' modes, and so the heading
marker of the HSI is aligned with the lubber line and
compass card to indicate the magnetic heading of 180°
The ADI indicates the level flight attitude.
Figure 9.17 Flying to a VOR
station.
HS,
~;~i
Amber
... ,,1•••'
l
0)
Heading marker
@
229
Position 2
Position 3
Position 4
Position 5
In order to fly onto the required VOR beam radial, the
course selector knob is rotated to set the course marker
and deviation bar at 090°. The VOR/LOC mode is then
selected on the mode controller, and the corresponding
light on the appropriate mode annunciator unit will be
illuminated~ e.g., on an approach progress display type of
unit, the VOR/LOC light will illuminate amber. At the
same time, , he deviation bar will respond to the signals
transmitted by the VOR station, and it will be deflected to
the extreme right of the course marker to give advance
indication of the fact that after turning towards the station,
the aircraft must subsequently 'fly right' to attain the
selected course.
The beam is captured, the annunciator light illuminates
green, and magnetic heading control is automatically
switched out since control is taken over by the VOR. The
course datum synch10 in the HSI supplies a 'fly left'
command to the ADI command bars. At beam capture, the
course error and beam deviation signals are summed in
the computer roll channel circuit so that as the aircraft
turns, the selected course, i.e. the beam centre, is
intercepted at a specific angle which, typically, is 20°.
The to-from arrow in the HSI is activated to indicate
flight to the station.
The HSI indicates the turn to the left by relative
movement between the lubber line a.nd compass card, and
the deviation bar deflection is progressively reduced to
indicate that the aircraft is flying towards the beam centre.
The ADI command bars are centred to indicate that the
correct bank angle for the turn is established.
The aircraft is approaching beam centre, and the ADI
command bars are deflected to the right, commanding th.1t
the aircraft be rolled out in this direction to a wings-level
attitude.
The HSI and ADI indicate the final courze situation.
If the aircraft encounters a cross-wind when 'on course', the effects
are compensated by a 'wash-out' circuit in the computer, and the
crab angle necessary to stay on course is established and indicated on
the HSI.
LOC mode
Operation in this mode relates to the approach of an aircraft to an
airport runway along the beam transmitted by the ILS localizer
transmitter. The mode is selected on the appropriate mode controller;
Figure 9. I 8
--- --- --- ---
Localizer
approach commands.
~-=-=-=--
- -- --
LOC
TX
AT 090°
BEAM CENTRE
®
ADI COMMAND BARS
DEFLECTED DOWN TO
COMMAND DESCENT
ONG/S
®
, \
\ \ I I I .,
-=, IVORILOC I:=
CD
[email protected];~;~~"
ADI
I
A/C HOG
1ao•
HSI
in the case of the unit shown in Fig. 9.12, the selector knob is set at
the 'AUTO APPR' position. Although it is called a LOC mode, it
also provides the control commands for capture of, and flight along,
the GS beam since on selection it automatically 'tunes in' to the GS
transmitter.
Figure 9 .18 illustrates, also in pictorial form, typical localizer
approach display sequences, and their similarity to those of the VOR
mode is very noticeable. The principal differences are: (i) the
changeover at position 1 from 'HDG' mode to 'AUTO APPR' mode;
(ii) the localizer capture point (position 2) varies with rate-of-change
of beam deviation, whereas capture of a VOR beam radial is at some
fixed value, typically 5° deviation; and (iii) the •ALT HOLD' mode
is automatically switched out when the GS beam is captured, for the
obvious reason that the aircraft must subsequently commence its
descent for landing. At position 5, the ADI command bars are
deflected to a 'fly down' command position in order to descend on
the glide slope.
The profile of the GS stage of approach is shown in Fig. 9 .19.
Prior to beam capture, 'ALT HOLD' is on, and the GS annunciator
231
Figure 9.19 Glide slope
approach profile.
9
T----
LA AA 1500 FT SIGNAL
STARTS GAIN PROGRAM
ND. 1
MIDDLE MARKER OR
LRAA 200 FT
SIGNAL STARTS
GAIN PROGRAM
NO. 2
AUTO Af'P SELECTEO
ALT HOlD ON
GS AMBER
INPUT COMMAND SIGNALS
ARE ALTITUDE ERROR (CAOC)
AND PITCH ERROR (VG)
,,,. /
.,,- .r
..,,.GO AROUND
GA GREEN
INPUT COMMAND
SIGNALS ARE PITCH
UP (FD COMPUTER) AND
PITCH ERROR !VG)
GLIDE SLOPE
TRANSMITIER
GS· CAPTURE 17.Sl'a
AlT HOLD OFF
GS GREEN
INPUT COMMAND SIGNALS
ARE GS DEVIATION (VHF NAV)
:Z0 PITCH DOWN (FD COMPUTER)
ANO PITCH ERROR (VG)
NOTE:
NAVIGATION RECEIVER
OUTPUT SCALE FACTOR
• 214 µ1\/DEGREE
light illuminates amber. At beam capture, the light illuminates green,
'ALT HOLD' is automatically disconnected, and a pitch-down bias
signal from the computer pitch channel is supplied to the ADI to
deflect the command bars 2° down, thereby commanding that the
aircraft's attitude be changed from that established at the time.
During the descent along the beam, any deviations from it are
provided by the GS pointers in the HSI and ADI, and they are
corrected by flying the aircraft in the directions commanded by the
pointers.
As can be seen from Fig. 9.19, the GS beam converges towards
the runway and this means that as an aircraft gets closer to landing,
less pitch control is required to counteract deviations from the beam.
To allow for convergence, therefore, the gain or response to GS
deviation signals as a function of time is automatically reduced by
gain-programming circuits in the pitch channel of the FDS computer.
Programming is carried out in two stages, and is activated by signals
from an aircraft's radio altimeter. The first stage is activated at about
1500 ft radio altitude, and activation of the second ·stage is at about
200 ft. Gain reductions against time (based on a typical system) are
shown pictorially in Fig. 9.20.
The go-around manoeuvre indicated in the diagram is established
for_ the reasons, and in the manner, already described (see page 227).
232
GP1
Figure 9.20 Gain
programming.
I
I
43
#
21
GP2
I
I
z
~
~·~-J
1 ~,,_
TIME
Figurt 9.21 · Altitude hold.
---- -~- -
--- -----
8
8
8
2
1
Altitude hold
In a typical FDS, this mode may be utilized when the system is
selected for operation in either .the 'HOG', 'VOR/LOC' or 'AUTO
APPR' modes.
When an aircraft is flying at the altitude it is required to maintain,
as at position I in Fig. 9.21, the altitud~ switch is placed in the 'ON'
position and is held there by a solenoid. At the same time, a 'hold'
signal is supplied to the ADC. In the case of an analog type of ADC,
the signal energizes a clutch between the altitude module sensing
Fi11ure 9.23 FDS with
auxiliary VGU.
VE RT I CAL
GYRO
NO. 1
l'ITCH & ROLL ATTITUDE
~
NORMAL
t
CAPT
ON
AUX
I
._
CAPT
ADI
VG XFR
Al Y NO. 1
AUX VO
AUXILIARY
PITCH & ROLL
VERTICAL 1--:..;;ATT.:..:.:.;IT;..;:U.;;.D.::.E- GYRO
INSTRUMENT
COMPARATOR
SYSTEM
VG XFR
SN
L-,
I
I
I
F/0
ON
PITCH & ROLL
ri~~~CAl l.-A~T~Tl!:IT~U~D'.::E~~~~~-1---1~,_A_u_x~~~~~~~t--~~~---+1
t
NO. 2
NORMAL
VG XFR
RELAY NO. 2
I
F/0
ADI
LIGHT
DIRECTOR
YSTEM • 2
these lights are placarded 'HDG', 'PITCH', 'ROLL', 'GS', 'LOC'
and 'ALT'. Failure of a comparator power supply is indicated by a
red 'PWR/MON' light which is also connected in parallel with a
corresponding light on the annunciator panels. A momentary 'PUSH
TO TEST' switch is provided for checking the serviceability of all
the lights (except the 'PWR/MON' light). The switch is paralleled
with a test switch on a flight deck panel to permit similar checks to
be carried out by the crew.
The annunciator panels are connected in parallel and they each
contain seven amber lights placarded as indicated in Fig. 9.25. If a
light illuminates, it can be dimmed by pressing its cover.
Operation
The operation of the annunciator panel lights is as follows:
l. HDG
236
The comparator supplies 26 V ac excitation to differential
resolver synchros in the captain's HSI and ADI. The
captain's HSI transmits heading signals to the first
officer's HSI. If the positions of the compass cards in both
indicators do not agree, the first officer's HSI transmits an
error signal to the comparator. This, in turn, produces a
signal to illuminate the 'HDG' lights on both annunciator
panels.
Figure 9.24 VHF nav. transfer
TEST
,witching.
~1------~•..
TEST SWITCH
EXCIT
HOG
EXCIT
PITCH
PITCH
ROLL
ROLL
F/0 ADI
GS CAPTURE
F/D STRG CMPTR•1
GS CAPTURE
FLT INSR
ACC UNIT
F/D STRG CMPTR•2
V/L DEV
I
of
I
NAV TRANSFER
RELAY ND. 1
-------D
GS DEV
GS S FLAG
1.....;;;28;..v_oc_,_Ls_....
VHF NAV 1
BOTH ON 2
TO INSTRUMENT
COMPARATOR AND
ANNUNCIATORS
I
_.}----1--:---a:+
NORM
--IN_T_E_R_N_A-LS
SAME AS 1_
VHF NAV 2
I
0 I
I
I
o, I
BOTH ON 1
NAV TRANSFER
RELAY NO. 2
fliure 9.25 Warning
comparator and annunciator
panels.
CAPTAIN
HEADING
ATTITUDE
VHF NAV
RAD.ALT
F/0
237
2. ROLL The captain's ADI transmits roll attitude signals to the
first officer's ADI. If the attitude indications do not
agree, the latter ADI transmits an error signal to the
comparator and so produces a signal to illuminate the
'ROLL' lights on both annunciator panels. Since a roll
attitude change initiates a change in heading, a separate
roll attitude signal input is supplied to the comparator to
modify the threshold at which the comparison between
heading indications operates (the greater the roll angle the
larger the heading threshold). This helps to reduce
'nuisance' warnings due, for example, to precession of an
MHRS directional gyroscope unit, and system tracking
during manoeuvring of an aircraft.
3. PITCH This light is illuminated in a simlar manner to that for
'ROLL'.
4. LOC
Illumination of this light occurs whenever ILS localizer or
VOR deviation signals from navigation receivers do not
agree. A 28 V de (ILS) signal is also supplied to the
comparator by the receivers, in order to activate the
comparison between deviation signals.
5. GS
This light is illuminated whenever ILS glide slope
deviation signals from the navigation receivers are not in
agreement. In addition to the 28 V (IL8) signal noted
above, a 'GS flag' signal is supplied from both receivers
to activate the comparison between GS deviation signals.
6. ALT
This light is illuminated whenever the altitude signals
from dual radio altimeter transmitter/receivers do not
agree.
Typical threshold levels required to illuminate the annunciator lights'
are given in Table 9.1.
If the excitation voltage from the comparator unit is lost, the
'HDG', 'ROLL' and 'PITCH' annunciator lights may also illuminate
simultaneously. In such a case the comparison between the signals
from the reference sources is, of course, unreliable.
FDS interfacing with
inertial navigation
systems {INS)
As in the case of an MHRS (see page 198), an FDS can also be
interfaced with an INS so th;it its ADI can then utilize the attitude
references established by th·; gyroscopically-stabilized inertial
platform, instead of it being dependent on a VGU. The
interconnection arrangement is, as far as attitude error signal
transmission is concerned, basically similar to that described earlier,
and as may be seen from Fig. 9.26.
· An HSI and its interconnections, however, differ somewhat froin
Table 9.1
Conditions
light
HDG
PITCH
&
ROLL
Threshold level
Before GS capture at
0° roll
6·
Before GS capture at
20° roll
9•
After GS capture at 0°
roll
4•
Before GS capture
4•
After GS capture
3•
LOC
At O dots deviation
At 2 2/3 dots deviation
30m V
54m V
GS
At O dots deviation
At 2 2/3 dots deviation
40m V
64m V
The altitude threshold levels increase from 5 ft at touchdown in
accordance with the graph below.
ALT
Threshold S5
(ft)
25
5 ..__ _ _ _ _ _ _ _.....
0
500
Altitude (ft)
Figure 9.26 INS attitude
references.
-,
r
I
I
ROLL SYNCHRO
ADI
PITCH SYNCHRO
L_ _ _ _ _ _ _ _ _
INS l"lA TfORM
J
Figure 9. 2 7 HSI interfaced
with INS.
MfLEACE tNOtCATOR
DRIFT ANGLE INDICATOR
HEADING REFERENCE
GROUND SPEED tNON:ATOA
HAVtGATION WAflMNG Fl.AG
OESIREO TRACK/COUA:$.E
POltfTER
VOA IL.$-fNS lNOtCATOR
- - - - - - TOIF'ROM POINTER
~-t,,lf-FtXEO UFERENCE ~
:!~: SlOP£ POINTER
&
CROSS TRAClC DISTANCE/
COURSE BAR
UIS AU.RT LtGHT
AZtMUTH CARO
H!At»NG POfNTtA
eot.tRSEffAACI( OIVIATIOtl
SCALE
the normal in that in addition to the navigational data supplied by an
MHRS and VHF navigation receivers, they can also display the data
computed by the navigation unit of an INS, as shown in Fig. 9.27.
The display elements, therefore, serve a dual role which, in the case
of the system utilizing this indicator, can be selected under mode
designations 'RADIO' and 'INS'. Details of the data displayed
appropriate to each mode are given in Table 9.2.
The HSI incorporates three warning flags: (i) navigation warning
flag which comes into view when there is a Joss of input from an
operating radio navigation receiver or INS; (ii) a 'HEADING' flag
which obscures the heading reference whenever there is an invalid
heading reference (magnetic or true); and (iii) a 'GS' flag which
obscures the glide slope scale whenever the relevant signal is invalid.
Data transfer switching
The ADI and HSI are part of two independent FD systems, i.e. one
for the captain and the other for the first officer, and so they are also
interconnected to ensure that in the event of failure of a data input
:;ource, a display of data will still be available. The data input
transfer switching arrangements for the ADI and HSI are shown in
Figs 9.28 to 9.31; these are based on the.application to some series
of Boeing 747 which has triple AFC and IN systems. The switches
related to each transfer function are mounted on panels located at
each side of the main instrument panel.
Attitude data (Fig. 9.28)
Under nonnal operating conditions, the captain's transfer switch is set
at position I so that his ADI is supplied with attitude data from INS
Table 9.2
Dara displayed
Display element
Azimuth card
Heading pointer
'RADIO' mode
'INS' mode
Heading referenced to
magne!ic North
Heading selected on
mode controller
Heading referenced to
true North
Biased to six o'clock
position and does not
function
illuminates when aircraft
is within 2 minutes of a
waypoint while navigating
along an INS track. Does
not function at ground
speed less than 250 knots
!NS 'Alert' light
Data source indicator
Shows source of data. e.g. System l or 2 Radio Nav. or INS
VOR/ILS indicator
'VORIILS'
Mileage indicator
Distance to DME station
Distance to next waypoint
Drift angle indicator
Drift Angle
Drift Angle
Heading reference
'MAG'
'TRUE'
'INS'
Ground sp.:ed ,n knots as
computed bv !NS
Ground speed indicator
Desired track/course
Course selected on
mode controller
Desired trad.
Cross track distance/
course bar
Displacement from VOR
or !LS course
Di,placemen! from track
Course/track
deviation scale
Deviation in degrees
from VOR or localizer
beam
Cros~ track distance
in rr iles (2 dots =
miles)
71
No. 1. The first officer's transfer switch is set at its position 2 to
supply his ADI with data from INS No. 2.
In the event of failure of data from INS No. l, the data established
by INS No. 3 can be transferred to the captain's ADI by setting the
switch to position 3 and energizing its associated relay. Similarly, the
first officer can also transfer data from INS No. 3 to his ADI by
resetting his transfer switch from the No. 2 to the No. 3 position.
Radio navigation data (Fig. 9.29)
For the transmission of this data, the main selector switches
placarded 'RADIO/INS' must each be selected to the 'RADIO'
position. In nc>rmal operation.of both FD systems, the captain's radio
transfer switch is set at position 1 so that his ADI and HSI are
supplied with the relevant deviation- signals from the VHF navigation
receiver No. 1. The first officer's transfer switch is set at its position
241
Figure 9.28 Altitude data
switching.
A
A
PITCH &ROLL
PITCH & ROLL
l____ J~ ,---f-13
I
ATTITUDE
ATTITUDE
r-------· LJ1
.•
PITCH & ROLL
INS NO. 3
i
I
I
I
I
INS NO 2
I
I
PITCH & ROLL
!
LJ 2"'1-·-·-·-r
PITCH & ROLL
INS NO. 1
!
PITCH & ROLL
i
PITCH & ROLL
I
A/P-F/0
A/P-F/0
A/P-F/D
COMPUTER A
COMPUTER C
COMPUTER B
2 and so his ADI and HSI are supplied with the signals from the No.
2 navigation receiver.
In the event of failure of receiver No. I, the captain's transfer
switch is set to the No. 2 position so that the No. 2 receiver can then
supply his ADI and HSI with deviation signals in addition to the first
officer's indicators. Similarly, the first officer's indicators can be
supplied from receiver No. l by the selection of his transfer switch to
the No. l position.
The transfer switches are electrically interlocked such that once a
transfer of data has been selected from one side, a transfer from the
other side cannot be effected.
Heading data (Fig. 9.30)
Heading data may be displayed on each pilot's HSI as either magnetic
or true depending on the setting of the 'RADIO/INS' selector
switches. Both" RMis always indicate magnetic heading since they are
part of each MHRS (see Chapter 8).
The transfer of data between MHR and IN systems is accomplished
by means of two switches (placarded 'COMPASS' and 'INS'} in the
242
Figure 9.29 Radio navigation
data switching.
... .......
I
I
I
I
I
I
Oil DEV
LOC DEV
i
I
I
I
I
I
I
-•
I
I
I
VHF NAV
N0.1
input circuits to each HSI. If, in normal operation, it is required that
magnetic heading be displayed on both HSls, then both 'RADIO/INS'
selector switches are set at the 'RADIO' position. The captain's
'COMPASS' switch is set at its No. I position so that his HSI and
RMI are supplied with magnetic heading data from MHRS No. l,
and the first officer's 'COMPASS' switch is set at its No. 2 position,
enabling his HSI and RMI to be similarly supplied from MHRS
No. 2.
In the case of failure of either MHRS No. l or No. 2, then the
captain's or first officer's transfer switches respectively would be set
at their No. 2 and No. l positions.
If, under normal operating conditions, it is required that one HSI
should display true heading and the other magnetic heading, then the
respective 'RADIO/INS' switches are selected at 'INS' and 'RADIO'
The diagram illustrates the case in which true heading is to be
displayed on the captain's HSI, and magnetic heading on that of the
first officer. True heading data is supplied from INS No. 1 via the
captain's 'INS' transfer switch when set at its No. 1 position. The
heading pointer of the HSI is oiased to the 6 o'clock position, the
Figure 9.30 Heading data
swilchinj!.
.
0
!.---r--·~--.J
:
•
~
:
-1
.------,
:
I
•
t'
:
•
AJP FID
..
-
COMPUTER C
:
· , • .,,,..,,,.,"""
•
:
l '·'
-
I1
coi:i~oo··-·- ...
A/P
I
•
1
I
A
:
•
• '
'···'t,f·· +······ ............... --- +····· .. 1•• /
MH
l
•
,
·,
·~/
,COMPASS COUPLERr----------JG
1
: :
f
NO 1
::.-.::
~'°:S: NO. 1
/
,,,,' '·,
TH
FLUX VIJ..VE
!
!
~
~
!
1
,' I•
I
I
-·-·-·-·-·.:
2
j
•
tNa
I co,,,0;-: .... ·•
co,,,0-2•·-· :L.
I
A/P F/0
·-·-·-·-·-·-· -·-.
l
1
~----·~~MANO
COMPUTER A ...
.
0
~··t ,
2•·-·,
............
1
'-'-·-·i-·-·-··
i
l \........ ,..,. . . . . . . . . . . . ..,.. ;·-·-·-·.l r-·-·-·-·-·-,
l
COMMAND
AIP
i
i
r----''---,
AIP F/0
COMPUTER B •·1
---,,----,
!
.
I
+
.
I
·,·-i·--·-·-·4·-·-·-·-·-·-·-·-·-·-·t-·-t,tAN_:
;
!
MHAIP
·,
TH
·
, ·-·-·-·-·
10 0
COMPASS COUPLER
I
r;ti
~
'..."'.:.".':.:
::...,~"°·a
~
L~
'VOR/ILS-INS' indicator displays 'INS' and the heading reference
indicates 'TRUE'.
Magnetic heading data is supplied to the first officer's HSI from
MHRS No. 2 via the No. 2 position of the compass transfer switch.
The HSI heading reference in this case indicates 'MAG'. If it is
required for this indicator to display true heading, in addition to that
of the captain, the associated 'RADIO/INS' selector switch would
also be placed in the 'INS' position.
In the case of failure of data from either INS No. l or No. 2, the
captain's or first officer's transfer switches respectively would be set
at their No. 2 and No. I positions.
As in the case of radio navigation data transfer, the 'COMPASS'
and 'INS' transfer switches are also electrically interlocked.
INS data (Fig. 9.31)
In addition to true heading, the other data shown in the diagram is
also supplied to both HSis when the 'RADIO/INS' selector switches
are set to 'INS' (see also Table 9.2). The transfer of this data is
effected in the same manner as that described above.
Figure 9.31 INS data
swi1ching.
--.
...
i-
.
r-·-·
I
i
!
I f! !
I
I
I
I
I
I
I
I
!
!
I
i
I
I
I
I
1I
I
:
J
I
I
I
INS
INS
:L_ _________ l---{@)----4----: RADIO
:
~2-C·---
,.,,'
DA, GS, lm(.
,- - - - - - -- - __,
,
OSRTI<, DIST
S• t
NSN0.1
DA TKE &XTK
•
i
RAOIO i
-~<r--·e1
- -·-·-L·-·-(@)-·-j
1•--'j/
INS
:
•
l-·-·-·-·-·-·-i·-·-·-·J
I
•
'·-j·-·-·-·-·-·-·4
j
•
:
I
I
j
g~
XTK
DSRTK
DIST
TKE
,
2
'\as;-•-ll>-
' ·,
.
INS
,
•,
·-·-·,
SNS~-
OA,GS,XTI<,
·-·-----
DSRTK. DIST
~:~
Cross track
Desired track
Oistallce
Track error
j
2
DA, Tl<E,&XTI<
f
A/P·f./0
A/P•F/0
COMPUTER A
COMP\JTER B
1O Inertial navigation/
reference systems
(INS/IRS)
,...... The navigation requirement of an aircraft is quite simply that of
determining its position in relation to its point of departure and points
en. route in order to reach a known destination. In practice, however,
the fulfilme~t of this requirement is made somewhat complicated by
having to provide for a lot of basic data, principally the following:
time, speed, distance between points, longitude, latitude, magnetic
heading, wind speed and direction, and bearings relative to known
points on the earth's surface.
The provision of such data can be, and is in fact, made by a
variety of navigational aids, a number of which are dependent on an
external reference source of one form or another. Although these aids
provide reasonably accurate answers to navigational task problems,
there are certain limitations: for example, radio navigation aids that
provide direction-finding and position-fixing capabilities require
extensive networks of ground stations and are subject to both natural
and man-made interference. An INS or IRS overcomes such
limitations by utilizing operating principles that make them entirely
independent·of external references.
~ystems and units
As the title of this chapter infers, there are two classifications of
inertial systems which, although performing the same basic
navigational functions, vary extensively in the manner in which they
process data, and in their capability of satisfying the needs for such
data by other interfaced systems.
The INS, which is the forerunner of systems and is still currently
used in some types of aircraft, utilizes analog and digital signalprocessing techniques, mechanical arrangements such as gimballed
platforms, and synchronous servo transmission loops. A system
consists of the four principal units shown in Fig. IO .1, together with
their interconnection, data outputs arid the other aircraft systems with
which .it is generally interfaced.
Figure JO. I
HOG
Units of an INS.
O MSU
I
...__~_.__
,...____
__,
TRACK
CROSS TRACK DEV'N
TRACK ANGLE ERROR
TIME TO GO
DISTANCE TO GO
CDU
c=:::I c::::::l
0
0
0
0
0
0
0 0
0 0
00
DO
CROSS TRACK DEV'N } -
{
HOG
DES. TRACK
TO/FROM
TRACK
.
F/D
HSI
PITCH.AOLL
STEERING
BATTERy - UNIT
To A/P and
COMMANDS F/D ADI
INU
ADC
m"""
RADAR
PITCH, AOLL
COMMANDS
1. Inertial navigation unit (INU): This unit contains an inertial
section consisting of accelerometers, gyroscopes and gimballed
platforms, a digital computer and all associated circuit module cards,
and a battery charger unit.
2. Control and display unit (CDU): This allows all associated data
to be inserted into the computer, and to be read out from it by means
of segmented LED displays.
3. Mode selector unit (MSU): This unit controls all the modes in
which the system can be operated.
4. 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 sytem is interrupted.
Although this type of system is highly accurate, the levels of
accuracy demanded for the navigation of those aircraft that are
designed for operation under what may be termed the 'computer
chips with everything' philosophy preclude its application to such
aircraft in favour of its more sophisticated descendant, namely, the
inertial reference system (IRS). It performs the same basic
navigational functions as an INS, but, as its fully digital computer can
also be pre-programmed with other relevant reference data, there was
some justification in changing its name.
The system consists of only two principal units. The outputs are
supplied to a greater number of interfacing systems, and since the
247
majority of them are also individually controlled by digital
computers, signal transmissions are via an ARINC 429 data bus (see
Chapter 6) as opposed to conventional 'hard wiring'.
The inertial reference unit (IRU) also contains accelerometers,
gyroscopes and the computer, but here, its similarity with the INU
referred to earlier ends. The major differences are: (i) the gyroscopes
are of the ring laser type (see page 270) instead of the spinning rotor
type; (ii) the complex mechanical arrangement of a gimbal system
and synchronous transmission loops is replaced by a mathematical
equation program so that acceleration and attitude signals required
for navigation are directly computed; (iii) the unit is directly mounted
to an airframe, i.e. it is of the 'strapdown' type so that the aircraft
itself becomes the inertial platform (see page 279); (iv) magnetic and
true headings are derived from a program of known data related to
the position data loaded into the computer, so that headings can be
computed without the aid of MHRS flux detector units (in fact, these
units are no longer required in aircraft equipped with an IRS); and
(v) no battery "unit and charger is used.
The inertial mode reference panel (IMRP) combines the functions
of mode selection and control and display of data.
Multi-installations
In order to ensure on-going navigation capability, and operation of
interfacing systems, adequate 'back-up' must be provided to safeguard
against system failures. Dual or triple IN/IR systems, depending on
size and operating category of an aircraft, are therefore installed with
the appropriate system transfer switching arrangements.
Power supplies
Both ac and de power is required for system operation which must be
maintained in the event of failures occurring. The sources from which
power is derived can vary depending on the type of system, but a
common feature is that after starting up, lhe system can be
maintained in opetation from either of the sources. This is effected by
the integration of power supply monitor and conversion circuits in the
navigation unit of a system.
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 unit provides
auxiliary power for the initial start-up, and also the power to maintain
system operation in the event of ac power failure, or a reduction in
voltage level. Under these conditions, the battery unit will sustam
system operation in any operating mode for periods up to 15 minutes'
duration. Indication that battery power is in use is provided by
248
illumination of an amber 'BATT' light on the control and display
unit.
The battery unit has a direct connection to the system's mode
selector switch so that when this is set to the positions for initial
starting of the system, battery power is used momentarily for
energizing a relay, the contacts of which are connected in the circuit
from the aircraft's ac busbar. Thus, ac power is supplied to the
navigation unit via the relay which is then held in the energized state
by the de produced by the power conversion unit. The battery supply
remains on for a short period (typically 10 seconds), enabling it to be
checked during alignment of the system (see page 279). On
completion of this check, the battery is isolated from the system and
is on standby until there is an interruption of the ac power supply. In
the event that an external power source is disconnected from an
aircraft while the INS in on, battery power will automatically be
transferred to the system, and some warning of this is required in
order to protect the battery against an inadvertent discharge. In one
example of warning system, a horn is located in the nose wheel bay
of an aircraft, and is activated 30 ± IO seconds after power transfer,
thereby alerting the ground crew.
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 multi-system installations, and after interruption of an aircraft's
power supply to the systems, switching arrangements are provided
which enable battery units to be paralleled in order to sustain the
operation of one of the navigation units. For example, in a triple
installation the battery units of Nos l and 3 systems can be paralleled
to supply the navigation unit of the No. I system.
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.
Navigation
~
fundamentals
Before going into details of the principles involved in the operation of
these systems, it is useful· at this juncture to consider some aspects
relating to the form of the earth, and also to define some of the tenns
associated with navigation over its surface.
Form
or the earth
The earth is not a true sphere; its equatorial diameter of 6884
nautical miles exceeds its polar diameter by about 23 nautical miles.
249
The 'flattening' a\ the polar regions gives rise to a more precise
definition of the.earth's form, which is known as an oblate spheroid.
For practical navigation purposes, however, the earth can be
considered as a sphere.
Direction on the earth
This is measured in degrees clockwise from north, and when the
datum is the direction of the north end of the earth's axis, it is
referred to as true direction. North is one of four points known as
cardinal points; the other three are south, east and west.
North and south define the axis about which the earth rotates from
west to east. To avoid ambiguity, a three-figure group is always used
to indicate direction, e.g. north - 000°; south - 180°; east - 090°
and west - 270 °.
Great circle
This is a circle on the surface of a sphere whose centre and radius
are those of the sphere itself. Relating this to the earth, the equator
and all the lines joining the north and south cardinal points (the
earth's poles) are examples of great circles.
On a plane surface, the shortest distance between two points is: of
course, a straight line which joins them. On a sphere, the shortest
distance between two points is the smaller arc of the great circle
which passes through both points.
Small circle
This is a circle on the surface of a sphere whose centre and radius
are not those of the sphere. With the exception of the equator, all
lines of latitude are small circles; they do not represent the shortest
distance between two points.
Longitude and latitude
These fonn a reference system for the position of points on the
earth's surface, and in determining the in-flight position of an aircraft
with respect to the earth.
Firstly, the datum is established by a great circle through the north
and south poles which passes through Greenwich. That half of the
circle which passes through Greenwich is known as the prime or
Greenwich meridian and is 000°. The other half is called the antimeridian and is 180°. Other great circles in the form of meridians,
or lines of longitude as they are called, are established to the east and
to the west of the prime meridian.
The next step is to have a datum point for positions in the direction
north and south. This is obtained by dividing the earth by a great
~RALLELS OF_
TITUDE
Figure /0.2 Gra1icule.
1Sma!fcirclesf
s
circle midway between the poles. This circle is the equator and is 0°
latitude.
From the foregoing, we can derive more precise definitions of
longitude and latitude as follows:
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.
The whole network of meridians (longitude and parallels of
latitude),' imagined to cover the earth, is called a graticule. Thus, as
sh9wn in Fig. I0.2, meridians or lines of longitude start from the
prime meridian or 0° and go right round up to 180° E and 180° W.
Similarly, the parallels of latitude start from the equator as 0° and go
up to 90° N and 90° S.
When giving a position, it is always quoted in the sequence latitu.ie
and longitude; e.g. the latitude of London Heathrow is the arc of the
meridian between the equator and Heathrow, and is 51 degrees and
28 minutes N. Its longitude is the shorter arc of the equator between
the prime meridian and Heathrow, and is 00 degrees and 27 minutes
W. It is expressed as: 51 ° 28' N 00° 27' W.
Convergency
Because the meridians converge towards each other to the poles, then
any line or track cutting successive meridians will do so at different
251
Figull JO. 3 Convergency.
y
X
angles; this inclination is called convergency and it equals the angular
difference between the measurements of the line or track at each
meridian. If, as shown in the example of Fig. 10.3, a track passes
through a point 'A' on meridian 'X' at an angle of 062 ", then in
passing through point 'B' on meridian 'Y' it will cut this at an angle
of 118°. The convergency is, therefore, equal to 118 - 062 or
056°.
If two places are on the same latitude, convergency may be
obtained from the formula:
Convergency
= Change in longitude x
sine of latitude.
Convergency is O at the equator (the meridians cutting it at 90°)
and increases to maximum at the poles.
Change of longitude
This is the smaller arc of the equator intercepted between the
meridians of the reference points, and is named east or west
according to the direction of the change. It is abbreviated as 'ch long'
(E or W).
Change of latitude
This is the arc of the meridians intercepted between the parallels of
the two places and is named north or south according to the direction
of the change. It is abbreviated as 'ch lat' (Nor S).
252
Figure /0.4 Rhumb-line.
U
b_
Changing track angles
Constant track angles
Rhumb-line
The ideal line to fly would be a great circle, since the shortest
distance between any two places is along the circle. There are,
however, two disadvantages: (i) the great circle from one point to
another will cross the converging meridians at different angles, and
(ii) because meridians form the basis of track angle measurements,
continuous alterations to these angles would be necessary as a flight
progressed.
In order to overcome these disadvantages, a curved line is therefore
followed which joins points along it and crosses each meridian at a
constant angle; such a line is called a rhumb-line (see Fig. 10.4).
Distances between points along this line are greater than those along
a great circle.
The meridians and the equator are the only examples of great
circles which are also rhumb-lines. Parallels of latitude are rhumblines because they cut all meridians at 90°.
Distances on the earth
Nautical mile
The distance on the earth's surface
which subtends an angle of one minute
of arc at the centre of the earth.
One nautical mile (nm) equals one
minute of latitude and is an av~rage
distance of 6080 ft. I O latitude = 60
nm. A change of latitude from the
253
Statute mile
Kilometre
equator to a pole is therefore equal to
90 x 60 = 5400 nm.
Equal to 5280 ft.
I/ 10 000th of the average distance from
the equator to either pole and is
accepted as being equal to 3280 ft.
Navigation terms
The following definitions are of navigation terms associated with
INS/IRS operation; they are also shown pictorially in Fig. 10.5.
Heading (HDG)
Track (TK)
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,
designated as Hdg (T), Hdg (M), and
Hdg (C). Hdg (T) is the only one of
the three which is plotted.
The direction in which an aircraft is
m~vi~g over the earth; it is al;o
Figure 10.5 -N-a~viLga-t-io-n-te_nns_._-+----~-~..
~-~-H----t-'-w-,N~-W-IN-'O'---
WPT !FROM)
254
Des~ track (DSR TK)
Drift (DA)
Ground speed (G/S)
Wind direction (W/D)
Wind speed (W /S)
Position (POS)
measured in degrees from true or
magnetic north. Only true TK is
plotted. If there were no wind, there
would be no drift and TK would be the
same as HDG; also the case with a
direct head- or tail-wind.
The planned direction over the earth in
wjlich it is iµiend~ thct aiittaft shall ·
'
'
~
move.
,
,;The angle between HOO and TK due to
the effect of wind. The direction of
drift is always from HOO to TK. Each
may be true or magnetic but never
mixed. If TK is less than HOO, drift is
to the left, and if TK is greater than
HDG, it is to the right as shown in
Fig. 10.6.
The actual speed (in knots) of an
aircraft over the ground, i.e. speed
relative to the earth. If there were no
wind, GS would be equal to tru~ air
speed (TAS).
The angle, measured in degree!:
clockwise from true north, with respect
to the direction from which the wind is
blowing.
The speed, in knots, at which the air is
moving relative to the ground.
Air:
The position of an aircraft
Figure 10.6 Drift angle.
(a) TK less than HOO;
(b) TK greater than HOO.
(a)
(b)
relative to the air at a
particular time.
Ground: The position of an aircraft
relative to the ground directly
beneath it at a particular time.
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.
Cross track distanee (XTK) Is the distance in nm (left or right)
measured from the nearest point on the
DSR TK line to the aircraft.
Waypoint (WPT)
This is a point of navigational
significance on an air route. Typically,
routes are divided up into convenient
lengths or legs, defined at each end by
a WPT. The end WPT of one leg is the
beginning of the next leg as illustrated
in Fig. 10.5. The 'FROM' WPT is the
one defining the· beginning of the flight
plan leg currently being flown. The
'TO' WPT is the one defining the end
of the current leg of the flight plan.
The 'NEXT' WPT in this convention is
the one defining the end of the next leg
to be flown, i.e. after passing the 'TO'
WPT.
~ndamental
principle of a system
The operating principle is derived directly from the Newtonian laws
of mechanics relating to velocity, acceleration and inertia. The
relationships may be summarized as follows:
l. Velocity is the rate of change of displacement with respect to
time for a moving object, and so is composed of both speed and
direction. If the speed of an object is constant, but its direction is
changing, then its velocity changes.
2. A change in velocity, either in magnitude or direction of
motion, is an acceleration (or deceleration). A body accelerates (or
decelerates), i.e. changes its state of motion, only if it is acted upon
by an external force.
3. AU matter tends to return to its existing state of motion and
consequently resists any changes to that state; this property is known
as inertia. The rate of acceleration of a body is proportional tp the
magnitude of its inertia. The inertial force displayed by a body under
a given rate of acceleration gives a measure of the mass of that body.
Figure JO. 7 Computer input
and output data.
DEST'N
WPT
START
WPT
LAT. & LONG
OF START & DEST'N
--.....---ACCELERATION
WPT'S ENTERED
DSRTK
G/S
PRESENT POS'N
STEERING COMMANDS
INS
COMPUTER
Recentring (feedback)
Figure 10.8 Accelerometer
..------VEL.
G/S
operation.
1st
2nd
>--,---<,-.DISTANCE
MASS
Accelerometer
·,.
DISTANCE FLOWN
l:
START------!1--e,.{)()---;-----PRESENT POS'N
POS'N
1 - - - - - - D S R TK
L_ _ _ _
Computer
In order for an IN/IR system to navigate an aircraft, its computer
must first have knowledge of the latitude and longitude of the starting
point of its flight, its final destination and, where necesary, a number
of intermediate waypoints. This data is inserted into the computer at
the time of initially starting and aligning a system prior to
commencement of a flight. The outputs computed are shown in Fig.
10.7.
Since acceleration forms the whole basis of computing the in-flight
navigation parameters from the pre-set data, then appropriate sensor
units are required. Let us consider the operation of one such unit as
shown schematically in Fig. 10.8. In this example, the mass of the
unit is normally centred by two springs between two pick-off
transformers.
When the aircraft accelerates, the mass is displaced, and a phaserelated signal is induced in the pick-off transformers. This signal is
amplified. and is applied as feedback to the accelerometer via what is
termed a force generator, which consists of a coil which moves with
the mass between the field of two permanent magnets. The effect of
this device is to recentre the mass, and the current needed to achieve
257
this is, therefore, a measure of the acceleration. The use of this
principle gives rise to the term 'force rebalancing accelerometer'.
The 'acceleration signal' is supplied to an integrator circuit which
relates acceleration to time, and therefore produces a signal
corresponding to ground speed (G/S). The G/S signal, in turn, is
supplied to a second integrator circuit which then produces a signal
corresponding to distance flown. This double integration process
solves the basic equation relating distance s travelled in a time t by a
body moving at a velocity v as given by:
S
= J~ V dt
[1]
Since an input voltage (V;) is made proportional to the velocity of
motion, and an output voltage (V0 ) is proportional to the distance
travelled, then in terms of electrical integration, equation [I]
becomes:
1
I
V = - - f V.dt
1
O
CR, j O
The product CR 1 is referred to as the time constant of the circuit.
If the 'distance flown' signal is then applied to a point within the
computer, and summed with one corresponding to the aircraft's
starting point, the result is a signal which gives the aircraft's present
position. If the latter together with the desired track (DSR TK) are
known, the computer can develop steering commands to keep the
aircraft on the DSR TK /iO that it can reach any desired destination.
Dual-axis accelerometers
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, as shown in Fig.
10.9. The designations 'X' and 'Y' relate respectively to accelerations
in the· horizontal plane and to the east, and to similar accelerations in
the direction of the local or north meridian. The outputs are
vectorially added to determine a total acceleration (A,) which after
integration gives the actual direction and the distance flown.
The accelerometers are aligned relative to the pitch and roll axes of
an aircraft, and not oriented geographically N-S and E-W. Their
output signals, however, can be related to a NE coordinate system,
and while the computer has 'knowledge' of their orientation, it will
always determine an aircraft's present po.sition in terms of the latitude
and longitude of that position.
Three coordinate systems are involved in the computations
performed by an INS and these are shown in Fig. 10.10. The XY
Figure JO. 9 Dual-axis
accelerometers.
I
START
POSITION
' - - - - + - DIST.
L.1 ___ ___,
reORIZONTAL PLATFORM
--.. . ---~
_J
" - - - - - - - - - + - V E L . (GJS)
Vector summing
'---------------·---·
COMPUTER
Figure JO. JO Coordinate
systems.
TRUE
NORTH
V
\
~,~·
,,J
'\
~
,v.
',
9 = tan· ·v.
.....
,
._:;..__ _ _ _ _ _ _ _...___ _ _ _ _ _.,.
EAST
- - - - XV system
Velocity North • v. • V, si,:i W,t - Vy cos W,t
Velocity East •
v. •
V, cos w,1 + Vy sin W,t
where V, • signals from ·x· accelerometer
Vr • signals from •y• accelerometer
1 rev/min
drift of
gyroscope
t • platform rotational lime
w, •
*
·z·
--· -
VU system
- - - - - NE sysiem
[-:J
Accelerometers
coordinate system is established by the X- Y platform which, as will
be described later, is continuously rotated at 1 rev/min; all
acceleration sensing is done in this system.
The second system is the non-rotational vu coordinate system, and
since the computer always assumes that 'v' and 'u' are related to
velocities in the north and east directions respectively, it utilizes this
system to perform all its calculations. The system is established by
259
supplying the accelerometer signals to the computer in terms of their
sine and cosine components. These components are derived from a
resolver synchro controlled by the X- Y platform, and thus
correspond to the initial position of the platform (relative to an
aircraft's longitudinal axis) before it begins to rotate. Since the X-Y
platform can start rotating from any position, then the XY and vu
coordinate systems will be in error with respect to the earth's nonrotating coordinate system (NE) by an angle () which is determined
during the alignment mode of the INS.
The NE coordinate system is the third involved in computation and
is the one in which position data are finally displayed in terms of
latitude and longitude. In order, therefore, to attain final alignment of
the platform with this system, and as may be seen from Fig. 10.10,
the accelerometer signals are resolved from X and Y through an
angle W,t, and a correction factor equal to the angle() is applied.
The formulae relating to the coordinate systems are also given in
Fig. 10.10.
Gyro-stabilized platform
For precision operation of an INS, it is essential for the 'X' and 'Y'
acceleration-sensing axes to be maintained normal to a local vertical
with the earth's centre at all times. If this were not done, false
gravitational forces would be sensed, giving rise to errors in the
computed distance flown and in the present position of an aircraft.
These forces and errors are overcome by mounting the accelerometers
in such a way that any displacements are detected by gyroscopic-type
sensors.
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 platform is referred to as the X- Y
platform.
The principle of this arrangement, and also that of 'strapdown', is
based on the Schuler theory of a pendulum whose bob is at the centre
of the earth, and supsended from a point above its surface. If, then,
the suspension point were accelerated around the earth, the bob,
being at the centre of the earth's gravity, would always remain
vertically below its suspension point. A platform mounted on the
suspension point tangential to the earth's surface would, therefore,
also remain horizontal irrespective of acceleration. If, for any reason,
the pendulum bob became displaced from the earth's centre, it would
start to oscillate with a period of 84.4 minutes; this is the value
obtained by substituting the earth's radius (in feet) for the pendulum
length I in the basic formula for calculating the time period of a
Figure I 0.11
Alignment of
gyroscope axes.
FWD
~FWD
~
---- x,
X
x,
z,
East or ·x· Gyroscope
---·
Z,
North or 'Y' Gyroscope
X-X 1 SPIN AXES
Y-Y 1 INPUT AXES
Z-Z, OUTPUT AXES
pendulum. Thus, by mechanizing an INS platform to remain
horizontal, an analogue of the Schuler earth pendulum with a period
of 84.4 minutes is produced, and the platform is then said to be
Schuler tuned.
The gyroscopes are of the integrating rate type, meaning that they
sense movement about only one axis, and that the rate changes are
integrated to give distance changes. The input and output axes of the
gyroscopes are positioned so that they relate directly to the north and
east coordinate system of 'X' and •y• accelerometer positioning and,
as shown in Fig. 10.11, they are designated as north ('Y') and east
('X') gyroscopes. The 'Y' gyroscope has its input axis aligned with
an aircraft's roll axis, while that of the 'X' gyroscope is aligned with
the pitch axis; the manner in which they sense attitude changes
depends on aircraft heading. Thus, if an aircraft and its INS platform
are heading north, the 'X' gyroscope senses pitch attitude changes,
and the 'Y' gyroscope senses roll attitude changes. The converse of
this is true, however, when an aircraft and platform are heading east.
The gyroscope shown in Fig. 10.12 is, for explanatory purposes,
drawn to represent sensing of roll attitude changes. In such an
attitude, therefore, the inertial platform and the spin axis of the
appropriate gyroscope will be deflected causing precession about its
output axis. The angular movement of the gyroscope operates an
electri.cal signal pick-off element which then transmits a signal, via an
amplifier, to the corresponding torque motor which then drives the
platform back to its level position.
Figure 10.12 Attitude sensing.
Pick-off
output signal
POSITION
TX SYNCHRO
-~/
TORQUE
MOTOR
IA Input axis
OA Output axis
A TX synchro is mounted on the gimbal system, its purpose being
to provide position signals proportional to aircraft attitude change for
use by other systems, e.g. automatic flight control and flight director
systems.
For a pitch attitude change, sensing and platform levelling is
accomplished in a similar manner and through a second gimbal ring.
A third gyroscope is also provided and is mounted on a second
platform. Its purpose is to sense changes about the local vertical
(designated 'Z') and to keep the X-Y platform in the same position
relative to space and the N-E coordinates, i.e. to maintain its north
datum. This gyroscope and its platform are also designated 'Z'. Any
change of platform or azimuth relative to the inner gimbal ring
corresponds to an equivalent heading change, and so by connecting a
signal pick-off element to the 'Z' gyroscope signals corresponding to
such change can be produced. These signals are supplied to an
azimuth torque motor which rotates the 'Z' platform in the opposite
direction to the heading change, and at the same time positions the
rotor of a TX synchro. The output signals from this synchro are
transmitted to the HSI of a flight director system which will then
indicate true heading (see also Chapter 9).
It will be apparent from the foregoing that between the headings
north and east, both the 'X' and 'Y' gyroscopes will exercise control,
the magnitude of which must be determined by heading. This is
accomplished by connecting a resolver synchro between the 'X' and
'Y' gyroscopes as shown in Fig. 10.13, and then positioning the
synchro rotor by means of the azimuth torque motor so that the
attitude signals transmitted to the pitch and roll torque motors are
modified by heading-related error signals.
Figure JO. I 3 Heading control
of pitch and roll torque motors.
·z·
GYROSCOPEl--~~~~~~~~~--1
•y•
Azimuth
gimbal
Pitch
North
GYROSCOPE
gimbal
East
·x·
GYROSCOPE
Roll
gimbal
Transport rate and earth rate compensation
As we learned from Chapter 4, gyroscopes must be compensated for
the effects of the rate at which they are transported over the earth's
surface, and for those of earth's rotation (earth rate). This
compensation, which is necessary to maintain the required earth
reference orientation, is also relevant to the gyroscopes of gimballedtype INS platforms; in other words, they must be Schuler tuned (see
also page 260). The effects of both these rates is to cause a platform
to tilt from the required horizontal position with respect to the earth's
surface, thereby causing the accelerometers to produce false output
signals. The gyroscopes must, therefore, be subjected to equal and
opposite forces so that corresponding signals can be produced for
maintaining the platform level by means of its torque motor system.
The principle of compensation is illustrated in Fig. 10.14 which,
although drawn to represent the 'Y' accelerometer loop, applies
equally to that of the 'X' accelerometer. The signals from the
accelerometers are supplied to an electronic resolver circuit which
converts the signals to the 'vu coordinate' system of the computer.
The signals are then integrated to produce the N-S and E-W
velocity signals, and these, in turn, are divided by a value R + h,
where R = the mean radius of the earth, and h = aircraft altitude. In
other words. R + h is the distance from the centre of the earth to the
X - Y platform. and since it corresponds to the radial path flown by
263
Figure /0.14 Transport and
earth rate compensation.
From 'Y'
gyro
·v·
'X' Gyro torque signal
From 'X'
accelerometer accelerometer
Recentring
.-----W,t
ELECTRONIC
RESOLVER
Accel'n
i~
·y· gyro torque
signal
2
._____ To other
torque motor
3
Start position
N-S distance
latitude
------LI<.'"""------<
LATITUDE
Present
CORRECTION
latitude ------..------P-,e-s-en_l_la-ti-lu-de_ _ _ _ _ _ _ ___,;_~---'
longitude
changes
Present
Start position
longitude
longitude
4
Earth rate compensation
Earth rate
LATITUDE
= 15°/hr - - - c o R R E C T I O N I - - - - - - - - - - - - - - - - - - - - - - - - - '
an aircraft around the earth, then after dividing the coordinate
velocity signals by R + h, signals corresponding to radial velocity
are obtained. The value of R is a constant that ts pre-programmed
into the navigation computer, while h is, of course, a variable that is
supplied to this computer from an ADC.
The coordinate velocity signals are supplied to a transport rate
compensation generator circuit for conversion to transport rate
signals, which are then supplied via summing points l and 2 to a
second electronic resolver circuit. The purpose of this circuit is to
convert these signals so that they relate to velocities in the XY
coordinate system. They are then transmitted to the torquer coils of
the 'X' and 'Y' gyroscopes, thereby precessing them in order to
establish angular distance output signals in their respective pick-off
coils, which are connected to a coordinate resolver. This synchro,
therefore, separates the signals into pitch and roll components, and
supplies them to the corresponding levelling torque motors, which
then tilt the inertial platform back to a level position, through an
angle determined by the transport rate.
The effects of earth rate on a gyroscope are dependent on the
latitude in which it is operating at any one moment, i.e. its present
position. Compensation signals related to this latitude must, therefore,
be generated and, as will be noted from Fig. I0.14, this is
accomplished by supplying the corresponding signal from summing
point 3 to a latitude correction circuit. The compensating signal
output from this circuit corresponds to a rate equal to 15° cos of
latitude and is summed at points I and 2. The outputs from these
points to the electronic resolver, and the resulting signals supplied to
the torquer coils of the gyroscopes, and to the platform torque
motors, therefore provide for combined transport rate and earth rate
compensation.
Gimballed platform arrangement
The arrangement adopted in a typical IN system is shown in Fig.
I0.15.
The two platforms are supported by an inner roll (IR) gimbal
within a pitch gimbal which is, in turn, supported by an outer roll
(OR) gimbal; this gimbal supports the whole system within the
structure of the system's navigation unit. The reasonJgr~,ivingtwo
m!LgiJnheJ§.Qf!!eDted as shown)s to prevent the cpndition known as
•'gimRal loc~'Jrom occurring (see also page 106). The angular
movement of the IR gimbaLis limited to ± l 0°.
Fig11r,• /0.15
platform.
Gimballed
; _,../' AZIMUTH SYNCHRO
_,../' AZIMUTH IOAOU£ R
r--:::,-7.':'::::-----------
PLATFORM
HEAOIIIG
INNER
ROlL
ci°i-:,--------
PITCH
IIIS COMPUTER
TRUE
HEADING
ERROR
X•Y Pt AIf ORM AES0l VER
COOAOIIIA TE RFSOI V! R
265
Table 10."I
Torque motors:
Outer roll
Inner roll
Azimuth
Synchros:
Pitch
Outer roll
Azimuth
Resolvers:
Inner roll
X-Y platform
Coordinate
Function
location
Component
Drives OR gimbal.
Drives IR gimbal.
platform in a direction opposite to aircraft
Drives
heading change.
Fixed to INU casing
Fixed to pitch gimbal
Fixed to IR gimbal
·z·
Rotor
Stator
Fixed to OR gimbal
Fixed to !NU casing
Fixed to IR gimbal
Fixed to pitch gimbal
}
Fixed to OR gimbal
Driven by azimuth torquer
Provides pitch and roll attitude signals to AFC and
FD systems.
Provides true heading signals to HSI of FDS.
Fixed to pitch gimbal
Fixed to IR gimbal
Fixed to IR gimbal
Fixed to platform
Fixed to IR gimbal
Fixed to platform
Supplies roll error signals to OR torquer.
Supplies signals to computer which are related to
platform heading. and in terms of sine and cosine
components.
Receives signals from ·x· and •y· gyroscope pickoffs. separates them into pitch and roll components,
and supplies them to the pitch and IR torquers
respectively.
A synchronous-type servomotor is mounted on the 'Z' platform and
it rotates the 'X-Y' platform at the rate of l rev/min. The reas.onior
rotating the platform is to modulate errors generated by m.i~!!li~nment
of the X - Y accelerometers and gyf()SCQpes, scale errors of the
accelerorm::ters., anddnft ·ofthe.gyrosc~pes so that the computer
'sees' a minimal error condition. Since the platform is rotating, the
'X' and 'Y' gyroscopes sense pitch and roll attitude changes
alternately.
The locations and functions of the various torque motors, synchros,
and resolvers are detailed in Table 10. l.
Accelerometer construction
The constructiona: arrangement of a typical accelerometer is shown in
Fig. 10.16.
The mass and its force coil are suspended by two flat springs so
that they can move linearly along the acceleration-sensing axis. A
ferrite armature is fixed at each end of the mass and in the proximity
of two pick-off transformers supplied with 5 _V ~c at 12.8 kHz. When
the~centtecr:-fne"·space·oetween"ffie"armaiu;;;··~~;rtransformers will be equal and so the coupling between primary and
secondary in each transformer will also be equal. The secondaries are
connected in series, so that the signal induced in the secondary of one
transfonner is in-phase with primary voltage, while that in the other
Figure JO. 16 Typical
accelerometer arrangement.
SUSPENSION SPRING
FORCER COIL LEAD~
PERMANENT MAGNET
transformer's secondary is out-of-phase with primary voltage. Thus,
with equarcoupling, the net output from both secondaries will be
zero.
When the mass is deflected as a result of an acceleration, the
coupling between the armatures and transformers will no longer be
equal, and so the secondaries produce an output related to phase and
position of the mass. The 'mass position' output signal is supplied to
an amplifier from which it is fed back to the force coil which is also
deflected within the field of two permanent magnets. The current
flowing in the coil interacts with the field of the magnets and
establishes a force that recentres the mass and coH to the 'null'
position. The current required to do this is therefore directly
proportional to the force that originally caused deflection; in other
words, it is a measure of acceleration.
The accelerometer assembly is enclosed within an hermeticallysealed case filled with a low-density fluid to provide damping of the
pendulous mass. The temperature of the fluid is regulated by a
thermistor-controlled heater element.
In a number of systems currently in use, the accelerometers adopt
capacitance-type pick-offs. The design of their sensing element
simplifies the construction and, in addition, eliminates the need for
fluid damping. The arrangement of such a unit is shown in Fig.
10.17.
The sensing element is a ceramic disc which is suspended by four
flexible metal 'hinges', and by means of a mass subassembly the disc
is made pendulous so that it can move through an arc under the
influence of an acceleration force. The disc is coated on each side
with a metallized pattern to form a capacitance plate. The plates are
also used to form the suspension 'hinges' which, in turn, are used for
the fOnnection of electrical leads. The other capacitor plates are
formed by the surfaces of upper and lower permanent magnets which,
together with a coil that is also supported on the disc assembly,
267
Figure /0./7 Operation of
o:apacitance-type accelerometer.
UPPER MAGNET
STRUCTURE
CAPACITIVE
PICK-OFF GAP
CAPACITIVE
PICK-Off
PLATE~~~~!...
,,....._,,,_-THIN FI LH
PROOF MASS
SUBASSEIIBLY
PICK-Off ANO
TORQUE LEADS
MAGNET
LOWER
MAGNET
STRUCTURE
LEAD SUPPORT
POSTS
SENSOR
ELECTRONIC
CONNECTOR
forms an electrical force or torquer device. The two capacitors of the
pick-off are incorporated in an ac bridge circuit.
When subjected to an acceleration, the rotation of the sensing
element about its suspension axis causes the capacitance of one half
of the disc to be increased, while the capacitance of the other half is
decreased. The bridge circuit is, therefore, unbalanced, and a signal
is developed and supplied to a differentiating operational amplifier
which produces a phase-sensitive signal proportional to sensing
element displacement. This signal is then demodulated and filtered,
and after amplification it is supplied to the torquer coil which then
develops a current proportional to the acceleration force, and restores
the sensing element to a 'null' position. The coil current is measured
as a voltage drop across a precision scaling resistor.
The compensation circuit module is provided for the purpose of
making the accelerometer insensitive to temperature variations with
regard to acceleration force scale factors, thereby eliminating thermal
lag.
Gyroscopes and construction
In systems utilizing electrically-operated gyroscopes, two-phase
synchronous/hysteresis-type motors are adopted, which operate at
about 24000 rev/min from an -input of I 15 V ac at 1.2 kHz supplied
via rotary transformers. An example of a typical unit is shown in
Fig. 10.18.
Figur,• I 0.18 Floated
HEATER
BELLOWS
gyroscope.
CONCENTRIC
RINGS
BAR
MAGNET
WHEEL TRANSFORMER
SHAFT SECONDARY WINDING
THRUST
PLATE
f\ELLOWS
rLUID
I
ffOTARY TRANSf"ORME'fl
PRIMARY WIN'DfNG
FLOAT
flOiARY TRANSFORMER
SECONDARY W!Nt)INC
ROTOR
emu;
-. COSCENTTIIC
RINGS
l--o
t.:z
KHZ.
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SUPPLY
Each motor is mounted within a cylindrical float filled with helium
so that in operation the motor 'rides' on a thin layer of the gas. The
float assembly is, in turn, mounted within a cylindrical main case
filled with a dense viscous fluid which supports the weight of the
~~~Jl(OYide.s.-'1mn12ing,._a_ruLI!LQl~£!i211~ ~gai_nst yi):,ration or
shock loading. The float is suspended radially and axially (on its
output axis) by a magnetic field generated by permanent magnets and
by tuning the inductance of the rotary transformers located at each
end of the assembly. This 'floated gyro' concept therefore eliminates
the mechanical forms of suspension adopted for conventional
gyroscopes.
Any angular displacement of the float assembly with respect to the
case is detected by a position pick-off which then supplies an analog
output signal to torquer coils which return the float assembly to a
'null' position. This is accomplished as a straightforward application
of a force, and not of precession, and since the null position provides
'.:?G9
the datum from which any tilt of the ·inertial platform is sensed, then
in functional terms the torquer coils may be considered analogous to
a restraining spring used for the rate gyroscope of a turn-and-bank
indicator.
In respect of gyroscopes that are used for the stabilizing of an X Y platfonn, the torquer coils are also supplied with constant current
signals, the purpose of which is to compensate for the effects of
transport rate and earth rate (see page 263).
Signals resulting from precession of the gyroscopes during platform
attitude changes are also derived from the position pick-offs, and
after amplification they are supplied to the respective gimbal ring
torque motors.
Each gyroscope is maintained at an operating temperature of 76°C
by a 30 W blanket-type heating element wrapped around the main
casing. The heater circuits and gyroscope temperatures are controlled
by thermistor-type sensors.
~Ring laser gyroscope (RLG)
This type of unit is essentially a rate sensor, and it is only in this
context that it can justify the name 'gyroscope'. It has no rotating
mass or gimbal system, and therefore does not possess the
conventional characteristics of rigidity and precession. Since it has no
gimbal system it is, and can only be, used in a 'strapdown'
configuration (see page 279) so that its attitude output signals can be
supplied direct to the navigation computer, and it eliminates the use
of complex platform levelling systems. It has many other advantages
such as wide dynamic. range, allowing of very short alignment times,
and high reliability factors, and is adopted as standard in the inertial
reference systems (IRS) currently in use in many types of civil
aircraft.
The basis of a typical sensor is a triangular block of specially
fabricated glass ('Cervit' glass) that is extremely hard and does not
expand or contract under varying temperature conditions. By means
of computerized ultrasonic diamond-drilling techniques, a cavity is
formed within the whole block. A precision-made mirror is fitted at
each comer of the block, and a cathode and two anodes are located
as shown in Fig. 10.9. The mirrors serve as both reflectors and
optical filters, reflecting the light frequency for which they are
designed, and absorbing all others.
·
The cavity is filled with a lasing medium (typically helium-neon)
and when excited by an electrical potential across the cathode and
anodes, the medium is ionized and is transfonned into light in the
orange-pink part of the visible spectrum. By design, two light beams
resonating at a single frequency are emitted and are made to travel in
opposite directions around the cavity. Since the beams travel at the
same constant speed and are 'bounced off the mirrors, then in a
270
Figure JO. I 9 Ring laser
gyroscope.
t,
START
t,
BEAM 2
BEAM 1
STATIONARY
R = radius of beams
O = cavity rotatiQn
L = length of path
Phase change at r2 = L x X
where >. = wavelength of light
12 -t, is also equal
to 2rR!C,
where C • speed of light
static condition of the sensor block, they take the same time to
complete a closed path in inertial space around the cavity.
Although the frequency is determined by the gas that is 'lasing', it
can be varied somewhat by changing the path length over which the
light waves have to travel; for a given length there are an integral
number of waves occurring over the complete path. If the length is
altered, the waves will either be compressed or expanded, and this
results, respectively, in an increase or decrease of their frequency.
Both beams combine in an optical sensor, or readout detector, located
at one corner of the block. The sensor operates on the interferometer
principle, i.e. it contains a prism that deflects the beams so that they
'interfere' with each other in order to form what is termed a fringe
pattern.
As in the case of a gimballed platform-type of INS, three sensors
are required to be mounted in an aircraft so as to detect attitude
changes about the pitch, roll and yaw axes. When the aircraft is in
271
straight and level flight, all three RLG sensors are in a static
condition, and so the resonant frequencies of the beams are equal.
If now the aircraft's attitude is changed about, say, the pitch axis,
the corresponding sensor will also be rotated about its axis
perpendicular to the plane of the beams. Since the beams are
travelling at a constant speed on paths in inertial space, the bodily
rotation of the sensor will then be with respect to inertial space. This
means, therefore, that the beam t!'avelling from one mirror to the
next in the direction of sensor rotation will move through a greater
distance than the se,;ond beam that is travelling in the opposite
direction. Thus, the times taken for the beams to travel around the
cavity of the sensor will now differ. As already pointed out, a change
in path distance produces a change in frequency of wave propagation,
and so by measuring the frequency difference resulting from rotation
of an RLG sensor, the angular rate at which it does so can be
determined.
The spacing of the light and dark portions of the fringe pattern
referred to above depend on the angle between the interfering of the
beams and their frequency. When both beams are of the same
frequency, the pattern is stationary, and constant signals are produced
by two photo-diode type detectors which are spaced an odd number
of quarter wavelengths apart in the pattern. Thus, when rotation of
the sensor causes a difference between frequencies, the fringe pattern
moves across the detectors, and due to the spacing of the beams, one
detector will receive maximum light when the other is at half
intensity. Each detector then converts this fringe pattern movement
into signal pulses, the phasing of which give the direction of sensor
rotation, while the frequency is proportional to the angular rate of
rotation. The signals, which are in digital format, are transmitted to
the appropriate attitude computing software within the IR computer.
The relationship between the input rate of rotation and the output
frequency is a linear one, and ideally it should remain so throughout
a full rotation of an RLG sensor. At low input rotation rates tending
towards zero, however, the output frequency can become non-linear,
and at a certain threshold value can drop abruptly to zero. This
phenomenon is known as 'lock-in', and is due to small amounts of
energy from the beams being back-scattered into each other, the
energy causing the beam frequencies to be pulled together until
eventually the beams synchronize. Since extremely low rotation rates
(typically 0.001 °/hr) are required to be measured in IN/IR systems,
'lock-in' can result in undesirable errors. In order therefore to
circumvent these effects, a technique known as 'dither' is introduced,
and is effected by means of a piezo-electric motor. This motor is
mounted on the sensor in such a way that it vibrates the laser ring
about its input axis through the 'lock-in' region, thereby unlocking
the beams and enabling the optical sensor to detect the smaller
272
movement of the fringe pattern. The motions caused by the dither
motor are decoupled from the output of an RLG.
Mode selection
In order to control the modes of IN/IR system operation, mode
selector units are provided. In systems utilizing gimballed inertial
platforms, the unit is located on a flight deck panel, while in others it
is integrated with the control and display unit (CDU) to form what is
termed an inertial reference mode panel.
Figure 10.20 illustrates the controls of a separately located unit; the
modes that can be selected are as follows:
STBY
ALIGN
NAV
This mode, which is for ground use only, selects power
onto the system and allows it to 'warm up' and to run the
gyroscopes up to speed. The aircraft's present latitude and
longitude are inserted in the CDU when in this mode, and
an auto-alignment sequence commences. The INS is not
affected by movements of the aircraft while in this mode.
This mode allows the INS to align automatically to its
true north point. When alignment is completed the
'READY NAV' light illuminates green to indicate that the
system is ready to go into the 'NAV' mode. The aircraft
must remain stationary in the 'ALIGN' mode.
This is the normal in-flight operating mode in which the
data required for navigating the aircraft and stabilizing the
FROM/TO OISPLAV
Fi~1m• 10.20 Control and
display unit.
SAT LIGHT {AMBER)
HOLO KEV/LIGHT
WAVPOINT
SELECT
THUMBWHEEL
KEVIIOARO
Xl~
JKE
DATA SELECT _...:'--H.;.;DG;.;,..;
SWITCH
OA
OIMKNOII
AUTOJIAANITEST
WFT CHG KEY/1.IGHT
CLEAR KEV
SWITCH
273
Figure .10.21 Inertial reference
mode panel.
!RS IIODE SEL
HUMER!C D I S P L A Y ~
DISPLAY SELECT
AHO D!llll!NG
SWITCH
DISPLAY
SYSTEM
SELECT
SWITCH
TK/GS
,~11-12=-"""2--1--,2__,,,3-,I
HDG
SYS DSPL
C
~R
L
ALIGN w
ALIGN w
ALIGN w
OH DC a
OH DC a
ON DC a
DC FAIL a
DC FAIL a
DC FAila
FAULT a
FAULT a
1
1
----Of-F®' Off,®
ALIGN
NAV ATT
1
SELECT
IIODE
SI/ITCHES
ALIGN
NAV ATT
1
/j
HOOE AND STATUS
ANNUNCIATORS
F4ULT a
NAV
ALIGN
\
I
ATT
I
INS are computed. It must be selected before the aircraft
moves from its parked position. The 'READY NAV' light
is extinguished on selection. Initial track selected may be
made and started when in this mode.
ATT REF Selects pitch, roll and platform heading stabilization
outputs only; no display is presented on the CDU. This
mode disables the navigational capability of the computer
for the remainder of a flight, because once turned off, the
computer cannot be switched on again. Normally, this
selection is only made when a computer failure has
occurred.
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.
.The red 'BATT' light is illuminated when back-up de voltage is
being used and is less than the minimum required (typically 18 V) to
operate the system; the system is automatically switched off.
An inertial reference mode panel as used in a typical IR system is
shown in Fig. 10.21, from which it will be noted that the mode
selector switches appropriate to the normal triple systems installation
are grouped together instead of having separate space-consuming units
of the type just described.
The three modes that can be selected and their functions are the
274
same as those already outlined. A 'STBY' mode is not required for
the reason that the application ofcomprehensive digital signalprocessing techniques, and of ring laser gyroscopes, eliminates the
need to allow for 'warm-up' and gyroscope 'run-up'.
Four mode and status annunciator lights are provided for each
system as follows:
ALIGN
Illuminates white when a system is in the alignment mode.
In the event of alignment procedure failure, it flashes on
and off.
ON DC This illuminates amber to indicate that power to the
system has automatically changed over from the normal
115 V ac to 28 V de power from the battery system.
DC FAIL Illuminates amber when the battery power source drops
below 18 V.
FAULT Illuminates amber when failures in the system are
detected.
Longitude and latitude data from any of the three systems (Left Centre - Right) are selected as appropriate, for readout on the single
. display at the top of the panel.
Control and display unit (CDU)
This unit serves as the primary interface between the flight crew and
the inertial system computer, in that it contains the controls necessary
for the selection and display of all essential navigational data. The
panel layout of a unit can vary between systems, but the one
illustrated in Fig. 10.20, and used in conjunction with the computer
of a gimballed platform-type of INS, serves to illustrate control
functions and selection methods that are generally applicable.
Control switches
DATA SELECTOR
WPT SELECTOR
This switch selects the navigational data
listed in Table 10.2 for presentation in the
upper left and right numerical displays of
the unit.
This switch is of the thumbwheel type,
and when the 'DATA SELECTOR' switch
is set to 'WPT', it enables WPTs I to 9 to
be selected for latitude and longitude
insertion, or selection of WPTs O to 9 for
presentation of their coordinates on the
upper displays. It is also used for inserting
and. displaying latitude, longitude, altitude
and frequency of up to nine DME stations.
275
This contains ten push-button key switches
(0-9) for entering present position and
WPT coordinates, 'FROM/TO' WPTs,
desired XTK effect, and TK hold. Each
key illuminates white when pressed. For
the display of DME station latitude,
longitude, altitude and frequency, the
following key selections are made:
3 and 9 enable altitude and frequency to
be displayed.
2 or 8 enable altitude to be loaded.
4 or 6 enable frequency to be loaded.
7 and 9 enable latitude and longitude to
be loaded.
The 'FROM' display remains blank, and
the 'TO' display flashes the station number
when DME data is being used.
INSERT
Operation of this push-button switch
transfers entered data into the computer.
CLEAR
This push-button switch is used to erase
data loaded into displays but not yet
loaded into the computer; it is illuminated
white.
AUTO/MAN/TEST
In 'AUTO' the system makes automatic
sequential TK leg changes, and permits
manual TK leg changes. In 'MAN', TK
leg changes can only be initiated manually.
The 'TEST' position enables checks to be
made on the displays and annunciators.
WPI' (or TK) CHANGE Allows initiation of manual TK leg
change. Illuminates white when pressed.
and is extinguished when the 'INSERT' 01
'CLEAR' keys are pressed.
HOLD
This permits a position check and up-date
to be made, and also a display of
malfunction codes. It also illuminates
white when pressed.
REMOTE
Enables semi-automatic (loading and
insertion of WPT and DME data into
more than one INS) autofill operation,
remote ranging, and display of XTK
offset. It illuminates amber when pressed.
DATA KEYBOARD
Displays
There are three electronic displays of the segmented type: left and
right for displaying the parameters listed in Table 10.2, and a
27&
Table 10.2
Data selector
switch position
TKIGS
HOG DA
XTK TKE
POS
WPT
DIS/TIME
WIND
DSR TK/STS
left
data display
TK
True HOG (See Note 2)
XTK distance (See Note I)
Lat. of present position
Lat. of selected WPT
Lat. or alt. of a selected
DME station
Distance to WPT
Distance to selected DME
sta1ton
Wind direction
DSR RK
{ - Blank -
t
Right data disp/ay
GS
DA (See Note 2)
TKE (See Note l)
Long. of present po,i1ion
Long. of selected WPT
Long. or freq. of a selected
DME station
Time 10 WPT
- Blank Wind speed
- Blank Action code, and alignment status
during 'ALIGN' and 'NAV'
modes
Noles: I An ·R· or.'L' also displayed to indicate that present position is to right or lefl of
DSR TK. and present TK angle to right or left of DSR TK.
2 An ·R' or ·L' also displayed to indicate that present TK is 10 right or left of
aircraft"s HDG.
'FROM/TO' display to indicate the number of the 'FROM' and 'TO'
WPTs on the TK leg being navigated. The characters of this display
will flash, if the 'REMOTE' switch has been pressed, to indicate the
remote ranging leg, remote direc;t ranging leg, or remote ranging
along the flight path to the desired WPT.
Annunciators
There are three annunciators as follows:
ALERT This illuminates amber two minutes before reaching a 'TO'
WPT. Tbe operation of this annunciator also depends on the
settings of the 'AUTO/MAN/TEST' switch. Thus, if it is at
'AUTO', the annunciator is extinguished when the
'FROM/TO' display changes to the next two WPT numbers.
If the switch is at 'MAN', the light flashes as the aircraft
flies over the WPT. The annunciator is operable at a G/S
greater than 250 knots.
BATT Illuminated amber when the INS is operating on battery
power .
. WARN Illuminated red when a system malfunction occurs, or
during 'ALIGN' mode it flashes to indicate system
degradation, or that a.1 alignment failure has occurred. It
will not extinguish unless the fault is corrected or the INS is
switched off.
277
Fi11ure 10.22 Navigation data
display on an HSI.
GROUND SPEED
A certain amount of the data listed in Table 10.2 can also be
derived from the indications that are normally displayed by the HSI
of a conventional type of flight director system as illustrated in Fig.
10.22. They are interpreted as follows:
TK is the compass card reading when referenced against the
diamond-shaped 'bug' or cursor, 025° in this case.
DA is the reading of the scale above the compass card when
referenced against the diamond-shaped 'bug'. In this case DA is 020°
to the left.
XTK is indicated by the deflection of the deviation bar with respect
to its scale. Each dot corresponds to 3.75 nm.
'Miles to go' is always displayed on the top left-hand indicator
even if the CDU is displaying the miles between WPTs, or DME
miles.
DSR TK is the compass card reading referenced against the pre-set
course arrow, 075° in this case.
TRUE HDG is the compass card reading referenced against the
lubber line: 045° as shown. An annunciator flag above the DA scale
also displays 'TRUE'.
GS is always displayed on the indicator at the top right-hand corner
of the HSI.
'Strapdown'
configuration
~
Alignment sequencing
This applies to the installation of the reference unit of an IRS which,
in dispensing with the complex mechanical gimballed platform
arrangement for its accelerometers and ring laser gyroscopes, enables
them to become part of the unit's fixture to an aircraft's structure.
Thus, the aircraft itself becomes a platform for the sensors, and in
moving with it, the signals they produce are used by the computer to
extract aircraft attitude, and to resolve body axis accelerations into
navigation axis displacements. For this purpose, mathematical
equations are programmed into the computer which, in essence, are a
functional replacement of a gimballed platform.
The accuracy of an IN/IR system is dependent on precise alignment
of its inertial platform with respect to the latitude and longitude of
the ground position an aircraft is in at the time of 'starting up' the
system. The computer must, therefore, be programmed to carry out a
self-alignment calibration procedure over a specific time period before
the system is ready to navigate an aircraft. This procedure can only
be carried out on the ground, and during the actual alignment stage
the aircraft must not be moved.
The technique and the time period involved can vary between types
of system installed in an aircraft, but the following, which is based
on a gimballed-platform system, is generally representative:
1. The selector switch on the mode select unit is set to the 'STBY'
mode position, and the data select switch on the CDU is set to
'POS'. The latitude and longitude of the present position is then set
in the appropriate displays from the keyboard.
The 'INSERT' pushbutton on the CDU will illuminate, and remains
illuminated, until the present position is 'loaded' into the computer.
During this mode, heater power is applied to the navigation unit until
it attains its operating temperature, and the gyroscopes are run up to
speed; the platform and gimbal system are 'caged' with respect to the
aircraft's axes.
2. The mode selector switch is then set to the 'ALIGN' mode
position. The present position of the aircraft is loaded into the
computer by pressing the 'INSERT' switch button: the switch light is
then extinguished. The computer also carries out a check which
compares the present position with that last displayed.
For purposes of sequencing, the 'ALIGN' mode consists of submodes designated by numbers that decrease in sequence as follows:
9. Standby.
8. Course levelling.
7. Course azimuth.} ·
.
Gyrocompassmg
.
.
6 . F me a11gnment.
5. Gyrocompassing complete.
279
The numbers are displayed by the fourth digit of the CDU's righthand display, when the CDU select switch is at the 'DSR TK/STS'
position. At the same time, the first and fifth digits of the display are
a 'O' and a '5' respectively, to indicate that the system is not in the
'NAY' mode. This last digit is referred to as a performance index
(Pl) number.
If the system has not reached operating temperature the sub-mode
number 9 is displayed when the mode select switch is moved from
'STBY' to 'ALIGN'. During this sub-mode, the platform and gimbal
system remains caged.
When operating temperature is reached, and the gyroscopes are up
to speed, the alignment sequence goes into:
sub-mode 8 and the CDU display changes accordingly. During this
mode, the 'BATT' annunciator in the CDU will illuminate to indicate
battery power availability, the gimbal system is uncaged, and the
platform is aligned to the local horizontal utilizing the accelerometers
to detect any 'out-of-level' orientation. At the end of the mode time
period, the CDU display changes to indicate:
sub-mode 7. This mode provides an initial estimate of the azimuth
orientation of the 'X' and 'Y' accelerometers' input axes with respect
to true north. Local latitude is also computed and compared to the
present position loaded into the computer. At the end of the time
period, the CDU indicates that the sequence goes into:
sub-mode 6, during which the initial determination of true north is
refined, and the computer provides a bias signal for the gyroscopes
and corrects for earth rate at the present position latitude. This,
together with sub-mode 7, is normally referred to as 'gyrocompassing'. When completed, the CDU display again changes to
indicate:
sub-mode 5. The 'READY NAY' light on the CDU will illuminate
and the 'NAY' mode can then be selected on the mode select unit.
When selected, confirmation of entering the 'NA V' mode is indicated
by the first digit of the CDU display changing from a O to a I, and
by extinguishing of the 'READY NAY' light.
When a pre~fEght calibration of the 'Z' gyroscope is required, the
system can be left in 'ALIGN' after gyrocompassing is complete, i.e.
sub-mode 5 is displayed. As the calibration takes place, the PI
number referred to earlier decreases in sequence from 5 to O, the
latter indicating the best calibration.
As noted earlier, alignment time periods can vary; typically they
would be ten minutes for an IRS and nineteen minutes for an INS.
Slewing
280
This is a procedure that provides the capability of moving the
gimballed platform of an INS about the pitch and roll axes, thereby
simulating in-flight attitude changes, in order to observe the resulting
changes in the displays of the units and instruments interfaced with
the system. The procedure, which can be carried out without the use
of a tilt table or removal of the navigation unit from an aircraft,
allows attitude changes of ± 71 ° pitch and roll, ± 180 ° azimuth to be
made. The slew rate of pitch and roll is approximately 2 °/min.
Since attitude changes are transmitted to interfacing systems,
slewing can be used to verify the integrity of wiring, to calibrate a
weather radar antenna and flight director HSI and ADI displays, or to
verify such maintenance activities as replacement of HSls, ADis.
weather radar antennae and AFCS pitch and roll computers.
~tem malfuncti<>ns
In the event of any malfunctions occurring in flight, appropriate
annunciations are made in a manner which depends on the type of
system installed. In the case of an INS, a 'WARN' annunciator light
on the CDU is illuminated, and in order to determine the cause of
the warning, and what action is to be taken, a corresponding code
numbering system is programmed into the computer. The code
numbers are made to appear in the right-hand display of the CDU by
moving the display selector switch to the 'DSR TK/STS' position,
and also by operating the 'TEST' switch. There are two groups of
code numbers designated as: (i) action codes and (ii) ma/function
codes.
Action codes
These are always displayed first, after the display selector switch is
moved to the 'DSR TK/STS' position. The codes and their
interpretation are as follows:
01. Complete system is inoperative.
02. Failure of the computer; the INS is no longer used for
navigation of the aircraft, but may be used to supply attitude data.
03. Does not illuminate the 'WARN' annunciator, but is detected
by monitoring the INS display on the FD system HSI. The computer
,.feeds data to the CDU and to a digital-to-analog converter in the
navigation unit. The code indicates that the INS may be used, but the
HSI and the AFCS should not be used in the INS modes.
04. Indicates abnormality that may be eliminated if the system is
realigned (see page 279).
05. Data from the automatic data entry unit is not reasonable. The
system should be switched off, then back on and automatic data
loading again be attempted. If this causes warning again, the data
may be loaded manually.
Before complying with an action code, the malfunction causing the
281
warning must be determined by operating the 'TEST' switch on the
CDU. Successive operation of the switch causes all existing
malfunction code numbers to appear in sequence in the right-hand
display.
~alfunction codes
These appear in place of the action codes. After the last malfunction
code is displayed, either an action code number is again displayed or
the display remains blank. If the latter is the· case, the malfunction is
momentary and the corresponding logic circuits are reset; the
'WARN' light is also extinguished. If, however, the action code
number reappears, the recommended procedure associated with it
must be carried out.
Any malfunction that occurs is stored in the computer memory and
their code numbers, of which there are 31, can be retrieved for
maintenance check-out and rectification purposes.
In an IRS, malfunctions are indicated by an amber 'FAULT' light
on the IRMP (see Fig. 10.22), a yellow 'fault ball' on the inertial
reference unit, and also appropriate messages displayed on the screen
of the interfaced engine indicating and crew alerting system (EICAS).
An invalid 'bit' is also transmitted to the data busses connecting other
user systems so that they may also generate fault messages. For
example, messages relating to IRS malfunctions can be generated and
displayed on the screens of an ADI and HSI comprising the
electronic flight instrument system (EFIS) and also on a maintenance
control and display panel (MRCP) unit which is integrated with all
interfacing computer-controlled systems in an aircraft. The IR
computer stores the appropriate message status words in a nonvolatile memory, and by means of the controls on the MRCP, these
can then be extracted and identified for the purpose of ground testing
and fault isolation.
282
11 Electronic (CRT) displays
Displays of this type, which are based on the electron beam scanning
technique, have been in use in aircraft for very many years. For
example, during World War II military aircraft used equipment
developed from the then existing ground-based radar systems. With
the aid of such equipment, and depending on an aircraft's specific
operational role, crews were able to navigate by 'radar mapping' of
terrain, to identify ground target areas, and also to detect the
positions of hostile intercepting aircraft.
As far as civil aircraft are concerned, this display technology first
came into prominence in 1946, with the introduction of weather radar
systems to satisfy the operational requirements for transport category
aircraft, and it has continued to be an essential part of the 'avionics
fit' of this and other categories of aircraft.
The situation, however, of a weather radar display indicator
remaining as an isolated item of video equipment was to undergo·
considerable change, largely as a result of systems analysis,
exploration of the versatility of the CRT, and also investigation into
methods whereby not only weather data, but also that associated with
the many other utilities systems of an aircraft, could be programmed
into computers. These had reached such high levels of sophistication
and capacity for data processing that it became possible for a single
CRT display unit, under microprocessor control, to project the same
quantity of system status data which would otherwise have to be
displayed by a very large number of conventional-type instruments.
Furthermore, the introduction of CRTs and circuits capable of
producing a wide range of colours made it possible to differentiate
between significant parts of a display, and in particular, to lay
emphasis on information of an advisory, cautionary, or warning
nature.
The development of such multi-data display technology for both
civil and military aircraft was also influenced by the fact that by
integrating all computers via a data 'highway' bus, the scene was set
for the management of all aspects of in-flight operation to be fully
automated while still enhancing flight safety. This also led to
improvements in levels of systems' redundancy, changes in the
layouts of transport aircraft flight decks, and a reduction in crew
complement with the attendant changes in their role and workloads.
The fi~st of the 'new technology' transport aircraft (generally
dubbed as 'glass cockpit' aircraft) were the Boeing 757, 767 and
283
Figure JI. I Flight deck layout
of the Boeing 757.
Airbus A3 IO. All three were launched as design projects in 1978,
and both the B757 and B767 first entered commercial service in the
US in December 1982. The first A310 services were operated by two
European airlines in April 1983. These aircraft, and several of their
descendant types, are now in service world-wide, together with many
types of smaller aircraft, including helicopters, in which the foregoing
technology has also satisfied an operational need.
Figures 11. l and 11.2 show the flight deck layouts and CRT
display locations of the B757 and A3 l0 respectively.
Principle of the CRT
284
A CRT is a thermionic device, i.e. one in which electrons are
liberated as a result of heat energy. As may be seen from Fig. 11.3,
it consists of an evacuated glass envelope, inside which are positioned
an electron 'gun' and beam-focusing and beam-deflection systems.
The inside surface of the screen is coated with a crystalline solid
material known as a phosphor. The electron 'gun' consists of an
Figure II. 2 Flight deck layout
of the Airbus A310.
indirectly-heated cathode biased negatively with respect to the screen,
a cylindrical grid surrounding the cathode, and two (sometimes three)
anodes. When the cathode is heated, electrons are liberated and in
passing through the anodes they are made to form a beam.
The grid is maintained at a negative potential, its purpose being to
control the current and so modulate the beam of electrons passing
through the hole in the grid. The anodes are at. a positive potential
with respect to the cathode, and they accelerate the electrons to a
high velocity until they strike the screen coating. The anodes also
provide a means of focusing, which, as will be noted from Fig. l l.3,
happens in two stages.
285
Figure //.3
Cathode ray tube.
Screen
beam
Graphite coaling
(collects secondary electrons
to prevent screen becoming
negatively charged)
The forces exerted by the field set up between the grid and the first
anode bring the electrons into focus at a point just in front of the
anode, at which point they diverge, and are then brought to a second
focal point by the fields in. the region between the three anodes. A
focus control is provided which by adjustment of the potential at the
third anode makes the focal point coincide with the position of the
screen. When the electrons impact on the screen coating, the
phosphor material luminesces at the beam focal point, causing
emission of a spot of light on the face of the screen.
In order to 'trace out' a luminescent display, it is necessary for the
spot of light to be deflected about horizontal and vertical axes, and
for this purpose a beam-deflection system is also provided. Deflection
systems can be either electrostatic or electromagnetic, the latter being
used in the tubes applied to the display units of aircraft systems.
The manner in which an electromagnetic field is able to deflect an
electron beam is illustrated in Fig. 11.4. A moving electron
constitutes an electric current, and so a magnetic field will exist
around it in the same way as a field around a current-carrying
conductor. In the same way that a conductor will experience a
deflecting force when placed in a permanent magnetic field, so an
electron beam can be forced to move when subjected to
electromagnetic fields acting across the space within the tube. Coils
are therefore provided around the neck of the tube, and are
configured so that fields are produced horizontally (X-axis fields) and
vertically (Y-axis fields). The coils are connected to the signal
sources whose variables are to be displayed, and the electron beam
can be deflected to the left or right, up or down, or along some
resultant direction depending on the polarities produced by the coils,
and on whether one alone is energized, or both are energized
simultaneously.
286
Fi111m' //.4
Magnetic field
Electron beam
deflection.
I
,
,..,--L '
I
I
',
{l"'-.... \
_.
'
....
'
'
_,.;
/
-~-"
'\
I
~
/
/
Electron beam coming
out of paper
Vertically disposed magnetic coil
produces horizontal deflection of beam
<;::~;>
/'·
\\
Ii
\\
Horizontally disposed magnetic coil
produces vertical deflection of beam
Figure 11.5 Data cells.
Rho-theta; (b) X-Y
coordinate.
(a)
\."'-.>---. -· ~/'·J
./
---
Resultant deflection of beam
w ·EB·
(a)
Colour CRT displays
.\
Jj
(b)
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 in the types of aircraft referred to earlier. In these
display units weather data is also integrated with the other data
displays, and since there is a fundamental similarity between the
methods through which they are implemented, the operation of a
weather radar display unit serves as a useful basis for study of the
display principles involved.
The video data received from a radar antenna is conventionally in
what is termed rho-theta form, corresponding to the 'sweeping'
movement of the antenna as it is driven by its motor (see Fig.
1 l.5(a)). In a colour display indicator, the scanning of data is
somewhat similar to that adopted in the tube of a television receiver,
i.e. raster scanning in horizontal lines. The received data is still in
rho-theta form, but in order for it to be displayed it must be
converted into an X-Y coordinate format as shown in Fig. 1l.5(b)).
This format also permits the display of other data in areas of the
screen where weather data is not displayed. In addition it permits a
doubling-up of the number of data cells, as indicated by the dotted
lines in the diagram.
287
Each time the radar transmitter transmits a pulse, the receiver
begins receiving return echoes from 'targets' at varying distances
(rho) from the transmitter. This data is digitized to provide output
levels in binary-coded form, and is supplied to the indicator on two
data lines. The binary-coded data can represent four conditions
corresponding to the level of the return echoes which, in turn, are
related to the weather conditions prevailing at the range in nm
preselected on the indicator. The data are stored in memories which,
on being addressed as the CRT is scanned, will at the proper time
permit the weather condition to be displayed. The four conditions are
displayed as follows:
Blank screen:
Green:
Yellow:
Red:
Zero or low-level returns.
Low returns (lowest rainfall rate).
Moderate returns (moderate rainfall rate).
Strong returns (high-density rainfall rate).
Scan conversion
The principle of conversion from rho-theta form to an X - Y
coordinate scan is shown in Fig. l l.6. With a 'target' at point P, at a
range Rand an angle 8, it will have coordinates: X = R sin 8 and Y
= R cos 8. Thus, for an echo received at an azimuth angle of, say,
30° and a range of 235 nm, the coordinates will be: X = 235 sin
30° = 117.5 nm, and Y = 235 cos 30° = 203.5 nm. The
conversion is performed by a microprocessor on the indicator's
display circuit board.
Screen format
The coordinate system format of the screen is shown in Fig. 11. 7,
and from this example it will be noted that the screen is divided into
two halves representing two quadrants in the coordinate system. The
origin is at the bottom .centre, so that values of X are negative to the
left and positive to the right; all values of Y are positive. The screen
Figure I I. 6 Scan conversion.
--
Aircraft heading
11H
(al
288
!b)
Figure ll. 7 Screen format.
Horizontal
sweep
-""l'\.~- - - - - - - - - - - - - ; ' - : \,. - 1
1861 1127
...ii
Horizontal
counter
~
r-
1281
Upldown
b~~:-f1-------+--------n-
_ _ _ _ _ _ 75µs - ; . . - - - - - - " " 1 ._.,4µs
l
.
E
~
,.
Q.
f.
C""
.. C
c)l!
Line seen tima = 61 µs
- - - - L i n e blank time= 14µs---"'1
is scanned in 256 horizontal lines, and there are 256 'bits' of
information displayed on each line.
Each line is located by a value of Y and each bit by a value of X;
the screen therefore has a 256 x 256 matrix. The X and Y values
are used to address the memory and display the information stored
there as the appropriate time in the scan occurs. The memory for the
weather data is in two parts which store the bits of the data words
that represent the colours red, yellow or green and the corresponding
weather conditions. Each part of the memory contains one address for
every bit on every line in the display; each memory, therefore, is
also a 256 x 256 matrix, and allows the entire weather display to be
stored continuously.
As the screen is scanned, the memory is addressed at each point on
each line by two counters: a horizontal or X counter for addressing
the rows in the memory, and a vertical or Y counter for addressing
the columns. The X counter generates an output for each of the 256
bits on a line, and counting is started by a 'high' state output signal
from an up/down divider circuit. The counter is caused to count
down, i.e. left to right, from the number 186 to O at the centre of the
screen. When it reaches 0, the divider circuit changes to a 'low' state
output, thereby causing the counter to count up to the number 128 at
the end of the line, at which point a 'line blank' pulse of 14 µ.s
duration is generated. The line scan time is about 61 µ.s, and so the
total time for each line is 75 µs. The divider circuit again changes to
289
a 'high' state to cause the counter to start down for the next line, and
is a process that is repeated for all remaining lines.
An output from the X counter is also applied to the Y counter,
which counts to 256 (one for each line) plus eight counts for a scan
blank time to allow for the CRT beam 'spot' to return to the upper
left comer of the screen. This process is repeated, and since there are
256 lines in the display it takes 20 ms to scan the entire screen (19.4
ms for the 256 lines and 600 µs for the scan blank time). The
vertical and horizontal sweep circuits are synchronized by the
triggering of the line and scan blank pulses.
In addition to the foregoing raster scanning technique, which
produces sections of a CRT screen in 'solid' colour, a stroke
scanning technique is also used for producing displays of symbols and
of data in alphanumeric format. Details of this will be given later in
this chapter.
Colour generation
A colour CRT has three electron guns, each of which can direct an
electron beam at the screen which is coated with three different kinds
of phosphor material. On being bombarded by electron beams, the
phosphors luminesce in each of the three primary colours red, green
and blue.
The screen is divided into a large number of small areas or dots,
each of which contains a phosphor of each kind as shown in Fig.
11.8. The beam from a particular gun must only be able to strike
screen elements of one colour, and to achieve this a perforated steel
sheet called a shadow mask is accurately positioned adjacent to the
coating of the screen. The perforations are arranged in a regular
pattern, and their number depends on the size of screen; 330 000 is
typical.
Beams emitted from each gun pass through the perforations in the
mask and they cause the phosphor dots in the coating to luminesce in
the appropriate colour. For example, if a beam is being emitted by
the 'red' electron gun only, then only the red dots will luminesce,
and if the beam completes a full raster scan of the screen, then as a
result of persistence of vision by the human eye, a completely red
screen will be observed. In the display units of electronic instrument
systems, a number of other colours are also required and these are
derived by independent circuit control of the three guns and their
beam currents, so that as the beams strike the corresponding
phosphor dots, the basic process of mixing of primary colours takes
place (see Fig. 11.9). In other words, an electronic form of 'paint
mixing' is carried out.
Referring once again to the weather radar indicator application, the
data readout from the memory, apart from being presented at the
Figure I !.8
Colour CRT.
A
G
8
R
G
G
8
I
I
I
R
G
8
'
I
I
'
/1)
Shadow mask
Coating
MASK APERTURE
SHADOW MASK
PHOSPHOR
OOT SCREEN
appropriate location of the CRT screen, must also be displayed in the
colours corresponding to the weather conditions prevailing. In order
to achieve this, the data is decoded to produce outputs which, after
amplification, will tum on the requisite colour guns; the data flow is
shown in Fig. 11.10. The memory output is applied to a data
291
Figure II. 9 Weather data
.. ,·.-.
··.·· ...
display.
Range data
blue
Green
Q.··
,,,.,,L-_ _ _
1
Yellow
Red
wx
Mode
blue
~
0
.
10
,,
Table I I. I
Outputs
Inputs
M
L
A1
0
0
0
I
0
I
Col
A3
A2
Black (oft)
Green
Yellow
0
0
0
Red
Table 11.2
Outputs to guns
B 1 Green
0
0
0
Resulting colours
B0 Blue
B2 Red
I
0
0
I
0
0
0
I
0
Black (oft)
White
Yellow
Red
Ligh1 blue
Green
demultiplexer whose output corresponds to the most significant and
least significant bits (M and L) of the two-bit binary words and is
supplied to a data decoder. The inputs are decoded to provide threebit output words corresponding to the colours to be displayed, as
shown in Table 11. l . The outputs are then applied to the colour
decoder and primary er.coder circuit, and this in turn provides three
outputs. each of which corresponds to one of the colour guns as
shown in Table 11.2.
The 'low' state outputs turn on the guns, and from Table 11.2 it
Figure I I. JO Data Oow for
A,
M
gun operation.
Memory
Output
dala
DMUX
Dala
decoder
L
A,
A,
Green
Colour
decoder
& priority
encoder
Blue
Red
Gu., amplifters
Guns
can also be seen how simultaneous gun operation produces other
colours from a mix of the primary colours. Figure 11. 9 illustrates a
typicai weather data display together with associated alphanumeric
data, namely ranges in ·nm, and an operating mode which in this case
is WX signifiying 'weather' mode.
Alphanumeric
displays
The display of data in alphanumeric and in symbolic form is
extremely wide-ranging. For example, in a weather radar indicator it
is usually only required for range information and indications of
selected operating modes to be displayed, while in systems designed
to perform functions within the realm of flight management, a very
much higher proportion of information must be 'written' on the
screens of the relevant display units. This is accomplished in a
manner similar to that adopted for the display of weather data, but
additional memory circuits, decoders, and character and symbol
generator circuits are required.
Raster scanning is also used, but where datum marks, arcs or other
cursive symbols are to be displayed, a stroke pulse method of
scanning is adopted. The position of each character on the screen is
predetermined and stored in a memory matrix, typically 5 x 7, and
when the matrix is addressed, the character is formed within a
corresponding matrix of dots on the screen by video signal pulses
produced as the lines are scanned.
Figure 11.11 illustrates how, for example, the letters 'WX' and the
number '40' are formed. One line of dots is written at a time for the
area in which the characters are to be displayed, and so for a 5· x 7
matrix, seven image lines are needed to write complete characters
and/or row of characters. As will be noted from Fig. 11.9, the
characters are displayed in blue, so only the 'blue' electron gun is
active in producing them. Spacing is necessary between individual
characters and also between rows of characters, and so extra line
'blanking bits', e.g. three, are allocated to character display areas.
In the example of the weather radar indicator, the characters each
have an al.location of eight bits (five for the characters and three for
the space following) on each of 21 lines (14 for the character and
seven for the space below). The increase in character depth to 14
lines is derived from an alphanumeric address generator output that
writes each line in a character twice during line scanning. The
293
Figure./ l.11
display.
Alphanumeric
15141 3 12 11 10 9 8 7 6 5 4 3 2 1 0 0 1 2 3 4 5 6 7 8 9101112131415
0
I
~D
2
3
5
5J
""
rn
._nr
6
8
9
lD
10
11
i ..
. ,.
, _ __ _
16 characters---~--- 16 characters-----i.,l
character format in this case permits the display of 12 rows each of
32 characters.
The CRT display units of the more comprehensive electronic
instrument systems (see Chapters 12 and 16) operate on the same
fundamental principles as those described, but in applying them, more
extensive microprocessing circuit arrangements are required in order
to display far greater amounts of changing data in quantitative and
qualitative form.
The microprocessor processes information from the data 'highway'
bus and, from the memory circuits, it is instructed to call up subprograms, each. of which correspond to the individual sets of data that
are required to be displayed. Signals are then generated in the
relevant binary format, and are supplied to a symbol generator unit.
This unit, in turn, generates and supplies signals to the beam
deflection and colour gun circuits of the CRT, such that its beams are
raster and stroke scanned, to present the data at the relevant parts of
the screen, and in the required colour.
The displayed data is in two basic forms: fixed and moving. Fixed
data relate in particular to such presentations as symbols, scale
markings, names of systems, datum marks, names of parameters
being measured, etc. Moving data are in the majority, of course,
since they present changes occurring in the measurement of all
parameters essential for in-flight management. The changes are
indicated by the movement of symbolic pointers, index marks, digital
counter presentations, and system status messages, to name but a
few.
12 Electronic flight
instrument systems
As far as the pure basic functions and number of display units are
concerned, this system, which is generally referred to as 'EFIS'
(pronounced ee-fiss), may be considered as being similar to the types
of flight director system described in Chapter 9. However, 'since it is
full inte rated with digital com uter-ba ed na . .
stems, andJ.!tilizes colour CRT types of ADJ and HSI, then it is far more
sophisticatediiot only in terms of physical construction, but also in
the extent to which it can present attitude and navigational data to the
flight crew of an aircraft.
Units of
a system
As in the case of conventional flight director systems,~
installation is made up of left (Captain) and -right (First Officer)
systems. Each system m turn 1~ com rised of two displa_}'. units: an
electronic ·
·rector indicatg/(EADI} and an ~tronic
porizontal situation indicator, (~HSI1i a control pgnel, a symbol
generator (SG), and a remote light sensor unit, A.third (centre) SQ 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 powered from an
aircraft's de power supply via pilot-controlled switches. The interface
between EFIS units, data busses, and other systems is shown in Fig.
12. I.
l EFIS
Display units
Each display unit consists of the sub-units shown in Fig. 12.2. ~
power sup ly units provide the requisite eve s of ac and de wer
'necessary for overa operation; t ~pplies are automatically
re ulated and monitored for undervoltage and overvoltage co
ions. _)
The v1 eo mom or car contains a video control microprocessor,
video amplifiers and monitoring logic for the display unit. ,The mah!_
tasks of the processor and associated ROM and RAM memories are
~ulate ain factors for the three video amplifiers (red, blue and '
green), and perfm;m input and sensor an 1splay unit monit~
296
Figure 12. 1 EFIS units and
signal interfacing.
R
L
Control panel
Control panel
R
Attitude
director
ind!CStO!
L
Remote
ligl;!t sensor
Remote
1ight·sensor
Ambient light
sensor data
Light
sensor
data
Horizontal
Horizontal
situation
situation
indicator
indicator
Digital flight data
acquisition unit
L
R
Symbol
generator
vjR
OME
c{1!s
FCC
ILS
RAO. ALT R
L
WXR
{
IRS
fCC
FMCS
FMCS
t
t
TMC
L & R
Symbol
generator
Symbol
generator
1~s
RAO. ALT r!c {v!R
OME
{VOA
OME
WXR
IRS
FCC
C
{ IRS
FMCS
FCC
R
FMCS
R
ILS
.RAO ALT
t
c{IRS
FCC
t
TMC
L FMCS
Data busses
:::::::::>
Display unit
Dllllllllllllli
WXR
~'b5c
FMCS
drive signals ~
switched
drive signals
Lei! L
..,.
••
Right R Centre C
~ the input/output interface functions for the processor art'
provided by analog multiplexers, an AID converter and a D/ A
converter.
The function of the convergence card is to take X and Y deflection
signals and to develop drive signals for the three radial convergence
coils (red, blue and green) and the one lateral convergence coil (blue)
of the CRT. Voltage compensators monitor the deflection signals in
order to establish on which part of the CRT screen the beams are
located (right or left for the X comparator, and top or bottom for the
Y comparator).
Signals for the X and Y beam deflections for stroke and raster
scanning are provided by the deflection amplifier card. The amplifiers
for both beams each consist of a two-stage preamplifier and a power
amplifier. Both amplifiers use two supply inputs, 15 V de and 28 V
297
Figure 12.2 Display Uf1il.
115V 400Hz
HV Power supply
1-----------.
Light sensor from
other display unit
Remote light sensor - - - - - - - - - - - - - - - - - ,
OU brightness
Raster brightness
Red
Green
Blue
Beam test
Synchron,z,ng
O,gital line
receivers
Video
monitor card
Analog hne
receivers
Deflection
card
Intensity
Raster/stroke
Day/night
X deflection
Y deflection
Conyergence
card
de; the former is used for effecting most of the stroke scanning or
writing, while the latter is used for repositioning and raster scanning.
The interconnect card serves as the interface between the external
connector of a display unit and the various cards. Digital line
receivers for the signals supplied by the SGs are also located on this
card.
In a typical system[six colours are assigned for the displa.Y,jJ~Jhe
many symbols, failure annunciators, messages and other alphanumeric
information, and are as follows:
,-/'
White
yreen
.Qjsplay of [email protected]!ion information,/.
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, !£!_ive flight path lines.,
. Sky shading on an EADI and f(£:low-uriorit)". informatio!1
such as non-active flight plan map data.
Yellow Ground shading on an EADI, caution information disPJ.,ay
such as failure warning flap, limit and alert annunciators
Rdr~ult messages.
Red
~ display of h~lest. precipi~ion levels a s ~ ~
the weatner radar.
----
Symbol generators (SGs)
~ e provide the analog, discrete and digital signal interfaces
between an aircraft's systems, the displar units and the control.panel,
Figure 12. 3 Symbol generator
and card interfacing.
WXR data
WXR input
Main PROM
WXR Raster data
Main RAM
Display
unit
video
Raster
generator
WXR memory
2 • 16K RAMS
~
Main
processor
. .
;; "'
:,
.0
Display
controller
"'ii\ ::-:§
>!!!
u
C
5}
FMC
TMC
RAD. ALT
VOA
!!!
"a:,
0
5!
1/0
>,
!!!
processor 1
5}
0
EFIS control
panel
..,
~
0
"'.,
~
e
Stroke position
data
iii
IRS
L.C&R FCCs
ILS
OME
S.G. wrap~
Stroke
generator
Display
110
sequencer
processor 2
Display
around
Character data
uni1
raster/stroke
select
and they perform symbol generation, system monitoring, power
control and the main control functions of the EFIS overall. The
interfacing between the card modules of an SG is shown in Fig. 12.3,
and card functions are given in Table 12.1.
Control panel
tA_control panel is provided for each system, and, as shown in Fig.
12.4, the switches are rou
for the u ose of controllin the
displays of their respective EADI and EHSI units as 1sted in Table
12.2.
Remote light sensor
his '1$ a photodiode device which res nds to fli ht deck ambient
lig .
1 ions and automatically adjusts the brightness of the CRT
------·-·······---· ····· ···
displays acori'ipatible level.
.. ------......___
Display presentations
The EADI displays t~ional pitch and roll attitude indications
against a raster-scanned background, and as inay be seen from the
Table 12.1
Symbol generator card functions
Card
Function
l/0 I & 2
Supply of input data for use by 1he main processor
Main processor
Main data-processing and control for the system
Main.RAM
Address decoding, read/write memory and 1/0 functions for the syslem
Main PROM
Read-only memory for the system
Display controller
Master transfer bus interface
WXR inpul
Time scheduling and interleaving for raster. refresh, input and standby
functions of weather radar input data
WXR memory
RAM selection for single-input data. row and column shifters for
rotate/translate algorithm, and shift registers for video output
Display sequencer
Loads data into registers on stroke and raster generator cards
Stroke generator
Generates all single characters. special symbols. straight and curved lines
and arcs on display units
Raster generator
Generates master timing signals for raster, stroke, EADI and EHSI
functions
Display driver
Converts and multiplexes X and Y digital stroke and raster inputs into
analog for driver operation, and also monitors deflection outputs for
proper operation
example illustrated in Fig. 12.5, the upper half is in cyan and the
lower half in yellow. Attitude data is provided by an IR system. Also
displayed are flight director commands, localizer and glide slope
deviation, selected airspeed, ground speed, AFCS and autothrottle
system modes, radio altitude and decision height.
Figure 12.5 illustrates a display representative of an automaticallycontrolled approach to land situation together with the colours of the
symbols and alphanumeric data produced via the EFIS control panel
and SGs. 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 l 00 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
Figure 12.4 Control panel.
ADI
RST
@
1Di~E~ 10
a
HSI
BRT
@
40
0:::NE]
~~~~~~~-MAP~~~~~~~~
NAV AID
ARPT RTE DATA WPT
DJG][GJ[GJ[GJ
ARPT Airport
RTE DATA Route Data
WPT Waypoint
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.
301
Table 12.2
Switch
Funcrwn
EADI section:
BRT
DH SET
RST
EHSI section:
RANGE
MODE SELECT
BRT
WXR
MAP switches
Figure 12.5 EADI display.
Controls levels of display brightness.
Setting of decision height.
Manually resets decision height circuus after aircraft has passed
through decision heighL
Selects range for displayed WXR and navigation data.
Selects display appropriate to mode required.
Outer knob controls main display brightness: inner knob controls WXR
display.
When pushed in. WXR data displayed during all modes except PLAN.
Used in MAP mode. and when pushed in they cause their placarded
data to be displayed. Illuminate white.
Roll Pointer (W)
Pitch scales
(W)
Control panel swuch function,
Roll scale
fW)
- - - -....
Sky shading (C Raster)
-~~-+t----,.{..._
Selected decision
height (G)
Radio altitude
(W)
Speed error
Flight director
command bara
(M)
Aircraft symbol
scale (W)
Horizon
llne (W}
Glide slope
deviation Pointer
DH
~
Radio
altitude
(W)
(M)
Localizer deviation Slip indicator
scale (Wl
(ball in tube)
Localizer deviation
Pointer (Ml
Glide slope
deviation scale (W)
Autoland status (G)
Pitch engaged
mode(Gl
Roll engaged
mode(G)
Ground shading
(Y Raster)
Colours:
C Cyan
G 'Green
M Magenta
W White
Y Yellow
The EHSI present5 a selectable, dynamic colour display of flight
progress'ano"pfan view orientation,: Fo1ffprincip;:ilJiispJay,modes may
be selected on the control panel:_MAP..i,.lk~N,JL{LanclVOR. Figure
12.6 illustrates the normally-used MAP mode display which,j!}
conjunction with the flight plan data programmed into a flight
11111nagemen1,computer, displays fotorination agaTnsi'a~moving map
background w1tfa1f elt:mOl1lS posit1oneo to a common s<:a1e.
The symbol representing the aircraft is at the lower part of the
display and an arc of the compass scale, or rose, covering 30° on
either side of the instantaneous track is at the upper part of the
302
Figure /2.6
mode.
EHSI in 'MAP'
Display orientation
(G)
Estimated time
of arrival (W)
Distance to go
(W)
---.--+-- Magnetic
compass
rose(W)
Manually set
VOA course---1--r-
Manually selected
course (mag. hdg.)
(G)
(M)
Range scale
---1--t--;;:;
(W)
Range to
altitude---..J-__,_ _
Vertical deviation
scale (W)
(G)
Vertical deviation
pointer (M)
Wind speed
and direction
(W)
ILS frequency (G)
Lateral deviation
scale (W)
+
0
Waypoints: Active (M) on-e a1rcrat1 currently navigating to
Inactive (W) a navigation point making up selected active route.
Airports (C)
~ Navaids (C)
/
Wind direction (W) with respect 10 map display orientation
6
Off-route wayPoints (C)
and compass reference,
Colours:
C
Cyan
G Green
M Magenta
W White
display. Heading information is supplied by the appropriate IRS, and
the compass rose is automatically referenced to magnetic north (via a
crew-operated 'MAG/TRUE' selector switch) when between latitudes
73° N and 65° S, and to true north when above these latitudes.
When the selector switch is set at 'TRUE', the compass rose is
referenced to true north regardless of latitude .
.Tuned VOR/DME stations, airports and their identification letters,
and the flight plan entered into the flight management system
computer are all correctly oriented with respect to the positions and
track of the aircraft, and to the range scale (nm/in) selected on the
EFIS control panel. Weather radar 'returns' may also be selected and
303
fi'gure
12. 7 EHSI in 'PLAN'
Distance
(W)
Heading orientation (G)
mode.
ETA(W)
Active route (M)
Colours.
G Green
M Magenta
W White
displayed when required, at the same scale and orientation as the
map.
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, together with the display of track and heading information as
shown in Fig. 12.7. Any changes to the route are selected at the
keyboard of the flight management system display unit, and appear
on the EHSI display so that they can be checked by the flight crew
before they are entered into the flight management computer.
The VOR and ILS modes present a compass rose (either ewanded
or full) with heading orientation display as shown in Fig. 12.8.
Selected range, wind information and system source annunciation are
also displayed. If selected on the EFIS control panel, weather radar
returns may also be displayed, though only when the mode selected
presents an expanded compass rose.
Heading orientation (G)
Figure 12.8 VOR and ILS
mode displays.
Selected heading
marker (M)
Course line (M)
Course indicator
(M)
...."
Nav data source (G)
Expanded
Full
VOR mode
GIS pointer
(M)
ILS frequency (G)
Expanded
Full.
!LS moae
Colours
M Magenta
G Green
All other symbols. white
Failure annunciation
Failure of data signals from such systems as the ILS and radio
altimeter are 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.
305
Figure I 2. 9 Source selector
switch panel.
INSTR SOURCE SEL
FLT DIR
R
,~®)
@
"'~
~
~
EfI~
~
IRS~
AIR~
DATA~---
Data source selection
In the type of system described, means are provided whereby the
pilots can, and independently of each other, connect their respective
display units to alternate sources of input data, e.g. left or right
ADCs, flight management computers, flight control computers, and
standby IRS.
Each pilot has a panel of selector switches arranged as shown in
Fig, 12.9. The upper rotary t}'.pe of switch connects either the left,
centre or right flight control computer to the EADI as the source of
attitude data. The other switches are of the illuminated push type and
ar.e guarded to prevent accidental switching. In the normal operating
configuration of systems they r.'!main blank, and when activated they
are illuminated white.
Display of air data
In a number of EFIS applications, the display of such air data as
altitude, airspeed and vertical speed is provided in the conventional
306
Figure 12.10 Flight deck
layout of the Boeing 747-400
series aircraft.
manner, i.e. separate indicators servo-operated from ADCs are
mounted adjacent to the EFIS display units in the basic 'T'
arrangement (!iee page 21). With the continued development of
display technology, however, CRTs with much larger screen areas
have been produced ~nd, as may be seen from the Boeing 747-400
aircraft flight deck layout in Fig. 12.10, (see also front cover), such
displays make it unnecessary to provide conventional primary air data
instruments for each pilot.
307
13 Engine instruments
At the present time there are three principal types of engine in use,
namely, piston (unsuJercharged and turbocharged), turbopropeller,
and pure turbine, an i their selection as the means of propulsion for
any one type of aircraft depends on its size and operational category.
In each case there are certain parameters that are required to be
monitored to ensure that they are operated in accordance with their
designed performance ratings, and within specific limitations. The
parameters involved overall are listed in Fig. 13.l, the .. ctual number
required varying, of course, in accordance with the type of engine.
Monitoring is accomplished by means of specifically designed
instrument systems, the sensor units of which may, in very basic
form, be incorporated within an indicator, or be of the remote type
which transmit data in the form of electrical signals to 'clock'-type
indicators, or to electronic display units.
Certain of the parameters listed in Fig, 13.1 relate to the operation
of engine systems, e.g. the pressure and temperature of lubricating oil
systems, while others are more of a primary nature in that they relate
directly to the performance of engines in terms of their power and/or
thrust.
The instrument systems required for monitoring purposes may,
therefore, be broadly grouped into these two main areas, and so the
details of operating principles and construction typical of those
currently in use are accordingly arranged to forrn the subject of this .
chapter and also of Chapter 15. The monitoring of data by means of
CRT-type display units will be covered in Chapter 16.
Pressure
measurement
Figure 13. I Monitored
operating parameters.
Pressure is measured by instrument systems which in the majority of
applications are of the remote-indicating type, i.e. their sensor (or
PRESSURE
TEMPERATURE
I
I
OIL
OIL
FUEL
FUEL
I
i
ENG1;~E PRESSURE
RATIO
I
M,\NIFOLD
I
TORQUE
I
I
CYLINDER HEAD
I
EXHAUST GAS
FUEL FLOW
VIBRATION
transmitter) units are connected to a pressure source located at some
remote point, and they transmit data through an electrical
transmission circuit.
The sensor units contain elements which, depending on the
particular design and pressure ranges to be measured, are in the form
of either metal capsules, diaphragms or bellows. Another form of
element which may be mentioned at this point is the earliest type ever
to be adopted, namely, the Bourdon tube. In present-day aircraft,
however, its application, if required at all, is limited to certain
systems other than those of engines, in which direct-reading
indicators may be permitted for monitoring of operation.
The elements are mechanically connected to electrical transmitters
which, in some cases, are of the moving core or synchro type of
induction device. These, in turn, are connected to indicators which
incorporate either a moving coil mechanism, a synchro receiver, or a
synchro/servomotor mechanism, as appropriate to the particular
design.
Indicating systems
Figure 13.2 illustrates the arrangement of a type of turbine .erig.;ne oil
pressure indicating system that utilizes a TX synchro transmission
system (see Chapter 5). The rotor of the TX is in this case connected
via a mechanical quadrant and pinion to two bellows; one is sensitive
Figure 13. 2 Synchronous
transmission-type pressure
indica1ing system.
1-·
28
AC
OIL
PRESSURE
_L
AMBIENT
COWL PRESSURE~=======
TRANSMITTER
v
'-----·---·
---INDICATOR
Figure 13. 3 Servo-operated
pressure indicating system.
---------·-----·---__,,_"'---el
26 V AC _
POWER
SUPPLY
UNIT
J
__
_
____
I•
Feedback
CONTROL
o
I, .
Feedback
I•
I
~ - - PRESSURE CHANGE
SIGNALS
'.
_ _J
to oil pressure, the other to prevailing pressure of the ambient air in
the cowled area of the engine. Thus, the rotor is positioned so as to
produce output signals proportional to the difference between the two
pressures, i.e. gauge pressure. In responding to these signals, the TR
rotor positions the pointer over the indicator scale .via an appropriate
ratio gear train. In the event of failure of electrical power to the
system then, due to the gear train, the indicator pointer will remain
at the pressure value that was being measured at the time of failure.
An example of a system which employs a capsule and moving core
type of inductive sensor and a servo-operated indicator is shown in
Fig. 13.3. This system is used in some series of Boeing 747 for the
measurement of such parameters as engine oil and oil filter inlet
pressures, fuel pump inlet and discharge pressures, and engine
breather pressure.
In its application to oil pressure measurement, the indicator
contains dual servomechanisms connected to individual transmitters,
one of which senses oil system pressure, the other oil filter inlet
310
Figure I 3.4 Pressure
transminer.
RECTANGULAR AIR GAPS
SPRING
CAPSULE
STATOR COIL 2
ANNULAR
AIR GAPS
ARMATURE
STATOR COIL 1
pressure. The only difference between the two is their operating
pressure ranges, which respectively are from O to 100 psi and O to
200 psi.
When pressures are applied to the capsules they vary the position
of the inductor core, and thereby cause a change in reluctance
between the two windings so that output signals proportional to
pressure are produced. These signals are supplied to the amplifiers of
the indicator servomechanisms. The mechanism connected to the oil
pressure transmitter drives a 'double' pointer, one part of which
registers against a fixed outer scale. The mechanism connected to the
oil filter inlet pressure transmitter drives an inner disc with an index
marker that is also registered against the outer scale.
The inner disc- also has a scale which registers against the second
part of the double pointer so that, in operation, a continuous
indication of pressure difference is provided. In the example readings
shown in Fig. 13.3, the oil pressure is 50 psi, filter inlet pressure is
100 psi and so the difference indicated is 50 psi.
The servomotor of each mechanism drives a potentiometer which
provides feedback signals to balance out those from the transmitters.
A solenoid-operated flag is provided in the indicator, and comes into
view whene'ller there is a failure of the 26 V ac power supply to the
system.
Another form of ac inductor type of pressure transmitter is shown
in Fig. 13.4. It utilizes a capsule which positions an armature core
relative to air gaps in the core of a stator. With pressure applied as
indicated, the length of the air gaps associated with stator coil 1 is
decreased. while that associated with coil 2 is increased. As the
311
reluctance of the magnetic circuit across each coil is proportional to
the effective length of the air gap, then the inductance of coil I will
be increased and tliat of coil 2 descreased; the current flowing in the
coils will, respective;y, be decreased and increased. The output
signals are supplied to a moving coil type of indicator which operates
on the de ratiometer principle (see page 316).
In several types of 'n!w technology' aircraft, pressure transmitters
are, as far as internal mechanical arrangement is concerned, much
simpler in that they use piezoelectric-type sensors as in the case of air
data computers (see page 165).
Pressure switches
In the measurement of pressure, it is a requirement in the case of
some engine systems that the flight crew be given some positive
indication of pressure variations which could constitute hazardous
operating conditions. This requirement also applies, of course, to
other systems involving liquids and/or gases under pressure which are
used in the operation of aircraft. To meet this requirement, therefore,
pressure switches are installed in the relevant systems and are
connected to indicator or warning lights (in some cases aural warning
devices also) located on cockpit or flight deck panels.
The most common type of pressure switch consists of a diaphragm(or capsule)-type sensor which is exposed on one side to oil or fuel
pressure as appropriate, and on the other side to a local ambient
pressure. The displacements of the sensor in response to pressure
changes cause it to activate a switch, the contacts of which are
connected to the relevant annunciator light and remain open under the
normal high-pressure operating conditions. If the pressure should fall
to a value at which the contacts have been pre-set to 'make', then the
circuit is completed for illumination of the annunciator light.
In engine oil systems, valves are provided to bypass oil around
filters in the event of their becoming clogged, and as an indication of
this bypassing operation pressure switches can also be used. An
example of one such application is shown in Fig. 13.5. In this case,
one side of the switch sensor is exposed to filter inlet pressure, and
the other side to filter outlet pressure, and, as will also be noted, its
switch contacts are arranged the reverse way round to that just
described, i.e. they keep the annunciator light circuit open under low
pressure conditions. Thus, if the pressure difference across the filter
increases to the pre-set highest value, the contacts complete the
circuit to the annunciator light which illuminates to indicate that the
bypass valve is about to open so that the oil can flow around the
filter.
The pressure inlets of switch units are normally in the mounting
flange, and they may either be in the form of plain entry holes
312
28 V de
Figure 13.5 Pressure switch
operation.
ANNUNCIATOR
LIGHTS
LOW OIL
PRESSURE
OIL FILTER
BYPASS
1B
LOW
HIGH
11
-
FILTER INLET
-FILTER OUTLET
(a)
(b)
Figure /3.6 Typical pressure
,witch.
PRESSURE CONNECTION
directly over the pressure source, spigots with 'O' ring seals as in the
example shown in Fig. 13.6, or threaded connectors for flexible or
rigid pipe coupling.
Temperature
measurement
In most forms of temperature measurement, the variation of some
property of a substance with temperature is utilized. These variations
may be summarized as follows:
1. Most substances expand as their temperature rises; thus, a
measure of temperature is obtainable by taking equal amounts of
expansion to mdicate equal increments of temperature.
313
2. When subjected to a temperature rise, many liquids experience
such motion of their molecules that there is a change of state from
liquid to vapour. Equal increments of temperature may, therefore, be
indicated by measuring equal increments of vapour pressure.
3. Substances change their electrical resistance when subjected to
varying temperatures, so that a measure of temperature is obtainable
by taking equal increments of resistance to indicate corresponding
changes of temperature.
4. Dissimilar metals when joined at their ends produce an
electromotive force (thermo-emf) dependent on the difference in
temperature between the two junctions. Since equal increments of
temper_ature are only required at one junction, a measure of the
thermo-emf produced will be a measure of the junction temperature.
5. The radiation emitted by any body at any wavelength is a
function of the temperature of the body, and is termed its emissivity.
If, therefore, the radiation is measured and the emissivity is known,
the temperature of the body can be determined; such a measuring
technique is known as radiation pyrometry.
In relation to engine instruments, we are particularly interested in
variations 3 and 4, since they are utilized extensively for the
measurement of temperature of such liquids and gases as oil, fuel,
carburettor air, and turbine exhaust gases. In certain types of turbojet
engine, the radiation pyrometry technique is adopted for the
measurement of actual turbine blade temperature.
Indicating systems
These fall into two main categories: variable resistance and
thermoelectric, and so, respectively, they are directly related to the
variations 3 and 4 noted earlier.
Variable resistance systems
A system consists of a sensor unit (generally referred to as a 'bulb')
and an indicator, connected in a series circuit configuration, and
requiring de power which may be directly supplied from a relevant
busbar or, in some cases, by rectification of a single-phase ac supply.
Sensor units employ resistance coils of either nickel or platinum wire,
and the indicator units are of the moving coil type, having their
internal circuits arranged in either the basic Wheatstone bridge
configuration, or in the more commonly adopted ratiometer
configuration.
Sensor units The general arrangement of a sensor unit commonly
used for the measurement of liquid temperatures is shown
schematically in Fig. 13.7. The resistance coil is wound on an
insulated former and the ends of the coil are connected to a two-pin
314
Figure I 3. 7 Schematic
arrangement of a temperature
sensor.
UNION NUT
PLUG
RECEPTACLE
2-PIN SOCKET
CONTACTING STRIPS
I
CAUBRATlNli OR
BAI.ANCING COIL
socket (or a plug) via contact strips. The 'bulb'. which serves as a
casing to protect and to seal the coil assembly. is a stainless steel
tube closed at one end and secured to a union nut at the other. The
union nut is used for securing the complete unit to the connecting
point at which the temperature is to be measured.
It will be noted from the diagram that the coil is wound at the
bottom end of its former and not along the full length. This ensures
that the coil is well immersed in the hottest part of the liquid. thus
minimizing errors due to radiation and conduction losses in the
'bulb'.
A calibrating or balancing coil is normally provided so that a
standard constant temperature/resistance characteristic can be
obtained. thus permitting interchangeability of sensor units. In
addition the coil compem,ates for any slight change in the physical
characteristics of the sensor material. The coil. which may be made
from Manganin or Eureka wire. is connected in series with the sensor
coil and is pre-set in value during initial calibration by the
manufacturer.
The resistance of nickel and platinum increases with an increase in
temperature, and the corresponding values are given in Table 3 on
page 409.
Wheatstone bridge systems Figure 13.8 shows the arrangement of a
system in which measurements are based on the principle of the
Wheatstone bridge circuit. Although very old in concept, it is still
applied to temperature indicators adopted in some types of aircraft.
The temperature sensor forms the unknown resistance arm R, of
the bridge, while the other three are contained within the moving coil
type of indicator; resistance values of each of these three arms are
fixed. The moving coil is wound on a former which is pivoted so that
it can rotate within the field of a permanent magnet. and. by means of
two controlling hairsprings, it is connected across the .. points B and D
of the bridge. The de supply for the system is connected across
points A and C.
When the sensor R, is subjected to temperature variations, its
315
Figure 13.8 Wheatstone bridge
28 V de
system.
SENSOR UNIT
resistance will also vary. This upsets the balance of the bridge
circuit, and the value of R, at any particular temperature will govern
the amount of current flowing through the moving coil. Thus, for a
given value of R,, the out-of-balance current is a measure of the
prevailing temperature. The coil current produces a surrounding
magnetic field which, on interacting with that of the permanent
magnet, results in rotation of the coil, and movement of its pointer to
a scale position that indicates the temperature.
In this type of system, there is only one point at which the circuit
is balanced, and at which no current will flow through the moving
coil; this is its null point. It is usually denoted on the indicator scale
by a datum mark, against which the pointer registers when the power
supply is disconnected.
A bridge circuit has the disadvantage that the out-of-balance current
also depends on the voltage of the power supply. Hence, errors in
indicated readings can occur if the voltage differs from that for which
the system was initially calibrated.
Ratiometer system A ratiometer system also consists of a resistive
sensor and a moving-coil indicator which, unlike the conventional
type, has two coils moving together in a permanent magnetic field of
non-uniform strength. The coil arrangements and methods of
obtaining the non-uniform field depend on the manufacturer's design,
but three typical methods are shown in Fig. 13.9.
316
Figure /3.9 Ratiometer coil
and field arrangements.
(a) Crossed-coil; (h) parallel
coil; (c) twin former.
POLE
PIECE
MAGNETIC
FIELD
CORE PIECE
I
MAX-
MAGNETIC
FIELD
STRENGTH
t
MIN-
MAGNET
MAX-
MOVING
COIL
INDICATOR
POINTER
(al
(bl
WINDING 'A'
WINDING 'B'
CORE PIECE
(cl
Figure 13. IO shows the circuit in basic form, and from this it will
be noted that two parallel resistance arms are formed, one containing
a coil and a fixed calibrating resistance R 1, and the other containing a
coil in series with a calibrating resistance R2 and the sensor R,. Both
arms are supplied with either pure de or rectified ac, and the coils
are so wound that current flows through them in opposite directions.
As in any moving-coil indicator, rotation of the measuring element,
i.e. coil former and pointer, is produced by forces which are
proportional to the product of the current and field strength, and the
direction of rotation depends on the direction of current relative to
the field. In a ratiometer, therefore, it follows that the force produced
by one coil will always tend to rotate the measuring element in the
opposite direction to the force produced by the second coil, and
317
Figurr 13. ff) Basic ratiometer
A
+
circuit.
B
SENSOR
INDICATOR
UNIT
furthennore, as the magnetic field is of non-unifonn strength, the coil
carrying the greater current will always move towards the area of the
weaker field, and vice versa.
For purposes of explanation, let us assume that the basic circuit of
Fig. 13.10 employs an indicator which has the crossed-coil method
(see (a) of Fig. 13.9), that winding 'B' is in the variable-resistance or
sensor arm, and winding 'A' is in the fixed-resistance arm. The
resistances of the arms are so chosen that at the zef() position of the
indicator scale the forces produced by the currents flowing in each
winding are in balance. Although the currents are unequal at this
point, and indeed at all other points except mid-scale, the balancing
of the torques is always produced by the strength of the field in
which the windings are positioned.
When the temperature of the sensor R, increases, then in
accordance with the temperature/resistance relationship of the material
used for the sensor, its resistance will increase and so cause a
decrease in the current flowing in winding 'B' and a corresponding
decrease in the force created by it. The current ratio is therefore
altered, and the force in winding 'A' will rotate the measuring
element so that both windings are carried round the air gap; winding
'B' is advanced further into the stronger part of the field, while
winding 'A' is being advanced into a weaker part. When the sensor
temperature stabilizes at its new value, the forces produced by both
windings will once again balance, at a new current ratio, and the
angular deflection of the measuring element will be proportional to
the temperature change.
When the measuring element is at the mid-position of its rotation,
the currents in both windings are equal since this is the only position
where the two windings can be in the same field strength
simultaneously.
In a conventional moving-coil indicator, the controlling system is
318
Figure I 3. I I Ratiometer type
of temperature indicating
system.
26 V ac
'Sweep-off
magnet
Centring
pot.
---0--Crank
Cabin
temperature
compensator
Temperature sensor
made up of hairsprings which exert a controlling torque proportional
to the current flowing through the coil. Therefore, if the current
decreases due to a change in the power supply applied to the
indicator, the deflecting torque will be less than the controlling torque
of the springs and so the coil will move back to a position at which
equilibrium between torques is again established. The pointer will
thus indicate a lower reading. A ratiometer, on the other hand, does
not require hairsprings for exerting a controlling torque, since this is
provided solely by the appropriate coil winding and non-uniform field
arrangements. Should variations in the power supply occur they will
affect both coils equally so that the ratio of currents flowing in the
coiis remains the same, and tendencies for them to move to positions
of differing field strength are counterbalanced.
Having noted this point, however, a spring is, in fact, used in
practical applications, but its sole function is to return the measuring
element to -the 'off-scale' position when the power supply is
disconnected. Since it exerts a very much lower torque than a
conventional control spring, its effects on the indicating accuracy of a
ratiometer in response to power supply ehanges are very slight.
Figure 13. l l schematically illustrates the circuit arrangement of a
type of ratiometer used for the measurement of oil temperature in
some series of Boeing 737. It utilizes a twin-coil former system and
319
operates from a 26 V ac supply which is rectified within the
indicator. Two potentiometers are included in the circuit, and are preadjusted, one to set the range (expansion) and the other (centralizing)
to set the mid-scale point at which the currents in the coils should be
equal. A thermoresistor, referred to as a cabin temperature
compensator, is also incl_uded in one arm of the circuit; its purpose is
to compensate for the effects that changes of ambient temperature at
the indicator's location could otherwise have on the current ratios
produced by oil temperature changes.
Another feature of the indicator is the method adopted for returning
the coils and pointer to an off-scale position when electrical power is
disconnected. This is done by a 'sweep-off magnet and a crank
which mechanically moves the measuring element. Adjacent to the
magnet is a coil which is connected in the indicator circuit, so that,
when power is switched on, there is an interaction of magnetic fields
causing the crank to move clear of the measuring element, thus
allowing its normal range of movement.
1henno-emf systems
These systems play an important part in monitoring the structural
integrity of vital components of air-cooled piston engines and turbine
engines when operating at high temperatures. In the former class of
engine the components concerned are the cylinders, while in turbine
engines they- are the turbine rotor discs and blading. In basic form,
the systems consist of a thermocouple type of sensor which,
depending on the application, is secured to an engine cylinder head or
exposed ro exhaust gases, and an indicator connected to the sensor by
special leads .
.,;piermocouple principle Thermo-emf systems depend for their
operation on electrical energy which is produced by the direct
conversion of heat energy at the source of measurement. Thus, unlike
variable resistance systems, they are independent of any external
electrical supply.
The form of energy conversion known as the Seebeck effect is
based on the fact that when two wires made of dissimilar metals are
joined at their ends, so as to form separate junctions as in Fig.
13.12, a thermo-emf is produced, and if the junctions are maintained
at different temperatures, a current is caused to flow around the
circuit. This arrangement is called a thermocouple, the junction at the
measuring
higher temperature being conventionally termed the hot
junction, and that at the lower temperature the cold or reference
junction. In practice, the temperature sensor forms the hot junction,
and since it is a separate unit, it is generally regarded as the
thermocouple proper.
or
320
Figure J3. 12 Thermocouple
principle.
··---·
A
HOT JUNCTION
B
COI.O JUNCTION
---
----
HOTJUNCTJON
COlO JUNCTION
Table 13. l
Group
Thermocouple combinations
Metals and composition
Positive wire
Negative ll'ire
Maximum
temperature
Application
oc
(continuous)
Base metal
Rare metal
Copper (Cu)
Constantan (NI. 40%;
Cu, 60%)
400
Iron (Fe)
Constantan (Ni, 40%;
Cu, 60%)
850
Chomel (Ni, 90%; Cr. 10%)
Alumel (Ni, 90%; Al, 2%
+Si+ Mn)
1,100
Exhaust gas temperature
measurement
Platinum (Pt)
Rhodium-platinum
(Rh, 13%; Pt, 87%)
1,400
Not utilized in aircraft
temperature-indicating
systems
Cylinder head temperature
measurement
Cr. chromium; Ni, nickel; Al, aluminium; Si, silicon; Mn. manganese
.
Thermocouple materials and combinations The materials selected for
use as thermocouple sensors fall into two main groups, base metal
and rare metal, and are listed in Table 13.1. The choice of a
particular thermocouple combination is dictated by the maximum
temperature to be encountered in service.
In order to utilize the thermocouple principle for temperature
measurement, it is obviously necessary to measure the emfs generated
at the various temperatures. Typical values are listed i,n the Tables 4,
5 and 6 given at the end of the book. In the basic form of system,
measurement is accomplished by connecting a moving-coil
millivoltmeter, calibrated in degrees Celsius, in series with the
circuits so that it forms the cold junction. The introduction of the
indicator into the circuit involves the presence of additional junctions
which produce their own emfs and so introduce errors in
321
Figure 13.13 Thermocouple
sensors. (a) Surface contact;
(b) immersion or probe.
(a)
(b)
measurement. However, the effects are taken. into consideration when
designing practical thermocouple circuits, and any .errors resulting
from 'parasitic emfs', as they are called, are eliminated.
Types of thermocouple sensor The sensors employed are in general
of two basic types: (i) surface contact and (ii) immersion. Typical
examples are shown in Fig. 13. 13.
The surface contact type is designed to measure the temperature of
a solid component and is used as the sensor in air-cooled engine
cylinder head temperature indicating systems. The material
combinations used are either copper/constantan or iron/constantan,
and their junction may be in the form of a 'shoe' bolted in good
thermal contact with a cylinder head representative of the highest
temperature condition, or in the form of a washer bolted between a
cylinder head and a spark plug.
The immersion, or probe, type of thermocouple is designed for the
measurement of gas temperatures, and is therefore adopted as the
sensor in turbine engine exhaust gas temperature (EGT) indicating
systems. Since EGT is a parameter very closely associated with
others required for the monitoring and control of engine power.
further details of probes and EGT indicators will be given in
Chapter 15.
Cold junction temperature compensation As we have already
learned, the emf produced by a thermocouple depends upon the
difference between the temperature of the hot and cold junctions. It is
thus apparent that, if the ambient temperature of an indicator should
change while the hot junction temperature remains constant, then, by
virtue of the indicator being the cold junction of a circuit, a change
in emf will result causing the indicated readings to be in error. Since
it is essential for the readings to be representative of the hot junction
only, means must be provided for the automatic detection of cold
junction temperature changes, and compensation of resulting errors.
Before going into the details of detection and compensation, it is
useful to consider first how the changes in emf actually arise.
The various combinations of thermocouple materials specified for
use in aircraft conform to standard temperature/emf relationships, and
the associated indicators are calibra,ed accordingly. The emfs
322
obtained correspond to a cold junction temperature which is usually
maintained at either 0°C or 20°C (see Tables 4, 5 and 6 given at the
end of the book).
Let us assume, for example, that the cold junction is maintained at
0°C and that the hot junction temperature has reached S00°C. At this
temperature difference a standard value of emf generated by a
chromel/alumel combination, say, is 20.64 mV. If now the cold
junction temperature increases to 20°C while the hot junction remains
at 500°C, the temperature difference decreases to 480°C and the emf
equivalent to this difference is now 20.64 mV minus the emf at
20°C; as a standard value this corresponds to 0.79 mV. Thus, the
indicator measuring element will respond to an emf of 19. 85 mV and
move 'down scale' to a reading of 480°C.
A change, therefore, in ambient temperature decreases or increases
the emf generated by a thermocouple and makes an indicator read
high or low by an amount equal to the change of ambient
temperature.
A method commonly adopted for the compensation of these effects
in some types of moving-coil indicator is quite simple and is, in fact,
an adaptation of the bimetallic strip principle as applied to directreading air temperature indicators (see page 61). In this case,
however, a strip of dissimilar metals is coiled in the shape of a flat
spiral spring. It has one end anchored to a bracket which forms part
of the indicator measuring element support, while the other end (free
end) is connected by an anchor tag to one of the controlling
hairsprings to form a fixture for this spring.
When the indicator is on 'op~n circuit', i.e. disconnected from the
thermocouple system, the spring responds to ambient temperature
changes such that the indicator functions as a direct-reading air
temperature indicator. With the thermocouple system connected to
complete the circuit, then if the two junctions are at the temperatures
earlier assumed, namely 0°C and 500°C, the emf will position the
measuring element to read 500°C. If the temperature at the indicator
increases to 20°C, then, as already illustrated, the emf is reduced but
the tendency for the measuring element to move down scale is now
directly opposed by the compensating spring as it unwinds in
response to the 20°C temperature change. The indicator reading
therefore remains at 500°C, the true hot junction temperature.
The compensation of cold junction temperature changes can also be
accomplished electrically, but since this relates particularly to the
application of servo-operated EGT indicating systems, details of the
principle involved will be covered in Chapter 15.
Compensation of moving-coil resistance changes Changes of ambient
temperature can also have an effect on the resistance of the moving
coil of an indicator, and so this too must be compensated. This can
323
be accomplished by incorporating either a thermoresistor or a
thermomagnetic shunt within the indicator.
A thermoresistor, or thermistor as it is generally known, has a
large temperature coefficient which is usually negative. If, therefore,
it is connected in a moving-coil circuit, it will, under the same
temperature conditions, oppose coil resistance changes and maintain a
constant current appropriate only to that produced by the
thermocouple hot junction.
A thermomagnetic shunt is a strip of nickel-iron alloy which is
sensitive to temperature changes, and is clamped across the poles of
the permanent m.,gnet of a moving-coil type of indicator, so that it
diverts some of the airgap magnetic flux through itself. For example,
if the ambient temperature of an indicator increases, the reluctance of
the alloy strip will also increase so that less flux is diverted, or
'shunted', from the airgap. Since the deflecting torque exerted on a
moving coil is proportional to the product of current and flux, the
increased airgap flux counterbalances the reduction in coil current
resulting from the increase in ambient temperature; thus, a constant
torque and indicated hot junction temperature reading is maintained.
External circuit and resistance The external circuit of a thermo-emf
indicating system consists of the thermocouple and its leads, and the
leads from the junction box at an engine bulkhead, or 'firewall', to
the indicator terminals. From this point of view, it might therefore be
considered as a simple and straightforward electrical instrument
system. However, whereas the latter may be connected up by means
of the cables normally used in an aircraft, it is not acceptable to do
so in a thermo-emf system.
This may be explained by taking the case of a copper/constantan
thermocouple which is to he connected to a cylinder head temperature
indicator. If a length of normal copper twin-core cable is .connected
to the thermocouple junction box, then one copper lead will be joined
to its thermocouple counterpart, but the other one will be joined to
the constantan lead of the thermocouple. It is thus apparent that this
introduces another effective hot junction which will respond to
temperature changes occurring at the junction box, and so cause an
unbalance in the temperature/emf relationship and errors in hot
junction temperature readings. Similarly, all terminal connections
which may be necessary for routeing the leads through an aircraft,
and connections at the indicator itself, will create additional hot
junctions and so further aggravate indicator errors.
In order to eliminate these hot junctions, it is the practice m use
leads made of the same materials as the thermocouples themselves;
such leads are generally known as extension leads. It may sometimes
be the practice to use leads made of materials having similar thermoemf characteristics in combination; for example, a chromel/alumel
324
thermocouple may be joined to its indicator by copper/constantan
leads, the· latter then being referred to as compensating leads.
Additional hot junctions at indicators and junction boxes are
eliminated by making the terminals of the same materials as those of
the appropriate thermocouple combination, i.e. a positive terminal of
copper or chrome!, and a negative terminal of constantan or alumel.
Another important factor in connection with the external circuit is
its resistance, which must be kept not only low but also constant for
a particular installation. Indicators are, therefore, normally calibrated
for use with external circuits of specific resistance values, e.g. 8 n or
25 n, and they are identified accordingly. Therfl}ocouples, their leads,
and harnesses where appropriate, are made up in fixed low-resistance
lengths. Similarly, extension or compensating leads are also made up
in lengths and of uniform resistance to suit the varying distances
between thermocouple hot junction and indicator locations.
Adjustment of the total external circuit resistance following
replacement of thermocouples, leads or indicators is provjded by a
trimming resistance circuit connected in series with the extension or
compensating leads. Although the circuit introduces additional
junctions, the temperature coefficients of resistance of the materials
used for the resistors (typically Manganin or Eureka) ensure that the
circuit has negligible thermo-emf effect.
325
14 Fuel quantity indicating
systems
An indication of the quantity of fuel in the tanks of an aircraft's fuel
system is, of course, an essential requirement, and, in conjunction
with measurements of the rate at which the fuel flows to the engine
or engines, permits an aircraft to be flown at maximum efficiency
compatible with its specified operating conditions. Furthermore, a
knowledge of both parameters enables not only assessments of
remaining flight time to be made, but also comparisons between
present engine performance and past or calculated performance.
The method most commonly adopted for quantity indication is one
which measures changes in fuel level in terms of the changes in
electrical capacitance of special tank units or 'probes', which then
transmit corresponding signals to their associated indicators.
Capacitance-type
system
In its basic form, a system consists of a variable capacitor located in
a fuel tank, an amplifier and an indicator. The complete circuit forms
an electrical bridge which is continuously being rebalanced as a result
of differences between the capacitances of the tank capacitor and a
reference capacitor. The signal produced is amplified to operate a
motor, which positions a pointer to indicate the capacitance change of
the tank capacitor and, thus, the change in fuel level or quantity.
Before going into the· operating details of more practical systems,
however, it is first necessary to have some understanding of the
fundamental principles of capacitance and its effects in electrical
circuits.
Electrical capacitance
Whenever a potential difference is applied across two conducting
surfaces separated by a non-conducting medium, called a dielectric,
they have the property of storing an electric charge; this property is
known as capacitance. The device comprising the surfaces, or plates,
and a dielectric is called a capacitor
The flow of a momentary current into a capacitor establishes a
potential difference (pd) across its plates. Since the dielectr;c contains
no free electrons the current cannot flow through it, but the pd sets
Figure 14. I
principle.
-
Capacitor
o_
2_ _ _ _
I
I
,._ __ _
I
......
t
ll l l
A
I
I
I
+ +
t,t
B
I
I
I
I
I
I
I
'
I
I
It
I
~\
I
If
-------
~
f ..1 ----+
- - CHARGING CURRENT
- - - - DISCHARGING CURRENT
up a state of stress in the atoms comprising it. For example, in the
circuit shown in Fig. 14. l, when the switch is placed in position l .a
rush of electrons, known as the charging current, takes place from
plate A through the battery to plate B and ceases when the pd
between the plates is equal to that of the battery.
When the switch is opened, the plates remain positively and
negatively charged since the atoms of plate A have lost electrons
while those at plate B have a surplus. Thus, electrical energy is
stored in the electric field between the plates.
Placing the switch in position 2 causes the plates· to· be shortcircuited and the surplus electrons at plate B rush back to plate A
until the atoms of both plates are electrically neutral and no potential
difference exists between them. This discharging current is in the
reverse direction to the charging current, as shown in Fig. 14. l.
Units of capacitance
The capacitance or 'electron-holding ability' of a capacitor is the ratio
between the charge and pd between the plates and is expressed in
farads, one farad representing the ability of a capacitor to hold a
charge of one coulomb (6.24 x 10 18 electrons) which raises the pd
between its plates by one volt.
Since the farad is generally too large for practical work, a submultiple of it is normally used called the microfarad (1 µF = 10- 6F).
In the application of the capacitor principle to a fuel quantity
indicating system, an even smaller unit, the picofarad ( 1 pF
w- 12 F), is the standard unit of measurement.
Factors on which capacitance depends
Capacitance depends on the area a of the plates, the distanced
between the plates, and the capacitance (or absolute permittivity)
a unit cube of the dielectric material between the plates.
E
of
327
Permittivity E is usually quoted as being relative to that of a
vacuum. Relative permittivity, also called dielectric constant, is often
denoted by K which is the ratio of the capacitance of a capacitor with
a given dielectric to its capacitance with air between its plates. The
relative permittivitics of some pertinent substances are as follows:
Air
Water
Water vapour
Aviation gasolene
Aviation kerosene
l.OOu
81.07
1.007
1.95
2.10
Capacitors in alternating current circuits
As already mentioned, when direct current is applied to a capacitor
there is, apart from the initial' charging current, no current flow
through the capacitor. In applying the capacitance principle to fuel
quantity indicating systems, however, a flow of current is necessary
to make the indicator respond to the capacitance changes arising from
changes in fuel level. This is accomplished by supplying the tank
probes with an alternating voltage, because whenever such voltage
across a capacitor changes, electrons flow toward and away from it
without crossing the plates and a resultant current flows which, at
any instant, depends on the rate of change of voltage.
Basic indicating system
For fuel quantity measurement, the capacitors to be installed in the
tanks must, of course, differ in construction from those normally
adopted in electronic equipment. The plates therefore take ~he form
of two tubes mounted concentrically with a narrow airspace between
them, and extending the full depth of a fuel tank. Constructed in this
manner, two of the factors on which capacitance depends are fixed,
while the third factor, relative permittivity, is variable since the
medium between the tubes is made up of fuel and air. The manner in
which changes in capacitance take place is illustrated in Fig. 14.2,
and is described in the following paragraphs.
Figure I 4. i Capacitance
changes due to fuel and air.
-=
-=
I
': .': ---11H
-
- ------ --- ------------
,--
(a)
328
£A.•100pF
(b).K-2·1
(c)
..!:. - .l
..!:!.
2
I
At (a) a capacitance probe is fitted in an empty fuel tank and its
capacitance in air is 100 pF, represented by CAAt (b) the tank is filled with a fuel aving a K value of 2.1, so that
the capacitor is completely immersed and its sensing surfaces are
fully 'wetted'. As noted earlier, K is equal to he ratio of capacitance
using a given dielectric (in this case Cr) to that using air; therefore
[lJ
From equation [ lJ Cr = CA K, and it is thus clear that the
capacitance of the tank probe at (bJ is equal to 100 x 2.1, i.e.
210 pF. The in~rease of 110 pF is the added capacitance due to the
fuel and may be represented by rF. The tank probe may therefore be
represented electrically by two capacitors in parallel, and of a total
capacitance,
[2]
In Figure I4.2(c), the tank is only half full and so the total
capacitance is 100 + 55, or 155 pF. The added capacitance due to
fuel is determined by transposing equation [2], so that CF = CT- CA,
and by substituting CA K for CT, we thus obtain CF = CA K - CA,
which may be simplified as:
[31
the factor (K- 1) being the increase in the K value over that of air.
The fraction of the total possible fuel quantity in a 'linear tank' at
any given level is given by L/H, where L is the height of the fuel
level and H the total height of the tank. Thus by adding L/H to
equation [3] the complete formula becomes:
CF=.!::.._ (K-1) CA
H
[41
The circuit of a basic system is shown in Fig. 14.3. It is divided
into two sections or loops by a resistance R, both loops being
connected to the secondary winding of a power transformer. Loop A
contains the tank probe Cr and may therefore be considered as the
sensing loop of the bridge since it detects current changes due to
changes in capacitance. Sensing loop voltage Vs remains constant.
Loop B, which may be considered as the balancing loop of the
bridge, contains a reference capacitor CR of fixed value, and is
connected to the transformer via the wiper of a balance potentiometer
so that voltage V8 is variable.
The balance potentiometer is contained within the indicator together
with a two-phase motor which drives the potentiometer wiper and
indicator pointer. The reference phase of the motor is continuously
energized by the power transformer, and the control phase is only
energized when an unbalanced condition exists in the bridge.
329
Figure 14. 3 Basic capacitance
system.
The amplifier has two main stages: one for amplifying the signal
produced by bridge unbalance, and the other for discriminating the
phase of the signal which is then supplied to the motor.
Let us consider the operation of the complete circuit when fuel is
being drawn off from a full tank. Initially, and at the constant fulltank level, the sensing current ls is equal to the balancing current 18 ;
the bridge is thus in balance and no signal is produced across R.
As the fuel level drops, the tank probe has less fuel around it;
therefore the added capacitance (CF) has decreased. The capacitance.
of the probe decreases and so does the sensing current.Is, the latter
creating an unbalanced bridge condition with balancing current 18
predominating through R.
A signal voltage proportional to laR is developed across R and is
amplified and its phase detected before being applied to the control
phase of the indicator motor. The discriminator output signal is
converted from a half-wave pulse into a full-wave signal, and
supplied to the control phase winding of the motor. As in normal
two-phase motor operation, the control phase current will either lead
or lag that in the reference phase. Thus, in the condition we are
considering, the control phase current is lagging since balancing loop
current 18 is the predominating one. The motor and balance
potentiometer wiper are therefore driven in such a direction as to
reduce the current 18 . When this current equals ls, the bridge is once
again in balance, the motor ceases to rotate, and the indicator pointer
registers the new lower value of fuel quantity.
330
Figure 14.4 Temperature
+5 .--....,,.-,----,---,----,,
effects on fuel characteristics.
0
+55
TEMPERAl1JRE. 'C
Effects of fuel temperature changes
With changes in temperature the volume, density and relative
permittivity of fuels are affected to approximately the same degree as
shown in Fig. 14.4, which is a graph of the approximate changes
occurring in a given mass of fuel. From this it should be noted that
K - 1 is plotted, since for a system measuring fuel quantity by
volume, the indicator pointer movement is directly dependent on this.
It should also be noted that, although it varies in the same way as
density, the percentage change is greater.
Thus a volumetric system will be subject to a small error due to
variations in fuel temperature. Furthermore, changes in Kand density
also occur in different types of fuel having the same temperature. For
example, a system which is calibrated for a K value of 2.1 has a
calibration factor of 2.1 -1 = 1.1. If the same system is used for
measuring a quantity of fuel having a K value of 2.3, then the
calibration factor will have increased to 1.3 and the error in
indication will be approximately:
QI. I
X
100-100%
=
18%
,
i.e. the indicator would overreact by 18 per cent.
Measurement of fuel quantity by weight
A more useful and accurate method of measuring fuel quantity is to
do so in terms of its mass or weight. This is because the total power
developed by an engine, or the work_ it performs during flight,
depends not only on the volume of fuel but 'On the energy it contains,
i.e. the number of molecuies tmit can combine with oxygen in the
engine. Since each fuel molecule :has some weight, and also because
one pound ~r fuel has the 'Same number of molecules regardless of
temperature and therefore volume, the total number of molecules
(total available energy) is best indicated by measuring the total fuel
weight.
In order to do this, the volume and density of the fuel must be
known, and the product of the two determined. The measuring device
must, therefore, be sensitive to changes in both volume and density
so as to eliminate undesirable effects due to temperature.
This will be apparent by considering the example of a tank holding
1000 gallons of fuel having a density of 6 lb/gal at normal
temperature. Measuring this volumetrically we should of course
obtain a reading of 1000 gallons, and from a mass measurement,
6000 lb. If a temperature rise should increase the volume by IO per
cent, then the volumetric measurement would increase to 1100
gallons, but the mass measurement would remain at 6000 lb because
the density of the fuel (weight/volume) would have decreased when
the temperature increased.
For the calibration of systems in terms of mass of fuel, the
assumption is made that the relationship between the relative
permittivity (K) and the density (p) of a given sample of fuel is
constant. This relationship is called the capacitive index and is
defined by
K-1
p
An indicating system calibrated to this expression is still subject to
indication errors, but they are very much reduced. This may be
illustrated by a second example. Assuming that the system is
measuring the quantity of a fuel of nominal K = 2.1 and of nominal
density p = 0.779, then its capacitive index is:
2.1-1
0.779
=
1.412
If now the same system measures the quantity of another fuel for
which the nominal Kand p values are respectively 2.3 and 0.85, then
its capacitive index will increase; thus,
2.3 - I
0.85
= 1.529
However, the percentage error is now
529
1.
1.412
X
100-100%
=
8%
and this is the amount by which the indicator would overread.
Compensated indicating systems
Although the assumed constant. permittivity and density relationship
results in a reduction of the indicator error, tests on the properties, of
f'igure I 4. 5
Compensator
TANK UNIT
capacitor.
C.. __j
~$FOAMER {
TAPPINGS
...,.__ ___,vvv,,vv,.~--........j
CR
TO BA1.ANCE
_,,_ _ _-+fl-Uf--t-----'
POT£NTIOMmR
L,
l
BAl.ANCE
LOOP
CcoMP
(BOTTOM OF TANK UNIT)
fuels showed that, while the capacitive index could vary from one
fuel to another, the variation tended to follow the permittivity. Thus,
if an indicating system can also detect changes in the permittivity of a
fuel as it departs from its nominal value, then the density may be
inferred to a greater accuracy, resulting in an even greater reduction
of indication errors. Indicating systems currently in use are therefore
of the permittivity or inferred-density compensated type, the
compensation being effected by a reference capacitor added to the
balance loop of the measuring circuit and in parallel with the
capacitor CR, as shown in Fig. 14.5.
A compensator is similar in construction to a standard tank probe.
In some systems it is fitted to the bottom end of a standard probe so
that it is always immersed in fuel, while in others it forms a separate
probe. Located in this manner, its capacitance is determined solely by
the permittivity, or K value, of the fuel, and not by its quantity as in
the case of the standard tank probe. In addition to variable voltage,
the balancing loop will also be subjected to variable capacitance,
which means that balancing current 18 will be affected by variations
in K as well as sensing-loop current !5 •
Let us assume that the bridge circuit (see Fig. 14.3) is in balance
and that a change in temperature of the fuel causes its K value to
increase. The tank probe capacitance will increase and so current ls
will predominate to unbalance the bridge and to send a signal to the
amplifier and control phase of the motor. This signal will be of such
a value and phase that an increase in balancing-loop current is
required to balance the bridge; and so the motor must drive the wiper
of the balance potentiometer to decrease the resistance. Since this is
in the direction of the 'iank full' condition, the indicator will
obviously register an increase in fuel quantity. The increase in K,
however, also increases the compensator capacitance so that loop
current 18 is increased simultaneously with, but in opeosition to, the
increasing sensing-loop current fs. A balanced bridge condition is
therefore obtained which is independent of the balance potentiometer.
In practice, there is stiH an indication error due to the fact that the
density also varies with temperature, and this is not directly
measured. But the percentage increase of density is not as great as
that of K - 1, and so by careful selection of the compensator
333
capacitance values in conjunction with the reference capacitor, the
greatest reduction in overall indication error is produced.
Densitometers Although the application of compensator probes
provides for greater accuracy of measurement in terms of mass
measurement, the fact that this is based on an assumed relationship
between the K and p values of fuel results in indications that still fall
short of those desired, i.e. true weight of fuel. Such indications can,
however, be achieved by measuring variations in p separately, and
then applying the corresponding values as continuous correction
signals to the measuring and indicating circuits. This is a method
which is applied to fuel quantity indicating systems installed in
several types of current-generation aircraft, and from the point of
view of techniques involved, the one adopted in such aircraft as the
Boeing 757, 767 and 747-400 serves as an interesting example.
In addition to the normal probes and the compensator probes, a
device known as a densitometer is installed in each fuel tank. It is
comprised of two principal units: {i) an emitter which contains a
collimator, a radioactive capsule, and a microswitch, and (ii) an
electronics unit containing two detector tubes filled with xenon gas,
pre-amplifier, signal processor and power supply circuit card
modules. The complete unit is installed on the rear spar of an integral
tank such that the emitter is on the inside of the tank. The principle
of its operation is shown in Fig. 14.6.
Figure 14.6 Densitometer.
DRY SIDE
WET S!OE
TO
PROCESSOR
I
RADIOACTIVE CAPSULE
COLLIMATOR
I
ELECT!".ONICS
UNIT
1420 V de
The gamma radiation from the emitter passes through the
collimator and the fuel, the density of which will vary the level of
radiation and the paths taken from the collimator to the detector
tubes. The collimator directs the radiation to the tubes in a conical
pattern. One of the tubes detects radiation from what is termed a
short fuel path, while the other detects long fuel path radiation ~ that
fuel density is determined as a ratio of radiation.
Each tube is made up of a cathode and an anode which passes
through the centre, and is supplied with 1420 V de generated by the
power supply module of the densitometer electronics unit. As noted
earlier, the tubes are filled with xenon gas, which on being exposed
to gamma radiation causes electrons to be released. This, in turn,
generates a pulsed signal which, after amplification and processing, is
transmitted from the electronics unit to the processor unit that
controls the whole operation of the fuel quantity indicating system
adopted for the types of aircraft already mentioned. The number of
pulses per unit of tifi!e are counted, and are a function of fuel
density. In order to obtain the true fuel weight for display on the
quantity indicators, the system's processor unit then multiplies the
fuel volume values measured by the tank probes in the normal
manner, by the fuel density values received from the densitometers.
As will be noted from Fig. 14.6, the collimator, radioactive capsule
and microswitch are in a 'dry' and sealed section of a densitometer.
In the event that fuel should leak into this section, it causes a piece
of foam material, located under the switch, to 'swell' and so exert a
pressure on the switch plunger. The resulting actuation of the switch
contacts causes the densitometer signals to be transmitted to, and
stored in, a fault memory circuit within a built-in test section of the
quantity indicating system's processor unit.
Construction of probes
In the majority of applications, probes utilize tubes made of
aluminium suitably protected against the effects of corrosion, shortcircuiting and grounding. In some cases, the outer tube may be of
aluminium, while the inner tube is a non-conducting plastic material
that is coated with a metallized film on its outer surface to serve as a
capacitor plate.
The tubes are held apart concentrically by insulated cross-pins, and
for the purpose of initial calibration and characterization (see page
339) to suit individual tank shapes and sizes, the concentricity is
varied throughout the length of a probe. This is accomplished by
having an inner tube whose diameter varies along its length, or by
off-centering it by means of spacers which differ in length. The
spacers are made of teflon, and are colour-coded corresponding to
length. In some types of probe, characterization is effected by
Figure 14.7 Tank probes.
{a) Standard probe: I rubber
ring. 2 nylon sleeve. 3 outer
tube. 4 inner tube, 5 rubber
ring, 6 nylon sleeve,
7 insulating cross-pin,
8 bracket, 9 miniature coaxial
connector; (b) probe with
compensator: I rubber ring,
2 nylon sleeve, 3 outer tube,
4 insulating cross-pin, 5 rubber
ring, 6 inner tube, 7 nylon
sleeves, 8 reference unit,
9 bracket. lO miniature coaxial
connector.
varying the area of an inner tube's conducting surface at various
points over its length.
Mounting arrangements vary depending on whether the probes are
to be secured to the top and bottom sections of a tank, as in the case
of the probes shown as examples in Fig. 14. 7, or are to be secured
to tank baffles, ribs or spars; in the latter case, mounting brackets are
provided. Probes are connected to the indicating system wiring
harnesses via terminal blocks which, depending on design, provide
for either co-axial or screw-type terminal connections.
The probe illustrated in Fig. 14.8 is of the independent
compensator type containing three tubes. The outer and inner tubes
are of low impedance, and the middle one is of high impedance. A
'tuning' plate is provided for calibration of the probe, and also for
controlling what is termed capacitance fringing. Since the probe is
always immersed in fuel, there is a possibility of contaminant buildup which would cause the probe to sense a stagnant sample of fuel.
To prevent this, therefore, a tube is provided so that during refuelling
operations, a stream of fuel is directed up through the bottom of the
probe, thereby washing it out.
Location and connection of tank probes
In practical indicating systems, a number of tank probes are
positioned within the fuel tanks·as illustrated in Fig. 14.9, and are
Figure 14.8 Compensator
probe.
TERMINAL BLOCK
ASSEMBLY
TUNING PLATE
TUBES OR ELECTRODES
WASH TUBE
(FROM FUELLING MANIFOLD)
connected in parallel to their respective indicators. The reasons for
this are to ensure that indications remain the same regardless of the
attitude of an aircraft and its tanks, to take into account the effects of
wing flexing and, of course, to give a high total capacitance value.
This may be understood by considering a two-unit system as shown
in Fig. 14.10. If the tank is half-full and in a level attitude, each
probe will have a capacitance of half its maximum value; since they
are in parallel the total capacitance measured will produce a 'halffull' indication. When the tank is tilted, and because the fuel level
remains the same, probe A is immersed -deeper in the fuel by the
distance d and gains some capacitance, tending to make the indicator
overread. Probe B, however, has moved out of the fuel by the same
distance d and loses an equal amount of capacitance. The total
337
Figure 14. 9 Location of tank
probes.
-
PROBES & HARNESS
@
COMPENSATOR PROBES
i:jt'l"' PLUG CONNECTIONS
Figure 14.10 Attitude
compensation.
capacitance therefore remains the same as for the level-tank attitude
and the indication is .unchanged.
The number of probes required is governed by the fuel capacity
_i:equirements and tank configurations of any one type of aircraft. In
the one on which Fig. 14.9 is based, there are left and right wing
main tanks each containing 14 probes, and left and right centre
auxiliary tanks each containing four probes. In addition, one
compensator probe is installed in each of the tanks.
In the majotity of systems, the probe terminals and the wiring
harnesses interconnecting them are located within the tanks. The
harnesses terminate at single connector boxes, or 'bussing' plugs,
which are externally mounted and provide for the connection of
signal-processing and indicator units.
Characterization of probes
The fuel tanks of an aircraft may be separate units designed for
installation in wings and centre sections, or they may form an
integral sealed section of these parts of the structure. This means,
therefore, that tanks must vary in contour to suit their chosen
locations, with the result that the fuel level is established from
varying datum points.
Figure 14.11 represents the contour of a tank located in an
aircraft's wing and, as will be noted, the levels of fuel from points
A, B and C are not the same. When probes are positioned at these
points the total capacitance will be the sum of three different values
due to the fuel (Cp), and as the probes produce the same change of
capacitance for each inch of 'wetted' length, the indicator scale will
be non-linear corresponding to the non-linear characteristic or' the
tank contour.
The non-linear variations in fuel level are unavoidable, but the
effects on the graduation spacing of the mdicator scale can be
overcome by designing probes so that they can be calibrated. or
characterized; to measure capacitance changes proportional to tank
contour. Some examples of characterization methods have already
been described (see page 335).
In addition to tank contouring, account must also be taken of the
effects of wing flexing that occur in flight. These effects are normally
simulated in full-scale static fuel gauging test rigs, designed for the
initial calibration of indicating systems appropriate to the aircraft
type.
In indicating systems utilizing digital signal-processing techniques,
the need for varying the concentricity of probe inner tubes for the
purpose of characterization can be eliminated by incorporating a
software program in the processor circuit. The program takes into
account all known variables and provides more accurate calculation of
,Figure I 4. JI Characterization
of probes.
A
B
C
339
tank volumetric changes resulting from wing bending both on the
ground and in the air. The program may be selected to provide four
different ground characterizations, and three different airborne
characterizations.
'Empty' and 'Full' position adjustments
In systems based on the circuit arrangement shown in Fig. 14.3, it is
necessary during calibration to balance the current and voltage of the
sensing and balancing loops at the datum corresponding to emptyand full-tank conditions. This is achieved by connecting two
potentiometers into the circuit as shown in Fig. 14.12; they are
adjusted from the rear of the indicators.
The 'empty' potentiometer is connected at each end to the supply
transformer and its wiper is connected to the tank units via a balance
capacitor. When a tank is empty, due to the empty capacitance
of the probes current will still flow through them. The balance
potentiometer wiper will also be at its 'empty' position, but since it is
grounded at this point, no current will flow through the reference
capacitor. However, current does flow through the balance capacitor
and it is the function of the 'empty' potentiometer to balance this out.
The balancing signal from the potentiometer is fed to the amplifier,
the output signal of which drives the indicator servomotor and pointer
to the empty position of the scale.
The 'full' adjustment may be regarded as a means of changing the
position of the point on the balance potentiometer at which the
balance voltage for any given amount of fuel is found, and also of
determining the voltage drop across the potentiometer. Reference to
Fig. 14.12 shows that, if the 'full' potentiometer wiper is set at the
bottom, the maximum transformer secondary voltage will be applied
to the balance potentiometer. Therefore, the distance the
potentiometer wiper needs to move to develop a given balance
voltage can be varied.
Figure 14./2 'Emp1y' and
Full' adjus1ments.
REFERENCE
CAPAC!TOO
~->-----.ID-I-----'
!IA.l.ANCE
POTENTIOMETER
Electronic displays
These are adopted in systems which process fuel weight values in
digital signal format, and two examples based on the Boeing 767
system are illustrated in Fig. 14.13. The primary indication of weight
is given by the LCD indicators (diagram (a)), while data related to
system operation, e.g. fault occurrences, low fuel state, etc., are
displayed as advisory or cautionary messages on a CRT display unit
as in diagram (b). The display unit in this case is associated with an
engine indicating and control system, details of which. will be covered
in Chapter 15.
Fail-safe and test circuits
These circuits are incorporated in all indicating systems, and their
arrangements can vary dependent on the type of system. In those
which operate on the basic principle of bridge balancing, a fail-safe
circuit consists of a capacitor and a resistor, connected in the
indicator motor circuit so that the reference winding supply is
paralleled to the control winding. A small leading current always
flows through the parallel circuit, but under normal operating
conditions of the system, it is suppressed by tank probe signals
flowing from the amplifier through the motor control winding. If a
failure of signal flow occurs, the current in the parallel circuit
predominates and flows through the control winding to drive the
indicator pointer slowly downwards. to the 'empty' position.
A test circuit for this type of system incorporates a switch mounted
adjacent to its appropriate indicator. When the switch is held in the
'test' position, a signal simulating an emptying tank condition is
introduced into the indicator motor control winding, causing it to
drive the pointer towards zero if the circuit is functioning correctly.
When the switch is released, the pointer should return to its original
indication. In multi-indicator installations, a single switch serves to
test all indicators simultaneously.
In computer-controlled systems. circuits are provided within the
signal processor unit which automatically carry out self-test routines
at the time when power is initially applied to a system, and then on a
continuous basis. Any faults that occur are stored in memory circuits,
and since these can only be recalled when carrying out built-in-test
(BIT) checks, an appropriate maintenance message such as 'FUEL
QTY BITE' is displayed on the control and monitoring system
display unit. Faults are assigned specific status code numbers which
are displayed by a two-digit LED display on the processor unit test
panel when the 'BIT' switch is operated to recall faults in the order
they were stored in memory. When faults have been rectified by the
appropriate maintenance actions, the memory is cleared of stored data
by the operation of a 'reset' switch.
.. .. ..
-
Figure 14. I3 Electronic
displays. (a) LCD indicators;
(b) EICAS display units.
L
R
C
FUEL QTY
rilll
X 1000
TOTAL QTY
FUEL TEMP°C
(a)
LOW FUEL
FUEL CONFIG
Upper
FUEL QTY BITE
Lower
(b)
The circuits and software programming are such that adequate
redundancy is provided to ensure the display of valid quantity
indications in the event of failures of channel power supplies, tank
probes, or densitometers:···
Since the quantity indicators are of the LCD type, functional testing
is simply a matter of checking that each of the seven segments that
make up each digit is activated by operating a test switch on a flight
deck panel. Thus, all the digits will be displayed as '8's. A test
switch is also provided on the processor unit test panel for carrying
out similar checks on the fault status code LED display.
Totalizer indicators
These indicators are provided in some types of aircraft in addition to
the normal fuel quantity indicating system. The dial presentations of
two typical indicators are shown in Fig. 14. 14, and from this it will
be noted that in addition to the total quantity of fuel remaining, an
aircraft's gross weight is also displayed. Both indicators operate on
the same principle, the major difference between them being in the
method of display. The indicator at (a) utilizes mechanical-type digital
counters, while that at (b) utilizes a segmented electronic display.
Each indicator comprises a resistance network with the same
number of channels as there are primary fuel quantity indicators
appropriate to the particular aircraft system. Each channel receives a
signal voltage from a primary indicator, and the resistance network
apportions current according to the amount of fuel that each voltage
represents, and then sums the currents to represent total fuel weight
remaining. Thus, any change in the signal voltage from one or more
of the primary indicators causes a proportional change in 'total fuel
quantity current' and an unbalancing of circuit conditions which
drives the counter to the corresponding value.
The gross weight counter, which is pre-set to indicate an aircraft's
Figure 14. /4 Totalizer
indicators.
(a)
weight prior to flight, is also connected to the primary indicator
system, but in such a manner that as fuel is consumed the counter
continuously indicates a decreasing gross weight.
Refuelling and load
control
In the larger types of aircraft, the location of the primary fuel
quantity indicators is, of course, far removed from the tank refuelling
points, and so in order that ground handling crews can monitor more
precisely the necessary 'fuel-uplift' operations, refuelling and load
control panels are located adjacent to the refuelling points.
Figure 14.15 illustrates a panel which is used in conjunction with a
volumetric top-off (VTO) system, i.e. one which automatically
terminates refuelling at a pre-set fuel level. The indicators are
connected into the same sensing circuits as the primary indicators,
and so they provide a duplicate indication of the fuel quantity during
refuelling. The switches above the indicators control the positions of
their respective tank refuelling valves. The indicator lights, which are
normally blue, illuminate when the valves are in the open position.
If only a partial load of fuel is required, the indicators are
monitored and when the desired quantity has been uplifted, the
refuelling valves are closed by selecting this position of the
appropriate switch. When, however, a full load of fuel is required,
the valves are automatically closed by the VTO system.
A separate system is installed for each fuel tank, comprising a
VTO unit and a compensator type of tank probe. Since the system is
shared with the primary quantity indicating system, then the latter
records the increasing fuel quantity in the normal units of mass, and
transmits the corresponding signals to a bridge circuit in the VTO
unit. The signals from the VTO compensator probes relate to tank
Figure 14. I 5 Refuelling panel.
TANKN0.2
-
OFF¢-))
AUX. FUELLING
POWER CONTROL
G'
4
,
OPEN
CLOSED
CENTRE TANK
...,,::-.OPEN
11'r.J1
{}'CLOSED
TANK NO. 1
VAL\/!:
CONTROL
SWITCHES
Figure 14.16 Refuelling and
load select panel.
LOAD SELEC I
@@
,ur
~
f>i:SP
0)(kflL!,.
::=:=:f/
unit capacity in volumetric units, and these are also supplied to the
bridge circuit which then compares both signals. The resultant signal
is amplified and fed through a solid-state switching circuit to energize
the fuelling shut-off valve when the pre-set nominal full level, or
volume, of the tank is reached. Calibration of the whole VTO system
to the desired nominal 'tank full' volume is accomplished by
adjusting the current through the VTO compensator probes so that it
corresponds with the total current from the tank probes at the
particular shut-off level.
On completion of a refuelling operation, the control panel is
enclosed by a door which also actuates a switch that isolates all
electrical power to the fuelling system.
Figure 14.16 illustrates a refuelling control panel which is used in
conjunction with a computer-controlled quantity indicating system. It
operates in a similar manner to the one already described but. as will
be noted, it also provides for the pre-selection of a desired fuel
weight in any tank. The quantity indicators are also of the LCD type.
The desired weight to be loaded is set by means of three
thumbwheel rotary switches. each having ten pt,;,mvns and numbered
0 to 9; thus, the ranges are covered in tenths, units and tens. At each
numbered position a resistor of differing value is selected so that with
power on, and signal currents corresponding to the selected position,
multiples are produced and supplied to a· load select multiplexer in
the processor unit of the main quantity indicating system, A 'SET'
switch is provided on the refuellinr control panel below each of its
quantity indicators, and when pushed in it allows the signals from the
345
processor unit to pass to the lower LCD indicators, which then
display the load selected. When this load has been reached, the
fuelling valves are automatically closed.
The segments of the six LCD indicators can be checked
simultaneously by a push-button 'TEST' switch on the control panel.
15 Engine power and control
Instruments
The power of an engine refers to the amount of thrust available for
propulsion, and depending upon the type of engine, i.e. piston or
turbine, it is expressed as power ratings in units of brake or shaft
horsepower, or of thrust developed at an exhaust unit. Power ratings
of each of the various types of engine are determined, together with
operational limitations, during the test-bed calibration runs conducted
by the manufacturers for various operating conditions such as takeoff, climb, normal cruise, emergency and contingency.
The parameters associated with engine power ratings, and which
must be monitored by appropriate instrumentation, are listed in Table
15.1.
RPM measurement
The measurement of engine speed in terms of revolutions per minute
is relevant to the three main types of engine and, with the exception
of a simple type of unsupercharged piston engine, it is related to the
other parameters involved in power control. The power of an
unsupercharged engine is directly related to its speed, and so with the
throttle at a corresponding operational setting, an rpm indicator
system can also serve as a power indicator.
Indicating systems, also generally referred to as tachometers, are of
the electrical type and fall into two main categories: (i) generator and
indicator, and (ii) tac ho probe and indicator.
Table 15. I Power rating parameter,
Paramerer
Typt of EnHlnt
Plswn
Rpm
Manifold pressure
Torque
Exhaust gas temperature
Pressure ratio
Fuel now
Vnsuptrcharged
Suptrchar//ed
X
X
X
X
X
Turblnt
Turboprop
X
X
X
X
X
X
X
347
Generator and indicator system
A generator is of the ac type, consisting of a permanent magnet rotor
rotating within a slotted stator which carries a star-connected threephase winding. The rotor may be of either two-pole or four-pole
construction, and is driven by a splined shaft coupling; the generator
is bolted directly to a mounting pad at the appropriate accessories
drive gear outlet of an engine. In order to limit the mechanical loads
on generators, the operating speed of rotors is reduced by means of
either 4: I or 2:1 ratio gears in the engine drive system.
Two- or four-pole-type generators are utilized in conjunction with a
three-phase synchronous motor within the indicator. For the operation
of servo-operated indicators, the display of data by electronic CRT
indicators, and for supplying signals to automatic power control
systems, the appropriate data signals are supplied from generators
having 12-pole rotors; these produce a single-phase output at a much
higher frequency and sensitivity.
A typical indicator consists of two interconnected elements, a
driving element and a speed-indicating element. The driving element
is a synchronous motor having a star-connected three-phase stator
winding, and a rotor which is so constructed that the motor has the
self-starting and high torque characteristics of a squirrel-cage motor,
combined with self-synchronous properties of a permanent magnet
type of motor.
The speed-indicating element consists of a permanent magnet device
which operates on the eddy-current drag principle, and as indicated in
Fig. 15.1 it may utilize either a drag cup or a drag disc. In the
former version, the magnet is inserted into a drum so that a small
airgap is left between the periphery of the magnet and drum. The
drag cup is mounted on a shaft and is supported in such a way that it
fits over the magnet rotor to reduce the airgap to a minimum. A
calibrated hairspring is attached to one end of the drag cup shaft, and
at the other end to the mechanism frame. At the front end of the
shaft, a gear train is coupled to two concentrically-mounted pointers;
a large one indicating hundreds and a small one indicating thousands
of rpm.
System operation
As the generator rotor is driven round inside its stator, the poles
sweep past each stator winding in succession so that three waves or
phases of alternating emf are generated, the waves being 120° apart.
The magnitude of the emf induced depends on the strength of the
magnet and the number of turns comprising the phase coils.
Furthermore, as each coil is passed by a pair of rotor poles, the
induced emf completes one cycle at a frequency determined by the
rotational speed of the rotor. Therefore, rotor speed and frequency
Figure J5. 1 Principle of a
generator and indicator system.
PERMANENT-MAGNET
ROTOR
DRAG DISC
MAGNET PAIRS
\
TEMPERATUIIE
COMPENSATOR
SPACERS {3)
are directly proportional, and since the rotor is driven by the engine
at some fixed ratio, then the frequency is a measure of the engine
speed.
The generator emfs are supplied to the corresponding phase coils of
the indicator stator to produce currents of a magnitude and direction
dependent on the emfs. The distribution of stator currents produces a
resultant magnetic field which rotates at a. speed dependent on the
generator frequency. As the field rotates it cuts through the bars of
the squirrel-cage rotor, inducing a current in them which, in turn,
sets up a magnetic field around each bar. The reaction of these fields
with the main rotating field produces a torque on the rotor causing it
349
to rotate in the same direction as the main field and at the same
speed.
As the rotor rotates it drives the permanent magnet of the speedindicating unit, and because of relative motion between the magnet
and the drag cup. eddy currents are induced in the latter. These
currents create a magnetic field which reacts with that of the
permanent magnet, and since there is always a tendency to oppose
the creation of induced currents (Lenz's law), the torque reaction of
the fields causes the drag cup to be continuously rotated in the same
direction as the magnet. However, this rotation is restricted by the
calibrated hairspring in such a manner that the cup will move to a
position at which the eddy-current drag torque is balanced by the
tension of the spring. The resulting movement of the drag cup shaft
and gear train thus positions the pointers over the dial to indicate the
engine speed prevailing at that instant.
Indicators are compensated for the effects of temperature on the
permanent magnet of the speed-indicating element by a thermomagnetic shunt. device fitted adjacent to the magnet.
The drag disc version of the speed-indicating element consists of
six pairs of small permanent magnets mounted on plates bolted
together in such a way that the magnets are directly opposite each
other with a small airgap between pole faces to accommodate the
disc. Rotation of the disc as a result of eddy-current rlrag is
transmitted to the pointers in a similar manner to that already
described.
Percentage rpm indicators
The measurement of engine speed in terms of a percentage is adopted
for turbine engirie operation, and was introduced so that various types
of engine could be operated on the same basis of comparison. The
dial presentations of two representative indicators are shown in Fig.
15.2. The main scales are calibrated from O to 100 per cent in 10 per
cent increments, with 100 per cent corresponding to the optimum
turbine speed. In order to achieve this the engine manufacturer
chooses a ratio be~ween the actual turbine speed and the generator
drive so that optimum speed produces a specific value at the
generator drive. A second pointer or digital counter displays speed in
one per cent increments.
F/guN 15. 2 Percentage rpm
indicators.
350
Figure /5.3 Servo-operated
indicator.
Servo-operated indicators
A schematic diagram of the internal circuit arrangement of a typical
indicator is shown in Fig. 15.3, and its modular construction is
illustrated in Fig. 15.4; it is used in conjunction with an ac generator.
The generator signals are first converted to a square waveform by
a squaring amplifier within the signal-processing module, and in
order to obtain suitable positive and negative triggering pulses for
each half-cycle of the waveform, it is differentiated by a signalshaping circuit. The pulses pass through a monostable which then
produces a train of pulses of constant amplitude and width, and at
twice the frequency of the generator signal. In order to derive the
voltage signal to run the de motor to what is termed the demand
speed condition, the monostable output is supplied to an integrator via
a buffer amplifier.
The demand signal from the integrator is then applied to a sensing
network in a servo amplifier and monitor module, where it is
compared with a de output from the wiper of a position feedback
potentiometer. Since the wiper is geared to the main pointer of the
indicator, its output therefore represents indicated speed. Any
difference between the demand speed and indicated speed results in
an error signal which is supplied to the input and output stages of the
servo amplifier, and then to the armature winding of the motoi;; the
indicator pointer and digital counter are then driven to the demanded
speed position. At the same time, the feedback potentiometer wiper is
also repositioned to provide a feedback voltage to back-off the
fl> SIGAA!. SOOAAlNG
~ BUfFER
~ SERVO ORM!
~ FUO!IACK 0UfFER
3% MECHMICAI. lllllVE
351
Figure I 5. 4 Construction of a
servo-operated indicator.
demanded speed signal until the error signal is zero; at this point, the
indicator will then display the demanded speed.
The output voltage from the servo amplifier input stage is also fed
to a servo loop monitor, the purpose of which· is to detect any failure
of the servo circuit to back-off the error signal voltage. In the event
of such failure, the monitor functions as an 'on-off switch, and in
the 'off state it de-energizes a solenoid-controlled warning flag which
appears across the. digital counter display.
An overspeed pointer is also fitted concentrically with the main
pointer, and is initially positioned at the appropriate scale graduation.
If the main pointer exceeds this position, the limit pointer is carrieci
with it. When the speed has been reduced the main pointer will move
correspondingly, but the limit pointer will remain at the maximum
speed reached since it is under the control of a ratchet mechanism. It
can be returned to its initial position by applying a separately
switched 28 V de supply to a reset solenoid within the indicator.
Tacho probe and indicator system
This system has the advantage of providing a number of separate
electrical outputs additional to those required for sf>eed indication,
e.g. automatic power control and flight data acquisition systems.
Figure 15.5 Tacho probe.
POI.E PIECE
COILS
MAGNET
ELECTRICAi. CONNECTOII
AXIS OF POLARIZATION
Furthermore, and as will be noted from Fig. 15.5, it has the
advantage that there are no moving parts for subjection to high
rotational loads.
The stainless steel, hermetically-sealed probe comprises a
permanent magnet, a pole piece, and a number of cupro-nickel or
nickel-chromium coils around a ferromagnetic core. Separate
windings (from five to seven depending on the type and application
of a probe) provide outputs to the indicator and other processing units
requiring engine speed data. The probe is flange-mounted on an
engine at a station in the high-pressure compressor section so that it
extends into this section. In some turbofan engines, a probe may also
be mounted at the fan section for measuring fan speed. When in
position, the pole pieces are in close proximity to the teeth of a gear
wheel (sometimes referred to as a 'phonic wheel') which is driven at
the same speed as the compressor shaft or fan shaft as appropriate.
To ensure correct orientation of the probe, a locating plug is provided
in the mounting flange.
The permanent magnet produces a magnetic field around the
sensing coils, and as the gear wheel teeth pass the pole pieces, the
intensity of flux through each pole varies inversely with the width of
the air gap between poles and gear wheel teeth. As the flux density
353
Figure 15.6 DC torque motor
tachometer.
14 \I DC
changes, an emf is induced in the sensing coils, the amplitude of the
emf varying with the rate of flux density change. Thus, in taking the
position shown in Fig.· 15.5 as the starting position, maximum
intensity would occur, but the rate of density change .would be zero,
and so the induced emf would be at zero amplitude.
When the gear teeth move from this position, the flux density
first begins to decrease, reaching a maximum rate of change and
thereby inducing an emf of maximum amplitude. At the position in
which the pole pieces align with the 'valleys' between gear teeth, the
flux density will be at a maximum, and because the rate of change is
zero the emf is of zero amplitude. The flux density will again
increase as the next gear teeth align with the pole pieces, the
amplitude of the induced emf reaching a maximum ccincident with
the greatest rate of flux density change. The probe and gear teeth
may therefore be considered as a magnetic flux switch that induces
emfs directly proportional to the gear wheel and compressor or fan
shaft speed.
Figure 15.6 illustrates an example of indicator circuit used in a
tacho probe system. The probe output signals pass through a signalprocessing module and are summed with an output from a servo
potentiometer and a buffer amplifier. After summation the signal
passes through a servo amplifier to the torquer motor which then
rotates the pointer to indicate the change in probe signals in terms of
speed. The servo potentiometer is supplied with a reference voltage,
and gi::i.ce its wiper is also positioned by the motor, the potentiometer
will control the summation of signals to the servo amplifier to ensure
signal balancing at the various constant speed conditions, In the event
of a power s upply or signal failure, the pointer of the indicator is
returned to an 'off-scale' position under the action of a pre-loaded
helical spring.
In some
the torquer motor is of the ac
which
drives a digital counter in addition to the pointer. Indication of power
failure also differs in that a flag is energized to obscure the counter
display.
Manifold pressure indicators
These indicators, colloquially termed 'boost gauges', are of the
direct-reading type and are calibrated to measure absolute pressure in
inches of mercury, such pressure being representative of that
produced at the induction manifold of a supercharged piston engine.
Before considering a typical example, it is useful to have a brief
understanding of the principle of supercharging.
The power output of an internal combustion engine depends on the
density of the combustible mixture of fuel and air introduced into its
cylinders at that part of the operating cycle known as the induction
stroke. On this stroke, the piston moves down the cylinder, an inlet
valve opens, and the fuel/air mixture, or charge prepared by the
carburettor, enters the cylinder as a result of a pressure difference
acting across it during the stroke. If, for example, an engine is
running in atmospheric conditions corresponding to the standard sealevel pressure of 14.7 lbf/in 2 , and the cylinder pressure is reduced to,
say, 2 lbf/in2, then the pressure difference is 12.7 lbf/in 2 , and it is
this pressure difference that 'pushes' the charge into the cylinder.
An engine in which the charge is induced in this manner is said to
be normally aspirated; its outstanding characteristic is that the power
it develops steadily falls off with decrease of atmospheric pressure.
This may be understood by considering a second example in which it
is assumed that the engine is operating at an altitude of 10 000 ft. At
this altitude, the atmospheric pressure is reduced by an amount which
is about a third of the sea-level value, and on each induction stroke
the cylinder pressures will decrease in roughly the same proportion.
We thus have a pressure of about 10 lbf/in2 surrounding the engine
and 1.5 lbf/in2 in each cylinder, leaving us with a little more than 8.5
lbf/in2 with which to 'push' in the useful charge. This means then
that at IO 000 ft only a third of the required charge gets into the
cylinders, and since power is governed by the quantity of charge, we
can only expect a third of the power developed at sea-level.
This limitation on the high-altitude performance of a normallyaspirated engine can be overcome by artificially increasing the
available pressure so as to maintain as far as possible a sea-level
value in the induction system. The process of increasing pressure and
charge density is known as supercharging or boosting, and the device
employed is, in effect, an elaborate form of centrifugal air pump
fitted betwen the carburettor and cylinders and driven from the engine
crankshaft through step-up gearing. It pumps by giving the air a very
high velocity, which is gradually reduced as it passes through diffuser
355
30
Figure 15. 7 Principle of
manifold pressure indicator.
25
35
SEALED SPRING LOADED BELLOWS
Manifold pressure
vanes and a volute, the reduction in speed giving the required
increase in pressure.
In order to measure the pressure delivered by the supercharger and
so obtain an indication of engine power, it is necessary to have an
instrument which indicates absolute pressure. The mechanism of a
typical indicator is schematically illustrated in Fig. 15. 7. The
measuring element is made up of two bellows, one open to the
induction manifold and the other evacuated and sealed. A controlling
spring is fitted inside the sealed bellows and distension of both
bellows is transmitted to the pointer via a lever, quadrant and pinion
mechanism. A filter is located at the inlet to open the bellows, where
there is also a restriction to smooth out any pressure surges.
When pressure is admitted to the open bellows the latter expands
causing the pointer to move over the scde (calibrated in inches of
mercury) and so indicate a change in pressure from the standard sealevel value of 29.92 (zero 'boost'). With increasing altitude, there is
a tendency for the bellows to expand a little too far because the
decrease in atmospheric pressure acting on the outside of the bellows
offers less opposition. However, this tendency is counteracted by the
sealed bellows, which also senses the change in atmospheric pressure
but expands in the opposite direction. Thus a condition is reached at
which the forces acting on each bellows are equal, cancelling out the
effects of atmospheric pressure so that manifold pressure is measured
directly against the spring.
Torque monitoring
The monitoring of torque relates particularly to the power control of
certain types of piston engines, to turbopropeller engines, and also to
the control of engines in some types of helicopter. In all cases it
involves the use of a torquemeter which is essentially an engine
component, and is normally built in with the gear transmission
assembly between the main drive shaft and the propeller shaft or the
main rotor shaft in respect of a helicopter. The construction of
torquemeters depends on the type of engine, but in most ca~::s they
are of hydro-mechanical form, operating on the principle wh::reby
any tendency for some part of the gear transmission to rotate is
resisted by pistons working in hydraulic cylinders secured to the ge~r
casing. The principle as applied to a piston engine is shown in Fig.
15.8.
Oil, which is supplied from the engine lubricating system to the
Figure 15.8 Torquemeter
principle.
STATIONARY
RING GEAR
''...)==::!.!========::::::=J
DIRECTION IN WHICH RING
GEAR TENDS TO ROTATE·
-
::ff: Of CRANKSHAFT
----
01RECTlON Of PROPEUER
SHAFT ROTATION
357
~I= J. - ----·· •2-il~•1
..JlJUlJL .._-
Figure I 5. 9 Electrical type of
torque-indicating system.
l Power shaft, 2 sleeve,
3 sensor, 4 notched wheel,
5 toothed wheel. 6 sleeve.
Roror
end
I
I
6
2
•2
•2
e1
~
Low ... /'\...ll..A.
ror<1ue ( ~ \
,
-i
-•-
(
I
\ , ... ___I .... , ,'
'
/
e
-m..,..High
~
\
rorque
I
I
I-·-
\
I
' ' ..... ___, , ,
I
I
/
cylinders via a torquemeter pump, absorbs any loads due to
movement of the pistons. The oil is thus subjected to pressures which
are proportional to the applied loads or torques, and are transmitted
to the torque pressure-indicating system which is normally of the
synchronous transmission type.
The power, in this case brake horsepower (bhp), is calculated from
the formula
bhp == pNIK
where p is the oil pressure, N the engine speed (rev/min) and K is a
torquemeter constant derived from the reduction ratio between engine
and propeller shaft gearing, length of torque arm, and number and
area of pistons.
Turbopropeller engines are, as far as power is concerned, similar
to large supercharged piston engines: most of the propulsive force is
produced by the propeller, only a very small part being derived from
the exhaust unit thrust. They are, therefore, fitted with a torquemeter
and pressure-indicating system of which the readings are an indication
of the shaft horsepower (shp). The torquemeter pressure indicator is
used in conjunction with the rpm and exhaust gas temperature
indicators.
Figure 15.9 schematically illustrates an electrical torque-indicating
system which is used in one type of helicopter currently in service
for measuring the torsion of the main power shaft in relation to the
effects of engine torque and the drive resistance set up by the main
rotor. The torquemeter consists of two wheels: one wheel is notched
and is attached at the engine end of the power shaft by a sleeve, and
358
it interfaces with teeth on the second wheel which is also attached by
a sleeve at the rotor end of the power shaft. An electromagnetic
sensor is mounted in close proximity to the peripheries of both
wheels.
Under operating conditions, there is opposition between engine
torque and main rotor drive resistance, resulting in torsion of the
power shaft. This, in turn, results in relative displacement of the two
wheels and variation in the gap widths e 1 and e2 between teeth and
notches. As the gaps pass the sensor unit they cause variations in its
magnetic field which induce signal pulses the shape of which
represent gap width and, therefore, the torque. The signals are
transmitted to the signal processor unit which then produces a de
voltage signal proportional to the torque. After amplification, the
signal is supplied to the indicator for the purpose of driving its motor
and pointer drive mechanism.
Exhaust gas
temperature
The measurement of exhaust gas temperature (EGT) is based on the
thermo-emf principle already described in Chapter 13 (see page 320),
and requires the use of chromel/alumel thermocouple probes
immersed in the gas stream at the selected points appropriate to the
type of engine.
Types of probe
Probes are generally classified as stagnation and rapid response, their
application depending upon the velocity of gases. In pure jet engines
the gas velocities are high, and for this reason stagnation-type probes
as indicated in Fig. 15. lO(a) are employed. The gas entry and exit
holes, usually called sampling holes, are staggered and of unequal
size, thus slowing up the gases and causing them to stagnate at the
hot junction, thus giving it time to respond to changes in temperature.
Rapid-response thermocouple probes are normally adopted in the
EGT systems of turbopropeller engines since their exhaust gas
velocities are lower. As can be seen from (b) of Fig. 15.10, the
sampling holes are diametrically opposite each other and of equal
size; the gases can, therefore, flow directly over the hot junction
enabling it to respond more rapidly.
Typical response times for stagnation and rapid-response probes are
1-2 sec and 0.5-1 sec respectively.
Probes may also be designed to contain double, triple, and in some
cases up to eight hot junctions within a single probe. A triple
arrangement is shown in Fig. 15. lO(c). The purpose of such multiarrangements is to provide signals to other systems requiring exhaust
gas temperature d.ata. The thermocouple elements are insulated from
359
Figure 15.10 Types of
thermocouple probe.
(a) Stagnation; (b) rapid
response; (c) triple-element.
(a)
(b)
each other and maintained in position by a special compound
material, e.g. compacted magnesium oxide.
When the hot junctions of immersion-type probes are in contact
with the gas stream, then not only will the stream velocity be
reduced, but also the gas will be compressed by the expenditure of
kinetic energy, resulting in an increase of hot-junction temperature. It
is in this connection that the term recovery factor is used, defining
the proportion of kinetic energy of the gas recovered when it makes
contact with the hot junction. This factor is, of course, taken into
account in the design of thermocouples so that the 'heat transfer', as
we may call it, makes the final reading as nearly as possible a true
indication to total gas temperature.
Location of probes
The points -at which the gas temperature is to be measured are of
great importance, since they will determine the accuracy with which
measured temperature can be related to engine performance. The
ideal location is either at the turbine blades themselves, or at the
turbine entry, but certain practical difficulties are involved which
preclude the application of thermocouple probes at these locations.
Consequently, probes are installed at such locations as exhaust units,
Figure 15. /J
Probe locations.
Location
(b)
HP Turbine
LP Turbine
as at (a) of Fig. 15 .11, between high and low pressure turbines as at
(b), or in some engines at the leading edge of stator guide vanes
between turbine stages.
For accurate measurement it is necessary to sample temperatures
from a number of points evenly distributed over a cross-section of the
gas flow. This is because temperature differences can exist in various
zones or layers of the flow through the turbine section and exhaust
361
Figure 15./2 Probe grouping.
Steel tube
..,...,,.__ _ Long-reach
(a)
Main junction
box
(b)
unit, and so measurement at one point only would not be truly
representative of the conditions prevailing.
A measuring system, therefore, always consists of a group of
thermocouple probes suitably disposed in the gas flow, and connected
in parallel so as to measure a good average temperature condition.
The probes in a group may contain a single hot junction, or pairs of
junctions referred to as short reach and long reach from the extent to
which they reach into the gas stream; an example is shown in Fig.
15.12(a).
Probes and their chrome! and alumel cables are made up into a
harness assembly of a design appropriate to the type of engine and
number of probes required. An eight-probe arrangement comprising
16 hot junctions is shown at (b) of Fig. 15.12. The cables pass
through steel tubing and terminate at a main junction box which
serves as the 'take-off' point for the connection of indicators and
other units requiring EGT data. Terminal studs of junction boxes are
also made of chromel and alumel and, in order to ensure correct
polarity of cable connections, the diameters of alumel studs are large1
than those of the chromel studs.
Indicators
Depending on the instrumentation configurations adopted for a
particular type of aircraft, the indication of EGT, as in the case of
other power and control parameters, may be provided by servooperated indicators or by electronic display methods. The modular
362
arrangement of one type of servo-operated indicator is illustrated in
Fig. 15. l3(a).
The output from the thermocouple probes is supplied first to a
cold junction reference bridge circuit, the purpose of which is to
compensate for changes in ambient temperature of the indicator. The
circuit is shown in more detail in diagram (b). The thermocouple
harness and cables are connected to copper leads which are embedded
in close proximity to each other within a former which supports a
copper coil resistor R4 ; thus, together they form the effective cold
junction of the system. The bridge circuit is supplied with 7 V de
from a stabilized reference supply module within the indicator, and
the bridge output is supplied to a servo amplifier.
As we have already learned, the standard values of emf produced
by a thermocouple are related to a selected value of cold junction
temperature (see page 322). In this case, the bridge circuit is adjusted
by means of a variable resistor RV 1 so that an emf of the correct
sense and magnitude is injected in series with that of the
thermocouples such that, in combination, the emf is equal to that
which would be obtained if the cold junction temperature were OOC.
Since the ambient temperature of the indicator, and hence the cold
junction, will in the normal operating environment always be higher
than this, then the temperature difference will reduce the
thermocouple output. The resistor R 1 will, however, also be subjected
to the higher ambient temperature, but because under such conditions
the resistance of R 1 decreases, it will modify the bridge circuit
conditions so as to restore the combined emf output to the standard
value corresponding to a cold junction temperature of 0°C.
The output is termed the demand EGT signal and is compared with
a de output from the wiper of a positional fee,dback potentiometer,
and since the wiper is geared to the main pointer and digital counter
of the indicator, t.hen the de output which is fed back to the cold
junction reference circuit represents the indicated EGT. Any
difference between demanded and indicated EGTs results in an error
signal being produced by the reference circuit which then supplies the
signal to the servo amplifier as shown in (a) of Fig. 15.13. The
amplifier output is fed to the armature winding of the de servomotor
which then drives the pointer and digital counter, causing them to
display a coarse and fine indication respectively of the EGT. The
feedback potentiometer wiper is also repositioned to provide a
feedback voltage which backs-off the demanded temperature signal
until the error signal is zero; at this point the indicator will then
display the demanded temperature.
The output voltage from one stage of the servo amplifier is also fed
to a servo loop monitor, the purpose of which is to detect any failure
of the loop to back-off the error signal voltage. Should such failure
·occur, the monitor functions as an 'on-off switch, and in the 'off
363
Figure 15.13 Servo-operated
EGT indicator.
OVER-TEMPERATURE
WARNING LIGHT
I
OVER· TEMPERATURE LIMIT
/POINTER
THERMOCOUPLE
SIGNAL
LAMP SUPPLY '
28V {
o.c.
'
TEST SUPPLY
____________
_
IL_:SERVO
AMP. AND MONITOR MODUlEJ
115 V 400 Hz A.C.
G>----""
'-----------------;~;J~~:L
- - - - - - - - - - - - - - - - - - - . . . P O T E N T I O M E ER
28 V O.C.
f>....,_ ERROR SIGNAL ~ FEED8ACK
V
V .... BUFFER
[}:>,
SERVO DRIVE
lCDC MECHANICAL DRIVE
(a)
REFERENCE SUPPLY
- - - - - - - - - • - - - - - - O U T P U T TO POSITIONAL
FEEDBACK PQTENTIOMETER
------•----
COPPER LEADS
REFERENCE SUPPLY
VOLTAGE
THERMOCOUPLE HARNESS
I
AND
-----t>j
EXTENSION I.EADS
}
•
CHROMEL/ALUMELLEADS
(b)
364
TEST POINT
OUTPUT TO
DIFFERENTIAL
AND SERVO
AMPLIFIERS
state de-energizes a solenoid-controlled warning flag which appears
across the digital counter display. The flag will also appear in the
event of the 115 V ac supply to the indicator falling below 100 V.
An over-temperature warning light is incorporated in the indicator,
and is controlled by a relay, a comparator, and a solid-state switching
circuit. The function of the comparator is to compare the feedback
voltage from the positional potentiometer with a pre-set voltage the
level of which is equivalent to a predetermiend over-temperature limit
for the particular type of engine. In the event of this limit being
exceeded, the feedback voltage will exceed the reference voltage
level, and the switching circuit will cause the relay to energize,
thereby completing the circuit to a warning light. A separate supply
voltage may be connected to the Iighi by means of an 'override'
facility as a means of testing its filament at any point over the
temperature range of the indicat.or.
An over-temperature pointer is also fitted concentrically with the
main pointer, and is initially positioned at the appropriate scale
graduation. It operates in a similar manner to the over-speed pointer
of a servo-operated tachometer indicator (see page 351).
Examples of EGT indications by means of electronic display
systems will be covered in Chapter 16.
Engine pressure ratio
(EPR) measurement
EPR is an operating variable which, together with rev/min, EGT and
fuel flow, provides an indication of the thrust output of turbine
engines, and involves the measurement of the ratio between the
pressures at the compressor intake and the turbine outlet or exhaust.
In general, a measuring system consists of an engine inlet pressure
probe, a number of pressure-sensing probes projected into the exhaust
unit of an engine, a pressure ratio transmitter, and an indicator. The
interconnection of these components based on a typical system is
schematically shown in Fig. 15. l~.
The inlet pressure-sensing probe is similar to a pitot probe, and is
mounted so that it faces into the airstream in the engine intake or, as
in some power plant installations, on the pylon and in the vicinity of
the air intake. The probe is protected against icing by a supply of
warm air from the engine anti-ice system.
The exhaust pressure-sensing probes are interconnected by pipelines
which terminate at a manifold, thus averaging the pressures. In some
engine systems, pressure-sensing is done from chambers contained
within the EGT sensing probes. A pipeline from the manifold, and
another from the inlet pressure probe, are each connected to the
pressure ratio transmitter which comprises a bellows type of pressuresensing transducer, a linear voltage differential transformer (LVDT),
a two-phase servomotor, an amplifier and a potentiometer. The
365
Figure 15.14 EPR system.
REF. PHASE
_ - - - _ _ _ _ _ TRANSMITTER _ _ •
J
MECHANICAL LINKAGE
MANIFOLO
transducer bellows are arranged in two pairs at right angles and
supported in a frame which, in turn, is supported in a gimbal and
yoke assembly. The gimbal is mechanically coupled to the
servomotor via a gear train, while the yoke is coupled to the core of
the LVDT. The servomotor also drives the wiper of the potentiometer
which adjusts the output voltage signals to the indicator in terms of
changes in pressure ratio.
The indicator shown in the diagram is of the servo-operated type.
In electrnnic display systems (see Chapter 16) the transmitter output
ti;.e appropriate system computer.
signals are supplied direct
From Fig. 15.14 it will be noted that the intake pressure is
admitted to two of the bellows in the transducer, exhaust gas pressure
is admitted to the third bellows, while the fourth is evacuated and
to
sealed. Thus the system, together with its frame, gimbal and yoke
assembly, forms a pressure balancing and torsional system.
When a pressure change occurs, it causes an unbalance in the
bellows system, and the resultant of the forces acting on the
transducer frame acts on the yoke such that it is pivoted about its
axis. The deflection displaces the core of the LVDT to induce an ac
signal which is amplified and applied to the control winding of the
servomotor. The motor, via a gear train, alters the potentiometer
output signal to the indicator so that its pointer and digital counter are
servo-driven to indicate the new pressure ratio. Simultaneously, the
motor drives the transducer gimbal and LVDT core in the same
direction as the initial yoke movement, so that the relative movement
now produced between the core and coils starts reducing the servomotor drive signal, until it is finally 'nulled' and the system stabilized
at the new ratio.
The lower counter shown in the diagram is for the purpose of
indicating a reference EPR value; it is set manually by rotating the
setting knob.
If a circuit malfunction occurs, an integrity monitoring circuit
within the indicator activates a warning flag circuit, causing the flag
to obscure the digital counter display.
In some types of aircraft, a maximum allowable EPR limit
indicator is also provided, and is integrated with a TAT indicator (see
page 64) and also with an ADC; its purpose being to indicate limits
related to air density and altitude values from which thrust settings
have been predetermined for specific operating conditions. The
conditions are climb, cruise, continuous and go-around, and are
selected as appropriate by means of a mode selector switch connected
to a computing and switching circuit which generates a datum signal
corresponding to each selected condition. The signal is then supplied
to a comparator, which also receives temperature signals from the
TAT sensor and altitude signals from the ADC. These signals are
compared with the datum signal and the lower value of the two is
automatically selected as the signal representing the maximum EPR
limit for the selected operating condition. The comparator transmits
this signal to an amplifier and a servomotor which then drives a
digital counter to display the limiting value.
Fuel flow
measurement
Fuel flow measuring systems vary in operating principle and
construction, but principally they consist of two units: a transmitter
or flowmeter, and an indicator. Transmitters are connected in the
delivery Hnes of an engine fuel system, and are essentially electromechanical devices producing output signals proportional to flow rate
which in a basic system is indicated in either volumetric or mass
367
units. In many of the systems currently in use, an intermediate
amplifier/computer is also included to calculate a fuel flow/time ratio
and to transmit signals to indicators which can display not only flow
rate but also the .tmount of fuel consumed.
Basic system
Figure 15.15 is a sectioned view of a transmitter that forms the
measuring unit of a simple flow rate indicating system. It has a cast
body with inlet and outlet connections in communication with a
spiral-shaped metering chamber containing the metering assembly.
The latter consists of a vane pivoted so that it can be angularly
displaced under the influence of fuel passing through the chamber. A
small gap is formed between the edge of the vane and the chamber
wall which, on account of the volute form of the chamber, increases
in area as the vane is displaced from its zero position. The variation
in gap area controls the rate of vane displacement which is faster at
the lower flow rates (gap narrower) than at the higher ones. The vane
is mounted on a shaft carried in two plain bearings, one in each
cover plate enclosing the metering chamber.
At one end, the shaft protrudes through its bearing and carries a
two-pole ring-type magnet which forms part of a magnetic coupling
between the vane and the electrical transmitting unit, which m::.y be a
precision potentiometer or an ac torque synchro. The shaft of the
transmitting unit carries a two-pole bar-type magnet which is located
inside the ring magnet. The interaction of the two fields provides a
'magnetic lock' so that the potentiometer wiper (or synchro rotor) can
follow any angular displacement of the metering vane free of friction.
The other end of the metering vane shaft carries the attachment for
the inner end of a specially calibrated control spring. The outer end
Figure 15.15 Rotating vane
fuel tlowmeter.
of the spring is anchored to a disc plate which can be rotated by a
pinion meshing with teeth cut in the periphery of the plate. This
provides for adjustment of the spring torque during flowmeter
calibration.
Any tendency for the metering assembly to oscillate under static
flow conditions is damped out by a counterwejght and vane, attached
to the metering vane shaft, and operating in a separate fuel-filled
chamber secured to one side of the transmitter body.
When fuel commences to flow it passes through the metering
chamber and deflects the metering vane from its zero position and
tends to carry it round the chamber. Since the vane is coupled to the
calibrated spring, the latter will oppose movement of the vane,
permitting it to take up only an angular position at which spring
tension is in equilibrium with the rate of fuel flow at any instant.
Through the medium of the magnetic-lock coupling the vane will also
cause the potentiometer wiper, or synchro rotor, to be displaced. In
the former case, and with a steady direct voltage across the
potentiometer, the voltage at the wiper is directly proportional to the
fuel flow. The voltage is fed to an amplifier, whose output current
drives a milliammeter pointer to indicate the current in terms of fuel
flow in gal/hr or lb/hr.
In a system employing synchros, the current flow due to
differences in angular positions of the rotors will drive the indicator
synchro rotor directly to the 'null' position and thereby make the
pointer indicate the fuel flow.
In meters of this type it is also necessary to provide a bypass for
the fuel in the event of jamming of the vane or some other
obstruction causing a build-up of pressure on the inlet side. As may
be seen from Fig. 15.15, this is accomplished by a spring-loaded
valve incorporated in the metering chamber. The spring tension is
adjusted so that the valve lifts· from its seating and allows fuel to
bypass the metering chamber when the pressure difference across the
chamber exceeds 2.5 lbf/in 2 •
Integrated tlowmeter systems
An integrated flowmeter system may broadly be defined as one in
which a fuel consumed measuring element is combined with that of
fuel flow, thus permitting the display of both quantities in a single
indicator.
In order to accomplish this it is necessary to introduce an
integrating system to work out fuel consumed in the ratio of fuel flow
rate to time. Such a system may be mechanical, forming an integral
part of an indicator mechanism, or as in electronic fuel flow
measuring systems, it may be a special dividing stage within the
amplifier, or even a completely separate integrator unit.
Decoupling disc
Figure 15.16 Integrated
flowmeter system.
LVDT
------. J
Turbine
60 V ac
To
INLET
Engine
Fuel
passage
Fuel
passage
Reset
switch
.......
..__ _ _ _ _ _ _ _ _--4,,
The components of a typical system are shown in Fig. 15.16. The
transmitter comprises an impeller driven by a two-phase ac motor, a
turbine which is interconnected with a calibrated restraining spring,
and an LVDT sensor. A decoupling disc is located between the
impeller and turbine, its purpose being to prevent an 'hydraulic
transmission' effect on both units when operating at low rates of fuel
flow. The indicator is servo-operated with the drive to the flow rate
pointer being effected by means of an eddy-current drag type of
mechanism similar to that adopted in some rpm indicators (see page
350). The fuel consumed indicator is a digital counter which is·
mechanicaily integrated with the servomotor via a gear transmission,
the ratio of which is preselected to establish the requisite relationship
between the m.otor speed, which is proportional to flow rate, and
time.
The system is supplied with 115 V single-phase 400 Hz ac from an
aircraft's power system and this is utilized by a power supply unit
within the indicator, the primary coil of the LVDT in the transmitter,
and by a separate power supply unit (not shown in the diagram). This
unit contains a temperature stable oscillator connected to a
voltage/frequency converter which converts the main supply into a
two-phase 60 V 8 Hz output; this, in turn, is supplied to the
370
transmitter impeller motor. The rotating field set up in the rotor
windings interacts with its permanent magnet rotor which rotates in
synchronism and drives the impeller at a constant speed.
Fuel flow rate is, in the first instance, always established by an
engine's fuel control unit which is calibrated or 'trimmed' to control
rates commensurate with the varying operational conditions and the
other associated power parameters, i.e. rpm, EGT and EPR. When
fuel enters the transmitter it passes through passages in the impeller
which, on account of its rotation, causes the fuel to swirl at a
velocity governed by the flow rate. The fuel is then diverted around
the decoupling disc, and in passing through passages in the turbine, it
imparts a rotational force which tends to continuously rotate the
turbine in the same direction as the impeller. This tendency is,
however, restrained by the calibrated spring such that the rotation is
limited and balanced at an angular position proportional to the flow
rate of. fuel passing through the transmitter.
The movement of the turbine and its shaft alters the position of the
LVDT sensor core, so that a signal voltage (up to 5 V at maximum
flow rate) is induced in the secondary winding and supplied to the
indicator servomotor via the closed contacts of the reset relay and
amplifier. The servomotor rotates at a speed proportional to tht flow
rate, and by means of the eddy-current drag mechanism posi;irn,., the
pointer to indicate this rate.
The reset switch is separately located on a flighl deck panel. and
when pressed it energizes the relay in the indicator to supply l l 5 V
ac to the servo amplifier and motor, causing it to drive 'downscale'
rapidly in order to reset the fuel consumed counter display to zero.
Figure 15.17 illustrates the components of another type of
integrated system. It differs from the one just described in that the
transmitter utilizes two electromagnetic pick-off elements, and the
processing of signals relevant to flow .ate and fuel consumed is
carried within a separate electronic unit.
The transmitter consists of a light-alloy body containing a flowmetering chamber, a motor-driven impeller assembly, and the
externally-mounted coils of the pick-off elements. The impeller
assembly consists of an outer drum which is driven through a
magnetic coupling and reduction gear by a synchronous motor, and
an impeller incorporating vanes c>nd fuel passages to impart swirl and
angular velocity to the fuel flowing through the metering chamber.
The drum and impeller are coupled to each other by a calibrated
spring. The motor is contained within a fixed drum at the inlet end
and rotates the impeller at a constant speed (a typical value is 100
rpm). Straightening vanes are provided in the fixed drum around the
motor to remove any angular velocity already present in the fuel
before it passes through the impeller assembly. A point to note about
the use of a magnetic coupling between the motor and impeller
371
Figure I 5. 17 Electronic ··
integrated flowmeter system.
I Fixed drum, 2 fuel passages,
3 fllO!or shaft, 4 rotating drum,
5 fuel passages, 6 impeller,
7 calibrated spring, 8 pick-off
(drum), 9 magoets, 10 pick-off
(impeller), II magnetic
coupling, 12 motor.
TO ELECTRONIC
UNIT
SUl'l'I.Y TO TRAHSMITTER SYNCHRONOUS MOTOR
.--------0
2.
3.
ELECTRONIC UNIT
assembly is that it overcomes the disadvantages which in this
application would be associated with rotating seals. The motor and its
driving gear are isolated from the fuel by enclosing them in a
chamber which is evacuated and filled with an inert gas before
sealing.
Each of the two pick-off assemblies consists of a magnet and an
iron-cored inductor. One magnet is fitted to the outer drum while the
other is fitted to the impeller, thus providing the required angular
reference points. The magnets are so positioned that under zero flow
372
conditions they are effectively in alignment with each other. The coils
are located in an electrical compartment on the outside of the
transmitter body, together with solid-state circuit units which amplify
and switch the signals induced.
The electronic unit performs the overall function of providing the
power for the various circuits of the system, detecting the number of
pulses produced at the transmitter, and computing and integrating the
fuel flow rate and amount of fuel consumed. It consists of a number
of stages interconnected as shown in Fig. 15.17. The power supply
section (l) controls the voltage and frequency of the supply to the
transmitter synchronous motor, and consists of a transformer, crystal
oscillator, output and power amplifier units.
From the diagram it will be noted that section (2) is made up of
three stages: inhibitor, gate and divider. The respective functions of
these stages are: to suppress all transmitter signals below a certain
flow rate; to control or gate the pulse signals from the power supply
oscillator; and to produce output signals proportional to true flow
rate, and to provide the time dividing factor and output pulses
representing unit mass of fuel consumed. Section (3) is also made up
of three stages: signal comparator, modulator and servo amplifier.
The respective functions of these stages are: to compare the
transmitter output signals with time-base signals fed back from the
indicator; to combine the comparator output with 400 Hz ac and
produce a new output; and to provide an operating signal to the
indicator servomotor control winding.
The indicator employs a flow indicating section consisting of an ac
servomotor which drives a pointer, a digital counter display, and a
potentiometer wiper through a reduction gear train. The reference
winding of the motor is supplied with a constant alternating voltage,
while the control winding receives its signals from the servo amplifier
in the electronic unit. The potentiometer is supplied with de and its
wiper is electrically connected to a solid-state time-base circuit, also
within the indicator. Transmitter output signals are fed into the timebase circuit via a pre-set potentiometer which forms part of the
electronic unit's comparator stage. The difference between the timebase and the indicated fuel flow signal voltages is fed to the
servomotor, which operates to reduce the error voltage to zero and so
to correct the indicated fuel flow.
The fuel consumed section of the indicator consists of a solenoidactuated five-drum digital counter and a pulse amplifier. The
amplifier receives a pulse from the divider stage of the electronic unit
for each unit mass of fuel consumed and feeds its output to the
solenoid, which advances the counter drums appropriately. A
separately located reset switch is also provided for returning the
counter to zero; it operates in a similar manner to that described on
page 371.
373
A
Figure 15. 18 Operation of
pick-offs.
OUTER & INNER DRUMS
INTERCONNECTED BY
CALIBAATED SPRING
(8)
(b)
A
B&C
D
9
Pick-offs
Magnets
Stop
Lag angle at which
impeller and drum rotate
together
When electrical power is switched on to the system, the transmitter
impeller motor is, as already mentioned, rotated at constant speed.
Under zero fuel flow conditions, the magnets of the pick-offs are
effectively in line with each other, although in practice a small
angular difference (typically 3°-5°) is established to maintain a
deflection representing a specific minimum flow rate. This is
indicated in Fig. 15.18(a).
As the fuel flows through the metering chamber, a constant angular
velocity is imparted to the fuel by the rotating impeller and drum
assembly, and since the two are interconnected by the calibrated
spring, a reaction torque is created which alters the angular
displacement between the impeller and drum and their corresponding
magnets. Thus, angular displacement produces a time difference
between signal pulses in the pick-offs, both being proportional to
flow rate. Diagrams (b) and (c) illustrate the displacement for typical
cruising and maximum fuel flow rates respectively.
The position of each magnet is sensed by its own pick-off coil, and
374
the primary pulses induced as each magnet moves past its coil are fed
to the dividing stage in the electronic unit (see Fig. 15.17). The
output from this stage is fed to the control winding of the indicator
servomotor via section (3) of the electronic unit, and the indicator
pointer is driven to indicate the fuel flow. At the same time, the
motor drives the potentiometer wiper, producing a sigflal which is fed
back to the signal comparator stage for comparison with the output
produced by the transmitter. Any resultant difference signal is
amplified, modulated and power amplified to drive the indicator
motor, pointer and digital counter to a position indicating the actual
fuel flow rate.
The divider stage of the electronic unit also uses the transmitter
signals to produce pulse 'time' signals for the operation of the fuel
consumed counter of the indicator. During each successive revolution
of the transmitter impeller assembly the pulses are added and divided
by a selected ratio, and then supplied to the counter as an impulse
corresponding to each pound of fuel consumed.
Engine vibration
monitoring
Engine vibration is a feature of engine operation which cannot be
eliminated entirely even with turbine engines, which, unlike piston
engines, have no reciprocating parts. Thus, by accurate balancing of
such components as crankshafts, compressor and turbine rotor discs,
vibration must be kept down to the lowest levels acceptable under all
operating conditions. In respect of turbine engine operation, however,
there is always the possibility of these levels being exceeded as a
result of certain mechanical failures occurring. For example, a
turbine blade may crack or 'creep', or an uneven temperature
distribution around turbine blades and rotor discs may be set up;
either of these will give rise to unbalanced conditions of the main
rotating assemblies and possible disintegration. In order, therefore, to
indicate when the maximum amplitude of vibration of an engine
exceeds the pre-set level, monitoring systems, which come within the
control group of instrumentation, are provided.
A block diagram of a typical system is shown in Fig. 15 .19. It
consists of a vibration pick-off, or sensor, mounted on an engine at
right angles to its axis, an amplifier monitoring unit, and a movingcoil milliammeter calibrated to show vibration amplitude in
thousandths of an inch (mils).
The sensor is a linear-velocity detector that converts the mechanical
energy of vibration into an electrical signal of proportional
magnitude. It does this by mea,s of a spring-supported permanent
magnet suspended in a coil attached to the interior of the case.
As the engine vibrates, the sensor unit and core move with it; the
magnet, however, tends to remain fixed in space because of inertia.
375
Figure 15.19 Vibration
monitoring system.
r------------------1
INPUT & TEST - - - -
AMPLIFIER
& FILTERS
I
1
RECTIFIER
I
I
SUSPENDED
I
MAGNET
PICK-UP
WARNING
CIRCUIT
I
I
I
I
I
I
I
:
I
I
L_ __ --- --~:::!-==;=_==_!::_-:::!/_ - - - - - -
INDICATOR
_j
115V 400HZ
SINGLE· PHASE
SUPPLY
In other words, its function is similar to that of an accelerometer.
The motion of the coil causes the turns to cut the field of the magnet,
thus inducing a voltage in the coil and providing a signal to the
amplifier unit. The signal, after amplification and integration by an
electrical filter network, is fed to the indicator via a rectifying
section.
An amber indicator light also forms part of the system, together
with a test switch. The light is supplied with de from the amplifier
rectifying section and it comes on when the maximum amplitude of
vibration exceeds the pre-set value. The test switch permits functional
checking of the system's electrical circuit.
In some engine installations, two sensors may be fitted to an
engine: for example, in a typical turbofan engine, one monitors
vibration levels around the fan section, and the other around the
engine core section.
In systems developed for use in conjunction with LCD and CRT
display indicators, the vibration sensors are of the type whereby
vibration causes signals to be induced in a piezoelectric stack (see
also page 165). A CRT display of vibration is shown in Fig. 16.2 .
.376
Electronic instruments for
engine and airframe
systems control
CThe display of the parameters associated with engine performance ~~-~
air.f!:ame systems control by means of CRT-type display units has,
like those of flight instrument systems, become a standard feature of
many types of aircraft. The display units form part of two principal
systems designated as ~~gine indicating and crew alerting system
(EICAS) and electronic centralized aircraft monitorins_ (ECAM)
S.Y§_tem, which were fir.st introduced in Boeing 757 and 767 aircraft
!nd the Airbus A3 l0 respectivelOAt the time of their introduction,
there were differing views on the approach to such operating factors
as flight deck layouts and crews' controlling functions, the extent to
which normal, alerting and warning information should be displayed,
and in particular, whether engine operating data required to be
displayed for the whole of a flight, or only at various phases.
In respect of%J~~S, ~IJ&!!~ ..Q~~H.!!&.Q~~ls_Q~Pla.yed _on its CJlT
~ ~ ~ ~ ~ ~ ! ! ! ! ! } ~ . , The
data, as well as those relevant to other systems, are not necessarily
always on display but in the event of malfunctions occurring at any
t,ime, the flight crew's attention is drawn to them by an automatic
qisplay of messages in the appropriate colours, tfhe ECAM system,
on the other hand, gisplays systems' operation in checklist and
schematic forJ!1 and as this was a concept based on the view that
engine data need to be displayed during the whole of a flight,
traditional instruments were retained in the Airbus A3 l0. It is of
interest to note, however, that in subsequent types produced by this
manufacturer, e.g. A320, the ECAM system is developed to include
the display of engine data in one of its display units.
EICAS
\
£r~~~-~sterp comprises two display units, a control eanel, gn_d __
iwo computers supplied with analog and digital signals from engine
and system sensors as shown in the schematic functional diagram of
Fig. 16. l. ifhe computers are designated 'Left' and 'Right', and only
one is in control at a time; the o er ts on standby'. anct in the event
"'"~ failure it may be switched in either manually or automatically.\
377
f
E!CAS: (unctional
1gure /6. I
r
d iagram.
Discrete
caution &
warning
f,,o,.
~
~
lights
Aural
w arnmgs
~
Upper DU
w::::~;nss
Engine
ryfimary
&
d,splays
___, I
\.._
Standby
engine
indicators
:
\
0,splay
switching
:1
LJ
I
Engme secondary
or
status displays
or
mamtf?nance d1sptays
I
Lower OU
l
,
f
"
~
I
~
,;}
t
.
II
I
Engine sensors:
! 01! press
: 0,1 qty
! Oil temp
! Vibration
:
I
:
I
•I
Display select panel
11
FF
.
R computer
-
I
I.
N,
N,
N,
EPR
EGT
Maintenance
panel
\..
L computer
rl
Other system discrotes
Hyd. Qty & press
ADC Hyd syst. temp.
Control surface positions
Elect. syst volts. amps. freq.
FCC MCOP
TMC interlace
EEC interlace
FMC interlace
RAD Alt interlace
AOC interlace
ECS temps
APU EGT, RPM
Brake temp.
I
.
System sensors
Gen. drive temp
11 , " ;.
:;> Data bus
Operating in conjunction with the system are discrete caution and
warning lights, standby engine indicators and a remotely-located panel
for selecting maintenance data displays. The. system provides the
flight crew with information on pJimary engine parameters (fulltime), with secondary engine parameters and advisory/caution/
warning alert messages displayed as required.
Display units
These units provide a wide variety of information relevant to engine
operation, and operation of other automated systems, and they utilize
colour shadow mask CRTs and .associated card modules whose
functions are identical to those of the EFIS units described in Chapter
12. The units are mounted one above the other as shown in Fig.
16.2.
(The upper unit dis11Iays the primal): engine parameters N 1 ~.£1,
£9.I, a~ warning and caution messages,.. IJ1 some cases this unit can
a~~lay EPR depending on the type of engines installed1 and on
the methods of processing data by the thrust management control
system. ~ r unit displays secondary engine parameters, i.e. N2
378
-
~
-------·
-"·
Figure /6.2 EICAS: engine
data displays.
SECONDARY
s eed, fuel flow, oil quantity, pressure and temperature, and engine
vibration. In agdition, the status o non-engine systems, e.g. l!.!£_ht
control surface positions, hydraulic system, APO, etc., can also be
,displayed together with aircraft configuration ·and maintenance d ~
The rows of 'V's shown on the upper display unit only appear when
secondary information is being displayed on the lower unit.
(Seven colours are produced by the CRTs and they are used as
follows:
-
1.,White4
1 Red4
,All scales. normal operating range of pointer~
readouts.)
,Warning messages, maximum operating limit markL9n
scales, and digital readouts. ~
"'.
379
t Gree!JJ [hrust mode r7adouti and selected EPR/N I speed marks or
target cursors.
tTesting of system only\
tf.aution and advisory messages, caution limit mar~s on
~cales, digital readouts.)
rMagen_!S Qµring in-flight engine starting, and for cross-bleed
messages.i
~ ~ames of all parameters being measuredA(e.g. N 1 oil
pressure, TAT, etc.) and status marks or cues.
The displays are selected according to an appropriate display
selection mode.
f Display modes)
EICAS is designed to categorize displays and alerts according to
function and usage, and for this purpose ~here are th~ ~ s of
displaying information: (i) ~ n a l , (ii) status, and (iii)
nuJi.!W:nance. Modes ,{i) and (ii) are 1!~lectedby the flight ~te_w.._pJL the
display select an , while mode (iii) is selected on t~....m~11ance
~ w~1c is for_the u s e _ ~ .
Operational mode
tThis mode displays the engine operating information and any alen_s
required to be actioned b the crew in flight. Normally only thi,
up~r 1splay unit presents information; t
ower one remains blank
and can be selected to display secondary inf.Q~n ~ and wh_en
r~uired. \
Status mode
When selected this mode displays data to determine the dispatch
readiness of an aircraft, and is closely associated with details
contained in an aircraft's Minimum Equipment List.lThe dis(!la)'._
shows sitions of the flight control surfaces in the form of pointers
es, selecte su system parameters, and
registered against vert1c
5uipment status messages 9n the lower display unit. Selection is
normally done on the ground either as part of pre-flight checks of
dispatch items, or prior to shut-down of electrical power to aid the
flight crew in making entries in the aircraft's Technical Log.
Maintenance mode
This mode provides maintenance en ineers with informaiion in five
1 erent 1sp ay formats to aid them in trouble-shootin and
verificat10n testing o the major sub-system~· The displays, which are
presented on the lower display unit, are not available in flight.
Figure 16.3 EICAS: display
select panel.
DISPLAY
ENGINE STATUS
IDOi
8RT
COMPUTER
EVENT
RECORD
0
L~A
$
8RT
0
8Al
©
THRUST REF SET
'i"
MAX IND
RESET
0
Figure 16.4 Status mode
display.
Display select eand>
\__.
This panel, as indicated in Fig. 16.3, permits control of EICAS
functions and displays and can be used both in flight and on the
ground. It is normally located on the centre pedestal of an aircraft's
flight deck, and its controls are as follows:
l. r.§ngine display switch 'fhis is of the momentary-push t~or
rem.Q_vjng or presenting the display of secondary informatio11-0n the
lower display unit~
2. ,Status display swit<:!) Also of the.._momentary-push type, th!fu is
used to display the status mode information refyrred to earlier, ~the
low:r display un_!!, The display is known as a 'status page', an
example of which is shown in Fig. 16.4.
3. gyent record swit<;}z Th~is of the momentary-Qush type and is
used in the air or on the ground, toj!£tivate the recording of fault
data relevant to the environmental control~, e~ctrical Q2~er,
hydraulic system, performance and APU. Normally, if any
311
malfunction occurs in a system, it is recorded automatically (called an
'auto event') and stored in a non-volatile memory of the EICAS
computer. The push switch also enables the flight crew to record a
suspect malfunction for storage, and this is called a 'manual event'.
The relevant data can only be retrieved from memory and displayed
when the aircraft is on the ground and by operating switches on the
maintenance control panel.
4. [;_omputer select swi~h Jn the 'AUTO' position it selects the
left, or primary, computer and automatically switches to the other
computer in the event of failu3. The other positions are for the
man~! selection af left or right eompttte1s.
'
5. Display brighmess contra/. ,:he inner knob controls the intensity
q_f the displa~, aqd the outer knob controls brightness b..alance
betw~
6. Thrust reference set switch Pulling and rotating the inner knob
positions the reference cursor on the thrust indicator display (either
EPR or N 1) for the engine(s) selected by the outer knob.
7. Maximum indicator reset switg/z If any one of the meafil}red
parameters, e.g_:_o~s.s.ure.,..EGI.,--Shonld exceed normal o_p_~Jl!ing
limits, this will be automatically alerted on the displa)!'.....!!11jts . .:[he
purpose of the reset switch is to clear the alerts frorn_JbLdifil2.lay
when the excess limits no longer exi.§1\
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 messages appear on the top line at the left of the display
screen as shown in Fig. 16.5. In order to differentiate between a
caution and an advisory, the latter is always indented one space to the
right.
382
Figure 16.5 Alert message
levels.
A
B
C
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.
A message is automatically removed from the display when the
associated condition no longer exists. In this case, messages which
appear below the deleted one each move up a line.
When a new fault occurs, its associated message is inserted on the
appropriate line of the display. This may cause older messages to
move down one line. For example, a new caution message would
cause all existing caution and advisory messages to move down one
line.
If there are more messages than can be displayed at one time, the
whole list forms what is termed a 'page', and the lowest message is
removed and a page number appears in white on the lower right side
of the list. If there is an additional page of messages it can be
displayed by pushing the 'Cancel' switch. Warning messages are
carried over from the previous page.
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
383
Figure 16.6 Compact format.
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 ATUS' switch to be inhibited so that the
status page format cannot be displayed.
Display select panel failure
If this panel fails the advisory message 'EICAS CONTROL PANEL'
appears at the top left of the upper display unit together with the
primary information, and the secondary information automatically
appears on the lower display unit. The 'cancel/recall' switches do not
operate in this failure condition.
Standby engine indicator
This indicator provides primary engine information in the event that a
total loss of EICAS displays occurs. As shown in Fig. 16.7, the
Figure /6. 7 Standby engine
indicator.
0
OPERATING
LIMITS
111
SELF
SWIT
0
o I 1- 1
I 1- 1
0
DISPLAY
CONTROL
SWITCH
~
0
-. ,: ,-, EGT -, ,- ,-,
I _11_1
C•1 a_-· .t_1
I I
I :11_1
N2
Ct1 C
-. .,_1
,-,
information relates to N 1 and N2 speeds and EGT; the displays are of
the LCD type. Operating limit values are also displayed.
The display control switch has two positions: 'OW and 'AUTO'.
In the 'ON' position, the displays are permanently on. In the
'AUTO' position the internal circuits are functional, but the displays
will be automatically presented when the EICAS displays are lost clue
to failure of both display units or both computers.
The test switch has three positions, and is spring-loaded to a centre
off position. It is screwdriver-operated and when turned to the left or
right, it changes over power supply units within the indicator to
ensure that they each provide power for the displays. The test can be
performed with the display control switch in any position.
Maintenance control panel
This panel is for use by maintenance engineers for the purpose of
displaying maintenance data stored in system computer memories
during flight or ground operations. The layout of the panel and the
principal functions of each of the controls are shown in Fig. 16.8.
The five display select switches are of the momen~ry-push type,
and as each one is activated, a corresponding maintenance display
page appears on the lower display unit screen. The pages are listed
together with two example displays in Fig. 16.9. The upper display
unit displays data in the 'compact' format (see Fig. 16.6) with the
message 'PARKING BRAKE' in the top left of the screen.
System failures which have occurred in flight and have been
Figure 16.8 Maintenance
control panel.
Environmental control
systems and maintenance
message ·formats
Selects data from
auto or manUat event
in memory
Electrical and hydraulic
systems format
\_
r::--;::====:t=====::=:;:z=====;==t::-,
Configuration
and maintenance
control/display
·panel
Engine exceedances
BITE test switch
for self-test
routine
Records real-lime
data currently
displayed (in
manual event)
Erases stored data
currently displayed
Figure 16. 9 Examples of
maintenance mode displays.
(C)
(C)
,. ,. . \
. " .....
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I.W'1I
,. '" )"
''"'"
(C)
r•t•uut
(C)
o.n
•lUtl/1
-.u-.CJWT
Mf"1U
(C)
ri
_.,._
hU IUW
0.60
o.oo
...
'",o ,,'
.
"
tt•not
o.n o.ao
IJ;!!J9NOl
0.6.l 0.11
Auto e¥ent message (W)
0.7J
o.n
....'"'
....
t~ttf111c,e1tt
"(J11
Att;
OU•G~U
ttn
ftOlfJlO•
l',:Jl'IAlfll•Ulll
au.r( t00t.111G
uc.1.s ciu,,u,
lt.U:111 lUC
ft.T 11:(C Off
CAPT PJTOT _.,0,y
(C)
'"'°"'
IN.lilt
!!:;
0,7t
120
o.oo
,. ''""" "',,....'"
" '"" "
" "
0
••
"
.... ' "
1).12
JUO
$0
:::
0,00
0
0
----rT(C)
P•Gl 1
••• NfU
(C)
. ,:,
o.u
*""'-l •
J.l10
o.n.,
!140
'1
115
Auto event message (W)
Colours·
C Cyan
All other readouts in white
automatically recorded ('auto event') in computer memory, as also
data entered as 'manual event', can be retrieved for display by means
of the 'event record' switch on the panel.
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.
When the switch is momentar.ly pressed, a complete test routine of
the system, including interfa,e and all signal-processing circuits, and
power supplies, is automatically performed. For this purpose an
initial test pattern is displayed oa both display units. with a message
in white to indicate the system being tested, i.e. 'L or R EICAS'
depending on the setting of the selector switch on the display select
panel. During the test, the master caution and warning lights and
aural devices are activated, and the standby engine indicator is turned
on if its display control switch is at 'AUTO'.
The message 'TEST IN PROGRESS' appears at the top left of
display unit screens and remains in view while testing is in progress.
On satisfactory completion of the test, the message 'TEST OK' will
appear. If a computer or display unit failure has occurred, the
message 'TEST FAIL' will appear followed by messages indicating
which of the units has failed.
A test may be terminated by pressing the 'TEST' switch a second
time or, if it is safe to do so, by releasing an aircraft's parking
brake. The display units revert to their normal primary and secondary
information displays.
ECAM
The units comprising this system, and as originally developed for the
Airbus A310, are shown in the functional diagram of Fig. 16.10. As
far as the processing and display of information are concerned, it
differs significantly from EICAS in that data relate essentially to the
primary systems of the aircraft, and are displayed in check-list and
pictorial or synoptic format. Engine operating data are displayed by
Figure 16./0 ECAM system
functional diagram.
,
L.H. display unit
R.H. display unit
~
~
Display
driving
====:> Data bus
387
conventional types of instruments as noted in the introduction to this
chapter. Other differences relate to display locations and selection of
system operating modes.
I Display uni~
~ units are mounted side-by-side; the left-hand unit is dedicated
to information on the status of system~, warnings and correctiye
a£tion in a sequencerl check-list form'll, ~ e the right-hand unit is
de_dicated to associa•ed information in pictorial or synoptic forma~
-
Display mode!i
....
1Ihere are four display mode;1, hree of which are automatically
~ected and referred to as tflight phase-re ate , a v1sory (mode and
status), and failure-related mode!i. ~e fourth mode ~ a l and
permits the selection of diagrams related to any one of 12 of the
aircraft's systems for routine checking, and also the selection of
status messages provided no warnings have been 'triggered' for
display. !!he selections, are made by means o{_ illuminated push.:-b.11.!!~
~witches on the system control panel.,
Iµ normal operation the automatic flight 12b.ase-related m~js
~ a11d in this case the displays are appropriate to the current
phase of aircraft operation, i.e~ pre-flight, take-off, climb, cruise,
descent, approach. and after landin . An exariipleofa pre-flight
phase is shown in Fig. l . I; the left-hand dispiay uniidispla)::s an
advi~ memo mode, afidthe rj.gbt-band unit display~gram of
the a i r c r a ~ o o r s , ~i.!!g_()!__~ll,e e~cape slides
deployment system.
·
The failure-related mode takes precedence ~)Ver the other two
Figure 16. I I Pre-night phaserell!!_ed mode d i ~
!MEMO!
APU RUNNING
NO SMOKING ON
SEAT BELTS ON
PARKING BRAKE ON
,.____
DOOR
CA BIN----;11
F WO COMPT
CA 8 I N--•-•M111
L.H. d1Splay unit
0
ARM
R.H display unit
Examples: Doors locked: Door symbols green and name of door white
Doors unlocked: Door symbols and name of door ambar
Figure 16.12 Failure-related
mode display. ,.__ _ _-
(C)
~
~
TE MP HOT
- FANS ••••••••••••••••••.• O.N
-DELAY T.O. FOR COOL
.---
ROLL__.
---.
SPLR
.._ROLL
---==--SPD BR~---===---
L.H. Display un,t
RH Display unit
Colours: A Amber
C Cyan
G Green
Remainder of display white
modei._and the man~al ~ode. An example of a display associat;d
with this mode is s own in Fig. 16.12. In this case, while "Gxying
out for take-off,
........____.,__ the temperature of the brake unit on the rear - right
wheel of the left rn.ain_lanili!l.&_g~ar bogie has become exce~~ive. A
diagram of the wheel brake system is imfllediately displayed on the
left-hanl11nh-- displays
right-hand
display uni~ simultaneously-the
.___________
--··-----------·
corrective
action
to
be
taken
by
the
flight
crew.
In addition, an aural
-------··--'"-"•--·--•o--•-•••-·~.-,-• •• --•
-~---warning is sounded, and a light (placarded 'L/G WHEEL') on a
central warning light display panel is illuminated. As the corrective
action is carried out, the instructions on the left-hand display are
replaced by a message in white confirming
result of the action.
The diagram on the right-hand display unit is appropriately
'redrawn'.
In the example considered, the warning relates to a single system,
and by convention such warnings are signified by underlining the
system title displayed. In cases where a failure can affect other subsystems, the title of the sub-system is shown 'boxed', as for instance
in the display shown in Fig. 16.13. Warnings and the associated
lights are cleared by means of 'CLEAR' push-button switches on
either the ECAM control panel or a warning light display panel.
Status messages, which are also displayed on the left-hand display
unit, provide the flight crew with an operational summary of the
aircraft's condition, possible downgrading of autoland capability, and
as far as possible, indications of the aircraft status following all
failures except those that do not affect the flight. The contents of an
example display are shown in Fig. 16.14.
~
t1f
.,
Figure I 6. I 3 Display of
failure affecting a sub-system.
'
Y E,L LOW P U MP L O PRESS ... 0 FF
iYE.LLOWSYS LOPR!
Colours:
Figure I 6. I 4 Example of
(C)
A Amber
C Cyan
(G)
r::::======+========::::::-,
status display.
1ST A TUS]
limitations {
!C)
Systems/functions {
lost (A)
Information
(G)
LANO 3 I NOP
Autoland
capability
(A)
e..!!..Q£. SINGLE ENG OPER
PROC FOR APPR HYO SYS LO PR
PROC FOR APPR: I NCR LOG DIST
HYO BLUE SYS I NOP
GEN 1 I NOP
SPLR PART I ALLY ,I NOP
SLATS SLOW
Colours
A Amber
C Cyan
Green
G
tControl panel
l
The layout of this panel is shown in Fig. 16.15; all switches, with
the exception of those for display control, are of the push-button,
illuminated caption type.
l. i SGU selector switches) Control the respective symbol generator
units, and the lights are off in normal operation of the system.
tThe 'FAULT' caption is illuminated ambeg if a failure is detected
by an SGU's internal self-test circuit. Rel~asing a~witch i~lates
the, corresponding SGU, and causes the 'FAULT' G.aptiruLto
extinguisl.L and the 'OFF' caption ta illuminate white.
2. Synoptic display switches Permit individual selection of synoptic
diagrams corresponding to each of 12 systems, and illuminate
white when pressed. A display is automatically cancelled
whenever a warning or advisory occurs.
Figure 16. 15 Control panel.
DISPLAY "ON' AND
BRIGHTNESS CONTROL
MESSAGE
CLEARANCE
SWITCH
STATUS
MESSAGE
SWITCH
RECALL
SWITCH
SYSTEM SYNOPTIC
DISPLAY SWITCHES
3. tCLR switch., Light illuminated white whenever a warning or
status ,message is displayed on the left-hand_display unit. Pressed
to clear messages •
4. STS switch .l Permits manual selection of an aircraft status
£,nessage if no warning is displayed; il,luminated wlute. Pr~ng
the switch also causes t~ CLR switch to illuminate. A status
message is suppressed if a warning occurs or if the CLR switch
is pressed.
5. RCL switch En~les previously cleared warning messages tq be
recalled provided the failure conditions which initiated them still
exist~ Pressing the switch also causes the CLR switch light to
illuminate. If a failure no longer exists the message 'NO
WARNING PRESENT' is displayed on the left-hand display unit.
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,
391
BITE
SWITCH
Figure 16.16 Maintenance
panel.
DISPLAYS
SWITCH
FAULT
ANNUNCIATOR
LIGHTS
INPUTS
TEST
SWITCH
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.
Manual self-test checks for inputs and displays are carried out from
a maintenance panel shown in Fig. 16.16. When the 'INPUTS'
switch is pressed, a 'TEST' caption is illuminated white, and most of
the inputs to each computer are checked for continuity. Any incorrect
inputs appear in coded form on the left-hand display unit. The righthand display unit presents a list of defective parameters at the
system's data analog converter. The diagrams of systems app<;:ar on
the right-hand display unit with the caption 'TEST' beside the system
title, as each corresponding push-button switch is pressed. Caiibrated
outputs from the data analog converter are also displayed. Any
defective parameters are identified by a flag display.
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. The 'LOAD' caption is
illuminated each time a failure is memorized in the relevant test
circuits of the SGUs.
The annunciator lights on the maintenance panel illuminate white
simultaneously with a failure warning system light on the central
warning light display panel when a corresponding computer fails.
The 'INHIB OVRD' switch enables inhibited warnings to be
displayed.
17 Flight management
systems
Computerized systems designed in various forms to carry out
performance/advisory or comprehensive flight management functions
are an essential feature of a number of types of public transport
category aircraft, their development having stemmed from the need to
~sure the most efficient use of fuel. to reduce flight crew workload,
and to reduce operating costs overall)
In..performing an advisory function a system merely advises the
flight crew of the optimum settings of various control parameters,
such as engine pressure ratio (EPR) and climb rate under varying
flight conditions, in order to achieve the most economical use of the
available fuel. Such systems require adjustments of controls on the
part of the flight crew if they are to be utilized to maximum
advantage.
,A system performing a combined function is one in which the
computer and display units are interfaced with a thrust or power
mana ement control system and an automatic fli ht control s stem.
Thus, in isolating the flight crew rom the control loop, an integ.-ated
automatic flight management system (FMS) is formed to provide
greater precision of engine power and flight path control.
Early forms of such sy~tems, whether purely advisory or of
combined function, were limited to supervising control parameters
'![!'ecting the vertical flight pafu. In order to ensure maximum fuel
economy it is, however, also necessary to integrate this optimized
flight path management with the lateral flight path; (in other words, a
l system must also be provided with a navigation capabill!y_. This
requires interfacing the computer with such navigation systems as
Doppler; INS/IRS, DME and VOR.JThe inputs from these systems
c perrnit continuous monitoring of an Jircraft's track in relation to a
flight plan which may be pre-stored in the computer memory and an
immediate identification of deviations. Furthermore, it allows flight
plan variants to be constructed and evaluated. It is thus apparent that
by combining these inputs with those controlling the vertical flight
path parameters mentioned earlier, n FMS can inte rate the
\functions of nav1 ation
r om:ance mana ement, 1 t lanning an
·~-dimensional guidance and control along a pre-planned flight
l?'aIntlf.Jaddition to changing data inputs from such systems as those
393
mentioned above, an FMS also requires data bases for storing bulk
navigation data, and the characteristics of an aircraft and its engines,
in order that the system will operate in a full three-dimensional
capacity. The navigation data base is capable of storing the necessary
flight environmental data associated with a typical airline's ..:ntire
route structure, including pertinent navigation aids and waypoints,
airports and runways, published terminal area procedures, etc. The
memory bank also contains flight profile data for a variety of
situation modes such as take-off, climb, cruise, descent, holding, goaround and 'engine-out'. The cruise mode is also sub-divided into
sub-mode variants such as economy, long-range, manual and thrustlimited. The integration of all the foregoing data, plus other variable
inputs such as wind speeds and air traffic control clearances, permit
the automatic generation or modification of flight plans to meet the
needs of any specific flight operation.
Typical systems
Performance data computer system
This system provides advisory data in alphanumeric format on a CRT
display, in addition,to the positioning of target command 'bugs' on a
Macfi7airspeed indicator and EPR indicators, such indicators operating
on electrical servomechanism principles. Provision is also made for
interfacing the system with autothmttle and a.u!Qmatic flight control
·systems. A schematic diagram of the system, which consists primarily
of a control and display unit, computer and mode annunciator, is
shown in Fig. 17 .1.
Abbreviations are extensively used for the display of data by the
control and display units of this and other flight management
computer systems, and these abbreviations/acronyms and their
definitions are given in Appendix 3.
Control and display unit (CDV)
The CDU provides the major input link to the system and allows__Qie
.flight crew to make inputs to obtain EPR and airspeed displays and
.can also be used for obtaining decision-making data in relation to an
aircraft's flight profile. The CRT has a 2 in x 3 .iu_§creen and
enables data to be displayed&reral3 (column) ~ w ) matriis.J
The selection of EPR and airspeed data for various phases of flight is
accomplished by a flight mode select switch, the modes and
associated displays being as follows:
TO
CLB
394
Take-off EPR limits for the outside air temperature entered
by the flight crew
EPR and speeds for the desired climb profile: best economy,
maximum climb rate, or crew-selected speeds
Figure 17. I Performance data
computer system.
Max airspeed'-pointer
EPR pointer
(black & yellow)
Mach readout
Mach failure
flag
"'
Aclual EPR
ElCternat index
marker (lyp1cat)
Speed reference
marker
EPR command bug
setting
Airspeed cursor
INOP flag
VMO flag
Airspeed
failure flag
Airspeed
pointer
EPA set knob
Mode Annunciator
,\irspeed
Speed reference knob
manual sel (J>ull out)
perf. dala computer (push 1n)
readout
EPR Indicator
Mach Airspeed lnd1ca1or
Page No
Failure lights
L-----------r-:-;:r:~~I Light
lest
switch
Power i;upply
115 -V 400Hz
----->-l
Fhght index
number set
switches
Air data: altitude & airspeed
DME: distance to go----t""I
Targel EPA
:;~:ne air bleed----"'*R
Aclual EPA
S\Jmmation unit
Input/output data
total fuel quantity
From tank
sensing probes
CRT display
CRT
bnghtness
control
ECON
IAS selector
switch
Clear button
Enter button
Selecl bullon
!AS
MACH
WI ND
E P ft
-+
2.
I
Keyboard
& mode control
Computer
Total air temperature
Control & Display Un,t
CRZ
EPR and speeds for the desired cruise schedule: best
economy, long-range cruise, or crew-selected speeds
DES
Descent speed, time and distance for best economy
HOLD EPR, speed and endurance for holding
CON
Maximum continuous EPR limits for existing altitude,
temperature and speed
GA
Go-around EPR limit for existing altitude, temperature and
speed
The standby (STBY) position of the select switch is used for data
entry and for an automatic check.-out of the system.
The function of the 'ENGAGE' key is to couple the target
command 'bugs' of the Mach/airspeed indicator and EPR indicators
to computer command signals which drive the bugs to indicate the
speed and BPR values corresponding to those displayed on the CRT
screen. If the data is verified by the computer to be valid, engageable
and different from the data presently engaged, the engage key
illuminates and is extinguished after engagement talces place; at the
same time the appropriate light of the mode annunicator is
illuminated.
The·key marked 'TURB' is for use only in cruise and when
395
turbulent flight conditions are to be encountered. When pressed it
causes the CRT to display the appropriate turbulence penetration data,
i.e. airspeed in knots (also Mach number at high altitudes), pitch
attitude and the N I percentage rpm. In the turbulence mode, the
target command speed and EPR 'bugs' engage automatically. This
mode is disengaged-by pressing the key a second time or else
engaging another flight mode.
In order that- the flight crew may load keyboard-selected data into
the system, three push-button switches are provided above the.
keyboard for SELecting, CLeaRing and ENTERing data. In
connection with the selection and entering of data, question marks
and two symbols are displayed at the right-hand end of a data line: a
caret ( <) and an asterisk (*). The caret signifies that the computer is
ready to accept data, while the asterisk signifies that the data next to
it may be entered or changed if necessary.
The keyboard primarily serves a dual function in that it (1) permits
the flight crew to enter pure numeric data into the computer and (2)
permits desired performance function data to be called up from the
computer for display. The data appropriate to the keys is given in
Table 17. l and is displayed in the form of pages, each page being
numbered in the top right-hand comer. For example, the page shown
on the CDU in Fig. 17 .1 is page l of a set of four relating to
'economy fuel' in the cruise mode. In order to call up each of the
-
-
.
remaining pages the PAGE key is successively pressed. Similarly, the
PAGE key permits cycling of the pages in reverse order. When a
flight mode or performance function is first selected, the first page of
a set is always automatically displayed.
The RCL key is used whenever a performance function is being
displayed and if it is required to recall a display corresponding to a
selected flight mode.
The two switches in the upper right-hand corner of the CDU are
associated with autothrottle system operation. When the A/T
annunciator switch is pressed, an internal light is illuminated to
indicate connection of the auto-throttle system, and at the same time
an 'EPR'light in the mode annunicator is illuminated. The PDCS then
adjusts the throttles to track the EPR target values displayed on the
CDU and by th~ command bugs of the associated EPR indicators. In
order for the autothrottle system to adjust engine power in relation to
indicated airspeed, the second switch 'IAS SEL Annunicator' is
operated; the system then drives the throttles so as to track the speed
target values displayed on the CDU and by the command bug of the
Mach/airspeed indicator.
Computer
The computer is of the hybrid type, and the inputs, outputs and unit
interfaces are as .shown in Fig. 17. l. Program storage is by means of
396
Table 17. I Performance function data
Key
Data pages displayed
Outside air temperature, destination airport elevation, reserve and alternate fuel
and zero fuel weight for the intended flight plan
Flight level intercept data for use in solving time and distance problems in climb
and descent
Economy cruise and. long-range cruise speeds at appropriate flight level
Present ground speed computed from a known true airspeed (TAS) and wind
component
Total flight endurance as well as distan=eitime solutions to fuel reserve at any
flight level
Fuel for engaged cruise speed (economy, long-range and manually entered
speeds). Used only in the CRZ and CON modes
Ambient temps (TAT & SAT) TAS, and temperature deviation from ISA
Reference landing speeds for various flap setting,. based on aircraft's correct
gross weight
Optimum initial cruise flight level for inserted trip distances
Data in respect of automatically computed and manually inserted wind
components
D
Negative value data and also 'test pages' while in 'STBY' mode
a PROM and an additional non-volatile memory for retaining all
entered data during any interruption of the power supply. Built-in test
equipment circ 4its and software operate continuously to check all
critical circuits of the system. The fuel summation unit, which is a
397
component of the PDCS, develops an ac voltage signal that is
proportional to the total fuel on board the aircraft, the signal being a
combination of those produced by the fuel quantity indicating system
sensing probes which are located in the fuel tanks.
Failure lights on the front of the computer indicate whether a fault
is in the computer, CDU or input signals. The INDEX NUMBER
switches, which are of the rotary type, are used for programming a
flight index number from O to 200 into the computer so that
maximum economy flight modes are modified according to timerelated costs compared to fuel costs. The switches are guarded. to
eliminate the possibility of inadvertent changing of the index number.
Mode annunciator
This unit indicates the flight mode driving the command bugs of the
EPR and Mach/airspeed indic~tors. The legend appropriate to a flight
mode is illuminated when the mode is selected, and the 'ENGAGE'
or 'TURB' button switches are depressed.
T)pical display
The number of data pages associated with all the performance
functions and flight modes it is possible to select on the CDU is quite
considerable, and limitations on space do not permit a full description
of each to be given here. However, the fundamentals of presentation
and data entry methods may be understood from the example given in
Fig. 17.2.
Figure 17.2 Examples of
PDCS displays.
1a1
FU E L E CON
0 I S T_ NM
RSV+ALT
FOD
WI N 0
F U E L WT
r FU E L
0 I S T
E CON
NM
RSV+ALT
fbi
FOD
WI N 0
F U E L WT
rFUEL
0 I ST
E CON
N M
RSV+ALT
FO 0
WI N 0
IC)
398
FUEL
WT
1- 4
? ? ? ?
<
1 0. 2
- 10 •
3 4. 7
1 -4
1 5 00<
10 .2
1 0 .7
- 1 O•
3 4 •7
'
1- 4
1 5 00•
1 0. 2
10 • 7
- 1 s<
3 4. 7
In this case, the first page of the fuel performance function for the
maximum economy speed schedule is displayed, and the question
marks on the second line (Fig. 17 .2(a)) indicate that the computer is
asking for the distance to go from the present position of the aircraft.
The caret indicates the computer's readiness to accept the
information. The CLR button switch is then pressed and this causes
erasure of the question marks and the caret to flash on and off to
indicate that the numbers appropriate to the distance should be 'keyed
in' to the display from the keyboard - 1500 nautical miles in this
case (Fig. 17.2(b)). When this has been done the ENTER button
switch is then pressed, following which the computer goes through an
input validity check routine. If the input is valid, the caret will stop
flashing, thereby advising that the data has been accepted. An
INV AUD message is displayed if the input exceeds any limitation.
The computer also computes the fuel over destination (FOD) value
and causes it to be displayed. It the computer requires more data on
another line the caret drops to that line automatically. The asterisk
against the wind component of - 10 knots signifies that any change
to its value may be effected. If. for example, the wind speed has
increased to 15 knots (the minus sign indicates a headwind) then it
may be entered by first pressing the SEL key which causes the caret
and the asterisk to exchange positions (Fig. 17.2(c)). The CLR key is
then pressed, the new wind speed value is keyect' in and is finally
displayed by pressing the ENTER button switch.
) Flight management system
J
.A fli ht management system (FMS) is currently the most advanced of
systems, providing !lS it does full integrat10n o all the functions
referred to earlier (see page 393) and which are necessary to fly
O£timized flight profiles either in manual or fully automatic control
modes.~ The system is a union of autonomous and generally
asynchronous units interconnected by a network of ARINC data
busses to satisfy specific functional needs. In many cases redundant
units are present to meet requirements for functional availability,
flight safety or aircraft coverage.
Figure 17 .3 illustrates schematically the computing units which are
typically a formal part of an FMS and also how by means of the. data
busses they communicate with the principal elements of the system,
namely the flight management computer (FMC) and its associated,
control and display unit (CDU). By virtue of their communication
link these two units are together designated as a flight management
computer system and this provides the primary interface between the
flight crew and the aircraft. Inputs from other interfacing systems and
sensors are also transmitted to the data busses, but for reasons of
clarity have been omitted from Fig. 17.3.
399
Figure / 7, 3
interfacing.
Fhght crew
FMS and data
.
Controls &
,nd1caro,s
Control panels
EFIS
AFCS MODE
IRS MODE
ILS
EFIS
HSI
Ma,n1enance
ATC
ROM!
VOR,OME
AOF
SPWS
ADI
control &
WXA
AOC
;RS
TMC
symbol
generators
iransooncter
Data
Discrete inpu1s
lrom 1,rcraft
base
systems
•oad:ng
NOTE· Reier to App<,ndix 3 for defir\itions of abbreviations
\-ttf) ~
Fuel
Quan111y
&
fuel flow
dala
Cort1ro1
WXA
surface
S,e1VQ5
Antennae
ffiight management computer
~ ~ypically, a computer incorporates three different types of memo~: a
bubble memory·for holding the bulk navigation and aircraft
-----performance characteristics data bank; a C-MOS RAM for holding
~cific navigation and performance data, and the activ~ and
· gecondary flight plan, all 'down-loaded' from the bubble memory;
'and a UV-PROM for the operation program. which may be
reprogrammed at card 'tevel.
The data base which is used for all computations contains
numberous types of records in memory and these are given in Table
l7. 2. All the data are unique to each aircraft operator. depending on
the routes flown, and are initially programmed on magnetic tape. The
tape cartridge is inserted into a portable data base loader unit which,
after connection to the computer, is operated so as to transfer the
data to the bubble memory. Any subsequent changes in navigation
aids and procedures, and route structure changes, are also
incorporated in the data base by means of the data loader, in
accordance with a specified time schedl!!e, e . ~
Control and display unit
The CDU of one example of an FMCS which is currently in use is
shown in Fig. 17 .4. It is basically similar to that adopted for the
Table l7 .2 Records in data base memory
Record
How identified and defined
Radio-nav aids:
VOR, DME,
VORTAC, ILS,
TACAN
Waypoints
Identifier ICAO region, latitude and longitude, frequency, magnetic·
declination, class (VOR, DME, etc.), company defined figure of
merit,* elevation for DME, ILS category, localizer bearing
En-route airways
Airports
Runways
Airport procedures
Company routes
Each waypoint defined by its !CAO region, identifier, type (en-route,
tem1inal), latitude and longitude
Identified by route identifier, sequence number, outbound magnetic
course
Each identified by ICAO four-letter code, latitude and longitude,
elevation, alternate airports
Each identified by !CAO identifier, number, length, heading. threshold
latitude and longitude, final approach fix identifier, threshold
displacement
Each identified by its ICAO code, type (SID, STAR, profile descent,
ILS, RNA V), runway number/transition, path and termination code
Origin airport, destination airport, route number, via code (SID. airway,
direct, STAR, profile descent, approach), via identifier (SID, name,
airway identifier, etc.), cruise altitude, cost index
* The figure of merit is a number assigned to each navigational aid to indicate the maximum
distance at which it can be tuned.
Figure 17. 4 FMS control and
display unit.
keys
~ iDSPL! iOFST! iMSG!
Annunciators
Function keys
Alpha keys
performance data computing system described earlier, but in keeping
with the role to be played by an FMCS the unit is much· more
sophisticated in respect of data selection and corresponding displays.
The operation of the keyboard function keys is summarized in
Table 17.3. As in the case of a PDCS display unit, annunciators are
also provided but form an integral part of the unit.
401
Table 17 .3 Summary of operation of keyboard function key,
Key
Selection and data displayed
Returns display
position
NEXT
PHASE
EJ
[:]
10
show the active navigation leg page. i.e. the aircraft's present
Changes a navigation leg display to the beginning of the next phase of the flight
plan
Selection of performance pages
Permits direct entry of revisions to flight plan from present position to any
waypoint
Selection of fuel pages
Displays the navigation legs page which includes the next airport along the
current flight plan
Selection of headings to be flown automatically via the FMS
Displays data index pages relevant to: lateral, vertical. performance. key
waypoints, sensors. maintenance, navigation. aircraft configuration. history
To check or up-date aircraft's position
Selects 'START DATA REQUIRED' pages for flight crew to initiate and
construct flight plans
Presents performance data pages relating to engine out operation
Selects secondary flight plan facility for re-clearance or return-leg planning
To promote a temporary plan to active status. The bar illuminates when the
FMCS has enough data to create an active plan, and remains illuminated until
~he temporary plan has been executed. or cancelled by pressing PPOS"
Informs computer that any message displayed on CDU has been acknowledged
and the message will either be stored or erased
To delete incorrect scratch pad entries
402
While on the ground, the flight crew can construct a detailed flight
plan by inserting data in selected data pages and, in conjunction with
the comprehensive data base stored in the computer, the plan is
raised to active control status. In flight, the system receives data from
the relevant aircraft sensors and radio-navigational aids, and then as
the flight proceeds it presents the flight plan in a progressive 'scroll'
form. Pitch, roll and thrust demands are also computed, and in
communicating these to an AFCS and a TMS accurate control of the
flight profile and maximum fuel economy can be provided. Various
pages of data can be selected for review by the flight crew at any
stage of a flight, and predictions concerning its future phases can be
assessed by inserting detailed revisions, the future implications of
which are computed and displayed. In addition to its own display
unit, the FMCS also has the unique capability of presenting
navigation data in 'changing map' form on the display unit of an
electronic flight instrument system (see also Chapter 12).
The flat-faced CRT of the display unit gives a dual character size
presentation _with 24 stroke-written characters per 14 line Small-size
characters sigm
ata wit e au t va ues, or computer-predicted
values which can be changed by the flight crew when the data is
bei~ supplied by the computer; large-size characters signi(y 0·1ta
entered by the flight crew.
Data pages and flight plan construction
Many pages of data can be accessed and displayed, and space docs
not permit them all to be shown here. However, some iwlication of
character presentation and method of entering data in ger,eral may be
understood by considering the example of the display shown in Fig.
17 .5, which is used to initiate the construction of a flight plan when
the 'START' key is pressed.
The data lines adjacent to the line keys constitute the 'operational
area' of the display and can be accessed by line key selection for
entry or revision of flight plan information or for selection of
displayed options. The first 12 character spaces on the bottom line
(line 14) are used as a scratch pad for information entered by the
flight crew via the alphanumeric keys. The next IO spaces are
reserved for FMS messages to the crew and the last two spaces are
reserved for scroll cues signified by upward- and downward-pointing
arrows. When the appropriate sc:oll key on the keyboard is pressed
the display moves up or down for the purpose of reviewing the
display.
Arrows which appear against characters of a display indicate flight
crew options, which may be the choice of display data or may result
in some functional activity of the system such as aligning the inertial
reference system {IRS) as indicated in Fig. 17 .5. The choice of
option is signified by pressing the line key adjacent to an arrow; this
403
Figure 17.5 Example of a
page display.
Information required
on page displayed
Information required on
next page
r:::======:!!;::::::;z:==:::::::,
Page numbers
START DATA RORDl-TMPY 1/2
FROM-TO
FLT
NO
CRZ
ALT
L SGG LGA T
ALTERNATE
?
Fl 290
LGTS
GATE
AV
W/V
1-5
LAT
000/0
HISTORY
w,v-
N46°14·3
LO NG
E006°06·6
AISA/TROPO
0/ 36000
COST
-ALIGN IRS
Scratch pad/messages (line 14)
INDEX
20
H
Scroll cues
results in a change to the content of an existing page and erasure of
the arrow.
Requests for data entry are indicated by question marks. For
example, in Fig. 17.5 the request is for the flight number, and in
response this is first entered into the scratch pad by using the
alphanumeric keys and then pressing the line key adjacent to the
question mark. If the format of the entry is correct, the flight number
appears in large-size characters on the appropriate line and the
scratch pad clears. If the format is incorrect, the word ERROR
appears in the message space, the MSG annunciator illuminates and
the incorrect nu ..1ber remains in the scratch pad. After an incorrect
entry has been attempted the ERROR legend and MSG annunication
are acknowleged by pressing the MSG key. The CLEAR key is then
pressed to delete the entire scratch pad entry.
It is possible to change any data by over~writing with new values
from a scratch pad entry. If, for example, it is required to change the
displayed data base cruise altitude, the new value is 'keyed' into the
scratch pad and then transferred to the altitude line by pressing the
adjacent line key. Similarly, if the airport terminal gate from which
an aircraft is to depart is changed, the new gate number may be
over-written in the display. The FMC will automatically enter the
revised latitude and longitude values for the new gate to which the
IRS must be realigned.
The TMPY legend on the title line of start data pages signifies that
the data being entered by the flight crew relates to the construction of
a temporary flight plan.
Flight plan construction and associated changes of data pages are
essentially in two sections. First, the navigation section involves
insertion and acceptance of route number. airport codes, cruise
altitude, latitude and longitude data and IRS alignment. When this has
been accepted by the computer the bar in the EXEC key is
illuminated and this section of the plan is executed by pressing the
key. A page is then displayed requesting data needed to construct the
second section of the plan which relates to fuel on board, reserve
fuel, trip fuel and time, weights and centre of gravity position. When
this has been computed, a 'START COMPLETE' page is displayed
to show the relevant fuel, weight data and time components, and the
EXEC key is again pressed so that the whole of the flight plan is
raised to 'active' status.
Performance and in-flight displays
The next step is to initiate the display of data pages relevant to the
performance section of the flight plan, and this is done by pressing
the PERF key. These pages are concerned with management of
engine thrust, pitch attitude and other alternative modes of
performance control limited to the various phases of flight. The
normal mode is 'economy' by which the most economical climb,
cruise and descent speeds are computed. Each flight plan has its own
performance scroll of pages related to those of the navigation legs
scroll. The first page is the take-off performance page in which the
flight crew must enter the values of the take-off speeds V1 and V R·
After take-off and during climb at the appropriate climb speed, the
CLB annunciator is illuminated and the take-off performance page is
automatically replaced by a climb performance page. At the top of
the climb, the CRZ annunciator is illuminated and a cruise
performance page is then presented to display continuously updated
information related to optimum performance and destination arrival
fuel and time. During cruise descent a descent performance page is
displayed, but the point at which it is presented depends on the point
at which the descent is commenced. For example, if the descent is
commenced prior to the planned point in the vertical profile (by
authorizing it via the AFCS control panel) the cruise performance
page is first replaced by a cruise descent page and the aircraft
descends at the selected vertical speed. When the aircraft captures the
planned descent profile the display then changes to the descent
performance page and the DES annunciator is illuminated. If the
descent is commenced at the planned profile point, the cruise
performance page is replaced directly by the descent performance
page and the CR DES annunciator is illuminated. During approach,
and when the leading edge slats are extended, an approach
performance page replaces the descent performance· page and the
APPR annunciator is then illuminated.
If any revisions are made while in flight, the TMPY legend will
also appear on the appropriate data page being displayed to indicate
that a complete temporary flight plan is generated. The aircraft
405
continues under the control of the active flight plan until the
temporary plan is raised to active status by pressing the EXEC key.
If it is to be aborted, the P POS key is pressed.
Data index
An index of data is contained on a page which can be called up for
display by pressing the data key. When the line keys adjacent to the
titles are pressed, further pages are presented to display the
information noted in Fig. 17.6.
System configuration
Two FMC systems are installed in an aircraft, each having its own
CDU situated on the centre consol and each controlling its associated
automatic flight control system, autothrottle system and radionavigational aids. The basic configuration is shown in Fig. 17. 7. In
normal operating conditions, both computers operate together and
share and compare each other's information, i.e. they 'cross talk' by
means of an interconnecting data bus. The pilots can operate their
displays independently for review or revision purposes without
disturbance to the active flight plan and without affecting the other
CDU commands. Typically, a working arrangement would be for the
performance pages to be displayed on one pilot's CDU, while the
other pilot's CDU would display navigation legs.
When a temporary flight plan is created in one system the other
system (referred to as the 'offside' system) has no access to this plan,
but it can review the active plan in the normal way. In addition, the
offside system is inhibited from creating a temporary flight plan until
the previous temporary state is cancelled or raised to active.
Each computer has its own VOR/DME receiver and determines the
frequencies it requires for its own purposes. No interconnection
between systems is possible except when a lateral revision is effected
at the present position of the aircraft.
In the event of failure of one computer, each pilot has the means
whereby he can select his own CDU into the other system.
Figure /7.6 Data index page.
Information on present
lateral situation
Information on present
DATA
vertical situation
Information on present
performance situation
-
INDEX
VERTICAL
-------
Display time and distances
related to en·route way po,nts.
atso estimated fuel on board
----t- PERFORM
KEY WPTS
SENSORS
-
-
Information from system sensors
related to fuel quantity. flow.
AOC. IRS Radio nav .. clocks.
computers of AFCS. TM. FMS.
EFIS control panel
Information required to facilitate
fault diagnosis and test functio(ls
in association with maintenance manual
Displays tuned navigation aids, and priority
listing of aids which can b-e used for navigation
from aircraft's present position
NAV
CONFIG HISTORY
STORE
-
Aircraft and engine type, FMS programme
identification number, dates navigation data
valid
Data for recording in history file until
engine shut-down Sl'ld retention until
aircraft again airborne
Information (external to data base) on
waypoints. navaids, airports and routes
Figure 17. 7 FMS
configuration.
Captain'a
cou
Changeover
switch
Co-pilot's
CDU
Cha~r
switch
EFIS display &
auto control
outputs
EFIS display &
auto control
outputs
Inputs
Inputs
407
Tables
Table I Standard atmosphere
Altitude
Pressure
Temperature
Veloci1y of
sound
ft
-1.000
0
1,000
2,000
3,000
4,000
5,000
millibars
!050·41
1013.25
977.17
942.13
908.12
875. IO
843.07
in hg
31·019
29.921
28.856
27.821
26.817
25.842
24.896
lbf/in2
15·234
14.696
14.172
13.664
13.170
12.691
12.226
·c
+16·981
15.000
13.019
11.038
9.056
7.075
5.094
ft/sec
1119.9
1116.1
1112.3
1108.4
1104.5
1100.7
1096.7
6,000
7,000
8,000
9,000
I0,000
811.99
781.85
752.62
724.28
696.81
23.978
23.088
22.225
21.388
20.577
11.775
11.338
I0.913
10.502
10.104
3.113
1.132
-0.850
-2.831
-4.812
1092.8
1088.9
1085.0
1081.0
1077.0
11,000
12,000
13,000
14,000
15,000
670.20
644.41
691.43
595.24
571.82
19.791
I0.029
18.292
17.577
16.886
9.718
0.344
8.891
8.630
8.291
-6.793
-8.774
-10.756
-12.737
-14.718
1073.1
1069.1
1065.0
1061.0
1057.0
16,000
17,000
18,000
19,000
20,000
549.15
527.22
506.00
485.47
465.63
16.216
15.569
14.942
14.336
13.750
7.962
7.643
7.335
7.038
6.750
-16.699
-18.680
-20.662
-22.643
-24.624
1052.9
1048.8
1,044.7
1040.6
1036.5
21,000
22,000
23,000
24,000
25,000
446.45
427.91
4!0.00
392.71
376.01
13.184
12.636
12.107
11.597
11.104
6.472
6.203
5.943
5.692
5.450
-26.605
-28.686
-30.568
-32.549
-34.530
1032.4
1028.2
1024.0
1019.S
1015.6
26,000
27,000
28,000
29,000
30,000
359.89
344.33
329.32
314.85
300.89
10.628
10.168
9.725
9.298
8.885
5.216
4.991
4.773
4.563
4.361
-36.511
-38.492
-40.474
-42.455
-44.436
1011.4
1007.1
1002.9
998.6
994.3
31,000
32,000
33,000
34,000
35,000
287.45
274.49
262.01
249.99
238.42
8.488
8.106
7.737
7.382
7.041
4.166
3.978
3.797
3.622
3.455
-46.417
-48.398
-50.380
-52.361
-54.342
990.0
985.6
981.3
976.9
972.5
Table I cont'd
Altitude
Pressure
ft
36,000
37,000
38,000
39,000
40,000
millibars
227.29
216.63
206.46
196.77
187.54
in hg
6.7.12
6.397
6.097
5.811
5.538
lbf/in2
3.293
3.139
2.991
2.851
2.717
41,000
42,000
43,000
44,000
45,000
178.74
170.35
162.36
154.74
147.48
5.278
5.030
4.794
4.569
4.355
2.589
2.468
2.352
2.241
2.136
46,000
47,000
140.56
133.96
127.67
121.68
115.97
4.150
3.956
3.770
3.593
3.425
2.036
1.940
1.849
1.762
1.679
48,000
49,000
50,000
Temperature
Velocity of
sound
oc
-56.323
ft/sec
968.1
-56.500
967.7
J
Table 2 Mach no./airspeed relationship
Mach number
Speed of
sound
Heighr
/ft)
Abs
remp
oc
ft/sec
mph
0.3
0.4
0.5
288.0
278.1
268.2
258.3
248.4
238.5
228.6
218.7
216.5
761.6
748.4
734.9
721.2
703.3
691.1
678.5
663.7
660.3
228.5
224.5
220.5
216.4
212.2
207.9
203.5
199.1
198.1
304.6
299.4
294.0
288.5
282.9
277.2
271.4
265.5
264.1
380.8
374.2
367.5
360.6
353.7
346.5
339.2
331.8
330.2
0.7
0.8
0.9
533.1
523.9
514.4
504.8
495.1
485.2
474.9
464.6
462.2
609.3
598.7
587.9
577.0
565.8
554.5
542.8
531.0
528.2
685.4
673.6
661..4
649.1
636.6
623.8
610.6
·597.3
594.3
0.6
Airspeed in mph
0
5000
10000
15000
20000
25000
30000
35000
36000
457.0
449.0
440.9
432.7
424.4
415.9
407.1
398.2
396.2
Table 3 Temperature/resistance equivalents
Nickel sensing elements
•c
0
JO
20
30
40
50
60
70
80
'}()
Minus
Plus
100
200
100.0
152.2
218.8
95.4
104.7
158.2
226.4
90.8
109.5
164.4
234.3
86.4
114.4
170.6
242.3
82.1
119.4
176.9
250.6
77.8
124.6
183.4
259.0
73.6
129.9
190.1
267.7
69.5
135.3
197.0
65.4
140.8
204.1
61.4
146.4
211.4
409
Table 3 cont'd
Platinum sensing elements
oc
Minus
Plus
100
200
300
400
500
0
JO
20
30
40
50
60
70
80
90
110.0
150.3
189.4
227.3
264.1
299.5
105.9
114.l
154.2
193.2
231.0
267.6
303.0
JOl.8
118.1
!58.2
197 0
234.7
271.2
306.5
97.7
122.2
162.1
200.9
238.4
274.8
309.9
93.5
126.3
166.1
204.7
242.1
278.4
313.4
89.4
130.3
170.0
208.5
245.8
281.9
85.2
134.3
173.9
212.3
249.5
285.5
81.1
138.3
177.8
216.0
253.1
289.0
76.9
142.4
181.7
219.8
256.8
292.5
72.7
146.3
185.6
223.5
260.4
296.0
Table 4 Temperalure/millivolt equivalent of typical iron v constantan thermocouples
Cald junction at 0°C
oc
0
JO
20
30
40
50
60
70
80
90
JOO
0
100
200
300
400
500
600
700
800
0
5.02
10.28
15.62
20.93
26.28
31.75
37.77
44.97
0.48
5.54
10.80
16.16
21.51
26.32
32.32
38.39
0.96
6.06
11.33
16.70
22.04
27.36
32.90
39.01
1.46
6.59
11.86
17.24
22.57
27.90
33.48
39.63
1.96
7.12
12.33
17.79
23.10
28.44
34.06
40.25
2.46
7.65
12.93
18.33
23.62
28.98
34.66
40.87
2.97
8.18
13.47
18.87
24.15
29.52
35.28
41.49
3.48
8.71
14.01
19.40
24.68
30.07
35.90
42.11
3.99
9.24
14.55
19.93
25.21
23.62
36.53
42.73
4.50
9.76
15.08
20.45
25.74
31.18
37.15
43.35
5.02
10.28
15.62
20.98
26.28
31.75
37.77
43.97
Table 5 Temperature/millivolt equivalents of typical copper v constantan thermocouples
Cold jum·tion at 0°C
oc
0
JO
20
30
40
50
60
70
80
90
JOO
0
100
200
300
400
500
0
4.27
9.25
14.75
20.68
26.88
0.37
4.75
9.77
15.33
21.30
0.75
5.23
10.29
15.91
21.92
1.16
5.71
10.83
16.49
22.54
1.58
6.19
11.37
17.08
23.16
2.01
6.69
11.92
17.68
23.78
2.44
7.20
12.47
18.28
24.40
2.89
7.71
13.03
18.88
25.02
3.34
8.22
13.60
19.48
25.64
3.80
8.73
14.17
20.08
26.26
-~.27
9.25
14.75
20.68
26.88
Table 6 Temperature/millivolt equivalents of typical chrome! v alumel thermocouples
Cold junction at 0°C
410
oc
0
JO
20
30
40
50
60
70
80
90
/0()
0
100
200
300
400
0
4.10
8.13
12.21
16.39
0.40
4.51
8.53
12.62
16.82
0.80
4.92
8.93
13.04
17.24
1.20
5.33
9.34
13.45
17.66
1.61
5.73
9.74
13.87
18.08
2.02
6.13
10.15
14.29
18.50
2.43
6.53
10.56
14.71
18.93
2.85
6.93
10.97
15.13
19.36
3.26
7.33
11.38
15.55
19.78
3.68
7.73
11.80
15.97
20.21
4.10
8.13
12.21
16.39
20.64
Table 6 cont'd
Cold junction at
·c
0
IO
20
30
500
900
20.64
24.90
29.14
33.31
37.36
21.07
25.33
29.56
33.71
37.76
21.49
25.75
29.98
34.12
38.16
1000
1100
1200
1300
1400
41.31
45.14
48.85
52.41
55.81
41.70
45.52
49.21
52.75
42.08
45.89
49.57
53.10
600
700
800
o•c
4.()
50
(j()
70
80
90
/()()
21.92
26.18
30.40
34.53
38.56
22.34
26.60
30.82
34.94
38.96
22.77
27.03
31.23
35.35
39.35
23.20
27.45
31.65
35.75
39.75
23.62
27.87
32.07
36.16
40.14
24.05
28.29
32.48
36.56
40.53
24.48
28.72
32.90
36.96
40.92
24.90
29.14
33.31
37.36
41.31
42.47
46.27
49.94
53.45
42.86
46.64
50.29
53.79
43.24
47.01
50.65
54.13
43.62
47.38
51.00
54.47
44.00
47.75
51.36
54.81
44.38
48.12
51.71
44.76
48.48
52.06
55.48
54.14
48.85
52.41
55.81
55.15
Table 7 Nominal dielectric constants and densities of fuels
Dielectric constant
Fuel
Type
91/98
100/130
115/145
JP-I
JP-3
JP-4
Density D
(lb/gal)
K
Temperature °C
+55
0
-55
+55
0
-55
1.914
1.912
1.895
2.071
2.017
2.007
1.990
1.991
1.971
2.145
2.098
2.083
2.066
2.070
2.047
2.219
2.179
2.159
5.636
5.597
5.517
6.493
6.100
6.160
6.025
5.988
5.913
6.835
6.464
6.520
6.414
6.379
6.30SJ
7.177
6.827
6.880
411
Prlncipal symbols and
abbreviations
Symbols for quantities are in italic type, and abbreviations for the names of units
(unit symbols) are in ordinary type.
A
a
B
bhp
C
·c
els
F
ft/min
ft/h
ft/s2
g
G
H
Hz
I
in Hg
K
k
L
lb/gal
lbf/in 2
M
m
mA
mb
mV
mph
mmH20
412
ampere
speed of sound
magnetic flux density
brake horsepower
capacitance
radiation constant
degree Celsius
cycle per second
farad
foot per minute
foot per hour
foot per second per
second
acceleration due to
gravity
conductance
magnetic field strength
hertz (frequency)
electric current
moment of inertia
inch of mercury
Kelvin
coefficient of heat
transmission
inductance
pound per gallon
pound force per square
inch
Mach number; torque
magnetic moment
mass
milliampere
millibar
millivolt
mile per hour
millimetre of water
N
newton (force)
number of turns of a coil
Nm
Newton Metre (torque)
pascal (N/m2 )
Pa
pF
picofarad
resistance
R
gas constant
radian
rad
rev/min revolution per minute
period, periodic time
T
V
volt
V
velocity
w
weight
weber (magnetic flux)
Wb
X
reactance
impedance
z
temperature coefficient of
a
resistance
angle
µ
permeability
µF
microfarad
p
density
resistivity
magnetic flux
~
heat current or flow
ohm
0
angular velocity
w
(radians per second)
thermal conductivity
A
permittivity
E
dielectric constant
conductivity
i'
cubic coefficient of expansion
ratio of specific heat
N
Appendix 1
Conversion factors
1. Pressure
multiply by
To convert
into
Atmospheres
29.921
inches Hg (0°C)
406.9
inches H20
kilogrammes per sq cm
1.0333
millibars
1013.25
millimetres Hg (0°C)
760.00
piezes
101.331
pounds per sq in
14.6%
101325.000
*pascals
Inches Hg
atmospheres
inches H20
kilogrammes per sq cm
millibars
millimetres Hg
pounds per sq in
Kilogrammes per
sq cm
0.03342
13.60
0.03453
33.8639
25.40
0.4912
atmospheres
inches Hg
kilogrammes per sq cm
millibars
pounds per sq in
pascals
2.458 X 10-3
0.07355
2.540 X 10-3
2.490
0.03613
249.089
atmospheres
inches Hg
millimetres Hg
piezes
pounds per sq in
0.000987
28.%
735.54
l.0197
14.223
"The pascal (Pa) is an SI unit and is the pressure produced by a force of I newton applied,
uniformly distributed, over an area of I square metre.
Note: It is common practice to refer to a pressure as so many 'pounds per square inch·. Since,
however. pressure is more exactly pounds-weight or force, acting per square inch. the symbol
·
'lbf/sq in' is now adopted in lieu of 'lb/sq in'.
413
Millibars
2. Velocity
0.02953
0.7450
0.0145
100.00
Millimetres Hg
inches Hg
kilogi:ammes per sq cm
pascals
pounds per sq in
0.03937
0.0013596
133.322
0.019337
Pounds per sq in
atmospheres
inches Hg
inches H20
kilogrammes per sq cm
millibars
millimetres Hg
pascals
Pascals
pounds per sq in
To convert
into
Feet per minute
feet per second
kilometres per hour
knots
metres per minute
miles per hour
0.01667
0.01829
35.524
0.3048
0.01136
Feet per second
kilometres per hour
knots
metres per minute
miles per hour
1.0973
0.5921
18.29
0.6818
Kilometres per hour
feet per minute
feet per second
knots
metres per minute
miles per hour
54.68
0.9113
0.5396
Knots
414
0.01450
atmospheres
inches Hg
millimetres Hg
pounds per sq in
pascals
feet per minute
feet per second
kilometres per hour
miles per hour
0.06804
2.03596
27.68
0.0703
68.9476
51.713
6896.55
0.0001450
multiply by
16.67
0.6214
101.34
1.689
1.8532
l.l516
3. Volumetric
Metres per minute
feet per minute
feet per second
kilometres per hour
knots
miles per hour
3.281
0.05468
0.06
0.03238
0.03728
Miles per hour
feet per minute
feet per second
kilometres per hour
knots
88.00
1.4666
1.60934
0.8684
To convert
into
Cubic cemtimetres
cubic feet
cubic inches
imperial gallons
litres
pints
quarts
Cubic feet
cubic centimetres
cubic inches
imperial gallons
litres
pints
quarts
US gallons
Cubic inches
cubic centimetres
cubic feet
imperial gallons
litres
pints
quarts
Imperial gallons
cubic centimetres
cubic feet
cubic inches
litres
US gallons
multiply by
3.531 X 10-s
0.06102
2.1997 X 10-4
0.001
1.7598 X 10-3
8.7988 X 10-4
28316.85
1728.00
6.2288
28.32
59.84
29.92
7.481
16.39
5.787 X 104
3.6047 X 10-3
0.01639
0.5688
0.2844
4546.087
0.160544
277.42
4.54596
1.201
415
4. Angular Measure
416
Litres
cubic centimetres
cubic feet
cubic inches
imperial gallons
pints
quarts
US gallons
1000.00
0.03532
61.025
0.21998
1.7598
0.8799
0.2642
Pints
cubic centimetres
cubic feet
cubic inches
litres
quarts
568.26
0.02007
34.68
0.5682
0.5
Quarts
cubic centimetres
cubic feet
cubic inches
litres
1136.522
0.04014
69.3548
l.13649
US gallons
cubic centimetres
cubic feet
cubic inches
imperial gallons
li!res
3786.44
0.1337
231.00
0.8327
3.785
To convert
into
Degrees
minutes
quadrants
radians
Minutes
degrees
quadrants
radians
Quadrants
degrees
minutes
radians
90.00
5400.00
1.571
Radians
degrees
minutes
quadrants
57.2957
3437.75
0.6366
multiply by
60.00
0.0111
0.0175
0.0166
1.852 X J0-4
2.909 X 10-4
5. Angular Velocity
To convert
inro
Degrees per second
radians per second
revolutions per second
revolutions per minute
Radians per second
degrees per second
revolutions per second
revolutions per minute
~
multiply by
Revolutions per second degrees per second
radians per second
revolutions per minute
Revolutions per minute degrees per second
radians per second
revolutions per second
6. Temperature
~F
°C
= (°C
= (°F
X 1.8)
+
- 32)/1.8
Absolute zero (K)
=
=
32
or
or
= (°C
= (°F
-273.15°C 0°C
-459.67°F 0°F
0.01745
2.788 X 10-3
0.1667
57.30
0.1592
9.5493
360.00
6.283
60.00
6.00
0.10472
0.01667
x 9/5) + 32
- 32) x 5/9
= 273.15K
= 459.67K
411
Appendix 2
Logic gates & truth tables
~
;::0-£
a
~
~
NANO
~
B
B
B
~
a
a~
a
NEGATED AND
~
~
B
NOR
NEGATED OR
Q!
L{>-! [ff)
~
INVERTER
418
~
B
011
101
110
~
B
I
EXCLUSIVE OR
EXCLUSIVE NOR
010
100
111
A,
A
B
B
010
100
110
~~
WIRED GATES
Appendix 3
Acronyms and
abbreviations
This Appendix, which is by no means exhaustive, is intended as a
guide to the meanings of acronyms and abbreviations found in the
documentation dealing with the description, operation, logic signal
functions and maintenance of instruments and integrated systems.
ACARS
ACAS
ACCEL
ACQ
AID
ADC
ADEU
ADF
ADI
AFCS
AFS
AGC
AGS
AHRS
AIDS
ALPHA
ALU
ANN
AP, A/P
ARINC Communications Addressing and
Reporting System
Airborne Collision and Avoidance System
Accelerometer
Acquire (prefixed by a condition, e.g. ALT ACQ)
Analog to Digital
.Air Data Computer
Automatic Data Entry Unit
Automatic Direction Finder
Attitude Director Indicator
Automatic Flight Control System
Automatic Flight System
Automatic Gain Control
Automatic Gain Stabilization
Attitude and Heading Reference System
Airborne Integrated Data System
Angle of Attack
Arithmetic Logic Unit
Annunciator
Autopilot (suffixed by condition, e.g. ENG,
DISC)
APFDS
APMS
APPR OC
APS
ARINC
ARM
AS, AIS
ASA
Autopilot and Flight Director System
Automatic Performance and Management System
Approach On Course
Altitude Preselect
Aeronautical Radio InCorporated
Armed (prefixed by condition, e.g. LOC ARM,
VOR ARM)
Airspeed
Autoland Status Annunciator
419
ATT ERR
AUTO APPR
Autothrottle
Automatic Test Equipment
Abbreviated Test Language for Avionic Systems
Automatic Test-Oriented Language
Austin Turnbull Radio (formerly Air Transport
Radio)
AutoThrottle System
AutoThrottle/Speed Control
Attitude (may be followed by condition, e.g. ATT
:iOLD)
Attitude Error
Automatic Approach
BIA
BARO
BB
BIB
BIC, BC, BICRS
BID
BITE
BRG
Banlc Angle
Barometric
Bar Bias
Back Beam
Back Course
Bottom of Descent
Built-In Test Equipment
Bearing
CADC
CAP
Central Air Data Computer
Capture (prefixed by a condition, e.g. LOC CAP,
NAV CAP)
Collins Adaptive Processor System
Caution And Warning Panel
Collective Bar Bias
Control and Display Unit
Course Error
Character Generator
Clock
Command (prefixed by another abbreviation, e.g.
FD CMD)
Compensation, Compass, Comparator
Controller
Control Panel
Coupled (prefixed by condition, e.g. ROLL,
PITCH, APPR)
Central Processor Unit
Course
Cathode Ray Tube
Control Systems Electronic Unit
Control Transformer
Caution and Warning
Control Wheel Steering
AT
ATE
ATLAS
ATOL
ATR
ATS
AT/SC
ATT
CAPS
CAWP
CBB
CDU
CE
CG
CLK
CMD
COMP
CONT
CP
CPL
CPU
CRS
CRT
CSEU
CT
cw
cws
420
DIA
.DAD
DADC
DAIS
DDI
DDS
DES
DEVN
DFDAU
DFDR
DG
DH
DI
DIFCS
DISC
DISPL
DME
DMLS
DMM
DMUX
DRC
DSR TK
DTG
DU
DVM
Digital to Analog
Data Acquisition Display
Digital Air Data Computer
Digital Avionics Information System
Dual Distance Indicator
Digital Display System
Desired (suffixed by condition, e.g. DES TRK,
DES CRS)
Deviation
Digital Flight Data Acquisition Unit
Digital Flight Data Recorder
Directional Gyroscope
Decision Height
Digital Interface
Digital Integrated Flight Control System
Disconnect
Displacement
Distance Measuring Equipment
Doppler Microwave Landing System
Digital Multi-Meter
Demultiplexer
Dual Remote Compensator
Desired Track
Distance-To-Go
Display Unit
Digital Volt-Meter
EXT
Electronic Attitude Direction Indicator
Electronic Centralized Aircraft Monitor
Environmental Control System
Electronic Data Processing System
Electronic Engine Control
Electronic Flight Control Unit
Electronic Flight Instrument System
Exhaust Gas Temperature
Electronic Horizontal Situation Indicator
Electro-Hydraulic Servo Valve
Engine Indicating and Crew Alerting System
Engage
Easy-On
Engine Pressure Ratio
Expanded Localizer
Extend
FAC
FADEC
Flight Augmentation Computer
Full Authority Digital Engine Control
EADI
ECAM
ECS
EDPS
EEC
EFCU
EFIS
EGT
EHSI
EHSV
EICAS
ENG
EO
EPR
EX LOC
FAWP
FWC
Final Approach Waypoint
Flight Control Computer
Flight Control Electronic System
Flight Control Electronic Unit
Flight Control Unit
Flight Director
Flight Data Entry Panel
Flight Guidance System
Fault Isolation Monitoring
Flight Instrument System
Flight Level CHange
Flight Mode Annunciator
Flight Management Computer
Flight Management Computer System
Flight Management Computer Unit
Flight Management System
Fiber-Optic Data Transmission System
Fuel Performance Computer
Flap Position Module
Fast Slew
Flap/Slot Electronic Unit
Force Trim Release
Frequency-to-Voltage Converter
Flight Warning Computer
GA, G/A
GPWS
GS, G/S
Go-Around
Ground Proximity Warning System
Glide Slope
HARS
HDG
Heading and Attitude Reference System
Heading (<;:an be suffixed by condition, e.g. HDG
HOLD, HOO SELect)
Hold
Horizontal Situation Indicator
Head-Up Display
High Voltage Power Supply
FCC
FCES
FCEU
FCU
FD, F/D
FDEP
FGS
FIM
FIS
FLCH
FMA
FMC
FMCS
FMCU
FMS
FOOTS
FPC
FPM
FS
FSEU
FTR
FVC
HLD
HSI
HUD
HVPS
IAS
IAWP
ICU
ILS OC
IMU
INC-DEC
INS
INTGL
INTLK
422
Indicated Airspeed
Initial Approach Waypoint
Instrument Comparator Unit
Instrument Landing System On Course
Inertial Measuring Unit
Increase- Decrease
Inertial Navigation System
Integral
Interlock
INWP
IRMP
IRS
IRU
ISA
ISS
IVS
Intermediate Waypoint
Inertial Reference Mode Panel
Inertial Reference System
Inertial Reference Unit
International Standard Atmosphere
Inertial Sensing System
Instantaneous Vertical Speed
LAU
LBS
LNAV
LOC
LRRA
LRU
LSU
LVDT
Linear Accelerome · wt t.J nit
Lateral Beam Sensor
Lateral NAVigation
Localizer
Low-Range Radar Altimeter
Line Replaceable Unit
Logic Switching Unit
Linear Voltage Differential (also Displacement)
Transformer
MALU
MAN
MAP
MAWP
MCDP
MCP
MCU
MDA
MIP
MM
MPU
MSU
MTP
MUX
MWS
Mode Annunciation Logic Unit
Manual
Mode Annunciator Panel
Missed Approach WayPoint
Maintenance Control Display Panel
Mode Control Panel
Modular Concept Unit
Minimum Descent Atitude
Maintenance Information· Printer
Middle Marker
Maximum Operating Mach No.
Microprocessor Unit
Mode Selector Unit
Maintenance Test Panel
Multiplexer
Master Warning System
NAV
NC
NCD
NCU
ND
NDB
NM
NOC
Navigation
No Connection or Normally Closed
No Computed Data
Navigation Computer Unit
Navigation Display
Non-Directional Beacon
Nautical Mile
NAV On Course
OAT
OC, 0/C
Outside Air Temperature
On Course
MMo
423
OD
OM
ONS
oss
Out of Detent (may be prefixed, e.g. CWS OD)
Outer Marker
Omega Navigation System
Over Station Sensor
PAFAM
PAS
PATT
PBB
PCA
PCB
PCPL
PCWS
PDCS
PDU
PECO
PFD
PHOLD
PIU
PMS
PNCS
PRAM
PSAS
PSM
PSO
PSYNC
Performation And Failure Assessment Monitor
Performance Advisory System
Pitch Attitude
Pitch Bar Bias
Power Control Actuator
Printed Circuit Board
Pitch Coupled
Pitch Control Wheel Steering
Performance Data Computer System
Pilot's Display Unit
Pitch Erection Cut-Off
Primary Flight Display
Pitch Hold
Peripheral Interface Unit
Performance Management System
Performance Navigation Computer System
Programmable Analog Module
Pitch Stability Augmentation System
Power Supply Module
Phase Shift Oscillator
Pitch Synchronization
'
RA, RIA
RALU
RBA
RBB
RCPL
RCVR
RCWS
RDMI
REF
REV/C
RG
R/HOLD
RLS
RMI
RN,RNAV
RN/APPR
RSAS
RSU
R/T
Radio (Radar) Altimeter
Register and Arithmetic Logic Unit
Radio Bearing Annunciator
Roll Bar Bias
Roll Coupled
Receiver
Roll Control Wheel Steering
Radio Distance Magnetic Indicator
Reference
Reverse Course (same as Back Course)
Raster Generator
Roll Hold
Remote Light Sensor
Radio Magnetic Indicator
Area Navigation
Area Navigation Approach
Roll Stability Augmentation System
Remote Switching Unit
Receiver /Transmitter
RTE DATA
RVDT
R/W
Route Data
Rotary Voltage Differential Transmitter
Read/Write
SAI
SAM
SAS
SAT
SCAT
SCM
SEL
SELCAL
SFCC
SID
SG
SGU
SPD
SRP
SSEC
STAR
STBY
STCM
STS
Stand-by Attitude Indicator
Stabilizer Aileron Module
Stability Augmentation System
Static· ~ir Temperature
Speect'Co"r~arid of Altitude and Thrust
Spoiler/Speedbrake Control Module
Select
Selective Calling
Slat/Flap Control Computer
Standard Instrument Departure
Symbol Generator (Stroke Generator)
Symbol Generator Unit
Speed (Airspeed or Mach hold)
Selected Reference Point
Slow Slew
Static Source Error Correction
Standard Terminal Arrival Route
Standby
Stabilizer Trim Control Module
Status (prefixed by a function, e.g. TRACK STS)
TACAN
TAS
TAT
TIC
TCC
TCS
T/D
TET
TGT
TK CH
TKE
TMC
TMS
TMSP
TRP
TTL
Tactical Air Navigation
True Air Speed
Total Air Temperature
Top of Climb
Thrust Control Computer
Touch Control Steering
Top of Descent
Turbine Entry Temperature
Turbine Gas Temperature
Track Change
Track Angle Error
Thrust Management Computer
Thrust Management System
Thrust Mode Select Panel
Thrust Rating Panel
Tuned to Localizer
VAR
VBS
VDU
VGU
Variable
Vertical Beam Sensor
Visual Display Unit
Vertical Gyro Unit
ss
~
425
VLD
VMo
VNAV
VOR
VOR APPR
VOROC
VORTAC
VREF
vs
vscu
WO, W/0
WXR
Washout
Waypoint
Weather Radar transceiver
XTK DEV
XTR
Cross Track Deviation
Transmitter
YD, YID
YDM
Yaw Damper
Yaw Damper Module
WPT
426
Valid (usually suffixing a condition, e.g. VG
VLD, FLAG VLD)
Maximum Operating Airspeed
Vertical Navigation
Very-high-frequency Omnidirectional Range
VOR Approach
VOR On Course
VOR T ACtical (Air navigation)
Reference Speed
Vertical Speed
Vertical Signal Conditioner Unit
Exercises
Chapter 1
1.
2.
3.
4.
5.
6.
7.
8.
9.
Chapter 2
1.
2.
3.
4.
5.
6.
Explain the difference between quantitative and qualitative displays,
and quote some examples of instruments to which they are applied.
What is the difference between static and dynamic counter displays?
To which types of instrument is a director display applied?
Describe the operating principle of an LED, and by means of a
diagram show how the principle is applied to produce a segmented
numeric display.
Describe the operating principl9 of an LCD.
What type of display is formed when the elements are arranged in,
say, a 4 x 7 configuration, and to which display element does it
specifically apply?
Name the flight instruments that comprise the basic 'T' layout, and
state their respective positions. Does this layout also apply to
electronic displays?
What do you understand by the term 'head-up' display? With the aid
of a diagram describe how the required basic flight data are
displayed.
What is the significance of the coloured markings and/or 'memory
bugs' applied to certain instruments?
Define the following: (i) troposphere, (ii) tropopause, and (iii)
stratosphere.
What do you understand by the term 'ISA'? State also the
assumptions made.
The pressure of the atmosphere:
(a) increases non-linearly with height;
(b) decreases non-linearly with height;
(c) decreases linearly with height.
1 inch of mercury is equal to:
(a) 14.7 millibars;
(b) 2.49 millibars;
(c) 33.87 millibars.
What are the principal components and instruments which comprise a
basic air data system?
With the aid of a simple diagram, describe the construction of a
combined pitot and static pressure sensing probe.
427
7.
8.
9.
IO.
11.
12.
13.
14.
15.
16.
17.
18.
19.
20.
21.
Chapter 3
1.
Explain what is meant by the 'PE' of an air data system, and how its
effects are minimized.
How are alternative sources of pitot and static pressure normally
provided for, and connected to the appropriate instruments?
Explain the principle of pitot pressure measurement and how the
l/2pV 2 law is derived.
Describe the construction arid operation of a typical pneumatic type
of airspeed indicator.
Computed airspeed is:
(a) calibrated airspeed corrected for PE;
(b) indicated airspeed compensated for the square-law response of the
airspeed sensor;
(c) indicated airspeed corrected for PE.
Define the term Mach number, and describe how it is indicated by
measurement in terms of the ratio (p, - p.)lp,.
Describe how the functions of a Machmeter and an airspeed indicator
can be combined to provide indications of Vmo·
What is the difference between 'pressure altitude' and 'indicated
altitude'?
When setting the BP counters of an altimeter to the pressure
prevailing at a particular airport, the corresponding 'Q' code is
known as:
(a) QFE and the altimeter will read zero;
(b) QFE and the altimeter will read the airport height above
sea-level;
(c) QNH and the altimeter will read zero.
With the aid of a diagram explain the operating principle of a VSI.
An IVSI provides more rapid indications of climb and descent
because it utilizes a vertical acceleration pump:
(a) instead of 1 metering unit;
(b) which is connected between the pressure sensing capsule and a
metering unit;
(c) which is connected directly to the metering unit.
Why is it customary to sense air temperature in terms of a total
value? Briefly describe the construction and operation of a typical
sensing probe us-xi for this purpose.
Describe the operation of a typical Mach warning system.
Explain the operating sequence of an altitude alerting unit when an
·
aircraft descends to a preselected altitude.
Explain the operation of a stall warning and stick-shaker system.
Define the following: (i) magnetic meridian, (ii) magnetic variation,
(iii) isogonal lines, and (iv) agonic lines.
2.
3.
4.
5.
6.
7.
8.
9.
Chapter 4
I.
2.
3.
4.
5.
6.
7.
8.
9.
10.
11.
The angle the lines of magnetic force make with the earth's surface is
called:
(a) deviation;
(b) dip;
(c) variation.
Describe the construction of a typical direct-reading compass.
Define the two principal errors that can occur in the readings of a
compass under aircraft operating conditions.
What do you understand by the terms 'hard-iron' and 'soft-iron'
magnetism of an aircraft?
Name the components of hard-iron magnetism and the aircraft axes
about which they are effective.
Which of the hard-iron and soft-iron components are associated with
the deviation coefficients B and C?
Express the formulae used for the calculation of deviation coefficients
A, Band C.
Briefly describe how a deviation compensating device neutralizes the
fields due to aircraft magnetic components.
Define the two fundamental properties of a gyroscope, and state the
factors on which they depend.
How are the properties defined in Q. l utilized in flight instruments?
What are the input and output axes of a gyroscope?
As the speed of a gyroscope rotor increases, the rate of precession
for a given torque:
(a) remains constant;
(b) decreases;
(c) increases.
With the aid of a diagram, explain how a gyroscope precesses under
the influence of an applied force.
How are the spin axes of gyroscopes arranged for the detection of
pitch and roll attitude changes, and for establishing directional
references?
What is meant by 'earth rate', and how would the input axis of a
gyroscope have to be aligned to exhibit apparent drift equal to this
rate?
What is meant by 'transport wander', and does it have the same
effects on horizontal-axis and vertical-axis gyroscopes?
Briefly describe some methods of controlling drift and transport
wander.
What do you understand by the terms 'gimbal lock' and 'gimbal
error'?
Describe how pitch and roll attitudes are displayed by a gyro
horizon.
429
12.
How is the gyroscope of an electrically-operated gyro horizon erected
to, and maintained in, its normal operating position?
13. Describe the operation of an electrolytic type of levelling switch.
14. What effects do acceleration and turning of an aircraft have on the
indications of a gyro horizon, and how are they compensated?
15. Assuming that an aircraft accelerates while in straight and level
flight, the effect on a gyro horizon utilizing a levelling torque motor
system would cause it to indicate a:
(a) pitch-up attitude;
(b) pitch-down attitude;
(c) pitch-up and a left bank attitude.
16. With the aid of a diagram, describe the operation of an erection cutout system.
17. Describe how the rate gyroscope principle is applied to indicate the
rates at which an aircraft turns.
18. With the aid of diagrams, describe how a ball type of bank indicator
displays (a) a correctly-banked turn, and (b) a turn to starboard in
which the aircraft is overbanked.
19. Describe how a rate gyroscope may be utilized -to sense both banking
and rate of turn.
Chapter 5
1.
2.
3.
4.
5.
6.
Chapter 6
I.
2.
3.
4.
Explain the operation of a torque synchro system.
In what type of synchro system is a control transformer utilized?
Explain how a synchro is applied to systems involving the
measurement of the sine and cosine components of angles.
What do you understand by the term 'electrical zero' as applied to a
synchro system?
The letters 'TDX' designate a:
(a) combined torque and differential synchro transmitter element;
(b) differential synchro used in a torque synchro system;
(c) torque synchro transmitter element when used with a differential
synchro.
Explain the operating principle of a synchrotel.
What are the elements that constitute the hardware and software of a
digital computer?
Name the principal sections of a CPU and explain their functions.
Briefly explain the functions of the busses comprising a computer
highway.
Which of the highway busses are bidirectional?
5.
6.
7.
8.
9.
10.
Chapter 7
1.
2.
3.
4.
5.
6.
7.
8.
9.
Chapter 8
l.
What is the name given to the digital code through which a computer
carries out instructions?
What is the difference between a RAM and a ROM?
A 16K memory has a bit storage capacity of:
(a) 16 000;
(b) 32 000;
(c) 16 384.
How is data in analog form converted to binary-coded format?
In the ARINC 429 format of data transfer, how is data identified
according to function?
How are any errors in the codes used in the transmission of data
detected?
Explain some of the reasons why ADC systems are used in aircraft.
Desuibe the operation of a piezoelectric type of pressure transducer.
Explain how an output in terms of vertical speed can· be obtained
from the altitude module of an ADC.
What inputs are required to obtain signals whose values correspond to
TAS?
How are square- law characteristics and PE compensated and
corrected by an ADC?
What warning and indicating flags are provided in a typical servooperated Mach/airspeed indicator?
What is the function of the capsule-type sensor element in a
pneumatic/servo-operated type of altimeter?
How are indications of SAT derived from the signals produced by a
TAT sensing probe?
Explain how BP settings are made in a servo-operated altimeter
which is supplied with signals from a digital type of ADC.
Draw a diagram to show the basic coil and core arrangement of a
MHRS detector unit, and explain how fluxes and voltages are
2.
3.
induced.
Explain how the magnitude of the voltages induced in a detector unit
is used as a measure of aircraft heading.
The sensing element of a detector unit is pendulously mounted within
its casing, and has limited freedom in:
(a) pitch only. so as to reduce the effects of acceleration;
(b) pitch and roll so as to reduce the effect of the earth's magnetic
field component Z;
(c) azimuth only so as to reduce errors due to turning.
4.
5.
6.
7.
8.
9.
10.
Chapter 9_
l.
2.
3.
4.
5.
6.
7.
8.
9.
10.
Explain what is meant by a 'pre-indexed' type of detector unit, and
how it is mounted in an aircraft.
With the aid of a diagram explain how a detector unit monitors a
DGU and an RMI.
What is the purpose of the syrtchronizing and annunciator system?
When a monitored gyroscope type of MHRS is operating in the
'slaved' mode, any drift of the gyroscope is controlled by:
(a) manually resetting the compass card of the RMI;
(b) setting the known latitude of the aircraft on the scale of a latitude
corrector unit;
(c) the slaving and servo synchro loops between the DGU and RMI.
How are slaving and servo signal transmission circuits controlled in
an MHRS which is integrated with an INS?
Under what conditions is the DG mode of operation selected?
The purpose of the 'SET HDG' facility is to:
(a) mechanically set a heading 'bug' to the required beading, and
also to position the rotor of a CX synchro which provides
heading signals to other systems;
(b) mechanically rotate a compass card into alignment with a heading
'bug';
(c) position the stator of a servo CT and so provide a s~:rvo drive to
a compass card.
Draw a diagram to illustrate the display of a typical ADI.
With the aid of a schematic diagram, explain how primary attitude
changes -are displayed by an ADI.
What is the significance of the dots against which the localizer and
glide slope elements of an ADI and HSI are registered?
State the purpose of the ADI command bars, and explain how they
are positioned.
What is the difference between 'course' and 'heading'? Explain how
they are selected and displayed on an HSI.
How is an aircraft's position with respect to a VOR station
displayed?
GS and VOR/LOC deviation indicators are operated by:
(a) synchros;
(b) servomotors;
(c) de meter movements.
What is the purpose of 'TO-FROM' indicators, and in which of the
FDS indicators are they incorporated?
Explain how 'TO-FROM' indicators are activated.
Under what conditions is the 'GA' mode of operation selected?
Explain how selection is carried out and the effect it has on the ADI
display.
11.
12.
13.
14.
15.
16.
17.
Chapter 10
1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
11.
12.
13.
14.
At which stage of a flight profile are the amber and green lights of
an annunciator panel illuminated? Describe their operating sequence.
Describe how the convergence of a GS beam is allowed for during
the 'AUTO APPR' mode of operation.
What is the purpose of an instrument comparator and warning
system?
What warning and indicating flags are provided in an ADI and an
HSI?
What heading information is displayed on the HSI when an FDS is
interfaced with an INS and operates in the 'INS' mode?
In addition to heading, what other navigational data can be displayed
on an HSI when the FDS operates in the 'INS' mode?
In dual FD systems, how is it ensured that data displays are available
in the event of failure of data supply sources?
State the functions of the principal units that comprise an INS.
From the operational point of view, what are the differences between
an INS and an IRS?
Define the terms 'longitude' and 'latitude'.
What is meant by convergency?
What do the abbreviations TK, TKE, and XTK signify?
Draw a simple diagram to illustrate the significance of the
abbreviated term 'WPT'.
If the value of an aircraft's TK is geater than that of its heading:
(a) no drift will occur;
(b) drift would be to the left;
(c) drift would be to the right.
What are the fundamental laws of mechanics on which INS operation
depends?
What data must first be entered into an INS computer in order for the
system to navigate an aircraft?
With the aid of a schematic diagram, describe how accelerometer
signals are integrated to provide information on an aircraft's present
position.
Describe how the X and Y accelerometers are aligned, and state also
the coordinate system to which the output signals are related.
What do you understand by the term 'Schuler tuning'?
The gyroscopes utilized in a stabilized platform type of INU are of
the:
(a) displacement type;
(b) rate-integrating type;
(c) laser type.
How are the input axes of gyroscopes positioned with respect to an
aircraft's axes, and what attitude changes do they sense?
15.
16.
17.
18.
19.
20.
21.
22.
23.
24.
25.
26.
27.
Chapter 11
I.
2.
3.
4.
5.
6.
7.
What is the function of the Z gyroscope of a stabilized platform type
of INU?
Describe how the gyroscopes are compensated for earth rate and
transport rate.
What are the functions of the pitch, outer roll, and azimuth synchros
connected to a gimballed platform?
Describe the constructional arrangement of a laser gyroscope.
Describe how a laser gyroscope senses aircraft attitude changes in
terms of an angular rate.
How are attitude changes sensed and resolved when an inertial
reference unit is installed in a 'strapdown' configuration?
What are the principal modes in which an IN/IR system operates, and
how are they selected?
How are waypoints selected and inserted?
What is the purpose of an •ALERT' annunciator?
Briefly describe the alignment procedure essential for IN operation.
What do you understand by the term 'gyrocompassing'?
Under what conditions would a 'slewing' procedure be carried out?
What is the purpose of the code numbering system that is
programmed into an IN computer? State how the codes are displayed.
What are the principal elements of a CRT?
In a CRT, the electrons are produced and emitted by:
(a) passing signal currents through deflection coils;
(b) the heating of an element called a cathode;
(c) the heating of an element called an anode.
When an electron beam strikes the inside surface of a CRT screen,
how is it made to produce a spot of light?
How is the spot of light made to tr.ace out a line or a pattern on the
screen of a CRT?
The three primary colours of a CRT are produced by:
(a) electron beams that are coloured red, blue and green;
(b) passing the electron beams from three electron 'guns' through a
shadow mask containing red, blue and green screens;
(c) directing the beams from three electron 'guns' through a
perforated shadow mask so that they strike three different kinds
of phosphor coatings.
Describe how colours other than the primary ones are produced.
Describe the raster scannin¥ and stroke scanning techniques, stating
the types of display provided.
Chapter 12
I.
2
3.
4.
5.
6.
7.
8.
9.
iO.
Chapter 13
I.
2.
3.
4.
5.
6.
7.
8.
9.
10.
11.
12.
13.
What are the main units that comprise an EFIS?
What are the functions of an SGU?
To which areas of the indicator displays is the raster scanning
technique applied?
State the colours assigned to the displays and the type of information
to which they correspond.
What is the function of a remote light sensor?
In dual EFIS installations, how is it ensured that the failure of an
SGU will not affect the display of data?
State the modes that can be selected for display on the EHSI.
How are failures of data signal sources displayed?
Under what conditions do the GS and LOC deviation pointers of the
EADI change from white to amber?
Can weather radar ·returns' be displayed in all modes of EBSI
operation?
What types of sensing elements are used for pressure measurement?
Describe a method of measuring engine oil pressure based on the
principle of synchronous transmission.
Explain how an inductor type of pressure transmitter produces the
varying currents required for the operation of a ratiometer.
For what purposes are pressure switches required?
With the aid of a diagram, explain how a pressure switch is made to
give a warning of a pressure falling below its normal operating value.
Describe how temperature changes can cause variations ii1 the
properties of substances. Which of these variations are utilized in
engine temperature indicating systems?
Name the materials most commonly used for variable resistance-type
sensing elements, and describe the construction of a typical element.
Describe how a Wheatstone bridge circuit may be utilized for the
measurement of temperature.
Describe the construction and operation of a ratiometer type of
temperature indicating system.
What would be the effect of an 'open circuit' between the sensing
element and indicator of the ratiometer type?
Explain the operating pririciple of a thermo-emf system, and state the
engine parameters measured by sJch a system.
The metal combinations used in an EGT sensing probe are:
(a) copper and constantan.
(b) chrome! and alumel.
(c) iron ,!nd constantan.
What effects can changes in cold junction temperature have on the
indications of thermo-emf instruments? Describe a method of
compensation.
435
14.
15.
Chapter 14
l.
2.
3.
4.
5.
6.
7.
8.
9.
IO.
11.
12.
13.
14.
15.
Chapter 15
1.
2.
3.
What is the difference between extension leads and compensating
leads?
The 'external circuit' of a thermo-emf system is that part which
extends from the thermocouple probes to the:
(a) ends of the harness only.
(b) junction at which conductors and extension leads enter an
aircraft's fuselage.
(c) terminals of the indicator.
Explain the operating principle of a capacitor and state the factors on
which it depends.
Define: (a) the 'Units in which capacitance is expressed,
(b) permittivity.
Describe how the capacitance principle is applied to the measurement
of fuel quantity.
Why is it necessary to install a number of sensing probes in a fuel
tank system?
What effects do temperature changes have on the fuels used, and how
are they compensated?
Why is it preferable to measure fuel weight rather than fuel volume?
Explain why densitometers provide greater accuracy in fuel weight
measurement than compensator probes.
Describe the construction and operation of a densitometer.
What do you understand by the term 'characterization' as applied to
tank sensing probes?
What adjustments are normally provided in a capacitance type of
system?
Explain the function of the test switch incorporated in some fuel
quantity indicating systems.
How are checks carried out in computer-controlled fuel quantity
indicating systems?
Briefly explain how total fuel remaining and aircraft's gross weight
can be displayed.
Why are refuelling control panels provided on some types of aircraft?
Describe a method of refuelling control.
Describe the construction and operation of a tacho-generator type of
rpm-indicating system.
Why is the speed of turbine engines measured as a percentage?
What is the purpose of a tacho probe? Describe the operation of an
indicating system which utilizes such a sensor.
4.
5.
6.
7.
8.
9.
10.
11.
12.
Chapter 16
l.
2.
3.
4.
5.
6.
7.
8.
9.
10.
11.
12.
Briefly explain the principle of supercharging, and how the increase
in induction manifold pressure is measured.
To which types of engine does torque monitoring relate? Describe a
method of torque measurement.
What do you understand by the terms 'stagnation' and 'rapid
response' as applied to EGT sensing probes?
How is it ensured that an EGT indicating system measures good
average temperature conditions of the exhaust gases?
Describe the operation of an electrical method of CJ compensation.
Describe the operation of a basic type of fuel flow indicating system.
What is the purpose of an EPR indicating system? Explain the
fundmental principle of measuring the pressure ratio.
What is meant by an 'integrated' flowmeter system? Describe a
method of achieving integration.
Describe the operation of an EVI system.
What are the meanings of the acronyms 'EICAS' and 'ECAM'?
What are the principal units that comprise an EICAS installation?
Which engine parameters are classified as primary and secondary'
information for purposes of display?
Primary information is displayed on the:
(a) lower display unit;
(b) standby engine indicator;
(c) upper display unit.
Name the three modes of EICAS operation, and state how they are
selected.
What data are displayed in each of the three operating modes?
What is meant by 'auto event' and 'manual event', and how is the
associated data called up for display?
With an aircraft on the ground, and with electrical power 'on', is it
necessary to switch on the display units from the display select panel?
Name the three levels of alert messages, and state on which display
unit they are presented.
If several alert messages were displayed at the same time, and it was
noted that one of them was indented one space, what level of
message would it signify?
If the condition that generated an alert message no longer exists, the
message is removed from the display:
(a) by pressing the 'cancel' switch on the DSP;
(b) automatically;
(c) by pressing the 'cancel' switch located near the display units.
What happens to the display of secondary information in the event ,of
failure of the lower display unit?
13.
14.
15.
16.
17.
18.
19.
20.
Chapter 17
I.
2.
3.
4.
5.
6.
7.
8.
9.
10.
11.
12.
13.
The standby engine indicator displays are permanently on when its
display control switch is in the:
(a) 'on' position and the test switch is at the left or right position;
(b) 'auto' position;
(c) 'on' position only.
Under what conditions can an EICAS self-test routine be carried out,
and from which panels is it controlled?
What are the principal differences between EICAS and the ECAM
system?
Name the four mod;s of an ECAM display, and state which one of
these is used in normal operation of the system.
Which display mode takes precedence over the others?
On which display unit are 'STATUS' messages displayed and what
information do they provide?
In order to carry out an automatic test routine, is it necessary for a
test switch to be operated first?
During a manual test of ECAM, when do diagrams appear on the
right-hand display unit?
What are the modes that can be selected on the CDU of a
performance data computer system?
What is the significance of the caret ( <) and asterisk symbols
displayed on the CDU?
What is the purpose of the mode annunciator of a PDCS?
How are the records for an FMS computer data base entered into the
memory?
What do the two sizes of characters displayed by the CDU signify?
What is the purpose of the line keys either side of the CDU screen?
For what purpose is the bottom line of the display used?
The 'CLEAR' key of the CDU is used to remove:
(a) any of the data displays;
(b) all records from the computer data base;
(c) only incorrect entries from the scratch pad.
Under what conditions is data presented in 'scroll' form?
What is the purpose of the flight phase annunciators, and what
happens to data display pages when the annunciators illuminate?
What is the function of the 'EXEC' key, and at what stage does the
bar illuminate?
If an arrow appears against a data line of the display, what does it
signify?
Describe a typical dual FMS configuration.
Solutions to exercises
Chapter 2
3 (b)
4 (c)
IO (c)
Chapter 3
Chapter 4
Chapter 5
Chapter 6
Chapter 8
15 (a)
17 (b)
2 (b)
4 (b)
15 (a)
5 (b)
7 (c)
7 (c)
8 (a)
Chapter 9
Chapter IO
7
6
13
Chapter 11 2
Chapter 13 12
(c)
(c)
(b)
(c)
(b)
15 (c)
Chapter 16
4
ll
12
Chapter 17 8
(c)
(b)
(c)
(b)
439
Index
Absolute pressure 356
Acceleration errors
compasses 83
gyro horizons 120
Accelerometer 258, 266
Aclinic line 8 J
Active display 13
Address bus 153
Agonic lines 80
Aircraft magnetism and components
hard-iron 87
soft-iron 88
total magnetic effect 90
Air data alerting and warning systems
68
Air data computers
analog 161
digital 177
Air data instruments 39
Air data system 28
·-"'"'Airspeed indicators
pneumatic 39, 40, 42
servo-operated 39, 168
Air temperature indicators
SAT 61, 68, 176
TAT 64, 176, 367
Air temperature sensing 60
Alpha angle 73
Alphanumeric displays 14, 283
Alpha sensor 74
Alternate pressure sources 36
,- Altimeters
pneumatic 39, 49
servo-operated 174, 180
Altimeter setting region 54
Altitude 55
Altitude alerting system 71
Analog/digital converter l 56
Angle of attack indicators 77
Angle of attack sensing 73
Angle of dip 81
Angular momentum 98
Annual change 80
Apparent drift 102
ARINC 429 158
ARINC 629 160
Arithmetic logic unit 153
Atmospheric pressure 25
Atmospheric temperature 27
Attitude director indicator (ADI) JO,
210
Ball type erection unit 115
Barometric pressure setting 51
'Basic six' layout 20
·Basic T' layout 21
B/H curve 185
Binary-coded format 152
Byte 153
Calibrated airspeed (CAS) 41
Capacitance type fuel quantity indicating
systems
adjustments 340
compensator systems 332
densitometer 334
effects of fuel temperature 331
electronic displays 341
fail-safe and test circuits 341
fuel mass measurement 331
principle of 328
probes 328, 335, 336, 339
refuelling and load control 344
totalizer indicators 343
Cartesian coordinates see Synchro
systems
CDX synchro see Synchro systems
Central processing unit (CPU) 153
Characterization of tank probes 339
Chemosphere 25
Circular scale 1
'Clock' type display
Coercivity 186
Cold junction temperature compensation
322, 363
Compass coupler 198
Compensating leads 325
Compensator tank probe 333
Component 'P' 88
Component 'Q' 88
Component 'R' 88
Computed airspeed 41
Computer languages 155
Computer µrogram 152
Control bus 153
Control of drift 105, 196
Control of transport wander 105, 263
Control synchro see Synchro systems
Control transformer see Synchro systems
CRT displays 283
Data bus 153
DATAC see ARINC 629
Data highway 152, 157
Data transfer 156
Declination 79
Densitometers 334
Deviation coefficients 91
Deviation compensation 93, 205
Dielectric constant 328
Differential synchros see Synchro
systems
Digital/analog converter 156
Digital computer 152
Digital display 5
Direction indicator
construction 125
control of drift 127
erection devices 127
gimbal errors 127
Directional gyroscope unit 192, 195
Direci memory access (DMA) 155
Director displays 10
Direct-reading compasses
acceleration errors 83
aperiodic 81
compass safe distance 83
construction 81
dip compensation 81
liquid expansion compensation 83
location 83
turning errors 85
Displays
digital 5, 13
director 7
dual-indicator 7
electronic 12, 13, 15, 283
head-up 17
high-range long-scale 3
qualitative l
quantitative 8
Diurnal change 80
Dot-matrix display 13
Dual-indicator displays 7
Dynamic counter display 7
Earth gyroscope 105
Earth rate 102, 264
Earth's atmosphere 25
Earth's total force 81
-ECAM
components of 387
display modes 388
system testing 391
Eddy-current drag principle 348
--"-EFIS
data source selection 306
display of air data 306
display presentations 299
failure annunciation 305
units of 296
EICAS
alert messages 382
components of 377
display modes 380
failures 383
maintenance control panel 385
Electrical capacitance
governing factors 327
in a.c. circuits 328
principle of 326
units of 327
Electrical zero see Synchro systems
Electrolytic type levelling switch 119
Electronic ADI 299
Electronic displays 12, 296, 377
Electronic HSI 302
Elevation 55
Engine pressure ratio system 365
Engine vibration monitoring 375
Equivalent airspeed (!AS) 41
Erection errors see Gyro horizon
Exhaust gas temperature 359
Exosphere 25
Extension leads 324
External circuit resistance 324
Fast erection systems see Gyro horizon
Flight director systems
comparator and warning system 235
componen~s of 208
dual systems 234
flight mode annunciators 225
go-around switches 227
interfacing with INS 238
mode controllers 223
operating sequences 228
Flight levels 55
Flight management system 399
Flight mode annunciators 225
Flight mode controllers 223
'Floated gyro' 269
Flux detector elements 182
Force-balance tram,ducer 164
- Fuel flow measurement
basic system 367
integrated system 369
..;
Gauge pressure
310
Geographic poles 78
Gimbal error 106. 127
Gimbal lock 106
Gimballing effects 121, 127
Graduation marks 2
Gyroscopic properties 97
Gyroscope reference datums 10 I
Gyroscopic flight instrument operation
pneumatic I 06
[
electric 108
l-Gyro horizon
acceleration and turning errors 120
electric 112
erection cut-out 122
441
erection rate 120
erection systems 114
fast erection system 119
pitch-roll erection l23
pneumatic 110
principle of 108
standby attitude indicator 113
Gyro-stabilized platform 260
Linear scale 2
Limiting speeds
M,,., 42, 171
v,,.., 42, 171
Liquid crystal display (LCD) 15
Location of tank probes 336
Logarithmic scale 2
Long-reach probe 362
Longitude 250, 252
LVDT 365, 370
Hard-iron magnetism see Ai:-craft
magnetism and components
Hardware 152
Mach/airspeed indicator 46. 172
.,.... Machmeter 44
Head-up displays 17
Height 55
Mach number 45
High-range long-scale displays 3
Mach speed indicator 170
Horizontal situation indicator (HSI) 216 Mach warning system 69
Hysteresis loop 186
Magnetic dip 80
Magnetic equator 80
Magnetic foci 79
Immersion type thermocouple 322, 359
Magnetic heading reference system
Indicated altitude 51
(MHRS)
- Indicated airspeed (IAS) 41
detector elemen1s J82
Indicated/computed airspeed indicator
48
deviation compensation 205
Inertial navigation/reference systems
dual systems 206
accelerometers 258, 266
integration with INS 198
action codes 281
monitored gyroscope system 192
operating modes 203
alignment sequencing 279
components of 246
synchronizing 201
displays and annunciations 275
Magnetic meridian 78
gimbal platform arrangement 265
Magnetic variation 79
Manifold pressure indica1ors 355
gyro-stabilized platform 260
malfunction codes 282
Manometric system 28
mode selection 273
Memory
principle of 256
capacity 155
types 154
ring laser gyroscope 270
slewing 280
'Memory bugs' 8
Inferred density systems 333
Mercury type levelling switch 117
Input/output (l/0) ports 155
Microfarad 327
Instantaneous vertical speed indicator
M"", see Limiting speeds
Moving-tape display 5
(IVSI) 59
Instrument grouping
Navigation
flight instruments 20
fundamentals 249
power plant instruments 21
terms 254
Instrument panels 19
Non-volatile memory see Memory
Integrated rate gyroscope 261
International standard atmosphere (ISA)
Operation code 152
28
Operand 153
Interrupts 155
Operational range markings 7
Ionosphere 28
Isoclinals 81
Passive display 13
lsodynamic lines 81
Pendulosity errors see Gyro horizon
Isogonal lines .so
Pendulous vane unit ll4
Percentage rpm indicator 350
Kilobit see Memory
Perf'onnance data computer system 394
Permalloy 184
Label (digital code) 158
Permeability 186
Lapse rate 27
Permittivity 327
Latitude 250, 252
Phase quadrature 118
Light-emitting diode (LED) 13
442
Picofarad 327
Piezoelectric sensor 165, 312
Pipelines and drains 37
Pitot pressure 28, 40
Pilot probe 31
Pitot probe heating 32
Pitot-static probe 29
Polar coordinates see Synchro systems
Position error 34; 161
Precession 99
Pre-indexed detector element 190
Pressure altitude 51
Pressure error see Position error
Pressure measurement
indicating systems 309
pressure switches 312
'Q' code 54
Qualitative displays 8
Quantitative displays I
Radio magnetic indicator I%
Ram rise 60
Random-access memory (RAM) see
Memory
Rapid response probe 359
Raster scanning 287
Rate gyroscope 129
Ratiometer system 316
Read-only memory see Memory
Real drift 103
Recovery factor 62, 360
Refuelling and load control 344
Remote-indicating compass see Magnetic
heading reference system
Resolver synchro see Synchro systems
Rho-theta 287
Ring laser gyroscope 270
'Rolling digit' configuration 15, 24
RPM indicating systems 347
SAT indicators see Air temperature
indicators
Saturation point 186
Scale base 2
Scale length 3
Scale marks 2
Scale range 4
Schuler pendulum 260
Secular change 80
Seebeck effect 320
Segment matrix display 13
Shadow mask 290
Short-reach probe 362
Slewing 280
Soft-iron magnetism see Aircraft
magnetism and components
Software 153
Space gyroscope 97
Square-law compensation 43, 171
Stagnation point 40
Stagnation type probe 359
Stalling angle 73
Stall warning systems 73
Standard altimeter setting 54
Standard atmosphere 27
Standby attitude indicator 113
Standby engine indicator 384
Static ai.r temperature (SAT) 60
Static counter display 7
Static pressure 28
Static source error correction 165
Static vents 35
'Stick-push' 74, 76
'Stick-shaker' 74
Straight scale 5
'Strapdown' configuration 279
Stratopause 25
Stratosphere 25
Supercharging 355
Surface contact thermocouple 322
'Swinging' 93
Synchronizing (MHRS) 201
Synchro systems
cartesian coordinates 146
control synchro 135, 140, 192
control transformer 140
differential synchros 142, 199
electrical zero 136
polar coordinates 146
resolver synchro 145
synchrotel 148
torque synchro 135, 136
Synchrotel see Synchro systems
Tacho probe 352
TAT indicators see Air temperature
indicators
TAT probes 61
TDX synchro see Synchro systems
Temperature indicating systems
external circuit 324
ratiometer 316
thermo-emf 320, 359
Wheatstone bridge 324
Terrestial magnetism 78
Thermistor 324
Thermocouples
materials 320, 410
locations 322, 360
Thermo-emf system '320, 359
Thermomagnetic shunt 324
Torque indicating system 357
Torque motor and lev.elling switch
system 116, 196
Torque synchro see Synchro systems
Total air temperature ff AT) 61
Transition altitude 55
Transport wander 104, 263
Tropopeuse 25
Troposphere 2S
,,._ True airspeed (TAS) 42, 174
'Tuning' spring 44
Tum and bank indicator 129
Tum coordinator 133
Turning efn?rs
compasses 85
gyro horizons 120
Vertical speed indicator SS
V.., see Limiting speeds
Volatile memory see Memory
Volumetric top-off 344
Wheatstone bridge system 31S
Word IS3
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