Aircraft Instruments
Aircraft Instruments
By the same author
Aircraft Electrical Systems
3rd edition
Illustrated 179 pages
ISBN O 582 98819 5
Aircraft
Instruments
Principles and Applications
Second edition
EH J Pallett
I Eng, AMRAeS
with a foreword by
Air Cdre Sir Vernon Brown
PEARSON
Authorized Licensed Edition of the original UK edition, entitled Aircraft /11sflw11ents, Second
Edition by Pallett, authors published by Pearson Education Limited,© EH J Pallett 1972, 198 1
This Licensed Edition is published by Dorl.ing Kindersley (India) Pvt. Ltd,
Copyright© 2009 by arrangement with Pearson Education Ltd, United Kingdom.
All rights reserved. This book is sold subject to the condition that it shall not, by way o f trade
or otherwise, be lent, resold, hired out, or otherwise circulated without the publisher's prior
written consent in any form of binding or cover other than that in w hich it is published and
without a similar condition including this condition being imposed on the subsequent purchaser
and without limiting the rights under copyright reserved above, no part of this publication may
be reproduced, stored in or introduced into a retrieval system, or transmitted in any for m or by
any means (electronic, mechanical, photocopying, record ing or otherwise) , without the prior
written permission of both the copyright owner and the above-mentioned publisher of this
book.
ISBN 978-8 1-317-28 13-0
First Impression, 2009
This edition is manufactured i11 l11dia and is authorized for sale only i11 I11dia, Bangladesh,
Bhutan, Pakistan, Nepal, Sri Lanka and the Maldives. Circulation of this edition outside of
these territories is UNAUTHORIZED.
Published by Dorling Kindersley (India) Pvt. Ltd., licensees of Pearson Education in South Asia.
Head Office: 7th Floor, Knowledge Boulevard, A-8(A), Sector-62, Noida - 20 1309, U.P, India.
Registered Office: 11 Community Centre, Panchsheel Park, New Delhi 110 Ol 7, India.
Digitally Printed in India by Repro India Ltd. in the year of 20 15.
Contents
Foreword vii
Preface to the second edition
viii
Preface to the first edition
ix
Historical background xi
1
2
3
4
S
6
7
8
9
10
11
12
13
14
1S
16
Tables
Requirements and standards
Instrument elements and mechanisms
6
Instrument displays, panels and layouts
19
Pitot-static instruments and systems
50
Primary flight instruments (attitude indication)
116
Heading indicating instruments
l 59
Remote-indicating compasses
184
Aircraft magnetism and its effects on compasses
206
Synchronous data-transmission systems 225
Measurement of engine speed 242
Measurement of temperature
257
298
Measurement of pressure
313
Measurement of fuel quantity and fuel flow
Engine power and control instruments 338
Integrated instrument and flight director systems
358
Flight data recording
373
Principal symbols and abbreviations
392
Conversion factors
393
I. Standard atmosphere
399
2. Pressure/airspeed equivalents
400
401
3. Mach number/airspeed relationship
4 . Temperature/resistance equivalents 402
5. Temperature/millivolt equivalent of typical iron v. constantan
402
thermocouples
6. Temperature/millivolt equivalent of copper v. constantan
thermocouples
403
Temperature/millivolt
equivalent of typical chrome! v. alumel
7.
thermocouples 403
404
8. Nominal dielectric constants and densities of fuels
9. ARINC standard ATR case sizes
404
Typical pressure, temperature and rotational speed notations
Solutions to numerical questions 406
Index 409
yj
405
Foreword
by Air Cdre Sir Vernon Brown CB, OBE, MA,
CEng, FRAeS, HonFSLAET
The ~eroplane on which I took my Aero Club ticket had as its
instrumentation an engine rev. counter and an oil pressure gauge, an
altimeter which-recorded up to 16,000 ft (and mighty slow it was)
and an airspeed indicator on which the scale showed up to about 75
m.p.h. (the top speed of a Maurice Fannan Longhorn was plus/minus
60 m.p.h. and its best climbing speed about 28 m.p.h.). This·airspeed
indicator consisted of a small cup on the end of a rod at the end of
which was a bell crank lever which actuated a pointer. As it was
mounted on the outer strut, to see one's speed at all from the front
tandem seat meant turning one's head more than 90° to the left!
A lot of water has flowed under the world's bridges since early
1915, and with the advent of the cabin type of aircraft and the
turbine engine the aeroplane has become a very complicated vessel
indeed, the main purposes of which are either to carry enonnous loads
of highly flammable fuel at great speed and height with as big a
payload as possible, or simply to destroy an enemy before he destroys
you. Naturally its instrumentation has become proportionally difficult
to install and maintain.
This book by Mr Pallett has been written with the aim of helping
those whose job is to keep aircraft in the air to understand their
instruments and, of course, the fundamental principles underlying
their design and application to flight.
Every time I am invited to an aircraft pilot's cabin, I thank my
stars I no longer fly except as a passenger and my sympathy goes out
to the licensed aircraft engineers who are responsible for ensuring
that everything works according to plan. Theirs is a very difficult
job, and if Mr Pallett's book will help them, as indeed I feel sure it
must, then he will have done a good job and will be the recipient of
very many blessings.
vii
Preface to the
Second Edition
The continuing demand for a book of this nature has been most
encouraging, and it has been particularly gratifying to meet the
publisher's requirements for the production of a second edition.
During the seven-year period of the book's existence, many new
types of instruments and associated systems have, inevitably, been
developed. As is generally the case however, such developments have
been in the methods of applying established principles. It has been
hoped that coverage of these principles in the first edition and its
subsequent reprints has been of help to those readers who, having
encountered practical examples of the newer 'applications technology'
in the course of their work, have had to gain a deeper understanding
of operating fundamentals.
In order to include details of representative examples of new
instruments in this second edition, some 'surgery' has been
performed on details appertaining to instruments which have either
become obsolete or obsolescent.· It has also been necessary to
re-arrange the sequence and titling of chapters, considerably enlarging
their content in several instances. As a result of suggestions from a
number of readers, to whom I am indebted, the opportunity has been
taken to expand the first edition treatment of certain operating
principles; this applies particularly to the principles of the gyroscope,
and of synchronous transmission systems. A section on conversion
factors and tables appropriate to the parameters measured by
instruments has also been included as a new feature.
Valuable assistance in the illustrating of the text associated with
subjects new to this book, has also been given by Smiths Industries,
Collins Radio Company of England Ltd., Normalair-Garrett Ltd.,
British Airways, and Laker Airways (Services) Ltd., and I hereby
gratefully acknowledge such assistance and the permission granted
for reproduction of a number of photographs and diagrams .
Copthorne
Sussex
EP
1981
Preface to the
First Edition
The steady growth in the number and scope of aircraft instruments
has run parallel with the complex growth of aircraft themselves, and
in th'e development of methods of detecting, processing and presenting
relevant control information, the design and constructional patterns of.
instruments have likewise grown in a corri:-lex fashion. As a result,
instruments are often associa ted with a science veiled in considerable
'electrickery' contained with;_n numerous black boxes produced by
obscure persons posing as disciples of Pandora!
There is, of course, no denying the abundance of black boxes (they
are invariably grey these days anyway!) and the many functional and
constructional changes, cut as in the evolutionary processes of most
technological fields, complexities arise more often than not in
developing new method s of applying old but well-established
principles. For example, in the measurement of altitude, airspeed
and turbine engine thrust, the appropriate pressures are detected and
measured by the deflections of capsule- and diaphragm-type detecting
elements just as they were in some of the first instruments ever
dP,veloped for pressure measurement in aircraft. Similarly, instruments
employed for the measurement and control of engine temperatures
still depend for their operation on the changes in electrical resistance
and thermo-elec tric characteristics of certain metals under varying
temperature conditions.
Thus , in preparing the material for this book, emphasis has been
placed on fundamental principles and their applications to flight,
navigation and engine performance-monitoring instruments. It has
not been possible to include every type of instrument, but it is
considered that the coverage is representative of a wide range of
current aircraft instrument installations, and should pr0vide a firm
foundation on which to base further study .
The material is arranged in a sequence which the author has found
useful in the implementation of training programmes based on
relevant sections of the various examination syllabuses established
for aircraft mainte nance engineers, and the various ratings of pilots'
licences. In this connection it is therefore hoped that the book will
prove a useful source of reference for the experienced instructor as
well as for the student. A selection of questions are given at the end
of each chapter and the author is indebted to the Society of Licensed
ix
Aircraft Engineers and Technologists for permission to reproduce
questions selected from examination papers.
Valuable assistance has been given by a number of organizations
in supplying technical data, and in granting pennission to reproduce
many of the illustrations. Acknowledgement is hereby made to the
following: Smith's Industries Ltd , Aviation Division; Sperry Rand
Ltd, Sperry Gyroscope Division; Sangamo-Weston Ltd; Thom Bendix;
British Overseas Airways Corporation ; British Aircraft Corporation
(Operating) Ltd; R. W. Munro Ltd; Dowty Electrics Ltd; Negretti and
Zambra (Aviation) Ltd ; Hawker Siddeley Aviation Ltd .
EP
X
Historical background
In the days of the first successful aeroplanes the.problems of operating
them and their engines according to strict and complicated procedures,
of navigating over long distances day or night under all weather
conditions, were, of course, problems of the future. They were, no
doubt, envisaged by the then enthusiastir pioneers of flight , but were
perhaps somewhat overshadowed by the thrills of taking to the air,
manoeuvring and landing.
Such aeroplanes as these pioneers flew were rlither 'stick and
string' affairs with somewhat temperamental engines, the whole
combination being manoeuvered by a pilot lying, sitting or crouching
precariously in the ope'1 , for the luxury of a cockpit was also· still to
come. Instruments designed specifically for use in an aeroplane were
also non-existent ; after all, what instrument manufacturer at the time
had had the necessity of designing, for example, an instrument to show
how fast a man and a machine could travel through the air?
It is a little difficult to say exactly in what sequence instruments
were introduced into aeroplanes. A magnetic compass was certainly
an early acquisition as soon as pilots attempted to fly from A to B,
and flying greater distances would have required information as to
how much petrol was in the tank , so a contents gauge was fitted ,
usually taking the form of a glass sight gauge. Somewhere along the
line the clock found its place and was useful as a means of calculating
speed from a time/distance method , and as an aid to navigation. With
such supplementary aids a pilot was able to go off into the third
dimension flying mainly by his direct senses and afterwards boasting
perhaps that instruments would not be needed anyway!
As other pioneers entered the field many diverse aeroplane
designs appeared, some of which were provided with an enclosure
for the pilot and a wooden board on which the then available
instruments could be mounted . Thus the cockpit and instrum ent
panel were born .
Shortly before the outbreak of World War I, some attention was
given to the development of instruments for use on military and
naval aeroplanes, and the first principles of air navigation were
emerging with designs for instruments specially adapted for the
purpose. Consequently, a few more instruments appeared on the
xi
dashboards of certain types of aeroplane including an altimeter,
airspeed indicator and the first engine instruments - an r.p.m.
indicator and an oil pressure gauge.
During the war years very few new instruments were provided in
the many types of aeroplanes produced. A requirement did arise for
the indication of an aeroplane's pitch and bank attitude, which led
to the introduction of the fore-and-aft level and the cross-level. The
former instrument consisted of a specially constructed glass tube
containing a liquid which moved up and down against a graduated
scale, and the latter was a specially adapted version of the simple
spirit level.
The main progress of the war years as far as instruments were
concerned was in the development of the existing types to higher
standards of accuracy, investigation of new principles, and the
realization that instruments had to be designed specifically to withstand vibration, acceleration, temperature change, and so on. It was
during this period that aircraft instruments became a separate but
definite branch of aviation.
After the war, aviation entered what may be termed its second
pioneering stage in which ex-wartime pilots flew air routes never
before attempted. In 1919, for example, Alcock and Brown made
the first non-stop Atlantic crossing; in the same year the Australian
brothers Keith and Ross Smith made the first flight from England to
Australia. Flights such as these and others carried out in the 1920s
were made in military aircraft and with the aid of the same
instrument types as had been used in wartime. Although these flights
laid the foundation for the commercial operation of the aeroplane,
it was soon realized that this could not be fully exploited until
flights could be safely carried out day and night and under adverse
weather conditions. It had already been found that pilots soon lost
their sense of equilibrium and had difficulty in controlling an
aeroplane when external references were obscured. Instruments
were therefore required to assist the pilot in circumstances which
became known as 'blind flying conditions'.
The first and most important step in this direction was the
development of the tum indicator based on the principles of the
gyroscope. This instrument, in conjunction with the magnetic
compass, became an extremely useful blind-fiying aid, and when a
bank indicator was later added to the turn indicator, pilots were able,
with much patience and skill, to fly 'blind' by means of a small group
of instruments.
However, progress in the design of aeroplanes and engines
developed to a stage where it was essential to provide more aids to
further the art of blind flight. An instrument was required which
could replace the natural horizon reference and could integrate the
xii
information hitherto obtained from the cross-level and the fore-andaft level. It was also necessary to have some stable indication of
heading which would not be affected by acceleration and turning
-manoeuvres which had for long been a source of serious errors in
the magnetic compass.
The outcome of investigations into the problem was the introduction
of two more instruments utilizing gyroscopic principles, namely the
gyro horizon and the directional gyro, both of which were successfully
proved in the first ever instrument flight in 1929. At this time the
sensitive altimeter and the rate-of-climb indicator had also appeared
on the instrument panel ('dashboard' was now rather a crude term!),
together with more engine instruments. Engines were being supercharged and so the 'boost' pressure gauge came into vogue; temperatures
of oil and liquid cooling systems and fuel pressures were required to be
known, and consequently another problem arose - instrument panels
were getting a little overcrowded. Furthermore, these instruments,
essential though they were, were being grouped on the panel in a
rather haphazard manner, and this made it somewhat difficult for
pilots to assimilate the indications, to interpret them and to base on
them a definite course of action.
Thus, by about the middle 1930s grouping of instruments became
more rationalized so that 'scanning distance' between ins_truments
was reduced to a minimum. The most notable result of rationalization
was the introduction of the separate 'blind flying panel' containing
the airspeed indicator, altimeter, gyro horizon, directional gyro, rateof-climb indicator (vertical speed) and tum-and-bank indicator. This
method of grouping the flight instruments has continued up to the
present day.
Continuing developments in military and commercial aviation
brought about faster and bigger aeroplanes, multi-engine arrangements, retractable landing gear systems, electrical systems, etc., all of
which called for more and more instruments. The instrument designer
kept pace with these developments by introducing electrically
operated instruments and systems of remote indication, but the fitting
of knobs, switches and handles for the operation of other systems
imposed restrictions on space. Pilots were therefore once again facing
problems of assimilating the indications of haphazardly placed
instruments. In multi-engined aircraft the problems.were further
aggravated .
The greater ranges of multi-engined aircraft meant longer periods
in the air and presented the problem of pilot fatigue: a problem not
unknown to the pioneers of long-distance flying. This was alleviated
by equipping bomber and long-range commercial aircraft of the late
1930s with an automatic pilot, a device which had been successfully
demonstrated as far back as 1917. With the automatic pilot in
xiii
operation, pilots were able to devote more attention to instrument
monitoring, and to the involved navigational and radio communicatior
techniques whicb had also been introduced.
A further step in relieving the pilot's work load was made when
navigator, radio operator and flight engineer stations were introduced,
becoming standard features during World War II. It thus became
possible to mount the instruments appropriate to the crew member's
duties on separate panels at his station, leaving the pilot with the
instruments essential for the flight handling of the aircraft.
One of the most outstanding developments resulting from the war
years was in the field of navigation, giving rise notably to the fullscale use of the remote-transmitting compass system in conjunction
with such instruments as air and ground position indicators, and air
mileage units. Flight instruments had been improved, most instruments for engine operation were now designed for electrical operation,
and as an aid to conserving panel space more dual-type im:truments
had been introduced.
Another development which took place, and one which changed
the picture of aviation, was that of the gas turbine engine. As a
prime mover, it opened up many possibilities: more power could be
made available, greater speeds and altitudes were possible, aircraft
could be made 'cleaner' aerodynamically; and being simpler in its
operation than the piston engine, the systems required for its
operation could also be made simpler. From the instrument point
of view the changeover was gradual and initially did not create a
sudden demand for completely new types of instrument. The
rotational speeds of turbine engines were much higher than those of
piston engines and so r.p.m. indicators had to be changed
accordingly, and a new parameter, gas temperature, came into
existence which necessitated an additional thermometer, but apart
from these two, existing engine and flight instruments could still be
utilized.
This state of affairs, however, lasted for only a few years after the
war. As the ·aircraft industry geared itself up into its peacetime role,
various designs of gas-turbine powered aircraft went into production
and flew alongside the more conventional types. Demands for
greater speeds and altitudes meant bigger engines, and as new hightemperature materials and newer systems were developed, the
required power was produced, but the 'jet' was losing a lot of its
simplicity in the process!
Once again instruments had to be modified and new designs
introduced. For example, the turbine engine produced far less
vibration at the instrument panel, and as a result there was a tendency
for the slight inherent static friction of an instrument mechanism to
'stick' the pointers. The mechanisms were therefore designed to
make them function correctly and continuously in the absence of
xiv
Blackburn monoplane I 1910)
Lock head - I 011 (Tristar).
(by Courtesy of British
Airways)
xv
vibration. Increases of speed brought about an effect known as
compressibility, and to warn pilots of their approach to a dangerous
flight situation, the speed indicator known as the Machmeter was
introduced.
·
The higher altitudes made possible by the turbine engine, and the
fast controlled descent procedures later adopted, necessitated the
use of altimeters having extended ranges, and scales which could be
read easily and without ambiguity. In order to meet these requirements and to help pilots avoid subsequent 'mis-reading incidents,'
the altimeter has passed, and continues to pass, through various
stages of modification.
In the aircraft electrical and electronic field advances were also
being made so that new measJJring technqiues were possible for
instruments; for example, the me~surement of fuel quantity by
means of special types of capacitor located in the fuel tanks,
electrically-operated gyro horizons and tum and bank indicators
providing for greater stability and better performance at high
altitudes than their air-driven counterparts.
As a result of rapid growth of radio aids to navigation, specific
'radio aid' instruments were also introduced to present additional
information for use in conjunction with that provided by the standard
flight and navigation instruments. Although these were essential for
the safe operation of aircraft, particularly during the approach and
landing phases of flight, the pilot's workload was increased and it was
foreseen that eventually the locating of separate instruments on
panels would once more become a problem . It was therefore natural
for an integration technique to be developed whereby the data from
a number of instrument sources could be presented in a single
display. Thus, integrated flight instrument and flight director systems
were evolved, and are now a standard instrumentation feature of
many types of aircraft currently in service, not only for the display
of primary flight data, but also for the monitoring of advanced
1utomatic flight control systems.
The display of more varied data covering the performance of
engines and systems also became necessary with the growing complexity of engines, and so the number of separate indicators increased.
However, the application of miniaturized electrical and electronic
components, and of micro-circuit techniques, has permitted large
reductions in the dimensions of instrument cases thereby helping to
keep panel space requirements within reasonable bounds. Furthermore, it has led to the up-dating of an early data presentation method,
namely the vertical scale, for the engine instruments of a number of
today's aircraft. Another technique now applied as a standard feature
of turbine engine instrumentation is based on one successfully
developed in the late 1940s, i.e. the control of engine speed and gas
temperature by the automatic regulation of fuel flow. In this
xvi
technique th e signals generated by the standard tachom eter
generators and by thermocouples are also processed electronically
and are used to position the appropriate fuel control valve system .
From this very brief outline of instrument development it will be
particularly noted that this has for the most part been based on
'a quart into a pint pot' philosophy , and has been continuously
directed to improving the methods of presenting relevant data.
This has been a natural progression, and development in these areas
will continue to be of the utmost importance, paralleling as it does
the increasing complexity of aircraft and their systems. As the
controller of a 'man-machine loop' and in developing and operating
the machine , man has been continually reminded of the limitations
of his natural means of sensing and processing control information.
However, in the scientific and technical evolutionary processes,
instrument layout design and data presentation methods have become
a specialized part of ergonomics, or the study of man in his working
environment ; this, together with the rapid strides made in avionics,
culminates in the provision of electronic display ins.truments,
computerized measuring elements, integrated instrument and flight
control systems for fully automatic control, enabling man to deal with
an expanding task within the normal range of human performance.
xvii
1 Requirements and
standards
The complexity of modem aircraft and all allied equipment, and the
nature of the environmental conditions under which they must
operate, require conformity of design, development and subsequent
operation with established requ.irements and standards. This is, of
course, in keeping with other branches of mechanical and transport
engineering, but in aviation requirements and standards are unique
and by far the most stringent.
The formulation and control of airworthiness requirements as they
are called, and the recommended standards to which raw materials,
instruments and other equipment should be designed and manufactured, are established in the countries of design origin, manufacture
and registration, by government departments and/or other legally
constituted bodies. The international operation of civil aircraft
necessitates international recognition that aircraft do, in fact, comply
with their respective national airworthiness requirements. As a result,
international standards of airworthiness are also laid down by the
International Civil Aviation Organization (ICAO). These standards
do not replace national regulations, but serve to define the complete
minimum international basis for the recognition by countries of
airworthiness certification.
lt is not intended to go into all the requirements - these take up
volumes in themselves - but rather to extract those related
essentially to instruments; by so doing a useful foundation can be
laid on which to study operating principles and how they are applied
in meeting the requirements.
Requirements
Location, Visibility and Grouping of Instruments
1. All instruments shall be located so that they can be read easily by
the appropriate member of the flight crew.
2. When illumination of instruments is provided there shall be
sufficient illumination to make them easily readable and
discernible by night. Instrument lights shall be installed in such a
manner that the pilot's eyes are shielded from their direct rays and
that no objectionable reflections are visible to him.
1
3. Flight, navigation and power-plant instruments for use by a pilot
shall be plainly visible to him fro~ his station with the minimum
practicable deviation from his normal position and line of vision
when he is looking out and forward along the flight path of the
aircraft.
4. All flight instruments shall be grouped on the instrument panel
and, as far as practicable, symmetrically disposed about the
vertical plane of the pilot's forward vision.
5. All the required power-plant instruments shall be conveniently
grouped on instrument panels and in such a manner that they
may be readily seen by the appropriate crew member.
6. In multi-engined aircraft, identical power-plant instruments for
the several· engines shall be located so as to prevent any misleading
impression as to the engines to which they relate.
Instrument Panels
The vibration characteristics of instrument panels shall be such as
not to impair seriously the accuracy of the instruments or to <lama~~
them. The minimum acceptable vibration insulation characteristics
are established by standards formulated by the appropriate national
organization.
Instruments to be Installed
Flight and Navigation Instrument1,
1. Altimeter adjustable for changes in barometric pressure
2. Airspeed indicator
3. Vertical speed indicator
4 . Gyroscopic bank-and-pitch attituqe indicator
5. Gyroscopic rate-of-tum indicator (with bank indicator)
6. Gyroscopic direction indicator
7. Magnetic compass
8. Outside air temperature indicator
9. Clock
Pilot-static System
Instruments 1, 2 and 3 above form part of an aircraft's pi tot-static
system, which must also conform to certain requirements. These
are summarized as follows:
a The system shall be air-tight, except for the vents to atmosphere,
and shall be arranged so that the accuracy of the instruments
cannot be seriously affected by the aircraft's speed, attitude, or
configuration; by moisture, or other foreign matter.
b The system shall be provided with a heated pitot-pressure probe
to prevent malfunctioning due to icing.
2
c Sufficient moisture traps shall be installed to ensure positive
drainage throughout the whole of the system.
d In aircraft in which an alternate or emergency system is to be
installed, the system must be as reliable as the primary one and
any selector valve must be clearly marked to indicate which
syst em is in use.
e Pipelines shall be of such an internal diameter that pressure lag and
possibility of moisture blockage is kept to an acceptable minimum.
f Where static vents are used, to obviate yawing errors they shall be
situated on opposite sides of the aircraft and connected together
as one system . Where duplicate systems are prescribed, a second
similar system shall be provided.
Gyroscopic Instruments
Gyroscopic instruments may be of the vacuum-operated or electrically
operated type, but in all cases the instruments shall be provided with
two independent sources of power, a means of selecting either power
source, and a means of indicating that the power supply is working
satisfactorily.
The installation and power supply system shall be such that failure
of one instrument, or of the supply from one source, or a fault in any
part of the supply system, will not interfere with the proper supply
of power from the other source.
Duplicate Instruments
In aircraft involving two-pilot operation it is necessary for each pilot
to have his own pilot-static and gyroscopic instruments. Therefore
two independent operating systems must be provided and must be so
arranged that no fault which might impair the operation of one is
likely to impair the operation of both.
Magnetic Compass
The magnetic compass shall be installed so that its accuracy will not
be excessively affected by the aircraft vibration or magnetic fields of
a permanent or transient nature .
Power Plant Instruments
1. Tachometer to measure the rotational speed of a crankshaft or a
compressor as appropriate to the type of power plant.
2. Cylinder-head temperature indicator for an air-cooled engine to
indicate the temperatu re of the hottest cylinger.
3. Carburettor-intake air temperature indicator.
4. Oil temperature indicator to show the oil inlet and/or outlet
temperature.
5. For turbojet and turbopropeller engine~ a temperature indicator
3
to indicate whether the turbine or exhaust gas temperature is
maintained within its limitations.
6. Fuel-pressure indicator to indicate pressure at which fuel is being
supplied and a means for warning of low pressure.
7. Oil-pressure indicator to indicate pressure at which oil is being
supplied to a lubricating system and a means for warning of low
pressure.
8. Manifold pressure gauge for a supercharged engine.
9. Fuel-quantity indicator to indicate in gallons or equivalent unit:
the quantity of usable fuel in each tank during flight. Indicators
shall be calibrated to read zero during cruising level flight, when
the quantity of fuel remaining is equal to the unusable fuel, i.e.
the amount of fuel remaining when, under the most adverse
conditions, the first evidence of malfunctioning of an engine
occurs.
10. Fuel-flow indicator for turbojet and turbopropeller engines. For
piston engines not equipped with an automatic mixture control a
fuel flowmeter or fuel/.air ratio indicator.
11. Thrust indicator for a turbojet engine.
12. Torque indicator for a turbopropeller engine.
Standards
4
In the design and manufacture of any product, it is the practice to
comply with some form of specification the purpose of which is to
ensure conformity with the required production processes, and to set
an overall standard for quality of the product and reliability when
ultimately performing its intended function. Specifications, or
standards as they are commonly known, are formulated at both
national and international levels by specialized organizations. For
example, in the United Kingdom, the British Standards Institution
is the recognized body for the preparation and promulgation of
national standards and codes of practice, and it represents the United
Kingdom in the International Organization for Standardization (ISO),
in the International Electrotechnical Commission (IEC) and in West
European organizations performing comparable functions.
Standards relate to all aspects of engineering and as a result vast
numbers are produced and issued in series form corresponding to
these aspects. As far as aircraft instruments and associated equipment
are concerned, British Standards come within the Aerospace GI 00
and G200 series; they give definitions, constructional requirements,
dimensions, calibration data, accuracy required under varying
environmental conditions, and methods of testing. Also in
connection with instruments and associated electronic equipment,
frequent reference is made to what are termed ARINC specifications.
This is an acronym for Aeronautical Radio Incorporated, an organi-
zation in the United States which operates under the aegis of the
airline operators, and in close collaboration with manufacturers.
One notable specification of the many which ARINC formulate
is that which sets out a standard set of form factors for the items
colloquially termed 'black boxes'. In the main, these factors cover
case dimensions, mounting racks, location of plugs and sockets, and
a system of index ing fouling pins to ensure that only the correct
equ ipment can be fitted in its appropriate rack position. The size of
box is based on a standard width dimension called 'one ATR' (yet
another abbreviation meaning Air Transport Rack) and variations in
simple multiples of this provide a range of case widths. Two case
lengths are provided for, and are termed long and short, and the
height is standard. TI1e various configurations are detailed in Table 9
on page 404.
5
2 Instrument elements
and mechanisms
Elements
From the operating point of view, we may regard an instrument as
being made up of the following fou r principal elements: (i) the
detecting element, which detects changes in value of the physical
quantity or condition presented to it ; (ii) the measuring element,
which actually measures the value of the physical quantity or
condition in terms of small translational or angular displacements;
(iii) the coupling element, by which displacements are magnified and
transmitted ; and (iv) the indica ting element, which exhibits the value
of the measured quantity transmitted by the coupling element , by
the relat ive positio ns of a pointer, or index, and a scale. The relationship between the fo ur elements is sh own in Fig 2.1.
Figure 2. J Elements of an
instrument.
(
m~~l-1+
-+
I
DETECTING ELEM~NT
I
I
I
I
MEASURING
iZ
I
1
- - --~-~~:i_
l - ~~ · N
--T
Mechanisms
6
~- ~ --
- - - -INDICATING ELEMENT
In the strictest sense, the term mechanism refers to all four elements
as a composite unit and coniai ned within the case of an instrument.
However, since the manner in which the functions of the elements
are performed and integrated is governed by relevant instru ment
operating principles and construction, this applies to only a very few
instruments. In the majority of applicatio ns to aircraft, a separation
of some of the elements is necessary so that three, or maybe only
two, elements form the mechanism, within the instrument case. The
direct-reading pressure gauge shown at (a) in Fig 2.2 is a good
example of a composite unit of mechanical elements, while an
example of separated mechanical elements as applied to an airspeed
indicator is shown at (b ). In this example the detecting element is
separated from the three other elements, which thus form the
mechanism within the case.
There are other examples which will become evident as we study
subsequent chapters, but at this stage it will not be out of place to
consider the operation of a class of mechanisms based on the principles of levers and rods. These are utilized as coupling elements
which follow definite laws, and can introduce any required input/
output relationship . In aircraft instrument applications, such lever
Figl4re 2.2 Instrument
mechanisms. (a) Direct·
reading pressure gauge;
(b) airspeed indicator con·
taining measuring, cou piing
and indicating elements.
DETECTING AND MEASURING
ELEMENT
(al
=
DETECTING
ELEMENT
I
I
I
L _ ___
I
I
I
J
I
I
L _ __ _ _ _ __ __ !
(bl
7
al'ld rod mechanisms are confined principally to direct-reading
pressure gauges and pitot-static flight instruments.
Lever Mechanism
Let us consider first of all the simple Bourdon tube pressure gauge
shown at (a) in Fig 2.3. The Bourdon tube forms both the detecting
and measuring elements, a simple link, lever, quadrant and pinion
forms the coupling element, while the indicating element is made up
of the pointer and scale. This mechanism is of the basic lever type,
the lever being in this case the complete coupling element. When
pressure is applied to the tube it is displaced, such displacement
resulting in input and output movements of the coupling and
indicating elem en ts, respectively, in the directions shown.
In connection with mechanisms of this type, two terms are used
both of which are related to the movement and calibration of the
indicating element; they are, lever length, which is ·the distanced
between the point of operation of the measuring element and the
pivoting point of the lever, and lever angle, which is the angle 8
between the lever and the link connecting it to the measuring element.
In order to understand what effects these have on the input/
output relationship, let us again refer to Fig 2.3 (a). The movement
of the indicating element is proportional to the lever length; thus, if
the lever is pivoted at its centre, this movement will be equal to the
input movement. Let us now assume that the pivoting point is
moved to a distance d 1 from the point of operation. The lever length
is now reduced so that for the same input movement as before the
output movement of the indicating element will be increased. From
this it will be clear that an increase of lever length to a distance d 2 will
produce a decreased output movement for the same input movement.
The effect of lever angle on the input/output relationship is to
change the rate of magnification since the lever angle itself changes
in response to displacement of the measuring element. This effect is
evident from Fig 2.3 (b ). If we assume that the line AB represents
the axis of the lever at its starting position, then the starting lever
angle will be 8 . Assume now that the measuring element is being
displaced by equal increments of pressure applied to it. The lin\c
attachment point C will move to C 1 and will increase the lever angle
in two stages; firstly when the link pivots about point B, and
secondly when the link pulls the lever arm of the coupling element
upwards from the starting position taking point B to point 8 1 • Thus,
the axis of the lever arm has moved to A1 B1 and the lever angle has
increased to a total angl~8 1• When the next increment of pressure
is applied, point C reaches C2 , point B reaches B2 and the axis AB
moves to A 2 B2 , so that not only has the lever angle been further
increased, but also the magnification, the distance from A 1 to A2
8
Figure 2.3 Simple lever
mechanism . (a) Effect of
lever length; (b) effect of
lever angle on magnification.
INPUT
MOVEMENT
REDUCED
I.EVER
LENGTH d1
LENGTH
di
DISTANCES AB. BC
EQUAL
(al
{-A.
.7'
-A;
(bl
being much greater than that from A to A 1 •
From the foregoing, it would appear that the two effects counteract each other, and that erratic indications would result. In all
instruments employing lever mechanisms, however, provision is
made for the adjustment of lever lengths and angles so that the
indicating element follows the required calibration law within the
limits permissible.
9
Rod Mechanisms
Unlike pure lever mechanisms, rod mechanisms dispense with pin or
screw-jointed linkages for the interconnection of component parts,
and rely on rods in contact with, and sliding relative to, each other
for the generation of the input/output relationship. Contact between
the rods under all operating conditions is maintained by the use of a
hairspring which tensions the whole mechanism.
These mechanisms, shown in Fig 2.4, find their greatest application
in flight instruments, and can be divided into three main classes
named after the trigonometrical relationships governing their
operation. They are: (i) the sine mechanism, (ii) the tangent
mechanism and (iii) the double-tangent mechanism.
The sine mechanism, Fig 2.4 (a), is employed in certain types of
airspeed indicator as the first stage Qf the coupling element, and
comprises two rods A and B in sliding contact with each other, and a
rocking shaft C to which rod Bis attached. In response to displacement of the measuring element, the input movement of rod A is in a
vertical plane, causing rod B to slide along it and at the same time to
rotate the rocking shaft. The point of contact between the two rods
remains at a constant radius r from the centre of the rocking shaft.
The rotation of the rocking shaft is given by the trigonometrical
relationship
h2
-
lz 1 = r(sin
82 -
sin 8 1)
where h is the vertical input movement of rod A and 8 the angle of
rod B. The usable range of movement (8 in the diagram) of rod B is
±60°, and the angle at which it starts within this range depends on
the magnification required for calibration. For example, if rod A
moves upwards from a starting angle at -60°, the magnification is at
first high and then decreases with continued movement of rod A.
When the starting angle is at or near the zero degree position, the
magnification rate is an increasing one.
A tangent mechanism is similar to a sine mechanism, but as will be
noted from Fig 2.4 (b), the point of contact between the two rods
remains at a constant perpendicular distanced from the centre of
the rocking shaft. The rotation of the rocking shaft is given by the
relationship
h 2 -h 1 =d(tan8: -tan8i).
The magnification rate of this mechanism is opposite to that of a
sine mechanism except at a starting angle at or near zero, where
sin 8 and tan 8 are approximately equal.
Figure 2.4 (c) illustrates a double-tangent mechanism, which is
employed where rotary motion of a shaft is to be transferred through
10
Figure 2.4 Rod mechanisms.
(a) Sine mechanism;
(b) tangent mechanism;
(c) double-tangent mechanism.
/
/
(bl
h
Rob A
h
ROCKING
SHAFT D
(cl
a right angle. A typical application is as the second stage of an airspeed indicator coupling element and for the gearing of the indicating
element.
As will be evident from the diagram, it is formed basically of two
tangent mechanisms in series so that the rotary motion of one shaft
is converted into rotary motion of a second. This is instead of
converting a linear motion into a rotary one as with a sine or a
tangent mechanism. The input/output relationship is a combined
one involving two trigonometrical conversions; the first is related to
the movement of the contact point between rbds A and Band is
11
given by
h = d(tan 8 02 - tan 8 o 1)
where d is the perpendicular distance between the axis of shaft D and
the plane of contact between A and B, and 8 D is the rotation of D.
The second conversion is given by
h
=/(tan 8c2 -
tan 8c 1 )
where f is the perpendicular distance between the axis of shaft C and
the plane of contact between A and B, and 8c is the rotation of shaft
C. When the planes of movement of rods A and B intersect at right
angles and the rods are straight ones, the combination o f the two
conversions gives the relationship :
d(tan 801 - tan 802) = /(tan 8c1 - tan 8c2).
A variation on the double tangent theme is the skew tangent
mechanism. In this, the rocking shafts are orthogonal but the planes
of the rods A and B do not intersect at right angles.
Gears
The coupling and indicating elements of many aircraft instruments
employ gears in one form or another, for the direct conversion of
straight-line or arc-like motion into full rotary motion , and for
increasing or decreasing the motion. Figure 2.5 illustrates in schemati1
form how gears are applied to an instrument utilizing a multi-pointer
type of indicating element. The sector gear and its meshing pinion
provide for the initial magnification of the measuring element's
displacement. The gear is a small portion of a large geared wheel,
and since it has as many teeth in a few degrees of a~c as the pinion
has completely around it, the sector need only tum a few degrees to
rotate the pinion through a complete revolution. The other gears
shown in Fig 2.5 are designed to provide a definite magnification
Figure 2.5 Gear assembly for
a multi-pointer indicating
element.
SECTOR GEAR ACTUATED
BY MEASUIUNG AND
COUPLING ELEMENTS
12
ratio of movement between their respective pointers and the pointer
actuated by the sector gear and pinion.
In applying gears to instruments and control systems, a problem
which has to be faced is that a gear can always tum a small amount
before it will drive the one in mesh with it. This loss of motion, or
backlash as it is termed, is unavoidable since the dimensioning of the
gear teeth must allow for a set amount of 'play' to avoid jamming
of the gears . Other methods must therefore be found to minimize
the unstable effects which backlash can create.
The method most commonly adopted in geared mechanisms is one
involving the use of a coiled hairspring. The hairspring usually forms
part of an indicating element and is positioned so that one end is
attached to the pointer shaft and the other to the mechanism frame.
In operation, the spring due to tensioning always has a tendency to
unwind so that the inherent play between gear teeth is taken up and
they are maintained in contact.
Another method, and one which is adopted in certain instrument
systems involving the transmission of data, is the anti-backlash gear.
This consists of two identical gears freely mounted face to face on a
common hub and interconnected with each other by means of two
springs so that, in effect, it is a split single gear wheel. Before the
gear is meshed with ijs partner, one half is rotated one or two teeth
thus slightly stretching the springs. After meshing, the springs always
tend to return the two halves of the gear to the static unloaded
position; therefore the face of all teeth are maintained in contact.
The torque exerted by the springs is always greater than the operating
torques of the transmission system so that resilience necessary for
gear action is unaffected.
Hairsprings
Hairsprings are precision-made devices which, in addition to the antibacklash function already referred to, also serve as controlling
devices against which deflecting forces are balanced to establish
required calibration laws (as in electrical moving-coil instruments)
and for the restoration of coupling and indicating elements to their
original positions as and when the deflecting forces are removed.
In the majority of cases, hairsprings are of the flat-coil type with
the inner end fixed to a collet, enabling it to be press-fit ted to its
relevant shaft, the outer end being anchored to an adjacent part of
the mechanism framework. A typical assembly is shown in Fig 2.6 (a),
from which it will be noted that the method of anchoring permits
a certain degree of spring torque adjustment and initial setting of the
indicating element.
In certain types of electrical measuring instruments, provision
must be made for external adjustment of the pointer to the zero
position of the scale. One method commonly adopted, and which
13
Figure 2. 6 Hairsprings.
COLLET ON POINTER
OR GEAR SHAFT
(a) Method of attachment;
(b) method of zero adjustment.
TAPER PIN
ANCHOR POST ON
MECHl<NISM FRAME
(a)
POINTER
JEWELLED
PIVOT SCREW
(b)
- - - ZERO ADJUSTING SCREW
LOCATEO IN BEZEL OR FRONT COVEP
illustrates the principles in general, is shown in Fig 2.6 (b). The
inner end of the spring is secured to the pointer shaft in the normal
way, but the outer end is secured to a circular plate friction-loaded
around the front pivot screw. A fork, which is an integral part of
the plate, engages with a pin eccentrically mounted in a screw at the
front of the instrument. When the screw is rotated it deflects the
plate thus rotating the spring, shaft and pointer to a new position
without altering the torque loading of the spring.
The materials from which hairsprings are made are generally
phosphor-bronze and beryllium-copper, their manufacture calling
for accurate control and grading of thickness, diameter and torque
loading to suit the operating characteristics of particular classes of
instrument.
Temperature
Compensation of
Instrument
Mechanisms
14
In the construction of instrument mechanisms, various metals and
alloys are used, and unavoidably, changes in their physical
characteristics can occur with changes in the temperature of their
surroundings. For some applications deliberate advantage can be
taken of these changes as the basis of operation ; for example, in
certain electrical thermometers the changes in a metal's electrical
resistance forms the basis of temperature measurement. However,
this and other changes in characteristics are not always desirable, and
it therefore becomes necessary to take steps to neutralize those which,
if unchecked, would introduce indication errors due solely to
environmental temperature changes.
The methods adopted for temperature compensation, as it is called,
are varied depending on the type of instrument to which they are
applied. The oldest method of compensation is the one utilizing the
bimetal-strip principle and is applied to such instruments as airspeed
indicators, altimeters, vertical speed indicators, and exhaust-gas
temperature indicators.
Bimetal-strip Method
A bimetal strip, as the name implies, consists of two metals joined
together at their interface to form a single strip. One of the metals
is invar, a form of steel with a 36% nickel content and a negligible
coefficient of linear expansion, while the other metal may be brass
or steel, both of which have high linear expansion coefficients. Thus,
when the strip is subjected to an increase of temperature the brass or
steel will expand, and conversely will contract when the strip is subjected to a decrease of temperature. The invar strip, on the other
hand, on account of it having a negligible expansion coefficient, will
always try to maintain the same length and being firmly joined to the
other metal will cause the whole strip to bend.
An application of the bimetal-strip principle to a typical rod-type
mechanism is shown in Fig 2. 7 (a). In this case, the vertical ranging
bar connected to the rocking shaft is bimetallic and bears against the
Figure 2. 7 Application of
bimetal strip.
INVAR SHIP
(a)
(b)
15
arm coupled to the sector gear of the indicating element.
The principal effect which temperature changes have on this
mechanism is expansion and contraction of the capsule, thus tending
to make the indicating element overread or underread. For example,
let us assume that the positions taken up by the mechanism elements
are those obtaining when measuring a known quantity at the normal
calibration temperature of l 5°C, and that the temperature is gradually
increased. The effect of the increase in temperature on the capsule
material is to make it more flexible so that it will expand further to
carry the ranging bar in the direction indicated by the solid arrows
(Fig 2.7 (b)). As the ranging bar is in contact with the sector gear arm
the indicating element has the tendency to overread. However, the
·increase of temperature has a simultaneous effect on the ranging bar
which, being a bimetal and on account of the position of the invar
portion, will sag, or deflect in the direction indicated by the dotted
arrow, thus counteracting the capsule expansion and keeping the
indicating element at a constant reading. When the temperature is
decreased the capsule material 'stiffens up' and contracts so that the
indicating element tends to underread; as will be apparent from the
diagram, a ccnstant reading would be maintained by the bimetal
ranging bar sagging or deflecting in the opposite direction.
In some instruments, for example exhaust-gas temperature
indicators, indication errors can be introduced due to the effects
of environmental temperature on the values of the electromotive
force produced by a thermocouple system.
Although such errors ultimately result from changes in an
electrical quantity, compensation can also be effected mechanically
and by the application of the bimetal-strip principle. As, however,
the operation of the method is closely connected with the operating
principles of thermo-electric instruments, we shall study it in detail
at the appropriate stage.
Thermo-resistance Method
For temperature measurements in aircraft, many of the instruments employec:I are of the electrical moving-coil type, and as the
coil material is usually either copper or aluminium, changes of
indicator temperature can cause changes in electrical resistance of
the material. We shall be studying the fundamental principles of
moving-coil instruments in a later chapter, but at this point we may
note that, as they depend for their operation on electric current,
which is governed by resistance, the effects of temperature can
result in indication errors which necessitate compensation.
One of the compensation methods adopted utilizes a thermoresistor or thermistor connected in the indicator circuit. A
thermistor, which is composed of a mixture of metallic oxides, has a
16
very large temperature coefficient of resistance which is usually
negative ; i.e. its resistaf.lce decreases with increases in temperature.
Assuming that the temperature of the indicator increases, the
current flowing through the indicator will be reduced because
copper or aluminium will characteristically increase in resistance ; the
indicator will therefore tend to underread. The thermistor
resistance will, on the other hand , decrease, so that for the same
temperature change the resistance changes will balance out to maintain a constant current and therefore a constant indication of the
quantity being measured.
Thermo-magnetic Shunt Method
As an alternative to the thermistor method of compensating for
moving-coil resistance changes, some temperature measuring
instruments utilize a device known as a thermo-magnetic shunt.
This is a strip of nickel-iron alloy sensitive to temperature changes,
which is clamped across the poles of the permanent magnet so that
it diverts some of the airgap magnetic flux through itself.
As before, Jet us assume that the indicator temperature increases.
The moving-coil _resistance will increase thus opposing the current
flowing through the coil, but, at the same time, the reluctance
('magnetic resistance') of the alloy strip will also increase so that less
flux is diverted 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 current to
maintain a constant torque and indicated reading. Depending on the
size of the permanent magnet, a number of thermo-magnetic strips
may be fitt ed to effect the required compensation.
Sealing of Instruments
Against Atmospheric
Effects
In pressurized aircraft, the internal atmospheric pressure conditions
are increased to a value greater than that prevailing at the altitude at
which the aircraft is flying. Consequently, instruments using
external atmospheric pressure as a datum, for example altimeters,
vertical speed indicators and airspeed indicators, are liable to
inaccuracies in their readings should air at cabin pressure enter their
cases. The cases are therefore sealed to withstand external pressures
higher than those normally encountered under pressurized conditions.
The external pressure against which sealing is effective is normally
15 lbf/in2 •
Direct-reading pressure measuring instruments of the Bourdon
tube, or capsule type , connected to a pressure source outside the
pressure cabin, are also liable to errors. Such errors are corrected by
using sealed cases and venting them to outside atmospheric pressure.
Many of the instruments in current use depend for their
17
operation on sensitive electrical circuits and mechanisms which must
be protected against the adverse effects of atmospheric temperature,
pressure and humidity. This protection is afforded by filling the
cases with an inert gas such as nitrogen or helium, and then
hermetically sealing the cases.
Questions
2.1
2.2
2.3
2.4
2.5
18
What are the four principal elements which make up an instrument?Define lever length and lever angle and state what effects they have
on an instrument utilizing a lever mechanism.
(a) What are the essential differences between a lever mechanism and
a rod mechanism? (b) State some typical applications to aircraft
instruments.
What do you understand by the term 'lost motion'? Describe the
methods adopted for minimizing its effects.
Describe a method by which instru·m ent indications may be automatically corrected for temperature variations.
3 Instrument displays,
panels and layouts
In flight, an aeroplane 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
the loop is that of controller, and the extent of the control function
is governed by the simplicity or otherwise of the machine as an
integrated whole . For example, in manually flying an aeroplane,
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 aeroplane and adjustments to essential systems
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
content and form of the data display.
The most common forms of data display applied to aircraft
instruments are (a) quantitative, in which the variable quantity being
measured is presented in terms of a numerical value and by the
relative position of a pointer or index, and (b) qualitative, in which
the information is presented in symbolic or pictorial form.
Quantitative Displays
There are three principal methods by which information may be
displayed: (i) the circular scale, or more familiarly, the 'clock' type
of scale, (ii) straight scale, and (iii) digital, or counter. Let us now
consider these three methods in detail.
Circular Scale
This may be considered as the classical method of displaying information in quantitative form and is illustrated in Fig 3. 1
The scale base, or graduation circle, refers to the line, which may
be actual or implied, running from end to end of the scale and from
19
Figure 3. 1 Circular scale
quantitative display
SCALE Oil
GRADUATION MARK
which the scale marks and line of travel of the pointer are defined.
Scale marks, or graduation marks, are the marks which constitute
the scale of the instrument. For quantitative displays it is of extreme
importance that the number of marks be chosen carefully in order
to obtain quick and accurate interpretations of readings. If there are
too few marks dividing the scale, vital infonnation may be lost and
reading errors may occur. If, on the other hand, there are too many
marks, time will be wasted since speed of reading decreases as the
number of markings increases. Moreover, an observer may get a
spurious sense of accuracy if the number of scale marks makes it
possible to read the scale accurately to, say, one unit ( the smallest
unit marked) when in actual fact the instrument has an inherent
error causing it to be accurate to, say , two units. As far as
quantitative-display aircraft instruments are concerned, a simple
rule followed by manufacturers is to divide scales so that the marks
represent units of 1, 2 or 5 or decimal multiples thereof. The sizes
of the marks are also important and the general principle adopted is
that the marks which are to be numbered are the largest while those
in between are shorter and usually all of the same length.
Spacing of the marks is also of great importance, but since it is
governed by physical laws related to the quantity to be measured,
there cannot be complete uniformity between all quantitative
displays. In general, however, we do find that they fall into two
distinct groups, linear and non-linear; in other words, scales with
marks evenly and non-evenly spa<:ed. Typical examples are
illustrated in Fig 3.2, from which it will also be noted that nonlinear 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. In an instrument having a centre zero, this rule
would, of course, only apply to the positive scale. 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
20
Figure J.2 Linear and nonlinear scales. (a) Linear;
(b) square-law; (c) logarithmic.
(b)
(a)
\
1
RATE OF
CLIMB.
(cl
inside or outside the scale base; the latter. method is preferable since
the numbers are not covered by the pointer during its travel over the
scale.
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. Under
ideal conditions and purely from theoretical considerations, it has
been calculated that the length of a scale designed for observing at a
distance of 30 in and capable of being read to 'l % of the total
indicated quantity, should be about 2 in (regardless of its shape).
This means that for a circular-scale instrument a 1 in diameter case
would be sufficient. However, aircraft instruments must retain their
legibility in conditions which at times may be far from ideal conditions of changing light, vibrations imparted to the instrument
panel, etc. In consequence, some degree of standardization of
instrument case sizes was evolved, the utilization of such cases being
dictated by the reading accuracy and the frequency at which
observations are required. Instruments displaying information which
is to be read accurately and at frequent intervals have scales about
7 in in length fitting into standard 3Y.i in cases, while those requiring
only occasional observation, or from which only approximate
readings are required, have shorter scales and fit into smaller cases.
High-Range Long-Scale Displays
For the measurement of some quantities, for example, turbine21
engine rev ./min. , 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 to the
difficulties encountered are illustrated in Fig 3.3. 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 mis-readirig, a presentation by two interconnected pointers of different sizes is much better. A pr~ctical
example of this presentation is to be found in some current designs
of turbine-engine rev./min. indicator. In this instance a large pointer
rotates against the outer scale to indicate hundreds of rev./min. and
at the same time it rotates a smaller pointer against the inner scale
indicating thousands of rev./min.
Figure 3.3 High-range long·
I
scale displays. (a) Concentric
scales; (b) fixed and rotating
scales; (c) common scale,
triple pointers; (d) split
pointer.
90
"'80
°
<;J;)
10
20"
IV
--70
~-40
60
' sp '
22
(a}
(bl
(c}
(d}
In Fig 3.3 (b), we find a method which is employed in a certain
type of ai rspeed indica tor; in its basic concept it is similar to the one
just described. In this design, however, a single pointer rotates against
a circular scale and drives a second scale instead of a pointer. This
rotating scale, which records.hundreds of miles per hour as the
pointer rotates through complete revolutions, is visible through an
aperture in the main dial.
A third method of presentation, shown at (c) , is one in which
three ..:oncentric pointers of different sizes register against a common
scale. The application of this presentation has been confined mainly
to altimeters, the large pointer indicating hundreds, the intermediate
pointer thousands and the small pointer tens of thousands of feet.
This method of presentation suffers several disadvantages the
principal of which are that it takes too long to interpret a reading and
gives rise to too frequent and too serious mis-reading.
Figure 3.3 (d) illustrates a comparatively recent presentation
method applied to airspeed measurement. It will be noted that an
outer and an inner scale are adopted and also what appears to be a
single pointer. There are, however, two pointers which move
together and register against the outer scale during their first
revolution. When this has been completed, the tip of the longer
pointer of the two is covered by a small plate and its movement beyond this point of the scale is arrested. The shorter pointer continues
its movement to register against the inner scale.
Angle of Observation
Another factor which has an important bearing on the choice of the
correct scale length and case size is the angle at which an instrument
is to be observed. It is important because, even though it would be
possible to utilize longer scales in the same relevant case sizes, the
scale would be positioned so close to the outer edge of the dial ptate
that it would be obscured when observed at an angle. For this reason,
a standard is also laid down that no part of an instrument should be
obscured by the instrument case when observed at angles up to 30°
from the normal. A method adopted by some manufacturers, which
conforms to this standard, is the fitting of instrument mechanisms
inside square cases.
When observing an instrument at an angle errors due to parallax
are, of course, possible, the magnitude of such errors being governed
principally by the angle at which the relevant part of its scale is
observed , and also by the clearance distance between the pointer and
dial plate. This problem like so many others in the instrument field
has not gone unchallenged and the result is the 'platform' scale
designed for certain types of circular display instruments. As may be
seen from Fig 3.4, the scale marks are set out on a circular platfonn
23
A
SCALE ~
~ [!¥&taw~
~
DIAL P L A T ~
ENLARGED SECTION
THROUGH A - A
A
FYgure 3.4 Platform scale
which is secured to the main dial plate so that it is raised to the same
level as the tip of the pointer.
Scale Range and Operating Range
A point quite often raised in connection with instrument scale
lengths and ranges is that they usually exceed that actually required
for the operating range of the system with which the instrument is
associated thus leaving part of the scale unused. At first sight this
does appear to be 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-~0 lbf/in 2 • 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 opt:rating range, also divided into a convenient number of parts as shown in Fig 3.5 (a). However, under
certain.operating conditions of the system concerned, it may be
FYgure 3.5 Reading accuracy.
(a) Equal scale length and
operating range; (b) scale
range exceeding operating
range.
17
\
, 29
(a)
24
(b)
essential to monitor pressures having such values as 17 or 29 lbf/ in 2
and to do this accurately in the shortest possible time is not very
easy, as a second glance at the diagram will show. Let us now
redesign the scale so that its length and range exceed the system's
operating range and set out the scale marks according to the rule
given on page 20. The result shown at (b) clearly indicates how much
easier it is to interpret the values we have considered it essential to
monitor.
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
Figure 3. 6 Comparison
between moving-tape and
circular scale displays.
3
ENGINE NO.
% R.P.M.
Engine
No.
E.G.T.
l
500
2
3
4
470
480
520
oc
EG T "C
% R.P.M
<100
. 10
%
R.P.M.
89
90
88
90
1 2
3 4
1 2
3 4
25
is from bottom to top or from left to right.
In the field of aircraft instruments there are very few applications
of the straight scale and pointer displays, as they are not suitable
for the monitoring of the majority of quantities to be measured.
However, they do possess characteristics which can contribute to the
saving of panel space and improved observational accuracy,
particularly where the problems of grouping and monitoring a large
number of engine instruments is concerned.
The development of these characteristics, and investigations into
grouping and monitoring problems, have resulted in the practical
application of another variation of the straight scale display. This is
~nown as the moving-tape or 'thermometer' display and is illustrated
in Fig 3.6 as it would be applied to the measurement of two parameters vital to the operation of an aircraft powered by four turbojet
engines.
Each display unit contains a servo-driven white tape in place of a
pointer, which moves in a vertical plane and registers against a scale
in a similar manner to the mercury column of a thermometer. As will
be noted there is one display unit for each parameter, the scales being
common to all four engines. 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 than from the classical ·
circular scale and pointer display . This fact, and the fact that panel
space can be considerably reduced, are also clearly evident from
Fig 3.6.
Digital Display
A digital or veeder-counter type of display is one in which data are
presented in the form of letters or numbers-alpha-numeric display,
as it is technically termed . In aircraft instrument practice, the latter
Figure 3. 7 Application
of digital diSplays.
DYNAMIC COUNTER
DISPLAY
I
STATIC COUNTER
DISPLAY
26
presentation is the most common and a counter is generally to be
found, operating in combination with the circular type of display.
Typical examples are shown in Fig 3.7. In the application to the
altimeter there are two counters; one pr~sents a fixed pressure value
which can be set mechanically by the pilot as and when required, and
is known as a static counter display ; the other is geared to the altimeter mechanism and automatically presents changes in altitude, and
-is therefore known as a dynamic counter display. 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 discussed on page 22. The counter of the TGT indicator is
also a dynamic display since it is driven by a servo transmission to
the main pointer (see also page 292).
Dual-Indicator Displays
Filurt 3. 8 Examples of
dual-Indicator displays.
Dual-indicator displays are designed principally as a means of
conserving panel space, particularly where the measurement of the
various quantities related to engines is concerned. They are n.ormally
of two basic forms: one in which two separate indicators and scales
are embodied in one case; and the other, also having two indicators
in one case, but with the pointers registering against a common scale.
Typical examples of display combinations are illustrated in Fig 3 .8.
MIASUR!MINT
PIIUENTATION
"·
TWO OIFFEAENT QUANTITIES
OF ONE SYSTEM
8.
SAME QUANTITY OF TWO
OIFFEREN T SYSTEMS
c.
SAME QUANTITIE S OF TWO
IDENTICAL SYSTEMS
27
Coloured Displays
The use of colour in displays can add much to their value; not, of
course from the artistic standpoint, but-as a means of indidating
specific operational ranges of the systems with which they are
associated and to assist in making more rapid assessment of
conditions prevailing when scanning the instruments.
Colour may be applied to scales in the form of sectors and arcs
which embrace the number of scale marks appropriate to the
required part of the range, and in the form of radial lines coinciding
with appropriate individual scale marks. A typical example is
illustrated in Fig 3.9. It is usual to find that coloured sectors,are
applied to those parts of a range in which it is sufficient to know
that a certain condition has been reached rather than knowing actual
quantitative values. The colours chosen may be red, yellow or green
depending on the condition to be monitored. For example, in an
aircraft oxygen system it may be necessary for the cylinders to be
charged when the pressure has dropped to below, say, 500. lbf/in2 •
The system pressure gauge would therefore have a red sector on its
dial embracing the marks from O to 500; thus, if the pointer should
register within this sector, this alone is sufficient indication that
recharging is necessary and that it is only of secondary importance to
know what the actual pressure is.
Figure 3. 9 Use of colour in
instrument displays. White
arc 75- 140; Green arc
95-225; Yellow arc 225255 ; Red radial line 255.
Arcs and radial lines are usually called range markings, their
purpose being to define values at various points in the range of a scale
which are related to specific operational ranges of an aircraft, its
power plants and systems. The definitions of these marks are as
follows:
RED radial line
YELLOW arc
GREEN arc
RED arc
28
Maximum and minimum limits
Take-off and precautionary ranges
Normal operating range
Range in which operation is prohibited
When applied to fuel quantity indicators, a RED arc indicates fuel
which cannot be used safely in flight.
Airspeed indicator dials may also have an additional WHITE arc.
This serves to indicate the airspeed range over which the aircraft
landing flaps may be extended in the take-off, approach and landing
configurations of the aircraft.
Range markings may vary for different types of aircraft and are
therefore added by the aircraft manufacturer prior to installation in
their production aircraft.
It may often be found that markings are painted directly on the
cover glasses of instruments - a method which is simple.r since it does
not require removal of an instrument mechanism from its case.
· However, the precaution is always taken of painting a white index or
register line half on the cover glass and half on the bezel to ensure
correct alignment of the glass and the markings over the scale marks.
In addition to the foregoing applications, colour may also be used
to facilitate the identification of instruments with the systems in
which they are connected. For example, in one type of aircraft
currently in service, triple hydraulic systems are employed, designated
yellow system, green system and blue system, and in order to identify
the pressure indicators of each system the scales are set out on dials
painted in the relevant colours.
Qualitative Disµlays
Figure:]. IO Qualitative
displays. (a) Engine
synchronizing; (b) position
of flight control surfaces.
(a)
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, the movement of a
component and so on. Two typical examples are shown in Fig 3.10.
The synchroscope at (a) is used in conjunction with a rev./min.
indicating system of an aircraft having a multiple arrangement of
(bl
29
propeller-type engines, and its pointers, which symbolize the propellers, only rotate to show the differences of speed between engines.
The display, shown at (b), is a good example of one indicating the
movement of components; in this case, flight control surfaces,
landing flaps, and air spoilers. The instrument contains seventeen
separate electrical mechanisms, which on being actuated by transmitters, position symbolic indicating elements so as to appear at
various angles behind apertures in the main dial.
Director Displays
30
Director displays are th0se which are associated principally with fligh
attitude and navigational data (see Chapter 15), and presenting it in a
manner which indicates to a pilot what control movements he must
make either to correct any departure from a desired flight path, or to
cause the aircraft to perform a specific manoeuvre. It is thus
apparent that in the development of this type of display there must
be a close relationship between the direction of control movements
and the instrument pointer or symbolic-type indicating element ; in
other words, movements should be in the 'natural' sense in order
that the pilot may obey the 'directives' or 'demands' of the display.
Although flight director displays are of comparatively recent origin
as specialized integrated instrument systems of present-day aircraft,
in concept they are not new. The gyro horizon (see page 127) which
has been in use for many years utilizes in basic form a director
display of an aircraft's pitch and bank attitude. In this instrument
there are three elements making up the display: a pointer registering
against a bank-angle scale, an element symbolizing the aircraft, and
an element symbolizing the natural horizon. Both the bank pointer
and natural horizon symbol are stabilized 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 the aircraft, and that of the moving element. Assuming
that the aircraft's pitch attitude changes to bring the nose up, then
the horizon display will be shown as in Fig 3.11 (a), thus directing
or demanding the pilot to 'get the nose down'. Similarly, if the bank
attitude should change whereby the left wing goes down, then the
horizon display would be as shown at (b), directing or demanding the
pilot to 'bank the aircraft to the right.' In both cases, the demands
would be satisfied by the pilot moving his controls in the natural
sense.
Another example of a director display is that utilized in an indicator used in conjunction with the Instrument Landing System (ILS);
this is a radio navigation system which aids a pilot in maintaining the
correct position of his aircraft during the approach to land on an
airport runway. Two radio signal beams are transmitted from the
Figure 3. 11 Examples of
director display. (a) 'Fly
down' directive; (b) 'bank
right' directive; (c) 'fly left'
and 'fly up' directive;
(d) response matches
directive.
NATURAL HORIZON
SYMBOLIC ELEMENT
(b)
fa)
GYRO HORIZON
GLIDE SLOPE
POINTER ---l-lllot----
-=--e..:::----1
•
•
AIRCRAFT
SYMBOL
LOCALISER
(RUNWAY!
SYMBOL
(C)
(d)
ground; one beam is in the vertical plane and at an angle to the runway to establish the correct approach or glide slope angle; while the
other, known as the localizer, is in the horizontal plane; both are
lined up with the runway centre-line.
A receiver on board the aircraft receives the signals and transmits
them to two meters contained within the indicator; one meter controls a glide slope pointer, and the other a localizer pointer. In the
majority of current types of aircraft, the ILS directive display is
always presented on two indicators comprising what is termed either
an Integrated Instrument System or a Flight Director System (see
Chapter 15). A typical presentation of one of these indicators
(referred to as an attitude indicator) is also sho_wn in Fig 3.11. As
will be noted, it combines a gyro horizon directive display, and thereby eliminates the need for a pilot having to monitor the indications
of two separate instruments.
When the aircraft is on the approach to land and is, say, below the
31
glide slope beam, the glide slope pointer of the instrument will be
deflected upwards as shown in Fig 3.11 (c). Thus, the pilot is
directed to 'fly the aircraft up' in o rder to intercept the beam.
Similarly,.if the aircraft is to the right of the localizer beam the
localizer pointer will be deflected to the left thus directing the pilot
to 'fly the aircraft left'. As the pilot responds to the instrument's
directives the pointers move back to their centre positions indicating
that the aircraft is in the correct approach position for landing.
(Fig 3.11 (d)).
It will be apparent from the diagram that as the aircraft is
manoeuvred in response to demands, the pointer movements are
contrary to the 'natural' sense requirements; for example, in
responding to the demand 'fly left' the localizer pointer will move to
the right. However, in turning to the left the bank attitude of the
aircraft will change into the direction of the tum, and as this will be
indicated directly by the gyro horizon display, the response to the
ILS demands can be readily cross-checked.
Head-Up Displays
32
From the descriptions thus far given of the various instrument
displays, we. have gained some idea of the development approach to
the problem of presenting data which are to be quickly and
accurately assimilated. The simplicity or otherwise of assimilation is
dictated by the number of instruments involv.ed, and by the amount
of work and instrument monitoring sequences to be performed by a
pilot during the various phases of flight. In the critical approach
and landing phase, a pilot must transfer his attention more frequent!
from the instruments to references outside the aircraft, and back
again; a transition process which is time-consuming and fatiguing as:
result of constant re-focusing of the eyes.
A method of alleviating these·problems has therefore been
developed in which vital flight data are presented at the same level a!
. the pilot's line-of-sight when viewing external references, i.e. when h
is maintaining a 'head-up' position. The principle of the method is t
display the data on the face of a special cathode-ray tube and to
project them optically as a composite symbolic image on to a
transparent reflector plate, or directly on the windscreen. The
components of a typical head-up display system are shown in Fig
3.12. The amount of data required is governed by the requirements
of the various flight phases and operational role of an aircraft, i.e.
civil or military, but the four parameters shown are basic. The data
are transmitted from a data computer unit to the cathode-ray tube
the presentation of which is projected by the optical system to
infinity. It will be noted that the attitude presentation resembles
that of a normal gyro horizon, and also that airspeed and altitude ar
presented by markers which register against linear horizontal and
DIRECTOR
DOT
WINDSCREEN
\ ___ =i
I
______. I
---i
I
«"'"
DATAINPVT
Figure 3.12 Head-up
display system.
vertical scales. The length. or range, of the scales is determined by
operational requirements, but normally they only cover narrow
bands of airspeed and altitude information. This helps to reduce
irrelevant markings and the time taken to read and interpret the
information presented. .
Figure 3.13 illustrates the head-up display of a system known as a
Visual Approach Monitor (V AM). The system is an airborne
equivalent of the Visual Approach Slope Indicator (VASI) and was
evolved principally as an aid to pilots when approaching airfields not
equipped with VASI, ILS, or other approach aids.
LENS
Figure 3.13 Visual approach
APPROACH ANGLE SCALES
monitor display.
3,f
+
+
+
6-
BAR SYMBOLIZING
AIR CRAFT
' ~ 3
+
AIRflELO RUNWAY
-6
SPEED CUE
C ON SPEED (GREEN)
F FAST !AMBERJ
S SL OW {RED)
33
The display unit is mounted on a sliding tray located at a glare
shield panel in front of the pilot, and when required, the tray is pullec
out to automatically raise the lens through which the display is
projected. The display provides the following cues : vertical approach
angle, horizontal attitude, and speed error. Approach angle is
displayed by two vertical scal·es, one each side of the lens, and
graduated in degrees above and below the fixed bar symbolizing the
aircraft. The required angle is selected on a control module to
position the scales relative to the bar. During an approach, the pilot
holds the bar symbol in his line of sight and controls !he aircraft so
that the symbol is aligned with the runway threshold or touchdown
zone, thereby ensuring the approach is at the selected flight path
angle.
The speed error cue is in the form of three symbols which are
coloured to indicate whether the aircraft approach speed is correct,
too fast, or too slow (see Fig 3. 13)°. The error between actual airspeed
and that selected on the control module is indicated by variations in
the light intensity of the three symbols. For example, if the approach
airspeed drops to eight knots or more below the selected airspeed,
this is displayed by the red symbol 'S' appearing at full intensity and
flashing on and off continuously .
Light-Emitting
Displays
In the continuing development of aircraft displays, the trend has
been to exploit the techniques applied to those instruments which are
taken so much for granted these days; namely, the pocket calculator
and the digital watch. The displays adopted in both these instruments
are of the 'light-emitting' type the basis of which, in its turn, has its
origin in the well-known cathode-ray display.
There are several ways in which numerical data can be displayed by
means of light-emission, but the o nes which are of interest in this
context are the liquid-crystal display and the light-emitting diode
display.
Liquid-Crystal Display (LCD)
The basic structure of an LCD (see Fig 3.14) consists of two glass
plates, coated on their inner surfaces with a thin transparent
conductor such as indit..m oxide. The conductor on the front plate is
etched into a standard display for~at of seven bars or segments each
segment forming an electrode. Each bar is electrically separate and
is selected by a logic/driver circuit which causes the bars to illuminate
in patterns forming the digit to be.displayed (diagram (b)). A mirror
image of the digits with its associated electrical contact is also etched
into the oxide layer of the back glass plate, but this is not segmented
since it constitutes a common return for all segments.
34
Figure3.14 Liquid crystal
FRONT PLATE
display (LCD).
\
(al
LIQUID CRYSTAL
LAYER
(TYPICAL SPACING
• 10 MICRONS)
BACK PLATE
MIRROR IMAGE
(NOT SEGMENTED)
COMMON RETURN CONTACT
-:I -=,
Inu I' L..
-'
(b)
NO. OF SEGMENTS
6
2
S
5
4 '-_, CJ~ , I 8I Cl-'
4
5
6
3
7
6
The space between the plates is filled with a liquid-crystal
material, referred to as a nematic material (from the Greek word
nemator meaning 'thread' ) by virtue of its thread-like molecules
being oriented with their long axes parallel. The complete assembly
thus constitutes a special form of capacitor. When low-voltage current
is applied to the segments, the molecular order of the liquid-crystal
material is disturbed and this changes its optical appearance from
transparent to reflective. The magnitude of the optical change (called
contrast ratio) is basically a measure of the light reflected from, or
transmitted through, the segment area, to the light reflected from the
background area ; a typical ratio is 15: I. The current applied to the
segments is of the alternating type to avoid undesirable electrolytic
effects. Energizing of the segments is accomplished by the simultaneous application of a symmetrical out-of-phase signal to the front and
back electrodes of a segment, and thereby producing a net voltage
difference. When two in-phase signals are applied the display
segments spontaneously relax to the de-energized state.
An LCD may be of either the dynamic-scattering type or the
field-effect type, and these, in turn, may_produce either a trans35
missive or a reflective read-out. The dynamic-scattering display
operates on the principle of forward light scattering, which is caused
by turbulence of the ions of the liquid-crystal material when
current is applied to the segment electrodes. For a transmissive
read-out, a back-light source is provided, the light being directed
down by a light-control film similar in its action to a venetian blind.
In the area defined by the energized segment, the light is then
scattered up toward the observer to produce a light digit or
character on a dark background. For a reflective read-out, the lightcontrol film is replaced by a mirror. It also depends on forward
light scattering, but the source of normal ambient light is used
exclusively to produce a light digit on a dark background.
A field-effect LCD incorporates additional plates called
polarizers on the front and back glass plates of the assembly. It also
contains a specially prepared inside glass surface which causes the
liquid-crystal molecules to orient themselves in a 90° 'twist' configuration between the glass plates. This molecular configuration causes
the plane of polarization (polarized after passing through the front
polarizer) to be reflected by 90° as it passes through the LCD. Then,
depending on the orientation of the polarizer behind the back glass
plate, the LCD can be made either transmissive or opaque in the
de-energized state.
Depending on the application of an LCD, colour effects can be
achieved by the proper placement of colour films on the front surface
of the display, between the back surface and artificial light source, or
by colouring the reflective surfaces.
Light-Emitting Diode (LED)
An LED is essentially a transistor and so, unlike an LCD, it is classified as a solid-state display; the construction is shown in Fig 3.15.
The heart of the display is a slice or chip of gallium arsenide phosphide (GaAsP) moulded into a transparent plastic covering which not
only serves to protect the chip, but also as a diffuser lens. The diode
leads are soldered to a printed circuit board to form the numerical
display required,' e.g. the digit segment already referred to. When
current flows through the chip it produces light which is directly
transmitted in proportion to the current flow. To provide different
colours, the proportion of GaP and GaAs is varied during manufacture of the chip, and also the technique of 'doping' with other
elements e.g. oxygen or nitrogen is applied.
In the normal 7-bar or segment display format, it is usual to
employ one LED per segment, but the number depends on the overall
size of the digits required for display and its appearance. Two
methods may be adopted for increasing the size and for improving
the appearance of a single LED per segment display. In one a core36
Figure 3. 15 Light-emitting
diode (LED).
CRYSTAL CHIP
PROTECTIVE COVER/DIFFUSER LENS
~
'
CONNECTIONS
Figure 3.16 LED vertical
scale display.
shaped reflecting cavity known as a 'light-pipe' is placed over each
LED with its -small end down. The whole assembly is cast inside a
housing using glass-filled epoxy which fills the light-pipe cavities.
When each LED is illuminated, the light is reflected off the glass
particles within the epoxy and off the cavity sidewalls, and through a
surface area which can be up to twenty times that of the LED at the
bottom of the cavity. In the second method, several small chips
(10-15 mm square) are covered with a metallized plastic reflector
having seven bar-shaped reflector cavities designed so that an LED
is at the centre of each cavity.
In addition to the foregoing digital readouts, LED's can also be
adopted for circular scale or vertical scale displays, thereby eliminating the use of conventional pointers or moving tapes. In this case,
large numbers of individual diodes (900 would be typical) are
arranged in groups to form illuminated bars. An example of a
vertical scale employing this technique for the measurement of
temperature is shown in Fig 3.16. Signals from the appropriate
sensing elements (thermocouples in the example shown) are first
converted into digital form and read into a digital computer using
microprocessors. The computer then applies the appropriate scale
factors, establishes the required illumination pattern, and then
produces an output signal to the display such that a small column
of illuminated bars moves up or down the scale and registers against
the scale. The light intensity of the LED column is graded, the
highest intensity being provided at the 'reading edge' of the column.
37
Instrument Panels
and Layouts
Figure J. I 7 Location of
instrument panels in a turbo·
jet airliner. (By courtesy of
Laker Airways (Services) Ltd.)
38
All instruments essential to th e operation of an ai rcraft are
accommodated on special panels the number and distribution of
which vary in accordance with the number of instnimerits, the size
of aircraft and cockpit layout. A main instrument panel positioned
in front of pilots'is a feature common ·to all types of aircraft, since it
is manda tory for the primary flight instruments to be installed within
the pilots' normal line of vision (see Fig 3.17.) Typical positions of
other panels are: overhead, at th e side , and on a control pedestal
located centrally between th e pilots.
Panels are invariably of light alloy of sufficient strength and
rigidity to accommodate the required number of instruments , and are
attached to the appropriate parts of the cockpit struc ture. The
attachment methods adopted vary, but all should conform to the
requirement that a panel or an individual instrument should be easily
installed and removed.
Main instrument panels, which may be of the single-unit type or
made up of two or three sub-panel assemblies, are supported on
shockproof mountings since they accommodate the flight instruments and their sensitive mechanisms. The number, size an.d
disposition of shockproof mountings required are governed by the
size of panel and distribution of the total weight.
All panels are normally mounted in the vertical position, although
in some current aircraft types the practice of sloping main instrument panels forward at about 15° from the vertical is adopted to
minimize parallax errors.
Instrument and all other control panels which for many years were
painted black, are now invariably finished in matt grey, a colour
which apart from its 'softer' effects provides a far better contrasting
background for the instrument dials and thus contributes to easier
identification.
Instrument Grouping
Flight Instruments
Basically there are six flight instruments whose indications are so
co-ordinated as to create a 'picture' of an airc raft's flight condition
and required control movements ; they are, airspeed indicator,
altimeter, gyro horizon, direction indicator, vertical speed indicator
and turn-and-bank indicator. It is therefore most important for these
instruments to be properly grouped to maintain co-ordination and to
assist a pilot t o observe them with the minimum of effort.
The first real attempt at establishing a standard method of grouping was the 'blind flying panel' or 'basic six' layout shown in Fig 3.18 (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 ·c ontrol 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 init iat ed by ball'king an aircraft, and
the degree of heading change is obtained from the direction indicator ;
this instrument therefore supports the interpretation of roll attitude
and is positioned directly below the gyro horizon. The turn-andbank 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,
fligh t instruments and integrated instrument systems, it became
necessary to review the functions of ce rtain instruments and their
relative positions within the group. As a result a grouping known as
the 'basic T' was introduced (Fig 3. 18 (b )). 111e theory behind this
method is that it constitutes a system by which various items of
related flight information can be placed in certain sta ndard locations
in all instrument panels regardless of type or make of instrument
used. In this manner, advantage can be taken of integrated instru
ments which display more than one item of flight information.
It will be noted that there are now four 'key' instniments, airspeed
indicator, pitch and roll attitude indicator, an altimeter forming the
39
Figure 3.18 Flight instrument grouping. {a) 'Basic
six'; {b) 'basic T'.
(al
(b)
horizontal bar of the 'T', and the direction indicator forming the
vertical bar. As far as the positions flanking the direction indicator
are concerned, they are taken by other but less specifically essential
flight instruments, and there is a certain degree of freedom in the
choice of function. From Fig 3 .18 it can be seen, for example,
that a Machmeter and a radio magnetic indicator can take precedence
over a tum-and-bank indicator and a vertical-speed indicator.
40
Border lines are usually painted on the panel around the flight
instrument groups. These are referred to as 'mental focus lines' ,
their purpose being to assist pilots in focusing their attention on
and mentally recording the position of instruments within the groups.
Power-Plant Instruments
The specific grouping of instruments required for the operation of
power plants is governed primarily by the type of power plant, the
size of the aircraft and therefore the space available. In a singleengined aircraft, this does not present much of a problem since the
small number of instruments may flank the pilot's flight instruments
thus keeping them within· a small 'scanning range'
The problem is more acute in multi-engined aircraft; duplication
of power plants means duplication of their essential instrnments. 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 instniments.
Figure 3./9 Instrument
grouping at a night engineer' s station. (By courtesy
of Laker Airways (Services)
Ltd.)
41
In some large types of public transport aircraft, a flight engineer's
station is provided in the crew compartment and all the power plant
instruments are grouped on the control panels at this station
(Fig 3. 19). Those instruments measuring parameters required to be
Figure 3.20 (a) Power plant
instrument grouping .
I
I
No.I
I
I
I
I
(!) C) . . C) d)
00~:=~00
I
I
I
I
~~·~~·~0
~
I
42
~M~RATU~
I
\;;J;))
I
Figure 3.20 (b)
Power plant instrument
grouping - Boeing 747.
(By courtesy of Smiths
Industries.)
ENGINE PRESSURE
RATIO
R.P.M. (N 2 )
EXHAUST GAS
TEMPERATURE
FUEL FLOW
known by a pilot during take-off, cruising and landing, e.g. rev./min.
and turbine temperature, are duplicated on the main instrument
panel.
The positions of the instruments in the power plant group are
arranged so that those relating to each power plant correspond to
the power plant positions as seen in plan view. It will be apparent
from the layout of Fig 3.20 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 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. A practical example of this method of grouping as
adopted in the Boeing 747 is shown in Fig 3.20 (b).
Methods of Mounting
Instruments
The two methods most commonly used for the panel mounting of
instruments are the flanged case method, and the clamp method.
The flanged case method requires the use of screws inserted into
locking nuts which, in some instruments, are fitted integral with the
flange.
Since flanged-type indicators are normally mounted from the rear
of the panel, it is clear that the integrally fixed locking nuts provide
43
for much quicker mounting of an instrument and overcome the
frustration of trying to locate a screw in the ever-elusive nut!
As a result of the development of the hermetic sealing technique
for instruments, the cases of certain types are flangeless, permitting
them to be mounted from the front of the instrument panel. In order
to secure the instruments special clamps are provided at each cut-out
location. The clamps are shaped to suit the type of case, i.e. circular
or square, and they are fixed on the rear face of the panel so that
when an instrument is in position it is locatec;l inside the clamp.
Clamping of the instrument is effected by rotating adjusting screws
which draw the clamp bands tightly around the case.
Magnetic Indicators
and 'Flow Lines'
44
In many types of aircraft nume.r ous valves, actuators and similar
devices are used in many of their systems to obtain the desired
control of system operation ; for example, in a fuel system, actuators
position valves which permit the supply of fuel from the main tanks
to the engines and aiso cross-feed the fuel supply.
All such devices are, in the majority of cases, electrically operated
and controlled by switches on the appropriate systems panel , and to
confirm the completion of movement of the device an indicating
system is necessary.
l11e indicating system could be in the form of a scale and pointer
type of instrument or an indicator light. Both methods, however,
have certain disadvantages. The use of an instrument is rather spaceconsuming, particularly where a number of actuating devices are
involved, and unless it is essential for a pilot or systems engineer to
know exactly the position of a device at any one time, instruments
are uneconomical. Indicator lights are of course simpler, cheaper and
consume less power, but the liability of their filaments to failure
without warning constitutes a hazard , particularly in the case where
'light out' is intended to indicate a safe condition of a system.
Furthermore, in systems requiring a series of constant indications of
prevailing conditions, constantly illuminated lamps can lead to
confusion, misinterpretation (frustration also!) on the part of the
pilot or engineer.
Therefore to enhance the reliability of indication, indicators
containing small electromagnets operating a shutter or similar moving
element are installed on the systems panels of many present-day
aircraft.
In its simplest form (Fig 3.21 (a)) a magnetic indicator is of the
two-position type comprising a ball pivoted ~>n its axis and springreturned to the 'off' position. A ferrous armature, embedded in the
ball, is attracted by the electromagnet when energized, and rotates
the ball through 150° to present a different picture in the window.
The picture can be either of the line-diagram type, representing the
Figure 3.21 Typical magnetic
indicators. (a) Section through
two-position indicator - I
armature, 2 return spring, 3
balance weight, 4 magnet
assembly, 5 moisture-proof
grommet, 6 terminal, 7 coil,
8 plastic ball, 9 spindle;
(b) prism-type indicator I terminal, 2 coils, 3 armature,
4 prisms, 5 rack, 6 mounting
spring, 7 control spring, 8
cable gcommet.
8
45
flow of fluid in a system (see Fig 3.22), or of the instructive type
presenting such legends as OFF, ON, OPEN, CLOSE.
Figu re 3.2 1 (b) shows a development of the basic indicator. It inco
porates a second electromagnet which provides for three alternative
indicating positions. The armature is pivoted centrally above the two
magnets and can be attracted by either of them. Under the influence
of magnetic attraction the armature tilts and its actuating arm slides
the rack horizontally to rotate the pinions fixed to the ends of prisms.
The prisms will then be rotated through 120° to present a new pattern
in the window. When the rack moves from the central 'rest' position,
one arm of the hairpin type centring spring, located in a slot in the
rack, will be loaded. Thus, if the electromagnet is de-energized, the
spring will return to mid-position, rotating the pinions and prisms
back to the 'off' condition in the ·window.
The pictorial representation offered by these indicators is further
improved by the painting of 'flow lines' on the appropriate panels so
that they interconnect the indicators with the system control switches
and essential indicating instruments and warning lights.
A typical application of magnetic indicators and flow lines is
shown in Fig 3.22.
Illumination of
Instruments and
Instrument Panels
When flying an aircraft at night, or under adverse conditions of
visibility, a pilot is dependent on instruments to a much greater
extent than he is when flying in daylight under good visibility
conditions, and so the ability to observe their readings accurately
assumes greater importance. For example, at night, the pil ot's
attention is more frequently divided between the observation of
instruments and objects outside the aircraft, and this of course results
in additional ocular and general fatigue being imposed on him.
Adequate illumination of instruments and the panels to which they
are fitted is therefore an essential requirement.
The colour chosen for lighting systems has normally been red
since this is considered to have the least effect on what is termed the
'darkness adaptation characteristic' of the eyes. As a result of
subsequent investigations and tests, however, it would appear that
white light has less effect, and this is now being used· in some current
types of aircraft.
Pillar and Bridge Lighting
Pillar lighting, so called from the method of construction and
attachment of the lamp, provides illumination for individual instruments and controls on the various cockpit panels. A typical assembly,
shown in Fig 3.23 (a), consists of a miniature centre-contact filamentlamp inside a housing, which is a push fit into the body of the
46
Figure 3.22 Application of
magnetic indicators and now
lines.
FUEL
RIGHT TANK
BOOSTER PUMPS
FWD
REAR
REAR
f\10
Off
ON
CROSSFEEO VAlVE
(,'\ ~ClOSE
LP
VALVE
V orEN
lP
VALVE
EHGIHE lOW PRESSURE
Figure 3.23 Pillar light
assemblies.
(b)
(a)
assembly . The body is threaded externally fo r attachment to the
panel and has a hole running through its length to accommodate a
cable which connects the positive supply to the centre contact. The
circuit through the lamp is completed by a ground t ag to connect to
the negative cable.
Light is distributed through a red filter and an aperture in the
47
lamp housing. The shape of the aperture distributes a sector of light
which extends downwards over an arc of approximately 90° to a
depth slightly less than 2 in from the mounting point.
The bridge type of lighting (Fig 3.23 (b)) is a multi-lamp development of the individual pillar lamp already described. Two or four
lamps are fitted to a bridge structure designed to fit over a variety of
the standardized instrument cases. The bridge fitting is composed of
two light-alloy pressings secured together by rivets and spacers, and
. carrying the requisite number of centre contact assemblies above
which the lamp housings are mounted . Wiring arrangem ents provide
for two separate supplies to the lamps thus ensuring that Joss of
illumination cannot occur as a result of failure of one circuit.
Wedge-Type Lighting
This method of instrument lighting derives its name from the shape
of the two portions which together make up th e instrument cover
glass. It relies for its operation upon the physical law that the angle
at which light leaves a reflecting surface equals the angle at which it
strikes that surface.
The two wedges are mounted opposite to each other and with a
narrow air-space separating them, as shown in Fig 3.24. Light is
introduced into wedge A from two 6 V lamps set into recesses in its
Figure 3.24 Wedge-type
lighting.
I
I
',
\
/
/
\
\
\
\
\
\
\
/
I
I
/
I
I
I
I
I
OUTER W EOGE (8)
\
\
\
\
\
\
/
DtAl. PlATE
/
I
I
I
I
I
I
I
I
--'I
I
48
----
wide end. A certain amount of light passes directly through this
wedge and onto the face of the dial while the remainder is reflected
back into the wedge by its polished surfaces. The angle at which the
light rays strike the wedge surfaces governs the amount of light
reflected ; the lower the angle, the more iight reflected .
The double wedge mechanically changes the angle at which the
light rays strike one of the reflecting surfaces of each wedge , thus
d istributing the light evenly across the dial and also limiting the
amount of light given off by the instrument. Since the source of
light is a radial one, the initial angle of some lights rays with respect
to the polished surfaces of wedge A is less than· that of the others.
The low-angle light rays progress further down the wedge before they
leave and spread light across the entire dial.
Light escaping into wedge Bis confronted with constantly
dec reasing angles and this has the effect of trapping the light within
the wedge and d irecting it to its wide end. Absorption of light
reflected into the wide end of wedge B is ensured by painting its
outer part black.
Questions
Describe two of the methods adopted for the display of indications
related t o high-range measurements.
3.2 Wh at is the purpose of a 'platform scale'? Describe its arrangement.
3.3 Name some of the ai rcraft instruments to which a digital counter
display is applied.
3.4 What is the significance of colou red markings applied to the dials of
certain instruments?
3.5 I f it is necessary· to apply coloured markings to the cover glass of an
instrument, what precautions must be taken? '
types of display would you associate with the following
What
3.6
instruments: (a) a synchroscope, (b) an altimeter, (c) a gyro hor izon?
3.7 What do you understand by the term "head-up display"? With the aid
of diagrams describe how required basic flight data is displayed to a
pilot.
3.8 Describe the 'basic T' method of grouping flight instruments.
3.9 Briefly describe one form of light-emitting d isplay.
3.10 Describe one of the methods of illuminating instrument dials.
3. 11 What is the function of a magnetic indicator? Explain its operating
principle.
3.1
49
4 Pitot•static instruments
and systems
Pitot-Static System
Figure 4.1 Basic pitotstatic system.
The pilot-static system of an aircraft is a system in which total
pressure created by the forward motion of the aircraft and the
static pressure of the atmosphere surrounding it are sensed and
measured in terms of speed, altitude and rate of change of altitude
(vertical speed). In other words, the system may be referred to as
a manometric, or air data system .
In its basic form the system consists of a pitot-static tube, or
probe, the three primary flight instruments-airspeed indicator,
altimeter and vertical speed indicator-and pipelines and drains,
interconnected as shown diagrammatically in Fig 4.1.
The complexity of a pitot-static system depends primarily upon
the type and size of aircraft, the number of locations at which
primary flight instrument data are required, and the types of instru-
AIRSPEED INDICATOR
ALTIMETER
VERTICAL SPEED INDICATOR
DRAINS
STATIC LINE
PITOT UNE
50
PITOT-STATIC PROBE
ment. The point about complexity is clearly borne out by comparing
Figs 4. 1 and 4.2.
Sensing of the total , or pitot pressure, and of the static pressure is
effected by the probe, which is suitably located in the airstream and
Figure 4.2 Pitot-static
system of a typical airliner.
F/ 0 STATIC
CAPTAIN'S
STATIC
PROBES
PROBES
I
I
I
;
•
I
-
- - -
-
...
- -
1
-1 -
·- -
- -
I
- ._ ·- -
· ~
..J _ - - - - - -
·~
I
ALTERNATE
STATIC PORT
/
I
- + -
I
I
I
I
I
r -
..
.. .
_.
I
'
.,'
I
- - - - STATI C PRE SSURE
I
-
-
PI TOT PRESSURE
I
RCSCM
-'
I CADC I CENTRAL AIR DATA COMPUTER
I RRCM I RUDDER RATIO CONTROL MODULE
~ MACH TRIM COUPLERS
~ A IRS PEED SWITCH
~ (FLAP POSITIONING!
~
0:§:J ELEVATOR FEEL COM PUTER
L..::.::'......
AIRSPEED MACH WARNING
SWITCH
@O CAB IN PRESSURE CONTROLLER
[JI]
FLIGHT REC ORDER
~ RATE CONTROLLERS ON STABILIZER
~ PRE SSURE ALTITUDE SWITCH
t.=::J
CONTROL MODULE
f"ssl STATIC SOURCE SELECTOR
L':J VALVE
51
transmits these pressures to the instruments. The probe (Fig. 4.3) is
shown in its simplest and original form to serve as the basis for
understanding pitot and static pressure measurement.
It consists of two forward-facing tubes positioned parallel to each
other and in a vertical plane. One of the tubes, the pitot tube, is open
at its forward end to receive the total air pressure resulting from the
aircraft's forward movement , while the other, the static tube, is closed
at its forward end but has a series of small holes drilled circumferen~
tially at a calculated distance from the forward end through which
the undisturbed air at prevailing atmospheric pressure is admitted.
Pressures are transmitted to the instruments through pipelines con·
nected to each tube.
ALTIMETER
AIRSPEED INDICATOR
VERTICAL SPEED INDICATOR
i ·-,----------
____.
STATIC PRESSURE.
:..
\ -~
p
y
t,-.----.
...
.'
---i
__.· ·-T
---------~-
' h-Ap•P-Pc
....
p
~~~~~~T--r--'
:
STAGNATION POINT
p1 = 'l, pV2 +p
Figure 4.3 Sensing and trans-
mission of pilot and static
pressures.
52
Pitut 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.probe placed in a fluid with its open end facing
upstream as shown in Fig 4.4 When the fluid flows at a certain
velocity V over the probe it will be brought to rest at the nose and
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 a pitot probe however, the
potential energy is neglected thus leaving the sum of the remaining
two terms as the constant. Now, in coming to rest at the stagnation
point, kinetic energy of the fluid is converted into pressure energy.
Figure 4.4 Pitot pressure.
h FEET
_______________
_____________________ -:-:-:-=-:-:-::: ::~- - - -- ::::::::
-------
------------
STAGNATION POINT
This means that work must be done by the mass of liquid and this
raises an equal volume of the fluid above the level of the fluid stream.
If the mass of the fluid above this level is m pounds then the work
done in raising it through a height h feet is given by
Work done= mg!z foot-pounds
where g is the acceleration due to gravity. In the British Isles at sealevel, g = 32.2 ft/s 2 = 9.81 m/s 2 •
The work done is also equal to the product of the ratio of mass to
density (p) and pressure (p):
m
Work done=- p .
p
The kinetic energy of a mass m before being brought to rest is
equal to Yun V2 , where Vis the SJ?eed, and since this is converted into
pressure energy,
!!2 p
p
= Yim
v
2
•
Therefore
p=Y2pV2
and is additional to the static pressure in the region of fluid flow.
The factor Yi assumes that the fluid is an ideal one and so does not
take into account the fact that the shape of a body subjected to fluid
flow may not bring the fluid to rest at the stagnation point. However, this coefficient is determined by experiment and for pitot-static
probes it has been found that its value corresponds almost exactly to
the theoretical one.
The Y2p V2 law, as it is usually called in connection with airspeed
measurement, does not allow for the effects of compressibility of air
and so other factors must be introduced. These effects and the
modified law will be covered in the section on airspeed indicators.
Pitot-Static Probes
As a result of aerodynamic 'cleaning up'· of aircraft it became
53
necessary for changes to be made in the design of pi tot-static probes
in order to measure the pressures more accurately and without
causing serious airflow disturbances around the probes. A further
requirement, and one which could not be met satisfactorily
employing separate tubes, was the provision of a heating system to
prevent the tubes from icing up when flying in icing conditions. The
required design changes resulted in the arrangement shown in basic
form in Fig. 4 .5
Figure 4. 5 Basic form of
pilot-static probe. 1 heating
element, 2 static slots, 3
pilot tube connection, 4
static tube connection, S
heater element cable, 6
external drain hole, 7 pitot
tube drain hole.
6
5
1
4
The tubes are mounted concentrically, the pitot tube being inside
the static tube, which also forms the casing. Static pressure is
admitted through either slots or small holes around the casing. The
pressures are transmitted from their respective tubes by means of
metal pipes which may extend to the rear of the probe, or at right
angles, depending on whether it is to be mounted at the leading edge
of a wing, under a wing, or at the side of a fuselage . .Locations of
probes will be covered in more detail under the heading of 'Pressure
Error' .
A chamber is normally formed between the static slots or holes
and the pipe connection to smooth out any turbulent air flowing
into the slots, which might occur when the complete probe is yawed,
before transmitting it to the instruments.
The heating element is fitted around the pitot tube, or in some
designs around the inner circumference of the outer casing, and in such
a position within the casing that the maximum heating effect is
obtained at the points where ice build-up is most likely to occur.
The temperature/resistance characteristics of some elements are
such that the current consumption is automatically regulated
according to the temperature conditions to which the probe is
exposed.
Figure 4.6 illustrates a type of pi tot-static probe which is fitted to
an airliner currently in service. The probe is supported on a mast
which is secured to the fuselage skin by means of a suitably angled
and profiled mounting flange. The assembly incorporates two sets
of static slots connected separately and independently to static
pressure pipes terminating at the mounting flange (see page 61 ).
Pitot pressure is transmitted to the appropriate connecting union via
54
Figure 4. 6 Pitot-static
probe. 1 Terminal block, 2
cover, 3 pilot pressure .
connection, 4 front-slot
static connection, 5 rearslot static connection, 6
rear heater, 7 mast, 8 static
slots, 9 drain holes, IO nose
heaters, 11 baffles, 12 mast
drain-screw.
10
11
I
_/
~
v9
12
r------
-~ -?~
"-'8
I
- -·--
!·
---
""
7 _....,_ __,----<.X
4
~~
3
2
a sealed chamber formed by the mast. To prevent the entry of water
and other foreign matter, the pitot pressure tube is provided with
baffles. Drain holes forward and to the rear of the baffles prevent
the accumulation of moisture within the probe assembly. Water
which might condense and accumulate in the mast can be drained by
removing a drain screw located in the position shown . It should be
noted that in addition to the pitot tube and casing, the mast is also
protected against ice formation by its own heating element.
Heating Circuit
Arrangements
The direct current for heating is controlled by a switch located on a
cockpit control panel, and it is usual to provide some form of
indication of whether or not the circuit is functioning correctly.
Three typical circuit arrangements are given in Fig 4.7.
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 sid e of the heater. If a failure of the
heater, or a break in another section of its circuit, occurs the relay
will de-energize and its contacts will then complete the circuit from
the second pole of the switch t o illuminate the red light which gives
warning of the failed circuit condition. The broken Jines show an
alternative arrangement of the light circuit whereby illumination of
an amber light indicates that the heater circuit is in operation.
The indicating devices employed in arrangements (b) and (c) are
respectively an ammeter and an on-off magnetic indicator. The
function in both cases is the same as at (a), i.e. they indicate that the
circuit is in operation.
Pressure (Position)
Error
The accurate measurement of airspeed and altitude, by means of a
pitot-static probe, has always presented 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 probe will not be affected by air di!;turbances due to the aircraft
55
Figure 4. 7 Typical heating
circuit arrangements.
(4) Light and relay;
(b) ammeter; (c) magnetic
indicator and relay.
o.c. Bl/S8AR
o.c. IIUS6NI
-
RED
·;·
WAICNMG
AMBER
:°~·' INOCATWG
UGlfT
c
r"
UGHT
r- - __I
I
~--,
RELAY
- -- ...,
I
•II
la)
-----'
HEATER
HEATER
(b)
O.C. BUS8AR
MAGNETIC
INDICATOR
___________
•II
"'_./\_./\_,.__,
-,,
______ __,
I
HEATER
RELAY
(c)
itself. These effects of such disturbances are greatest on the staticpressure measuring section of a pitot-static system giving rise to a
pressure or position error (PE) which is 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 altimeter
and an airspeed indicator can develop positive or negative errors.
The vertical speed indicator remains unaffected.
As far as airflow over the probe is concerned, we may consider
the probe and the aircraft to which it is fitted as being alike because
some of the factors determining air flow are: shape, size, speed and
angle of attack. The shape and size of the 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
56
down the flow 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 the probe.
One way of overcoming this is to decrease the casing diameter and
increase the dis'tance of the static orifices from the nose. Furthermore, a number of orifices may be provided along the length of the
probe casing so spaced that some will always be in a region of
undisturbed airflow.
A long and small-diameter probe is an ideal one from an aerodynamic point of view, but it may present certain practical
difficulties; its stiffness may not be sufficient to prevent vibration
at high speed; and it may also be difficult to accommodate the highpower heater elements required for anti-icing. Thus, in establishing
the ultimate relative dimensions of a prohe, a certain amount of
compromise must be accepted .
When a probe is at some angle of attack to the airflow, it causes
· air to flow into the static orifices 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
attac_k depend on the axial location of the orifices along the casing,
their positions around the circumference, their size, and whether the
orifices are in the form of holes or slots.
Static Vents
From the foregoing, it would appear that, if all these problems are
created by pressure effects only at static orifices, they might as well
be separated from the pressure head and positioned elsewhere on the
aircraft. This is one solution and is, in fact, put into practice on many
types of aircraft by using a pressure head incor'porating a pitot tube
only, and a static vent in the side of the fuselage. In some light aircraft the vent may simply be a hole drilled in the fuselage skin, while
for more complex aircraft systems specially contoured metal vent
plates are fitted to the skin. A typical pitot probe and a static vent
are shown in Fig 4.8.
Location of Pressure Heads and Static Vents
For aircraft whose operating ranges are confined to speeds below
that of sound some typical locations of pressure heads are ahead of
a wing tip, ahead of a vertical stabilizer, or at the side of a fuselage
nose section. At speeds above that of sound, a pressure head located
ahead of the fuselage nose is, in general, the most desirable location.
Independent static vents, when fitted , are always located in the skin
of a fuselage, one on each side and interconnected so as to minimize
dynamic pressure effects due to yawing or sideslip of the aircraft.
The actual PE due to a chosen location is determined for the
57
Figure 4.8 Pitot probe and
static vent. (a) Pitot probe l mast, 2 drain screw, 3
drain holes, 4 heater element,
S pressure tube, 6 baffles,
7 heater element, 8 flange
plate, 9 pipe adapter, 10
terminal block; (b) static
vent - 1 intercostal,
2 fuselage skin, 3 sealing
rings, 4 static vent plate,
5 static vents, 6 thiokol,
7 flanged adaptor boss.
10
7
6
/
4
(a)
(b)
appropriate aircraft type during the initial flight-handling trials of a
prototype, and is finally presented in tabular or graphical form thus
enabling a pilot to apply corrections for various operating conditions.
In most cases however, corrections are performed automatically and
in a variety of ways. One method is to employ aerodynamicallycompensated pitot-static 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 within
separate transducers described below, or within central air data
computers (see page I 06).
Pressure Error Correction Transducers
Figure 4.9 illustrates a typical transducer designed to supply correc58
Figure ./. 9 Pressure error
correction transducer.
~' - - - 26 V 400 Hz
' , ~ D.C. OUTPUT TO
/
FAILURE WARNING
400 Hz OUTPUT ;I'
TO SERVO ALTIMETER
DEVICE
tion signals to a servo type altimeter. The transducer is intended for
aircraft in which PE is essentially proportional to a function of Mach
number. The transducer consists of a static pressure-sensing capsule
and a dynamic pressure-sensing capsule, each caps1:c1le being connected
to a d.c. potentiometer. The PE corrections are also applied to the
potentiometers by a network of pre-adjusted resistors connected
across various tapping points on the potentiometers. Thus, their outputs are 'shaped' electrically to provide the signals necessary to
correct the known errors of a particular type of aircraft at various
speeds and altitudes. The outputs are then fed to a function generator
which produces an output as a function of the pressure ratio (p-s)/s;
in other words, the generator is an electrical equivalent to the mecha_nical dividing mechanism in a conventional Machmeter (see page 90).
The generator also amplifies the shaped signals which, after modulation, are finally supplied to the associated servo altimeter. The
power supply circuit of the transducer is energized from a 26 V
400 Hz source, and the circuit also provides a d.c. output for opera.ting a correction failure warning device. The device takes the form
of a solenoid-operated flag marked with a crossed 'C' and which
appears through an aperture in the dial of the altimeter (see Fig
4 . l 7(d). The shaping circuitry is mounted on a plug-in board, so that a
transducer may be interchanged between aircraft types simply by
changing boards.
Alternate Pressure
Sources
If failure of the primary pitot-static pressure source should occur, for
example, complete icing up of a probe due to a failed heater circuit,
then it is obvious that errors will be introduced in the indications of
th·e instruments and other -::omponents dependent on such pressure.
As a safeguard against failure, therefore, a standby system may be
installed in aircraft employing pitot-static probes whereby static
59
TO CAPTAIN'S
FllGKT INS'mVMENTS
···
'.·\
PITOl -STATIC
RH .
PROBE
J
~
I\ 1-:,·'c(
::'
D
ALITRNAT£
STATIC VENT
=-=
===
..
;
l'ITOT l'l!ESSUAE
NOflMAl STATIC
~ ALT£AMAT1: STATIC
Figure 4.10 Alternate static
pressure system.
60
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 flight
instruments, and located in the cockpit within easy reach of the
flight crew. Figure 4.10 diagrammatically illustrates 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 flight instruments are connected to the alternate source of static pressure by manually changing
over the position of the relevant selector valve.
The layout shown in Fig 4. I 1 is based on a system currently in use
and is one in which an al_temate source of both pi tot pressure and
static pressure can be selected. Furthermore , it is an example of a
system which utilizes the static slots of a pitot-static probe as the
alternate static pressure source. The valves are shown in the normal
position, i.e. the probes supply pitot pressure to the instruments on
their respective sides of the aircraft, and the static pressure is supplied
from static vent plates. In the event of failure of pitot pressure from
one or other probe, the position of the relevant selector valve must be
TO CAPTAIN'S FLIOHT
INSTRUMENTS
SELECTOR
VALVES
TO FIRST OFFICER'S FLIGHT
INSTRUMENTS
STATIC V£NTS
PITOT
{
PRESSURE
STATIC
PRESSUl!E
{
:a: L.H.
::rr:rr::: R.H.
==
NOAMAL
~ ALTERNATE
Figure 4.11 Alternate pitot
pressure and static pressure
system.
manually changed to connect the flight 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 4.11, it is a straightforward changeover function .
The probes emplc,yed in the system just described are of the type
illustrated in Fig 4.6, reference to which shows that two sets of static
slots (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 cross-connection
of the static slots and pipes; thus, the front slots are connected to the
rear slots on opposite probes. This balances out any pressure
differences which might be caused by the location of the static slots
along the fore-and-aft axis of the probes.
Drains
In order for a pitot-static 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
61
the lines. Such provision takes the form of drain holes in probes,
drain traps and drain valves in the system pipelines. Drain holes are
drilled in probe pitot tubes and casings (Fig 4.5), and are of such a
diameter that they do not introduce errors in instrument indications.
The method of draining the pipelines of a pitot-static system
varies between aircraft types because one aircraft manufacturer may
fabricate his own design of drain trap and valve, while another may
use a prefabricated component from a specialized .manufacturer.
Some examples of drain traps and valves are illustrated in Fig 4.12.
Drain traps are designed to have a capacity sufficient to allow for
Figure 4. J2 Pitot-static
system water-drainage
methods. (a) Water trap and
drain valve; (b) water trap
and drain plug; (c) drain
valve construction;
(d/ transparent water trap
and drain valve; (e) combined
sump and drain valve.
ADAPTER
PLUG
DRAIN VALVE
(a)
(b)
(c)
(d)
(e)
62
the accumulation of the maximum amount of water that could enter
the system between servicing periods.
The drain valves are of the self-closing type so that they cannot
be inadvertently left in the open position after drainage of accumulated water.
Pipelines
Pitot and static pressures are transmitted through seamless and
corrosion-resistant metal (light alloy and/or tungum) pipelines and
flexible pipes, the latter being used for the connection of components
mounted on anti-vibration mountings.
The chosen diameter of pipelines is related to the distance from
the pressure sources to the instruments (the longer the lines, the
larger the diameter) in order to eliminate pressure drop and time-lag
factors. There is, however, a minimum acceptable limit to the
internal diameter, namely !4 in. A smaller internal diameter would
present the hazard of a blockage due to the probability of a 'slug' of
water developing in such a way as to span the diameter.
Identification of the pipelines is given by conventional colourcoded tapes spaced at frequent intervals along the lines. In additiop
to this method, the lines in some aircraft may also be of different
diameters.
·
Measurement of
Altitude
The Earth's Atmosphere and Characteristics
The earth's atmosphere, as most readers no doubt already know, is
the surrounding envelope of air, which is a mixture of a number of
gases the chief of which are nitrogen and oxygen. This gaseous
envelope is conventionally divided into severaJ concentric layers
extending from the earth's surface, each with its own distinctive
features.
The lowest layer, the one in which we live and in which conventional types of aircraft are mostly flown, is termed the troposphere,
and extends to a height of about 28,000 ft at the equator. This height
boundary is termed the tropopause.
Above the tropopause, we have the layer termed the stratosphere
extending to the stratopause at an average height between 60 and
70 miles.
At greater heights the remaining atmosphere is divided into further
layers or regions which from the stratopause upwards are termed the
ozonosphere, 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.
63
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
earth's centre, so that atmospheric pressure decreases steadily with
altitude. The units used in expressing atmospheric pressure are:
pounds per square inch, inches of mercury and millibars.
The standard sea-level pressure is 14. 7 lbf/in 2 and is equal to
29.921 in Hg or 1013.25 mbar. Thus, a column of air one square
inch in section, extending from the earth's surface to the extremities
of the atmosphere, weighs 14.7 lb and so exerts this pressure on one
square inch of the earth's surface. Suppose now that this same
column of air is extended from points 5,000, 10,000 and 15,000 ft
above sea-level; the weight ..yill have decreased and the pressures
exerted at these levels are found to be 12.2, 10.1 arid 8.3 lbf/ in 2
respectively. At the tropopause, the pressure falls to about a quarter
of its sea-level value.
The steady fall in atmospheric pressure as altitude increases 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
characteristic. The air in contact with the earth is heated by
conduction and radiation, and as a result its density decreases and
the air starts rising. In rising, the pressure drop allows the air to
expand, and the expansion in tum causes a fall in temperature.
Under standard sea-level conditions the temperature is 15° C, and
falls steadily with increasing altitude up to the tropopause, at which
it remains constant at -56.5°C. The rate at which it falls (see also
page 65) is termed the lapse rate (from the Latin lapsus, slip). In the
stratosphere the temperature at fi~t remains constant at - 56 .5°C,
then it increases again to a maximum at a height of about 40 miles,
after which it starts to decrease reaching freezing point at about 50
miles. From this altitude there is yet a further increase reaching a
maximum of about 2200°C at about 150 miles.
Standard Atmosphere
In order to obtain indications of altitude, airspeed and rate of
altitude change it is of course necessary to know the relationship
between the pressure, temperature and density variables, and altitude.
Now, for such indications 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
64
factors. Such measurements, while not impossible, would; however,
demand some rather complicated instrument mechanisms. It has
therefore always been the practice to base all measurements and
calculations related to aeronautics on what is termed a standard
atmosphere, or one in which the values of pressure, temperature and
density at the 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 altimeters,
airspeed indicators and vertical-speed indicators are conce rned, the
inclusion of the assumed values of the relevant variables in the laws
of calibration permits the use of simple mechanisms operating solely
on pressure changes.
The assumptions are as follows: (i) the atmospheric pressure at
mean sea-level is equal to 1013.25 mbar or 29.921 in Hg; (ii) the
temperature at mean sea-level is 15° C (59° F) ; (iii) the air temperature decreases by l.98°C for every 1,000 ft increase in altitude (this
is the lapse rate already referred to) from 15°C at mean sea-level to
-56.5°C (69 . 7° F) at 36,089 ft. Above this altitude the temperature
is assumed to remain constant at -56.5°C.
It is from the above sea-level values that all other corresponding
values have been calculated and presented as the standard &tmosphere. These calculated values were originally defined as ICAN
(International Commission for Aerial Navigation) conditions, but in
19 52 a more detailed specification was finally established by the
International Civil Avic).tion Organization, and so the !CAO Standard
Atmosphere is the one now accepted.
Measurement of Atmospheric Pressure
The re are two principal methods of measuring the pressure of the
atmosphere, both of which are closely associated with pi tot-static
flight instruments. They are (i) balancing of pressure against the
weight of a column of liquid, and (ii) magnifying the deflectio n of
an elastic sensing element produced by the pressure acting on it.
The instruments employing these two methods are known respectively as the mercury barometer and the aneroid barometer.
Mercury Barometer
As shown in Fig 4.13 (a), a mercury barometer consists essentially of
a glass tube sealed at one end and mounted vertically in a bowl or
cistern of mercury so that the open end of the tube is submerged
below the surface of the mercury. Let us imagine for the moment
that the upper end of the tube is open as shown at (b) and that an
absolute pressure p I is applied to the mercury in the cistern, and an
absolute pressure p 2 to the column of m ercury within the tube. If
65
P2
Figure 4.13 Principle of.
mercury barometer.
p 1 is greater thanp 2 then obviously mercury will be forced down in
the cistern and will rise in the tube until a balance between the two
pressures is achieved. This balance is given by
(l)
where H is the difference in levels between the mercury in the cistern
and the tube, and p is the densi1Y of mercury.
But, as may be seen from the diagram, H =h + d, where h is the
distance the mercury has risen in the tube from the zero level, and d
is the corresponding distance the mercury has fallen in the cistern
from the zero level.
The same quantity of mercury has left the cistern and entered the
tube and is such that, if the sectional areas of the cistern and tube
are designated A I and A 2 , respectively, then
(2)
A 1 d=A~
so that
d ~:h.
(3)
=
Therefore
H =h+hA2
A1 .
(4)
Replacing H in eqn ( 1) by this expression,
Pi =p2 +h
(1 +~;) p.
In the practical case the upper end of the tube is sealed and a
vacuum (known as the Torricellian vacuum, after Torricelli, the
66
(5)
Italian inventor of the barometer) exists in the space above the mercury column, so that p 2 is zero and eqn (5) becomes
P1 = h
(1 + ~: P)
(6)
which forms the basic barometer equation, in which h, the height the
mercury rises in the tube, is a measure of the absolute pressure applied
to the mercury in the cistern.
Mercury barometers fall broadly into two-categories: (i) the Fortin
and (ii) the Kew. The Fortin is of the type in which settings are
made on both the base and the level of the mercury column. A
principal feature of this ·barometer is that before taking a reading, the
level of the surface of the mercury in the cistern is brought up to the
tip of what is termed a fiducial point, securely fixed in the top of the
cistern, and corresponding with the zero of the barometer scale. This
adjustment is achieved by means of a screw which, acting on a
leather bag forming the lower part of the cistern, raises the mercury.
The Kew barometer is of the 'fixed cistern' type, and provides
direct indications by means of a vernier-scaled cursor which is .set to
the level of the mercury column and is referenced against a scale
which extends virtually the height of the column. The small variations in the level of the mercury in the cistern, which are proportional to the changes in the height of the column, are automatically
allowed for by using a contracted but linearly divided scale, the
amount of the contraction being fixed by the effective crosssectional areas of both the tube and the cistern. The cistern is
provided with a connector which, on being coupled to a controlled
source of vacuum, permits the pressure acting on the mercury surface
to be decreased to .t he lowest value marked on •the barometer scale.
Thus, and in conjunction with a vacuum chamber, a Kew barometer
serves as a standard for the calibration of altimeters.
In order to use the barometer for altimeter calibration purposes
at any one location, corrections must be applied to the pressure
readings for the reason that the height of the mercury column is
dependent on local temperature conditions, height above sea-level,
the latitude and its corresponding value of gravity. Since the material
on which the scale is graduated expands and contracts with changes
in temperature, then the barometer reading will also depend on
scale temperature. A mercury-in-glass thermometer is therefore
mounted adjacent to the main scale. Details of temperature and
gravity corrections, and how they are derived, are given in British
Standard 2520 - Barometer Conventions and Tables.
Aneroid Barometer and Altimeter
In the general development of instruments for atmospheric pressure
67
mearnrement , certain practical applications demanded instruments
which would be portable and able to operate in various attitudes.
For example, for weather observations at sea it was apparent that a
mercury barometer would be rather fragile, and that under pitching
and rolling conditions the erratic movement of the mercury column
would make observations difficult. A more robust instrument was
therefore needed and it had to be one which required no liquid whatsoever. Thus the aneroid (from the Greek aneros, ' not wet'),
barometer came into being. The instrument has been developed to
quite a high standard of precision for many specialized applications,
but the simplest version of it, and one more familiar perhaps to the
reader, is the household barometer.
The pressure-sensing element of the instrument (see Fig 4 .14) is
an evacuated metal capsule. Since there is approximately zero
pressure inside the capsule, and assuming the instrument to be at sealevel, approximately 14.7 lbf/in 2 .on the outside, the capsule will tend
to collapse. This, however, is prevented by a strong leaf spring fitted
so. that one side is attached to the top of the capsule and the other
side to the instrument baseplate. The spring always tends to open
outwards, and a state of equilibrium is obtained when the pressure is
balanced by the spring tension.
Figure 4. 14 Aneroid barometer.
LEAF
SPRING
\
1-
l'
~ - -- ------- -------
CAPSULE
If now the atmospheric pressure decreases, the force tending to
collapse the capsule is decreased but the spring tension remains t~e
same and consequently is able to open out the capsule a little further
than before. If there is an increase in pressure, the action is reversed,
the pressure now collapsing the capsule against the tension of the
spring until equilibrium is attained.
The resulting expansion and contraction of the capsule, which is
extremely small, is transformed into rotary motion of the pointer by
means of a magnifying lever system and a very finely-linked chain.
From the foregoing description, we can appreciate that when it
first became necessary to measure the height of an aeroplane above
the ground, the aneroid barometer with change of scale markings
68
formed a ready-made altimeter. Present-day altimeters are, of
course, much more sophisticated, but the aneroid barometer principle still applies. The mechanism of a typical sensitive altimeter is
shown in Fig 4.15. The pressure-sensing element is made up of three
aneroid capsules stacked together to increase the sensitivity of the
instrument. Deflections of the capsules are transmitted to a sector
gear via a link and rocking shaft assembly . The sector gear meshes
with a magnifying gear mechanism which drives a handstaff carrying
14
24
Figure 4.15 Exploded view
of a typicai altimeter mechanism. 1 link, 2 calibration
arm, 3 spring-loaded balance
weight , 4 rocking shaft, 5
handstaff, 6 mechanism
adaptor plate, 7 cam-follower pin, 8 cam follower
(drives output wheel),
9 slotted cam, 10 millibar
counter, 11 baroscale adjust·
ing knob, 12 trace disc,
13 third pointer, 14 long
pointer, 15 intermediate
pointer, 16 dial, 17 top
mechanism gear train,
18 output wheel, 19 cam
gear, 20 spigot, 21 hairspring, 22 intermediate
pinion and gear wheel,
23 temperature-compensating U-bracket,
24 diaphragm unit.
23
22
21
20
19
18
17
16
a long pointer the function of.which is to indicate hundreds of feet.
A pinion is also mounted on the handstaff, and this drives a second
gear mechanism carrying second and third pointers which indicate
thousands and tens of thousands of feet respectively. In this particular instrument a disc is also attached to the third pointer gear and
moves ·with it. One side of the disc is painted i,vhite, and above
10,000 ft this becomes visible through a semicircular slot cut in the
main dial. Thus, the pointer movement is 'traced' out to eliminate
ambiguity of readings above 10,000 ft.
In order to derive a linear altitude scale from the non-linear
pressure/altitude relationship , it is necessary to incorpo rate some
form of conversion within the altimeter mechanism (see Fig 4.16).
Linearity is obtained by a suit able choice of material for the capsules
and corresponding deflection curve (2) and also of the deflection
characteristics of the variable magnification lever and gear system
adopted for transmitting deflections to the pointer system (curve 3).
The resultant of both curves produces the linear scale as indicated at
4. To cater for variations between deflection characteristics of
individual capsules, and so allow for calibration, adjustments are
always provided whereby the lever and gear system magnification
may be ma tched to suit the capsule characteristics.
The pressure-sensing element of the altimeter shown in Fig 4.15
is compensated for changes in ambient temperatures by a bimetal
69
INSTRUMENT INDICATION
WITH UNIFORMLY
HEIGHT
DIVIDED SCALE
(FT)
1 REVOLUTION • 1000 FT
PRESSURE - HEIGHT LAW
50.000
I"'
4
"' ""
I"'I
v~/
L
V
\
I'\
20.DOO
!"'-..
10
+-
+""
10,000
I
I
0
0
I/
V
\
30.000
2'o
3
\
40,000
0
"'
200 400: 800 800
1000
0
t
0.04
0.08
0.12
~/
I V
I/
r--
+- I/'
_L
V
0.16
0.20
VARIABLE MAGNIFICATION CAPSULE
PRESSURE-DEFLECTION
LEVER AND GEAR
DEFLF.CTION CHARACTERISTICS OF
SYSTEM
(INCHES)
CAPSULES
Figure 4.16 Conversion of
pressure/height relationship
to a linear scale.
70
ATMOSPHERIC PRESSURE
ON CAPSULES
(MILLIBARS)
2
EXAMPLE= 20 000 ft
1 PRESSURE s 466 mbar
2 CAPSULE DEFLECTION= 0.087 in
3 POINTER ROTATION= 20 revs
4 INDICATED HEIGHT= 20 000 ft
'U'-shaped bracket, the open ends of which are connected to the top
capsule by push rods. The tempe_rature coefficient of the instrument
is chiefly due t.o the change of elasticity of the capsule material with
change of temperature; this, in turn, varying the degree of deflection
of the capsule in relation to the pressure acting external to it. For
example, if at sea-level the temperature should decrease, the elasticity
of the capsule would increase ; in other words, and from the
definition of elasticity, the capsule has a greater tendency 'to return
to its 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 capsule will have decreased,
then by comparison, the capsule expansion becomes progressively
greater. The effect of a decrease in temperature on the 'U'-shaped
bracket is to cause the limbs to bend inwards, and by virtue of the
angular position of the pins, a corresponding downward force is
exerted on the capsule assembly to oppose the error-producing
expansion. The converse of the foregoing sequence will apply when
an increase of ambient temperature occurs.
A barometric pressure-setting mechanism, the purpose of which is
described on page 74, is mounted in front of the main mechanism.
It consists of a counter geared to the shaft of a setting knob. The
shaft also carries a pinion which meshes with a gear around the
periphery of the main mechanism casting. When th_e knob is rotated
to set the required barometric pressure, the main mechanism is also
rotated, and the pointers are set to the corresponding altitude change.
Linearity between the pressure counter and the pointers is maintained by a c·a m and follower (see page 75). The position of the
capsules under the influence of the atmospheric pressure prevailing
at the time of setting remains undisturbed. A spring-loaded balance
weight is linked to the rocking shaft to maintain the balance of the
main mechanism regardless of its attitude.
Altimeter Dial Presentations
The presentation of altitude information has undergone many
changes in recent years, principally as a result of altimeter misreading
being the proven or suspected cause of a number of fatal accidents.
In consequence, several methods are to be found in altimeters
currently iJ1 use, the most notable of which are the triple-pointer,
single-pointer and digital counter, and single-pointer and drum
presentations. The triple-pointer method is the oldest of presentations and is the one which really made it necessary to introduce
changes. This method is used in the altimeter shown in Fig 4.15:
the susceptibility of its predecessor to misreading of 1,000 ft and
I 0,000 ft, has been overcome to a large extent by giving the pointers
a more distinctive shape, and by incorporating the trace disc referred
to earlier. In addition, some versions incorporate a yellow and black
striped disc which serves as a low-altitude warning device by coming
into view at altitudes below 16,000 ft.
The counter pointer (Fig 4.17 (c)), and in some cases the drum
pointer presentations are used in servo altimeters and altimeters
forming part of air data computer systems (see pages 75 and I 06).
Errors Due to Changes in Atmospheric Pressure and Temperature
As we already know, the basis for the calibration of altimeters is the
standard atmosphere. When the atmosphere conforms to standard
values, an altimeter will read what is termed pressure altitude. Ir. a
non-standard atmosphere, an altimeter is in error and reads 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
1,013.25 mbar (29.92 in Hg) and indicate the pressure altitude of
zero feet. Similarly, at an airfield level of 1,000 ft, it would respond
to a standard pressure of 977.4 mbar (28.86 in Hg) and indicate a
pressure altitude of 1,000 ft. Assuming that at the sea-level airfield
the pressure falls to 1,012.2 mbar (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 mbar (29.95 in Hg), the altimeter in responding
to the pressure change will indicate that the airfield is approximately
30 ft below sea-level; an error of -30 ft:
71
Figure 4.17 Altimeter dial
presentations. (a) Triplepointer (10,000 ft pointer
behind, 1,000 ft pointer in
this view); (b) modified
triple pointer; (c) and (a)
counter/pointer.
LOW ALTITUDE
WARNING SECTOR
SITTING MARKERS
COARSE HEIGHT
INDICATOR 10 OOOFT POINTER
100 FT POINTER
(a)
100 FT POINTER
(b)
ALTITUDE ALERT I NG
FAILURE LIGHT
P.E.
CORRECTION
FAILURE
WARNING
FLAG
BAROMETRIC PRESSURE COUNTERS
BAROMETRI C PRESSURE SCALE
(C)
(d)
In a similac 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 st andard value.
For example, when an aeroplane flying at 5,000 ft enters a region in
which the pressure has fallen from the standard value of 842.98 mbar
to say , 837 i:nbar, the altimeter will indicate an altitude of approximately 5,190 ft.
The standard atmosphere also assumes certain temperature values
at all altitudes and consequently non-standard values can also cause
errors in altimeter readings. Variations in temperature cause
differences of air density and therefore differences in weight and
72
Figure 4.18 Effect of
atmospheric temperature
on an altimeter.
C
c,
e,
H 5000 FT
- - - ' - - A _ _.__ ___.__ A ---------'--A
1
--'--2
pressure of the air. This may be seen from the three columns shown
in Fig 4.18. At point A the altimeter measures the weight of the
column of air above it, of height AC to the top of the atmospheric
belt. At point H which is, say, at an altitude of 5,000 ft above A, the
weight or pressure on the altimeter is Jess by the weight of the part
AB of the column below H. If the temperature of the air in the part
AB rises, the column will expand to A 1 B1 , and at H the pressure on
the altimeter is now less by the weight of A 1 H. But the weight of
A 1 B1 is still the same as that of AB, and therefore the weight of A 1 H
must be less than that of AH, and so the altimeter in rising from A 1
to 5,000 ft will register a smaller reduction of pressure than when it
rose from point A to 5,000 ft . In other words, it will read less than
5,000 ft. Similarly, when the temperature of the air between points
A and H decreases, .the part AB of the column shrinks to A2 B1 and
the change of pressure on the altimeter in rising from A1 to 5,000 ft
will be not only the weight of A2 B2 (which equals AB) but also the
weight of B1 H. The altimeter will thus measure a greater pressure
drop and will indicate an altitude greater than 5,000 ft. The relationship between the various altitudes associated with aircraft flight
operations are presented graphically in Fig 4.19.
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 datums. In order, therefore, to compensate for altitude
73
PRESSURE AT THIS LEVEL.
19-00 IN HG
DENSITY AT THIS ALTITUDE
-0045 LB/FT 3
P'IIES&UIIE ALmUOE IN
STANDA/m ATMOlll'liEl'I£. 18250 FT
_l_ _
TRUE ALmUDE
15 000 FT
:. DENSITY ALTITUDE IN
STANDARD ATMOSPHERE -13 400 FT
_~Ai!s::~L ______L
ABSOLUTE ALTTTUDE
13 000 FT
AIRFIELD ELEVATION
2
FT
SEA LEVEL
Figure 4.19 Relation between
the various altitudes.
74
errors due to atmospheric pressure changes, altimeters are provided
with a manually operated adjustment device which allows the
pointers to be set to zero height for any prevailing ground pressure
so that the indications in flight will still be heights iri the standard
atmosphere above the ground.
The adjustment device consists basically of a scale or counter
calibrated in millibars or inches of mercury which is interconnected
between a setting knob and the altitude indicating mechanism in
such a way that the correct pressure/height relationship is obtained.
The underlying principle may be understood by referring to Fig 4.20,
which shows a scale-type device in very simple form. The scale is
mounted on a gear which meshes with a ·pinion on the end of the
control knob shaft and also with the pointer gearing. A differential
gear mechanism (not shown) allows the pointer to be rotated without disturbing the setting of the capsules.
In Fig 4. 20 (a) the al tir11eter is assumed to be subjected to standard
conditions; thus the millibar scale when set to 1,013 mbar positions
the pointer at the O ft graduation. If the millibar setting is changed
from 1,013 to 1,003 as at (b ), the scale will be rotated clockwise,
making the altimeter pointer turn anti-clockwise to indicate approximately - 270 ft. If now the altimeter is raised through 270 ft as at
(c), a pressure decrease of 10 mbar (1 ,013- 1,003) will be-measured
by the capsule and the pointer will return to zero. Thus whatever
pressure is set on the millibar scale, the altimeter will indicate zero
Figure 4.20 Principle of
barometric pressure
adjustment.
(cl
(a)
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 height indicated will be zero.
In practice, the method of setting is a little more complicated
because the relationship between pressure and altitude is non-linear.
Therefore, in order to obtain linear characteristics of the required
pressure and altitude readings, an accurately profiled cam correction
device _(an example is shown in Fig 4.15) is connected between the
pressure counter and main mechanism.
'Q' Code for Altimeter Setting
The setting of altimeters to the barometric pressures prevailing at
various flight levels and airfields is part of flight operating
techniques, and is essential for maintaining adequate separation
between aircraft, and terrain clearance during take-off and landing.
In order to make the settings a pilot is dependent on observed
meteorological data which are requested and tnansmitted from ground
control centres. The requests and transmissions are adopted
universally and form part of the ICAO 'Q' code of communication.
Three code letter groups are normally used in connection with
altimeter settings, and are defined as follows:
QFE
QNE
QNH
Setting the pressure prevailing at an airfield to make the
altimeter read zero on landing and take-off.
Setting the standard sea-level pressure of 1,013.25 mbar
(29.92 in Hg) to make the altimeter read the airfield
elevation.
Setting the pressure scale to make the altimeter read
airfield height above sea-level on landing and take-off.
Servo Altimeter
The mechanism of a typical servo altimeter is shown schematically
in Fig 4.21 from which it will be noted that the pressure-setting
capsule element is coupled to an electrical pick-off assembly instead
75
Figure 4. 21 Typical servo
altimeter mechanism.
of a mechanical linkage system as in conventional altimeters. The
inductive type of pick-off consists of a pivoted laminated I-bar
coupled to the capsules and positioned at a very small distance from
the limbs of a laminated E-bar pivoted on a cam follower. A coil is
wound around the centre limb on the E-bar and is supplied with
alternating current, while around the outer limbs coils are wound
and connected in series to supply an output signal to an external
amplifier unit. Thus, the pick-off is a special form of transformer,
the centre-limb coil being the primary winding and the outer-limb
coils the secondary winding.
A two-phase drag-cup type of motor is coupled by a gear train to
the pointer and counter assembly, and also to a differential gear which
drives a cam . . The cam bears against a cam follower so that as the cam
position is changed the E-bar position relative to the I-bar is altered.
The reference phase of the motor is supplied with a const.ant alternatin
voltage from the main source, and the control phase is connected to
the amplifier output channel.
Setting of barometric pressure is done by means of a setting knob
geared to a digital counter, and through a special rod and lever
mechanism to the differential gear and cam. Thus, rotation of the
setting knob can also alter the relative positions of the E-bar and
I-bar.
MOTOR
HEIGHT COUNTER S
OVERRUN
SWITCH
¢=::)
;::.~~;~~~~O
BAROMETRIC
. . . . . ~.?:.!~e\NT DUE TO ALTIT\JDE
76
A double-contact switch is provided within the case and is connected in the power supply circuit to interrupt the latter should the
servomotor overrun. A solenoid-operated warning flag comes into
view whenever an overrun occurs and for any other condition
causing an interruption of the power supply.
When the aircraft altitude changes the capsules respond to the
changes in static pressure in the conventional manner. The displacement of the capsules is transmitted to the I-bar, changing its angular
position with respect to the E-bar and therefore changing the airgaps at the outer ends. This results in an increase of magnetic flux
in one outer limb of the E-bar and a decrease in the other. Thus, the
voltage induced in one of the secondary coils increases, while in
the other it decreases. An output signal is ther~fo re produced at the
secondary coil terminals, ·which will be either in phase or out of
phase with the primary-coil voltage, depending on the direction of
I-bar displacement. The magnitude of the signal will be governed by
the magnitude of the deflection.
The signal is fed to the amplifier, in which it is amplified and
phase detected, and then supplied to the servomoto_r control winding.
The motor rotates and drives the pointer and height-counter
mechanism in the direction appropriate to the altitude change. At
the same time , the servomotor gear train rotates a worm-gear shaft
and the differential gear which is meshed with it. The cam and cam
follower are therefore rotated to position the E-bar in a direction
which will cause the magnetlc fluxes in the cores, and the secondarycoil voltages, to start balancing each other. When the E-bar reaches
the null position, i.e. when the aircraft levels off at a required
altitude, no further signals are fed to the amplifier, the servomotor
ceases to rotate, and the pointer and counters indicate the new
altitude.
When the barometric-pressure setting knob is rotated, the
pressure counters are turned and the lever of the setting mechanism
moves the worm-gear shaft laterally . This movement of the shaft
rotates the differential gear, cam and cam follower , causing relative
displacement between the E-bar and I-bar. An error signal is therefore produced which, after amplication and phase detection, drives
the servomotor and gear mechanisms in a sequence similar to that
resulting from a normal altitude change. When the null position of
the £-bar is reached, however, the pointer and counters will indicate
aircraft altitude with respect to the barometric pressure adjustment.
A typical altitude and pressure data presentation is shown in Fig 4.22.
Ca~in Altimeters
In pressurized-cabin aircraft , it is important that the pilot and crew
have an indication that the cabin altitude corresponding to the
77
Figure 4.22 Altitude altering system. (a) Servo altimeter; (b) alerting unit.
I
[ _ _ _ } POWER
SUPPLY
26V A.(
28V O.C
SYNCHRO I I
OUTPUT
I
BAROMETRIC PRESSURE
SETIING KNOB
(a) SERVO ALTlil/lETER
ALERTING LIGHT
TEST
eunoN
(b) ALERTING UNIT
ALTITUOE SELECTOR
KNOB
maximum differential pressure conditions is being maintained. To
meet this requirement, simple altimeters, calibrated to the same
pressure/altitude law as nonnal altimeters, are provided on the main
instrument panel, or on the pressurization system control panel,
their measuring elements responding directly to the prevailing cabin
air pressure.
Altitude Alerting Systems
In certain aircraft systems the control and operating conditions are
related to one specific altitude; for example, in a cabin-pressurization
system, the necessity arises for an indication of a possible increase of
cabin altitude above the desired level while th e air.::raft is at its normal
78
operating altittitL. Furthermore, and particularly as a result of the
introduction of altitude rep.orting systems; it is necessary for a pilot
to be warned of an approach to, and/or deviation from, a selected
operational altitude. To cater for the appropriate requirements
therefore, it is usual to employ altitude switching units or servotype altimeters capable of transmitting altitude signals to a separate
alerting unit via a synchronous transmission link.
Altitude switching units normally consist of an aneroid-capsule
measuring element similar to that used in altimeters, but in lieu of a
pointer actuating mechanism, the capsule is so designed that at a
pr~set altitude, its expansion actuates an electrical contact assembly
such as to complete a circuit to a warning light or aural warning
device.
Figure 4 .22 illustrates the components of an altitude alerting
system designed to give audio and visual warnings when an aircraft
approaches or deviates from a pre-selected altitude by more than a
pre-determined amount. The altitude is selected on the alerting unit,
and is indicated by a digital counter which is geared to the rotors of
a control transformer (CT) synchro, and a control transformer/
resolver synchro or transolver as it is called. The synchros are referred
to as coarse and fine synchros respectively, and are electrically connected to corresponding transmitter synchros within the servo altimeter. The rotors of the transmitter synchros are mechanically
positioned by a linkage system coupled to the altimeter capsule
assembly so that the output from the synchros is proportional to the
aircraft's altitude.
When an altitude is selected on the alerting unit, the selector knob,
in addition to rotating the digital counters, also rotates the rotors of
the unit synchros thereby developing a signal oorresponding to the
differencP. between the indicated and the selected altitudes. This
signal difference is supplied to an input section of the overall circuit
of the alerting unit, and at pre-determined values of both synchro
rotor voltages, two signals are produced and are supplied as inputs to
a logic circuit. The logic circuit comprises a timing network which
controls a remote audio warning device and the operation of the
warning lights in the servo altimeter and altitude alerting unit.
The sequence in which alerting takes place is illustrated in
Fig 4.23. As the aircraft descends or climbs to the pre-selected
altitude, the difference signal referred to earlier is reduced, and the
logic circuit so processes the two signals supplied to it, that at a preset outer limit H 1 (typically 900 ft) above or below the pre-selected
altitude, one signal actuates the audio warning device which
remains on for two seconds, and also illuminates the warning lights.
The lights remain on until at a further pre-set inner limit H2
(typically 300 ft) above or below the pre-selected altitude, the second
signal causes the circuit to the warning lights to be interrupted there79
AUDIO WARNING
ON FOR TWO SECONDS, ANO
WARNING LIGHTS ILLUMINATED
Hl (90 0 F T } - - - -
, ...... ......
' ' ,,
' , , ,~
H2 (300 FTI
RNING LIGHTS EXTINGUISHED
- - - - - - - - - - - - - -4iiall~ - -- - - - -- - - - - - -
~
SELECTED
ALTITUDE
Figure 4.23 Altitude
alerting sequence.
' ',....._
------
by extinguishing them. As the aircraft approaches the pre-selected
altitude the rotor voltages of the synchros approach their null and no
further warnings are given. If the aircraft should subsequently depart
from the pre-selected altitude by more than the inner limit H2 , the
logic circuit charJges the alerting sequence such that the warnings
correspond to those given during the approach through outer limit
H l, i.e. audio on for two seconds and warning light illumination .
A 'reset' circuit is incorporated in the alerting unit, its function
being to reset the logic circuit whenever the selector knob is operated to change the selected altitude by more than 100-ft and at a
rate greater than 8,000 ft per minute. The circuit utilizes a photoelectric cell which produces a signal of sufficient magnitude to override any signal present at the output of the logic circuit.
Functional testing of the audio and visual warning system is
accomplished by operating a test switch on the alerting unit while
rotating the altitude selector knob. In the event of failure of the
power supply to the system (26 V a.c. 400 Hz and 28 V d.c.) and
also of altitude signals, a solenoid is de-energized to actuate a
warning flag which obscures the altitude counters of the alerting unit.
Altitude Reporting System
The control of air traffic along the many air corridors in the vicinity
80
of major airports is dependent on rigid procedures for communication betwee·n individual aircraft and ground control stations in
order that traffic may be identified and assigned to requisite
separation levels. In addition to normal voice transmissions, the
communication procedure involves the use of an airborne transponder which, in response to interrogation signals from a radar
transmitter/receiver at the air traffic control centre , automatically
transmits coded reply signals to the centre. The signals are then
computer-processed, decoded , and then alpha-numerically displayed
to the air traffic controller on his primary radar screen.
Aircraft altitude is one of the important parameters required to
be known , and to further reduce time-consuming voice transmissions,
a method of automatically transmitting such data from an altimeter
was devised and also became a mandatory feature of the air-toground communication procedure.
The interrogation system as a whole forms what is termed Air
Traffic Control Secondary Surveillance Radar, and can operate in four
modes of interrogation: A, B, C and D. Modes A and B are used for
identification, Mode C for altitude reporting, while Mode Dis at
present unassigned. In each case the. interrogating signal, which is
transmitted on a frequency of 1,030 MHz from a rotating directional
antenna, comprises a pair of pulses PI and P3, and in order that the
airborne transponder can 'recognize' in which mode it is being
interrogated, the pulses are spaced at different time intervals. The
spacing is taken from the leading edge of the first pulse to the
leading edge of the second (see Fig 4.24). It will be noted from
diagram (a) that a third pulse P2 can also be transmitted from a
control antenna ; its purpose is to suppress side lobe radiation from
the interrogator antenna, and to ensure that the transponder replies
only to the main beam directional signal pulses. This is effected by a
'gating' circuit which compares the relative amplitudes of the pulses
and enables the transponder to determine whether the interrogation
is a correct one, or due to a side lobe.
When the transponder decodes the interrogating signal it will
reply by transmitting a train of information pulses on a frequency of
1,090 MHz, and in a coded sequence which depends not only on the
interrogation mode, but also on pre-allocated code numbers which,
for operation in Modes A and B, are selected by the pilot on the transponder control unit. In Mode C operation, code numbers are automatically transmitted by the transponder which also receives signals
corresponding to specific altitudes from an altimeter and in a manner
to be described later.
The train of information pulses from the transponder may consist
of up to twelve pulses spaced 1.45 µs apart, depending on the reply
code selected on the control unit (diagram (b) Fig 4.24 ). The pulses
lie between two additional framing pulses FI and F2, which are fixed
81
Figure 4.24 Pulse
structure; transponder
operation.
P2
MAIN INTERROGATION
BEAM PULS ES P1 AND P2
r---,
P,
TRANSPONDER
SUPPRESSED
T
1--~-~:
_J_
: TRANSPONDER
MAY OR MAY
I
1- -~~rePLY
9dB
P,
REPLIES
!
I
'
PULSE P2
SIDE LOBES
~
~
'
(a)
INTERROGATION BEAM AND PULSE SPACING
F,
C,
I
A,
c,
A,
c.
A,
X
e,
I
D,
MODE A
8
MOOE B
17 µSECS
'
MODE C
21 µ S E C S ~
MODE D
e,
µSECS~
25 µSECS
D,
e.
D,
F,
I
11.46
secsI
4
(b)
I
REPLY PULSE TRAIN
20.3
~secs
-I
X PULSE NOT USED IN
PRESENT CODING
at a spacing of 20.3 µs, and which are always transmitted in the reply.
In a twelve-pulse train the number of codes available is 212 = 4,096
the codes being numbered 0000 through 7777, the latter giving all
twelve pulses when four selector knobs of the control unit (Fig 4.25
(a)) are correspondingly set. Each control knob controls a group of
three pulses as sh9wn; the letters in this case designate pulse groups
and not interrogation modes. The first control knob controls the A
group, the second knob the B group, and so on. The subscripts to
each letter of a group are significant since their sum equals the digit
selected on the control unit, and this may be seen from the basic
code table and examples given in Fig 4.25 (b ). Since the whole
system is based on a digital computing process, then the encoding
and decoding of interrogation and reply signals is dependent on
logic variables and corresponding binary digits or 'bits' ·as they are
termed. In the example shown at (c) the code 2300 in Mode A has
been selected on the control unit and this produces the equivalent
binaries O 1 0, 1 I 0, 0 0 0 and O OO respectively. Since in logic
networks O signifies the absence of a signal, and I the presence of a
signal, then the selection of the digit 2 causes only pulse 2 of the A
82
group to be transmitted, while the selection of digit 3 causes the
transmission of pulse 1 plus pulse 2 of the B group . Thus, the reply
pulse train in this example would consist of only three pulses
spaced between the framing pulses at the intervals indicated in the
diagram.
Altitude reporting, as already stated, is a Mode C operation and
COOE INOICATORS ANO SELECTOR KNOBS
Figure 4.25 Reply pulse
coding.
OFF
ON
I / ALT
--o
(al
COOE NUMBER
SELECTED
PULSE
GROUP
PULSE
GROUP
A
B
8
C
0
A
8 C D
A
8 C
D
A 8 C
D
A
8
C
0
0
0
0
0
0
2
3
4
0
0
0
7
7
7
7 7
o,
", e, c,
A2 Iii C2
3
REPLY PULSES
CODE DIGIT
(bl
6
B1 C2
B2 C4 02
o.
D
NONE
e,
c,
o,
A,
2
A2
B2
c2
3
A1 A 2
0 1 e2
c,c,
4
A,
5
A1 A4
e.
e,e,
c,c.
6
A 2 A4
B 2 B,
C2 C4
7
A1 A 2 A 4
e, e2 e.
c,c,c.
01
D2
A• B, C, D,
NONE
1
L~ L~
3
PULSE ASSIGNMENT
B
C
0
I I I I 1 I
I I I 1 I I
,,secs
A.,
NONE
r1 r1 r1
5.8
A 1 B4
c2
A
02
o,o,
o,
o,o.
o,o,
o,o,o,
c,
n nn I
A
J
c, 02
NONE
F
~
D
A
FRAMING PULSES
ONLY
1
PULSE
GROUP
0
PULSE
GROUP
C
2
81
r1 r1
r1
I I I I
I I 1 1
~
L
1
11.6 ,,SECS
L~
82
I 1
I I
~
L
L
F,
r1 r1 r1
I 1 I I 1 I
I 1 1 I 1 I
~
L~
L~
L
.1
14.5 11SECS
(cl
83
when the 'ALT' position is selected on the control unit of the transponder, the latter will reply to the corresponding interrogation
signal as well as to a Mode A interrogation signal. However, whereas
in Mode A operation the pulse trains of reply signals are associated
with manually selected codes, in the altitude reporting mode, pulse
trains are produced automatically and supplied to the transponder
by an encoding altimeter or, in some cases, by the altitude measurint
section of an air data computer. The arrangement of an encoding
altimeter is shown in Fig 4.26, and from this it will be noted that
an encoder assembly is mechanically actuated by the aneroid
capsule assembly in addition to the conventional pointer and digital
counter display mechanism.
CONNECl
OSCILLATOR
'
,--::/
·:.
•: ,
\..
I
•
-,l
CAM
Figure 4.26 Encoding
altimeter.
DRUM
\
',
'
~\
.
FRAME
~~
COUii
~
%,
CYLINDRICAL
LENS
ENCODING
DISC
BANK OF
PHOTO CELLS
BAROSCALE COUNTERS
AMPLIFIER
The encoder assembly is of the optical type consisting of a light
source, light-collimating discs, a cylindrical focussing lens, encoder
disc, a bank of photo-electric cells, and an ampl_ifier. The encoder
disc (see Fig 4.27) is made of glass and is etched with transmitting
and non-transmitting segments arranged in eleven concentric rings
and spaced so as to produce binary-coded pulses corresponding to
100 ft increments of altitude. The coding is in accordance with that
established by ICAO and set out in Annex 10 - Aeronautical
Telecommunications. The disc makes a single turn to produce the
requisite number of counts compatible with the altitude range for
which the instrument is designed.
During operatiqn, light from the 14 V lamp passes through the
84
?
~
:"
"""
!T1
:,
0
0
0.
g
0.
~·
T
-. . --uu-- -- -- -- -- --uu-02
o,.
A,
Ai
A•
e,
e2
s,.
c,
c2
c,
-- -- ,. _ . . . . . . . -- -- n-o,
01
J
e,.
IIPULSE
1
0
O
O
O
O
O
1
0
BINARY
COOE
O
BINARY
CODE
O
O
O
A,
A,
I I I I
.J lJ
0
0
C2
A,
C,
PULSE
POSITION
I I II I I I I
U
LJ LJ
0
O
O
O
1
O
c..
II
I I
.J
L
1
0
C2
C4
ALTITUDE RANGE - 1000 TO -950 FT
A.
0
O
11
c,
I I POSITION
POSITION I I I I I I I I I I I I I I I I I I
~ ~ Li L.J L.J L.J L.J L.J L.J L.J L
O
11111
c,
I I I I I I I I I I I I
L_l L_l LJ LJ L_J
11 1-LflfLJlJ,1 11 11 11
II II
I I I I
__.I
1
0
L.J
I
L.J
BINARYOOO
CODE
BINARY
CODE
A,
,-Lflflj" II II IIUlJl
I I I I I
I
L.J
1
1
L.J
I
I I
LJ
0000
ALTITUDE RANGE 6850 TO 6950 FT
ALTITUDE RANGE 22550 TO 22650 FT
PULSE
POSITION
BINARY
CODE
II II
I ,
I I I I I I
n
.J LJ
0
LJ
0
L
0
II II ,1.I ,1
"
I I I I
I I I I I
..J
LJ
0
LJ
0
,-m
D
TRANSMITIING SEGMENTS
c,.
I
Al
L..J
0
LJ
0
I
I
LJ
0
ALTITUDE RANGE 14650 TO 14750 FT
U1
e2
111111111111
ALTITUOE RANGE 30450 TO 30550 FT
co
e,
I I I I
LJ
O
PULSE
A,
1111
_J
'
A,
A1
~1 1 1 1 1 1 1 ·
111
0
. . . NON· TRANSMITIING SEGMENTS
collimating discs which produce parallel rays that are then focussed
through the cylindrical lens onto the encoding disc in the fonn of a
sharp line of light. Depending on the aircraft's altitude at any one
time, the disc and segments along a particular radial will be at some
corresponding position with respect to the bank of photo-electric
cells. The cells will accordingly respond to the position of the transmitting and non-transmitting segments and produce outputs which
are then amplified and supplied to the transponder. Examples of
the pulse codings produced are shown in Fig 4. 27.
Airspeed Indicators
Airspeed indicators are in effect very sensitive pressure gauges
measuring the difference between the pitot and static pressures
detected by the pressure probe, in tenns of the Yip V1 fonnula given
on page 53.
Originally this formula served as the basis for the calibration qf
indicators, but as airspeeds began to increase a noticeable error began
to creep into the indications. The reason for this was that, as the
high-velocity air was brought to 'stagnation' at the pitot tube, it was
compressed and its density was increased. The formula was therefore
modified to take into account additional factors in order to minimize
the 'compressibility error'. Thus, present-day airspeed indicators are
calibrated to the law
p=Y1pV2
(1 +~;::)
where p = Pressure difference (mmH 1 0)
p = Density of air at sea-level
V = Speed of aircraft (m .p.h. or knots)
a0 = Speed of sound at sea-level (m.p.h.)
The numerical values to be inserted in the fonnula depend on
whether Vis expressed in m.p.h. or knots* :
p = O·O 12504 V2 (I + 0-43 V2 x 10- 6 ) with Vin m.p.h.
= 0·016580V1 (1
+ 0·57V2 x 10-6 ) with Vin knots.
Both fonnulae take into account the relative densities of air and
water.
The mechanism of a typical simple-type airspeed indicator is illustrated in Fig 4.28. The pressure-sensing element is a metal capsule
the interior of which is connected to the pitot pressure connector via
•
The knot is the unit commonly used for expressing the speeds of aircraft:
1 knot= 1 nautical mile per hour
= 1.15 m.p.h.
= 6,080 ft/h
86
Figure 4.28 Simple type of
airspeed indicator. 1 Pitot
connector, 2 pilot union,
3 capsule capillary, 4 cap·
sule plate, 5 locknut, 6
pointer movement assembly,
7 frame casting, 8 rocking
shaft assembly, 9 capsule
assembly, 10 static coMector.
a short length of capillary tube which damps out pressure surges.
Static pressure is exerted on the exterior of the capsule and is fed
into 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, gearing,
and a square-law compensating device to the pointer, which moves
over a scale calibrated in knots. The purpose of the compensator
is to provide a linear scale, and the principles of three commonly
used methods are described on page 88). Temperature compensation
is achieved by a bimetallic strip arranged to vai;y the magnification of
the lever system in opposition to the effects of temperature on
system and capsule sensitivity.
Square-law Cpmpensation
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 4.29
(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
87
Figure 4.29 Square-law
characteristics. (a) Effect
of linear deflection/
pressure response;
(b) effect of direct
magnification.
I
'150
t
200
0
250-
!I!
<I)
CAPSULE DEFLECTION - - -
(al
(bl
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 t·o overcome the effects of capsule
'drift' plus other mechanical irregularities as determined during
calibration.
There are two versions of this method in common use, the
principle being the same in both cases, i.e. 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.
The lever length referred to is the distance between the axis of the
main gear of the mechanism and the point of contact between its
rocking lever and the gear; this distance is indicated as d 1 and d 2 in
Figs 4.30 (a) and (b ) .
The method shown at (a) may be considered as a basic one and
serves as a very useful illustration of the operating principle. At the
starting position of the mechanism, the rocking lever and the sector
arm are in contact with each other at an angle preset for the ranging
of the instrument, therefore setting the distanced 1 • When pitot
pressure is applied to the capsule, the latter is deflected causing the
rocking shaft to rotate, and the rocking lever to move in a straight
line in the direction indicated. As may be seen from the diagram,
the rocking lever pushes the sector arm round and distance d I starts
increasing. The initial deflection of the capsule is of course small,
but this is magnified by the rocking lever contacting the sector arm
at the distance.e from the centre of the rocking shaft. Thus, a
magnified movement of the sector arm is obtained, and the sector
gear and pinion in turn provide further magnification to the pointer
so that it will travel through a large distance for the small pressure/
deflection characteristic of the capsule. Therefore the scale is
'opened up' at the low end of the speed range.
Assuming that the pitot pressure continues to increase, the
88
capsule will deflect at an increasing rate and the rocking lever
movement will follow the square-law deflections, but as its point of
-cohtact moves through the distance d 2 , its force acts further and
further along the sector arm and decreases the rotational movement
of the latter and also th e magnification of the pointer mechanism.
Therefore, & 'closing up' of the scale is obtained for increasing
airspeed, the initial settings of the whole arrangement being such
that, in following what may be termed a 'square-root law' , a linear
scale is produced throughout the speed range for which the
instrument is calibrated .
In the mechanism shown in Fig 4.30 (b ), the sector gear and arm
Figure 4.30 Methods of
square-law compensation.
(a) Rocking-lever/sector-arrn
mechanism.
SECTOR A R M - - - - . . . . ,
ROCKING L E V E R - - - - - - - - - i • - - , ,
I
0
I
'
/
/
I
I
I
I
I
I
I
.... __ _ , I
----J
DIRECTION OF
ROCKING LEVER MOVEMENT
{el
89
Figure 4.30 Methods
of square-law compen·
sation.
(b) 'Banana' slot.
"BANANA" SLOT
DIRECTION OF ROCKING
L£VER MOVEMENT
(bl
are replaced by a large circular gear and an integrally cut radial slot
which, for obvious reasons, is termed a 'banana slot'. The rocking
lever engages with this slot so that as the lever moves it rotates the
gear and changes the slot position, which decreases the magnification
in exactly the same manner as in method (a).
A third type of square-law compensating device is shown in
Fig 4.31. 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 the
speed and differential pressure increase, the spring rate increases
and its effective length is shortened; thus linearity is obtained directly at the capsule and not at the magnifying lever system as in
Fig 4.30.
Mach meters
90
With the advent of the gas turbine, the propulsive power available
made it possible for greater flying speeds to be attained, but at the
RANGING
SCREWS
Figure 4.31 Tuning spring
compensator. OX= effective spring length diminishing
as spring makes contact with
screws.
0
RANGING SPRING
same time certain limiting factors related to the strength of an airframe structure, and the forces acting on it, were soon apparent.
The forces which the complete structure. or certain areas of it,
experiences at high ~peed due to air resistance are dependent on how
close the aircraft's speed approaches that of sound. Since the speed
of sound depends on atmospheric pressure and density, it will vary
with altitude, and this suggests that in order to fly an aircraft within
safe speed limits a different airspeed would have to be maintained
for each altitude. This obviously is not acceptable and it therefore
became necessary to have a means whereby the ratio of the aircraft's
speed, V, and the speed of sound, a, could be computed from
pressure measurement and indicated in a conventional manner. This
ratio, V/a, is termed the Mach number (M), a parameter which has
now assumed great significance in practical aerodynamics, and the
instrument which measures the ratio M = V/a is termed a Machmeter.
Before dealing with the construction and operation of the instrument
we will review briefly some of the effects encountered by an aircraft
flying at high speed.
The passage of an aircraft through the air sets up vibratory
disturbances of the air which radiate from the aircraft in the form of
pressure or sound waves. At speeds below that of sound, termed
subsonic speeds, these waves radiate away from the aircraft in much
the same way as ripples move outward from a point at which a stone
is thrown into water. When speeds approach that of sound, however,
there is a d_rasti~ change in.the sound-wave radiation pattern. The aircraft is now travelling almost as fast as its own sound waves and they
begin to pile up on one another ahead of the aircraft, thus inceasing
the air resistance and setting up vibrations of the air causing turbu91
Figure 4.32 Machmeter.
1 Airspeed capsule, 2 altitude
capsule, 3 altitude rocking
shaft, 4 sliding rocking shaft,
5 calibration spring (squarelaw compensation), 6 calibration screws (square-law
compensation).
lence and buffeting of the aircraft, thus imposing severe stresses.
Shock waves are also developed which cause a breakdown in the airflow over wings, fuselage and tail unit and thus lead to stalling and
difficulties in movement of control surfaces. For regular flying in
this transonic range, and at supersonic speeds, aircraft must be
designed accordingly; for example, looking at some of to-day's highspeed aircraft, we note that the wings and horizontal stabilizer are
swept back; this delays the onset of shock waves and thus permits a
greater airspeed or Mach number to be attained.
It is possible for buffeting to occur when an aircraft is flying at
speeds below that of sound because the location, aerodynamic shape
and profile of parts of the structure may allow the airflow over them
to reach or exceed sonic speed. The Mach number at which this
occurs is referred to as the critical Mach number (Merit) of that particular aircraft, and being a ratio of airspeed and sonic speed, it will
be the same for any altitude.
A Machmeter is a compound flight instrument which accepts two
variables and uses them to compute tr.e required ratio. The
construction of the instrument is shown in Fig 4.32.
6
4
7--~
5
2
\·
~~~RE
~
~
-
I
--
PITOT
PRESSURE
MOVEMENTS OUE TO AIRSPEED CHANGES
- -
MD\/EMENTS DUE TD ALTITUDE CHANGES
The first variable is airspeed and therefore a mechanism based on
the conventional airspeed indicator is adopted to measure this in
terms of the pressure difference p - s, where p is the total or pitot
pressure ands the static pressure. The second variable is altitude ,
92
and this is also measure_d in the conventional manner, i.e. by means
of an aneroid capsule sensitive to the static pressures. Deflections
of the capsules of both mechanisms are transmitted to the indicator
pointer by rocking shafts and levers, 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 m.p.h. The speed of sound at sea-level
is app~oximately 760 m.p.h., 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 - s, 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, bu.t since it is governed by
static pressure conditions, the altimeter mechanism can do the next
best thing and that is to measures 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 V/a in terms of the pressure ratio (p - s)/s,.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 m.p.h ., and if an aircraft is
flown at 500 m.p.h. at this altitude, the Mach number will be
500/ 650 = 0. 7 5, a l 0% increase over its sea-level value. It is for this
·reason that critical Mach numbers (Merit) 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 functions in order to achieve this, by taking the case of
an aircraft having an Merit of say, 0.65. At sea-level and as based on
our earlier assumption, the measured airspeed would be 500 m.p.h.
to maintain Merit= 0.65. Now, if the aircraft'is to climb to and level
off at a flight altitude of 10,000 ft, during the climb the decrease of
static pressures causes a change in the pressure ratio. It affects the
pressure difference p - s in the same manner as a conventional
indic.ator is affected, i.e. the measured airspeed is decreased. The
airspeed mechanism therefore tends to r{iake the pointer indicate a
lo.w er Mach number.. However, the altitude· mechanism simultaneously responds to the decrease ins, 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
93
altered as the altitude mechanism divides p - s bys, 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
Machmeter.
Mach/Airspeed
Indicator
This indicator is one which combines the functions of both a
conventional airspeed indicator and a Machmeter, and presents the
requisite information in the manner shown in Fig 4.33. The
mechanism consists of two measuring elements which drive their
own indicating elements, i.e. a pointer anq a fixed scale to indicate
airspeed, and a rotating dial and scale calibrated to indicate Mach
number. A second pointer known as the velocity maximum
operating (Vmo) 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 (see page 92). 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 cf
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 jndicator. Two
external index pointers around the bezel may be manually set to any
MACH NO. SCALE
AIRSPEEO POINTER
Figure 4. 33 Mach/airspeed
indicator.
LIMIT SPEEO
IVMOI POINTER
COMMAND BUG
COMMAND BUG
SETTING KNOB
94
-
-
VMO SETTING
KNOB
desired reference speed, e.g., the take-off speeds V, 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 V mo 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.
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 4.33
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 pi tot and static pressures.
In addition to their basic indicating function, Mach/airspeed .
indicators can also be designed to actuate 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 autothrottle system, certain types of Mach/airspeed
indicator are designed to provide a speed error output which is
proportional to the d_ifference between the reading indicated by the
airspeed pointer and the setting of the command 'bug'. This is
accomplished by means of a CT/CX synchro 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 typt: of indicator is shown in Fig 4.34. It is very
similar in construction and presentation to the Mach/airspeed
indicator in that it employs pitof 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 page 106). 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 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 computed 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
95
Figure 4.34 Indicated/
computed airspeed
indicator.
MAXIMUM OPERATING
SPEED POlrHER
COMPUTED AIRSPEED COUNTER
COMMAND AIRSPEED
COUNTER
INDICATED AIRSPEED
POINTER
COMMAND AIRSPEED
SET KNOB
COMPUTED AIRSPEED
ON/OFF SWITCH
power to attain a commanded speed. In the example illustrated, the
command set knob mechanically adjusts a synchrotel (see page 240)
which also senses indicated airspeed. Thus, the synchrotel establishes
the airspeed error signal output required by the autothrottle computer. A read-out of the command speed is given on a digital
counter which is also mechanically set by the command speed knob.
Airspeed Switch
Units
Airspeed switch units, like their altitude counterparts, can be used
for a variety of warning applications; for example, in aircraft fitted
with a fatigue meter, a switch unit is employed to switch the meter
on and off at predetermined airspeeds; a unit may also be used to
operate an audio signal device and so give warning of an overspeed.
Whatever the application, the switch units are special adaptations of
conventional airspeed indicator mechanisms.
Two examples are shown in Fig 4.35. The switch mechanism of
Figure 4.35 Typical air·
speed switch units.
(a) Low-speed and high·
speed contact type;
(b) single-contact type.
DRUM SCALE FOR
ADJUSTMENT
(bl
96
the unit at (a) is housed in a standard airspeed indicator case and
consists of a capsule, a set of low-speed and high-speed con tacts and
a relay. A direct-current power supply of 28 Vis fed to the circuit
via a plug at the rear of the case.
When the differential pressure across the capsule increases it
expands until at a predetermined pressure it closes the low-speed
contacts. At a slightly higher predetermined differential pressure,
the high~speed contacts close and complete the circuit to the relay
coil thus energizing it. Operation of the relay then completes the
external circuit to the fatigue meter or other device connected to
the switch unit. In some applications, particularly those involving
audible warning devices, only one predetermined speed is required ;
in such cases, the circuit is arranged so that only the low-speed
contact or the high-speed contact completes the external circuit.
The switch unit shown at (b) basically employs the same number
of components as the one described in the preceding paragraphs. The
essential differences are the casing construction, single contacts
instead of double, plunger contact assembly instead of leaf contacts,
and an external scale for adjustment of the contact setting.
Combined Airspeed Indicator and Warning Switch
Figure 4.36 is a schematic of an airspeed indicator incorporating a
landing gear position warning system. The system comes into
operation when the approach speed of the aircraft is reached and the
Figure 4.36 Combined air-
speed indicator and warning
system. 1 Flag operating
coil, 2 contact 11ssembly,
3 capsule, 4 rocking shaft,
5 bimetal arm, 6 sector1
7 handstaff, 8 pointer,
9 rocking shaft flag
assembly.
I
)
I
,.
/
8
97
landing gear is not extended and locked in position. If this should
happen, a warning flag visible through an aperture in the dial,
adjacent to the approach speed graduations, commences to oscillate.
The system comprises a pair of contacts actuated by the capsule,
and connected to an operating coil and a relay. The relay, which is
supplied with 28 V d.c. via th~ aircraft's landing gear 'down' lock
system, controls the operating coil which, in turn, actuates the
warning flag by means of the rocking shaft assembly.
When the landing gear is retracted, the direct-current supply is
applied to the indicator, but while airspeeds are above the preset
value, the capsule holds the contacts open. As the airspeed decreases,
the capsule contracts until at the preset value the contacts close and
complete two parallel circuits. One circuit energizes the relay and
the other charges a capa~itor. Energizing of the relay causes its
contacts to change over, thus interrupting the supply to the relay
coil and also connecting a supply to the flag operating coil, causing
the flag to move into the dial aperture. At the same time the
capacitor is supplied with direct current and it starts charging. When
it discharges it does so through the relay coil and holds the contacts
in position until the discharge voltage reaches a point at ~hich it is
sufficient to hold the relay energized. The contacts change over
once again and de-energize the flag actuator coil, causing the flag to
move away from the dial aperture. The cycle is then repeated and
at such a frequency that the flag appears in the aperture at approximately.half-second intervals.
Mach Warning System
In many types of high-performance aircraft, a switch unit utilizing
the principle of the Machmeter is provided; its purpose being to give
an aural warning on the flight deck whenever the maximum
operating limit speed is exceeded. The mechanical and electrical
arrangement of a switch unit, based on that employed in the Boeing
747, is shown in Fig 4.37.
The switch contacts are actuated by the airspeed and altitude
capsule assemblies, and remain closed as long as the speed is below
the limiting value. The 28 V d.c. supply passing through the contacts
energizes the control relay which interrupts the ground connection
of the aural warning device, or 'clacker' as it is called from the sound
which it emits. When the limiting speed is exceeded, the switch
contacts open thereby de-energizing the control relay to allow d.c.
to pass through its contacts to activate the 'clacker' via the now
completed ground connection. The sound is emitted at a frequency
of7 Hz.
For functional checking of the system, a spring-loaded toggle switch
is provided. When placed in the 'Test' position it allows d.c. to flow t•
98
28 V
D.C. - - ; - - - - - - - - - - - - - - - - - - - - ,
'CLACKER'
STATIC PRESSURE
TEST SWITCH
SWITCH UNIT
Figure 4.37 Mach warning
system.
Vertical Speed
Indicators
the ground side of the control relay. thereby providing a bias
sufficient to de-energize the control relay and so cause the 'clacker'
to be activated.
These indicators, also known as rate-of-climb indicators, are the third
of the primary group of pitot-static flight instruments, and are very
sensitive differential pressure gauges, designed tp indicate the rate of
altitude change from the change of static pressure alone.
Now, it may be asked why employ a differential pressure gauge
which requires two pressures to operate it when there is only one
pressure really involved? Why not use an altimeter since it too
measures static pressure changes? These are fair enough questions,
but the operative clause is 'the. rate at which the static pressure
changes', and_as this involves a time factor we have to introduce this
into the measuring system as a pressure function. It is accomplished
by using a special air metering unit, and it is this which establishes
the second pressure required.
An indicator consists basically of three main components, a
capsule, an indicating element and a metering unit, all of which are
housed in a sealed case provided with a static pressure connection at
the rear. The dial presentation is such that zero is at the 9 o'clock
position; thus the pointer is horizontal during straight and level flight
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
99
the majority of applications indicators having a mechanism and 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 mark and so provides for better readability and for
more accurate observation of variations from level flight conditions.
The indicator mechanism, is shown in schematic form in Fig 4.38,
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 an airspeed indicator, i.e. it prevents
pressure surges reaching the capsule. It is, however, of a greater
length due to the fact that the capsule of a vertical speed indicator
is much more flexible and sensitive to pressure. 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: (a) level flight, (b) descent, and (c) climb.
In level flight, air at the prevailing static pressure is admitted to
the interior of the capsule, and also to the instrument case through
the metering unit. Thus, there is zero differential across the capsule
and the pointer indic;ates zero.
We will now consider the operation during a descent. At the
instant of commencing the descent the differential pressure will still
be zero, but as the aircraft descends into the higher static pressure
this will be applied at the static pressure connection of 'the instrument causing air to flow into the capsule and case.
As the capsule is directly connected to the static pressure
connection, the flow of air will create the same pressure inside the
capsule as that prevailing at the levels through which the aircraft
is descending. The pressure inside the case, however, is not going to
be the same because the metering unit is a specially calibrated leak
assembly designed to restrict the flow of air into or out of the
instrument case. Therefore, as far as the case pressure is concerned,
it is still at the same value which obtained at the original level flight
altitude, and cannot build up at the same rate as the pressure in the
capsule is increasing. The restriction of the metering unit thus
provides the second pressure from one source and establishes a
differential pressur~ across the capsule, causing it to distend and
make the pointer indicate a descent. This, of course, is just what is
required, but during the descent the case pressure must be maintained
lower than the capsule pressure and made to change at the same rate
in order to obtain a constant differential pressure.
The metering unit, being a restrictor, increases the velocity of the
air flowing into the static pressure connection, and as happens with
devices of this nature, increased velocity brings about a reduction in
pressure. In addition, the instrument case is of much greater volume
100
Figure 4.38 Principle of
vertical speed indicator.
(a) Level flight: zero
differential pressure
across capsule; (b) aircraft
descending: metering unit
maintains case pressure
lower than capsule
pressure, changing it at the
same rate and thereby
creating a constant differential pressure across the
capsule; (c) aircraft
climbing: metering unit
creates a constant differential pressure across capsule by maintaining case
pressure higher than capsule pressure.
METERING UNIT
\
F\
STATIC
PRESSURE-
\
CUMB
7,
CAP1UARY
DESCENT
(al
INCREASING
STATIC
PRESSURE
-
J
~
.--..
)
CLIMB
0
DESCENT
lb)
DECREASING
STATIC PRESSURE
OESCENT
(c)
than the capsule; consequently the flow of air into the case is going
to take some time to build up a pressure equal to that coming in at
the static pressure connection. By the time this is reached, however,
the aircraft will have descended to a new altitude and the static
101
pressure will have again changed. Thus, the metering unit introd uces
the required rate and time-lag factors, and differential pressure
across the capsule which posit ions the pointer to indicate the altitude
change in feet per minute. The design of a system is such that it
takes approximately four seconds for the case pressure to build up to
that in the capsule; but as the capsule always has an unrestricted air
flow to it , it will lead the case by four seconds and so there will be a
constant difference in pressure between them corresponding to four
seconds in time. The differential pressures produce_d are not very
large, a typical value being apr,,uximately 20 mm H2 0 at full-scale
deflection of the pointer.
During a climb, the metering unit will establish the required factors
and differential pressure, but as the static pressure under this condition is a decreasing one, and because the metering unit restricts the
flow out of the case, the case pressure leads the capsule pressure.
Apart from the changes of static pressure with changes of altitude,
which as we know are not constant, 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
air flow. From the theoretical and design standpoints, a vertical
speed indicator is therefore quite complicated, but the metering units
are designed to compensate for the effects of variables over the
ranges normally encountered.
The construction and operation of two typical units are described
in the following paragraphs.
Metering Units
The unit shown in Fig 4.39 (a) is known as a 'capillary-and-orifice'
type, the two devices in combination providing compensation for
the effects of the atmospheric pressure and temperature variables, as
shown at (b ).
The pressure difference across a capillary, for a constant rate of
climb, increases with increasing altitude and at a constant temperature. Thus, the use of a capillary alone would introduce a positive
error in instrument indications at altitudes above sea-level. With an
orifice, the effect is exactly the opposite. The primary reasons for
the difference are that the air flow through a capillary is a laminar
one while that through an orifice is turbulent; furthermore, the rate
of flow through a capillary varies directly as the differential
pressure, while that through an orifice varies as the square root of the
differential pressure. In combining the two devices we can therefore
obtain satisfactory pressure compensation at a given temperature.
The differential pressure across a capillary also depends on the
viscosity of the air, and as this is proportional to the absolute
temperature, it therefore decreases with decreasing temperature.
102
GASKETS
Figure 4.39 Vertical speed
AIR FILTER
indicator metering units.
(a) Capillary and orifice
type; (b) capillary and
orifice characteristics;
(c) ceramic type.
-AIRTO
CASE
fJt~C PRESSURE - ~ - - . . . - - 11:1..
(a)
-
CONNECTING
TUBE TO CAPSULE
(b)
500
40
50
x\000 FT
HEIGHT
SPRING-LOADED VALVE
SECOND CERAMIC LEAK
LEAK ASSEMBLY
Bl-METALLIC STRIP
FIRST CERAMIC LEAK
STATIC CONNECTION
DOME
FILTER
The differential pressure across an orifice varies inversely as the
temperature, and therefore increases with decreasing temperature.
Thus, satisfactory temperature compensation can be obtained by
combining the two devices. The sizes of the orifice and capillary are
chosen so that the readings of the indicator will be correct over as
wide a range of temperature and altitude conditions as possible.
The second unit, illustrated in Fig 4.39 (c),, is known variously
as the 'ceramic type' and 'porous-pot type', and is a little more
complicated in its construction, because a mechanical temperature/
viscosity compensator is incorporated.
It will be noted that the air from th~ static connection flows into
103
the capsule via a capillary tube, into the case via two ceramic
porous· tubes, and also through the valve of the viscosity
compensator. The valve opening is controlled by the effects of
temperature on a bimetallic strip.
Under low altitude conditions the effect of temperature on the
bimetallic strip is such that the valve is open by a certain amount,
so that after flowing through the first ceramic tube the air passes
into the case via the open valve. At higher al ti tu des the static air fed
to the instrument is at a lower temperature, and because its
viscosity decreases with decreasing temperature, the overall effect
is to reduce the pressure differential and give rise to errors.
However, the lower temperature also has its effect on the bimetallic
strip, causing it to bend and close the valve. The air must now also
flow through the second ceramic tube in order to get into the case,
and as they are in series and of calibrated porosity, the differential
pressure is increased and maintained. In practice, the valve takes
up positions between open and closed, but the calibration of the
metering unit as a whole gives a constant differential pressure for a
fixed rate of climb or descent at any altitude.
Typical Indicator
The construction of a typical vertical speed indicator employing an
orifice and capillary type of metering unit is shown in Fig 4.40. It
consists of a cast aluminium-alloy body which forms the support
for all the principal components with the exception of the metering
unit, which is secured to the rear of the indicator case. Displacements
Figure 4.40 Typical
vertical speed indicator
mechanism. l rocking shaft
assembly, 2 sector, 3 hand·
staff pinion, 4 gearwheel,
5 eccentric shaft assembly,
6 capsule plate assembly,
7 calibration springs, 8 capsule, 9 capillary tube,
10 calibration bracket,
11 static connection,
12 metering unit, 13 mech·
anism body, 14 hairspring,
15 link, 16 balance weight.
104
3
of the capsule in response to differential pressure changes are transmitted to the pointer via a link and rocking-shaft magnifying
system, and a quadrant and pinion. The magnifying system and
indicating element are balanced by means of an adjustable weight
attached to the rocking shaft. The flange of the metering u nit
connects with the static pressure connection of the indicator case,
and it also acts as a junction for the capillary tube.
Range setting of the instrument during initial and subsequent
calibrations is achieved by two calibration springs which bear on a
stem connected to the centre-piece of the capsule. The purpose of
these springs is to exert forces on the capsule and so achieve the
correct relationship between the capsule's pressure/ deflection
characteristics and the pointer position at all points of the scale.
The forces are controlled by two rows of screws, located in a calibration bracket, which vary the effective length of their respective
springs. The upper row of screws and the upper spring control the
rate of descent calibration, while the lower row of screws and lower
spring control the rate of climb.
A feature which meets a common requirement for all types of
vertical speed indicator is adjustment of the pointer to the zero
graduation. The fonn taken by the adjustment device depends on
the instrument design, but in the mechanism we have been
considering, it consists of an eccentric shaft coupled by a gearwheel
to a pinion on a second shaft which extends to the bottom centre
of the bezel. The exposed end of the shaft is provided with a screwdriver slot. When the shaft is rotated the eccentric shaft is driven
round to displace a plate bearing against the eccentric. The plate is
also in contact with the underside of the capsule, and as a result the
capsule is moved up or down, the movement b>eing transferred to
the pointer via the magnifying system and pointer gearing. The range
of pointer adjustment around zero depends on the climb and descent
range of the instrument but± 200 and ±400 ft/min are typical values.
Instantaneous Vertical Speed Indicators
These indicators consist of the same basic elements as conventional
VSls, but in addition they empl oy an accelerometer unit which is
designed to create a more rapid differential pressure effect , specifically at the initiation of a climb or descent.
The accelerometer COJJlprises two small cylinders or da!>hpots, containing pistons held in balance by springs and their own
mass. The cylinders are connected in the capillary tube leading to·
the capsule, and are thus open directly to the static pressure source.
When a change in vertical speed occurs initially, the pistons are
displaced under the influence of a vertical acceleration force, and
this creates an immediate pressure change inside the capsule, and an
105
instantaneous indication by the indicator pointer. The accelerometer
response decays after a few seconds, but by this time the change in
actual static pressure becomes effective, so that a pressure differential
is produced by the metering unit in the conventional manner.
Central Air Data
Computers
As we have already learned, the pressures on which the operation of
the primary flight instruments is dependent are transmitted througha system of pipelines. It should be apparent therefore, that the length
and quantity of the pipelines will vary according to the size of the
aircraft, and also on the number of stations at which indications of
the relevant air data are required. In order to minimize the 'plumbing'
arrangements, the concept of supplying the pressures to a special unit
at some centralized location, and then transmitting the air data
electrically to wherever required, was developed and resulted in the
design of units designated as central aid data computers (CADC). The
modular arrangement of a computer, and the methods by y.,hich signal
processing is carried out, can vary dependent on the number of parameters to be monitored, and also on the techniques adopted by any
one manufacturer. Basically however, a computer is an analogue
device that produces electrical signal equivalents of pitot and static
Figure 4.41 Modular
arrangement (If a CADC
ALTITUDE
MODULE
system.
-------,
log IS)
p- s
RATE OF
CLIMB
MODULE
log
p-s
p- s
$
TOTAL AIR
TEMPERATURE
I
MACH
MODULE
P.E.
CORRECTION
NETWORK
log
TAS
,---,.----,
I
TAS
TAS
~ ~ ~-~-------__j
====
=-
106
STATIC PRESSURE
PITOT PRESSURE
Figure 4.42 Example of a
force-balance pressure
transducer.
pressures by the combined operation of mechanical and synchronous
transmission· devices. The final computed output signals' are then
supplied to the appropriate indicators which, unlike their conventional counterparts, contain no pressure-sensing elements. The modular arrangement based on a representative CADC is shown schematically in Fig 4.41 .
Pressure sensing is accomplished by two pressure transducers, one
sensing static pressure within the altitude module, while the other
senses both pitot and static pressures within the computed airspeed
(CAS) module. The Mach speed module and true airspeed (TAS)
module are pure signal-generating devices, which are supplied with
altitude and airspeed signal data from the respective modules. Static
air temper:ature data required for T AS computation is sensed by a
probe (see page 269) located outside the aircraft at some predetermined position, and is routed through the Mach speed module.
An example of a pitot-static pressure transducer utilizing an 'E'
and 'I' bar type of inductive pick-off unit is shown in Fig 4.42 .
The transducer operates on what is termed the force-balance
principle, and comprises two capsules, the interior of one being
INPUT
I
I
'-A-MP_L_IFl~ER--'
)T PRESSURE
A
-
B
ITIC PRESSURE
I
PIVOTED BEAM
A.C. INPUT
[::,(> OUT OF BALANCE SIGNALS
I
__ _J
~
AMPLIFIED CONTROL SIGNAL
-
REFERENCE PHASE SIGNAL
107
connected to the pitot pressure source, while the other is connected
to the static pressures source. Both capsules are connected to a
pivoted beam which; in tum, is connected to the 'I' bar of the pickoff unit.
When a change in airspeed takes place, the capsules respond to the
corresponding change in differential pressure and the force they
produce deflects the pivoted beam thereby displacing the 'I' bar
relative to the limbs cf the 'E' bar. Thus, the air gaps are varied to
cawe out-of-balance signals to be induced in the outer limb coils in
the same manner as those induced in the servo altimeter pick-off
described earlier (see page 76). The signals are amplified and applied
to the control phase of the servomotor which drives an output shaft
and a lead screw. The lead screw is coupled to the pivoted beam via
a precision control spring so that as the screw rotates the spring
tension is varied, to balance the force exerted on the beam, and
therefore, to start 'backing-off the signal induced in the appropriate
outer limb coil. When a constant speed condition is attained,
equilibrium between capsule force and spring tension is established,
no further signals are fed to the amplifier and the servomotor ceases
to rotate. Since the output shaft is also rotated by the servomotor
then, by coupling the shaft to a CX synchro (see also Chapter 9), its
angular position can be measu_red in terms of the pressure differential
p - s applied to the transducer, and hence in terms of airspeed .
In the case of an altitude module which employs a force-balance
transducer, the construction is the same as that of an airspeed module, except that in place of the pitot pressure capsule, an evacuated
and sealed capsule is employed.
In some types of CADC altitude and airspeed transducers,
pressure sensing and transmission of corresponding signals is
accomplished by means of a silicon diaphragm containing piezoelectric elements. The elements function in a similar manner to a
strain gauge, in that their resistance changes as a function of the
strain imposed on the diaphragm under the influence of pressure.
The resulting voltage signal outputs are fed via integrated circuits to
the associated indicators.
Transmission of Airspeed and Altitude Data
The analogue signals from the CX synchros driven by the airspeed
and altitude transducers (see Fig 4.43) are transmitted to CT
synchros within the appropriate indicators, and the pointers are
motor-driven to their respective indicating positions in response to
the error signals produced in the CT synchros. At the same time,
the synchro rotors are driven to their corresponding ' null' positions.
As in the case of conventional airspeed indicators and altimeters,
compensation for the square-law characteristics and for the non108
g, ::?i
~-~·
.. . .
c:,.
~
~6
i3
~:
-
I
::s
MACH UNIT
- - - -- · - - ,
[! I I I I I I I ~ I I I I ~gRT~~Ci.EBN:1ACH OUTPUT
}
-,
P.E.C. POT"R
~~~ii
~~ii~P~ED I
TRANSDUCER
f~ r • • . • .. • • •
4
• .
I
~I
~I
I I I I I
.
,
I I I I I I I I I
~I
. .p
. . - -s
-- - ·
-
<
I I , I I I I I I I I I 11 I I 0
·
AIH:St'ttU o.;AM
·-
CAM
1
P.E.C. AS A
FUNCTION OF
MACH NO.
LOGS CAM ,'O
S
sj ~ ~~N~
SHAFT INPUT
FROM HEIGHT
TRANSDUCER
~ ~ I sI I I
E I I~
I I I I
~I
I I I
(' ~_d:M
ex
OUTPUT TO
VERTI CAL SPEED UNIT
OUTPUT TO AIRSPEED INDICATOR
-
SIGNAL FLOW
~
QI FFERENTIAL GEARS
•
-'V'.11/1/'-
I .. ",
P.E.C. POSITION
POT~NTIOMETER
. - ·-···.
I I I I • I •••• I •• I I I I I I I I I
LINEAR HEIGHT -
L_ _
0
ID
JJ
TO MACH
JCr.C MECHANICAL DRIVE
I
:~T~:~A~ I,. ~c ..
HEIGHT
CORRECTION
~~;~~~~~C~~6~uT
SQUARE-LAW
COMPENSATED
i
i
t
__J
LOG lp-s)-LOG S
LOG lp-s)
I
iPE.C.
I
ex
LI NEAR HEIGHT
CORRECTED FOR PE.-
~
OUTPUT TO Al TITUOE I NDICATOF
linear pressure/ height relationship is necessary. The principle of
compensation is basically the same as that described on page 87 , but
in lieu of a variable magnification lever system, scale linearity in a
CADC system is normally accomplished by means of cam s and cam
followers. The cams are driven by the output shafts of their
respective transducer units, and are so profiled that as they rotate ,
the cam followers are lifted at a controlled rate. Since the followers
are on shafts mechanically coupled to the rotors of the CX synchro.$,
then the rotors are turned through correspondingly controlled
angles so that 'linear-shaped' signals are transmitted to the CT
synchros within the associated indicators. The height correction
cam shown in Fig 4.43 is provided to correct for any inherent .
errors. in the shape of the height cam.
Transmission of Mach Number Data
In a CADC system, Mach number is computed by converting the
values of the pressure ratio p - s/sinto logarithmic values; such
conversion being carried out by means of cams and a differential
synchro (CDX). The mechanical arrangement is also shown in
Fig 4.43. The log (p - s) cam is rotated by the output shaft of the
airspeed transducer unit, and the angular movements of the cam
follower rotate the rotor of the CDX synchro. Th~ logs cam is
rotated by the output shaft of the altitude transducer unit, and its
cam follower rotates the CDX synchro stator. The signal output
from the synchro therefore corresponds to an angle proportional
to log (p - s) - logs or in other words, the signal output is the
logarithmic equivalent of the pressure ratio (p - s)/s.
·
The signals are transmitted to a CT synchro the output of which
is amplified to drive a servomotor and a cam which is profiled to
move its follower as a function of log [ (p - s)/s] . At the same time,
the cam follower drives the rotor of the CT to a position at which it
will 'null' the signal from the CDX. The motor also drives a
differential gear, and a position error correction cam, the follower
of which also provides a shaft input to the differential gear. Thus,
the rotation of the differential gear output shaft is proportional to
Mach number corrected for PE and this is transmitted to the
appropriate indicator via the signals generated by a CT synchro
transmission link:
Transmission of Vertical Speed Data
In a number of CADC systems, vertical speed is also transmitted by
way of a synchronous transmission link. This, generally, is
accomplished by means of a tachogenerator which is driven via the
output shaft of the height transducer; the tachogenerator thereby
produces signals equal to rate of altitude change. In order to ensure
110
that the signals relate to the linear height output to the altitude
indicator, the signals are modified by a potentiometer connected in
the output from the tachogenerator. The modified signals are then
transmitted to an amplifier and a servomotor which positions the
rotor of a TX synchro. The resultant error signals produced in the
stator of this synchro are supplied to the stator of its corresponding
TR synchro within the vertical speed indicator to produce a servo
drive to the pointer.
Pressure Error Correction
Correction for pressure error (see p_age 55) is effected by means of
cams profiled to suit the conditions applicable to particular types of
aircraft. The correction is applied as a function of Mach number
thereby correcting the indications of the Mach speed indicator, and
as will be noted from Fig 4.43 , the Mach unit servomotor drives a
second cam which, in conjunction with potentiometers, also corrects
the output to the altitude indicator for PE.
True Airspeed Computation
True airspeed (TAS) may be computed according to the formula:
TAS == aM
(I)
where
a == speed of sound in air
M = Mach number
The speed of sound, which cannot be measured directly , is proportional to the product of the square root of the absolute free air
temperature T and a constant C; thus a= C../T and by combination,
equation (I) becomes:
TAS = CM../T
(2)
Since it is not possible to directly measure T either, it is necessary
to obtain values from the adiabatic temperature formula:
T = Ti/(1 + 0.2KM 2 )
(3)
where
T; = absolute indicated stagnation temperature
K = a constant (the recovery factor of the temperature-sensing
element).
By combining equations (2) and (3):
(4)
111
This can be expressed as :
T AS = C . f(M) . ..,rr;-
(5)
where f(M) is the function of Mach number and is equal to
..jM 2 /1 + 0 .2 KM 2 • Equation (5) is therefore used as the basis for
TAS computation.
An example of a practical method of TAS computation is shown
in Fig 4.44. The total temperature is sensed by a resistance element
type of probe (see Fig 11 .7) which is connected to form part of a
MACH FUNCTION
INPUT
TEMPERATURE SIGNALS
1
I
TAS MODULE_
Figure 4.44 True airspeed
computation.
112
- --- - ----
bridge circuit. The other arms of the bridge are formed by a motordriven potentiometer, and two resistors of fixed ohmic value.
Changes in temperature cause the bridge to go out-of-balance and
the resulting signal is amplified and supplied to the control winding o:
the servomotor. The output shaft of the motor is geared to drive
the potentiometer in a direction that restores the bridge circuit to a
balanced condition. The calibration of the bridge circuit is such that
the motor is made to rotate through an angle proportional to log .../T;
and this temperature function is introduced as a shaft rotation
into a differential gear. The Mach function is also introduced into
the differential gear as a logarithmic value, and together with log .../T;
it produces a differential gear output proportional to log T AS.
This output is then converted to TAS by a cam which actuates a
follower coupled to a transmitting element that produces electrical
signals equivalent to TAS. In the example illustrated, the element is
in the form of a variable resistor which supplies varying voltages to a
servo-type digital counter indicator utilizing a d.c. torquer motor
(see also page 251 ).
Questions
4.1
4.2
4.3
4.4
4.5
4.6
4.7
4.8
4.9
4.10
4.11
4.12
4.13
4.14
4.1 S
4.16
4.17
4.18
4.19
4.20
4.21
What are the principal components and instruments which comprise
an aircraft pilot-static system?
Draw a line diagram of a dual pi tot and static system for port and
starboard instrument panels in an aircraft.
(a) Explain the principle of pilot pressure measurement and how
the \lzp V2 law is derived.
(b) Define the law to which current types of airspeed indicator are
calibrated.
Sketch and describe the construction of a pressure head.
How are the effects of turbulent air passing through the static slots
of a pitot-static probe neutralized?
Draw the circuit diagram of a typical pressure head heating system
and explain its operation.
What effects do the drain holes of a pressure head have on the
indications of the instruments connected to it?
(a) What is meant by the pressure error of a pilot-static system.
(b) How are its effects minimized?
(c) Explain why a vertical speed indicator is unaffected by pressure
error.
Define the following: (i) troposphere, (ii) tropopause,
(iii) stratosphere .
State two commonly used units of atmospheric pressure measurement.
What will be the effect on the density of an air mass if the temperature decreases but the pressure remains constant?
(a) What do you understand by the tenn 'standard atmosphere'?
(b) State the assumptions made by the ICAO Standard.
Describe how a mercury barometer measures atmospheric pressure.
What are the principal differences between a Fortin barometer and a
Kew barometer?
What is meant by a 'contracted inch' scale?
(a) Which type of barometer is used for the purpose of altimeter
calibration?
(b) What corrections must be applied to the pressure readings, and
why are they necessary?
With the aid of a diagram explain the operating principle of an
aneroid barometer.
Describe the construction and operation of an altimeter. Explain
any special features which improve its accuracy.
Discuss problems of the misreading of altimeters in association with
the instrument presentation. How are thes~ problems overcome?
What is the difference between 'pressure altitude' and 'indicated
altitude'?
Explain how an altimeter is compensated for errors due to atmospheric temperature changes.
113
4.22
4.23
4.24
4.25
4.26
4.27
4.28
4.29
4.30
4.31
4.32
4.33
4.34
4.35
4.36
4.37
4.38
4.39
4.40
4.41
4.42
4.43
4.44
4.45
4.46
4.47
114
Define the three principal Q codes used for altimeter pressure settings
For what purpose are altimeter switches provided in aircraft?
With the aid of a sketch, describe the construction and explain the
operation of a servo altimeter. State its advantages over a sensitive
altimeter.
Explain the sequence in which an altitude alerting unit operates when
an aircraft descends to a pre-selected altitude.
Which operating modes of an ATC SSR system are used for
identification and altitude reporting?
How does an airborne transponder 'recognize' the mode in which it
is being interrogated?
Assuming that for identification purposes the code 45 l O had been
selected, which reply pulses would make up the pulse train?
Describe how reply pulses are transmitted by an altitude encoder.
State the calibration law of ail airspeed indicator.
Describe the construction and operation of an airspeed indicator
(a) What is meant by 'square-law compensation' of an airspeed
indicator?
(b) With the aid of a diagram explain the operation of a typical
compensation method.
Define the following: (i) Mach number, (ii) critical Mach number.
Describe how Mach number is indicated by measuring in terms of the
ratio (p - Ps)fPs·
Describe how the function of a Machmeter and an airspeed indicator
can be combined to give an indication of maximum .safe airspeed.
Describe an electrical method of correcting position errors.
With the aid of a sketch, describe the construction of a vertical
speed indicator.
Explain the operation of a vertical speed indicator when the aircraft
in which it is installed goes from a level flight attitude into a climb
attitude.
With the aid of a diagram explain the construction and operation of a
metering unit incorporating a viscosity compensator valve.
Explain the operation of an instantaneous vertical speed indicator.
Explain some of th~ reasons why central air data computer systems
are used in aircraft.
Describe a typical method of sensing pi tot and static pressures in a
CADC system.
Explain how the signal output to a CADC airspeed indicator is compensated for the 'square-law'.
Describe how the logarithmic equivalent of the pressure ratio (p - s)/
is derived.
For what measurement would a tachogenerator be used in a CADC?
How are corrections for position error applied?
What is the formula used for the basis of true airspeed computation?
4.48
4.49
4.50
What types of synchros are used in a CADC and to which parameters
are they applied?
How are total air temperature signals sensed and applied to the
measurement of true air speed?
What is the purpose of the cam in a true airspeed module?
115
5 Primary flight
instruments (attitude
indication)
The Gyroscope and
its Properties
This chapter deals with the two flight instruments which provide a
pilot with the necessary indications of the pitch , bank and turn
attitudes of his aircraft. As both these instruments and the instruments covered in Chapters 6, 7 and 15 are dependent on gyroscopic
properties, this subject forms the opening to the present chapter.
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 spin and tilt axes (axis of veer ZZ 1 ).
The three degrees of freedom are obtained by mounting the rotor
in two concentrically pivoted rings, called inner and outer gimbal
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 5 .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.
Figure 5.1 _Elements of a
gyroscope.
ROTOR
,-~--.
:~.
TILTING
FREEDOM
116
INNER GIMBAL
'
lt~
SPINNING
FREEDOM
~,
The system will not exhibit gyroscopic properties unless the rotor
is spinning; for example, if a weight is hung on the inner gimbal ;ing,
it will merely displace the rings about axis YY I because there is no
resistance to the weight. When the rotor _is made to spin at high
speed 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, which means that tlie angular
momentum of a body about a given point remains constant unless
some force is applied to change it. Angular momentum is the
pro.duct 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 hung on the inner gimbal ring with the rotor
running, it will be found that the gimba-J ring will support the weight,
thus demonstrating the first fundamental property of rigidity. However, it will also be found that the complete gimbal system will start
rotating about the axis ZZ 1 , such rotation demonstrating the second
property of precession. Figure 5.2 illustrates how gyroscopic rigidity
may be demonstrated. If the frame and outer gimbal ring are tipped
about the axis YY 1 , the gyroscope maintains its spin axis in the
horizontal position. If the frame is either rotated about the axis
ZZ 1 or is swung in an arc, the spin axis will continue to point in the
same direction.
These rather intriguing properties can be exhibited by any system in
Figure 5.2 Gyroscopic
rigidity.
SUPPORT TIPPED:
GYRO MAINTAINS POSITION
t
I
I
I
f
I
/ -:;.·.:~:;'.'.:I,,
·--.
'
SUPPORT SWUNG IN ARC:
.....
t .. .GYRO CONTINUES TO POINT
IN Sl<ME DIRECTION
117
which a rotating mass is involved. Although it was left for man to
develop gyroscopes and associated devices, it is true to say that
gyroscopic properties are as old as the earth itself: it too rotates at
high speed and so possesses rigidity, and although it has no gimbal
system or frame on which external forces can act, it can, and does,
precess. There are, however, many mechanical examples around us
every day and one of them, the bicycle, affords a very simple means
of demonstration. If we lift the front wheel off the ground, spin it·
at high speed, and then turn the handlebars, we feel rigidity resisting
us and we feel precession trying to twist the handlebars out of our
grasp. The flywheel of a motor-car engine is another example. Its
spin axis is in the direction of motion of the car, but when turning a
corner its rigidity resists the turning forces set up, and as this
resistance always results in precession, there is a tendency for the
front of the car to move up or down depending on the direction of
the turn. Other familiar examples .are aircraft propellers, compressor
and turbine assemblies of jet engines; gyroscopic properties are
exhibited by all of them.
The two properties of an actual gyroscope may be more closely
defined as follows :
Rigidity. The property which resists any force tending to change
the plane of rotation of its rotor. This property 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 applied force, but
always at a point 90° away in the direction of rotation. The
rate of precession also depends on three factors: (i) the strength
and direction of the applied force, (ii) the moment of inertia of
the rotor, and (iii) the angular velocity of the rotoi:. 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 is in line with the plane of the applied
force and until the directions of rotation and applied force are coincident. At this point, since the applied force will no longer tend to
disturb the plane of rotation, there will be no further resistance to
the force and precession will cease.
The axis about which a torque is applied is termed the input
axis, and the one about which precession takes place in termed
the output axis.
118
Figure 5.3 Gyroscopic
precession. (a) Gyro
resists force ; (b) transmission of force;
(c) effect on rotor segments; (d) generation of
precession; (e) effect of
precession.
Determining the Direction of Precession
The direction in which a gyroscope will precess under the inOuence
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 deter-
2
PLANE OF
APPLIED FORCE
2
L
I
1
y '-
'
t--"
x--tf- I
x-j
l~- "-. I "
•
I
FORCE
Y,
F
21
\ PLAt~IN
2,
2
(a)
I
(b)
- X1
PLANE OF
APPLIED
2,
2
I
(c)
z
I
~-.a.i.
.17
::j
~
_/-y,
I
/ J
X -
21
21
(d)
I
PLANE OF SPIN
PLANE OF PRECESSION
(e)
119
Figure 5.4 Gy roscopic
precession. (a) Gyro resists
force; (b) transmission of
force: (c) effect on rotor
segments; (d) generation of
precession: (e) effect of
precession.
mining the direction in which a gyroscope will precess and also of
finding out where a force must be applied for a required direction of
precession. It is done by representing all forces as acting d irectly on
the rotor itself.
At (a) in Fig 5.3, the rotor of a gyroscope is shown spinning in a
clockwise direction and with a force F, applied upwards on the inner
gimbal ring. In transmitting this force to the rim of the rotor, as will
PLANE OF
APPLIED
FORCE
z
I
VJ-\ Ir" -,
PLANE OF
SPIN
<"
"-..
x-
I'
.
' ", J
I
z.
(b)
(a)
M
\
F
,.
/
---
- ---:-,
,,
F
X ----PRECESSION
~-'., , /
'\
/
(c)
M
z
z
I
: -'1
____ x, Y,
"-.._
>
--"'-..y
1
I
z,
(d)
120
(el
PLAN E OF
PRECESSION
be noted from (b), it will act in a horizontal direction. Let us
imagine for a moment that the rotor is broken into segments and
concern ourselves with two of th~rn at opposite sides of the rim as
shown at (c). Each segment has motion m in the direction of rotor
spin, so that when the force Fis applied there is a tendency for each
segment to move in the direction of the force. As the gyroscope
possesses rigidity this motion is resisted, but the segments will turn
abo-1t 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; therefore, when they are all joined to form
the solid mass of the rotor ·it will precess at an angular velocity
proportional to the applied force (see diagrams (d) and (e)).
In the example illustrated in Fig. 5 .4 (a), a force, F, is shown
applied on the outer ring; this is the same as transmitting the force
0n the rotor rim at the point shown in diagram (b ). As in the
1 ~vious example this ·results in the direction of motion changing to
the resultant of motion m and force F 1 • This time, however, the
rotor precesses about the axis YY I as indicated at (d) and (e).
References
Established By
Gyroscopes
For use in aircraft, gyroscopes must establish two essential reference
datums: a reference against which pitch and roll attitude changes
may be detected, and a directional reference against which changes
about the vertical axis may be detected. These references are
established by gyroscopes having their spin axes arranged vertically
and horizontally respectively , as shown in Fig 5.5.
X
I
z
VERTICAL AXIS
GYROSCOPE
DIRECTIONAL
y--~X1
-X-r»--Y,
21
HORIZONTAL AXIS
GYROSCOPE
Figure 5.5 References
established by gyroscopes.
121
Both types of gyroscope utilize the fundamental properties in the
following manner: rigidity establishes a stabilized reference unaffected
by movement of the supporting body, and precession controls the
effects of apparent and real drift thus maintaining stabilized reference
datums.
It will also be noted from Fig 5.5 that the pitch, roll, and directional attitudes of the aircraft are determined by its displacement
with respect to each appropriate gyroscope. For this reason, there- fore, the gyroscopes are referred to c1s displacement type gyroscopes.
Each one has the three degrees of freedom described on page 116, 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, vertical-axis and horizontal-axis gyroscopes are further classified as two-axis displacement gyroscopes.
Limitations of a Free
Gyroscope
Aircraft in flight are still very much a part of the earth, i.e. all
references must be with respect to the earth's surface. The free or
space gyroscope we have thus far considered would, however, serve
no useful purpose and must be corrected for drift with respect to the
earth's rotation, called apparent drift, and for wander as a result of
transporting the gyroscope from one point on the earth to another,
called transport wander.
Apparent Drift
The earth rotates about its axis at the rate of 15° per hour, and in
association with gyrodynamics, this is termed the earth rate (we).
When a 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 the orientation of its spin axis and its input axis (see
page 118, various components of the we 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 as it is normally termed.
This may be seen from Fig 5.6 (a) which illustrates a horizontal-axis
gyroscope at a latitude X. At 'A', the input axis is aligned with the
local N-S compon~nt of we; .therefore, to an observer at latitude X
the rotor and gimbal system of the gyroscope would appear to drift
clockwise (opposite to the earth's direction of rotation) in a horizontal plane relative to the frame, and at a rate equal to 15° cos X.
When the input axis is aligned with that of the earth ('B') drift would
also be apparent, but at a rate equal to we i.e. 15° per hour. If the
input axis is now aligned with the local vertical component of we
('C' in the diagram) the apparent drift would be equal to 15° sin X.
In order to further illustrate drift, we may consider diagram (b)
122
Figure 5. 6 Drift and
transport wander.
A
A
B
C
LOCAL NORTH
ALIGNEO WITH EARTH' S AXIS
LOCAL VERTICAL
w, EARTH'S RATE
X LATITUDE
15• COSAi\.?
I
Y,
'
'
(a)
X
--
AFTER 6
HOURS
~
"'
I
--x,---- -~
-
z
(c)
of Fig 5.6 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 have appeared 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.
123
If the same gyroscope we re to be positioned so that its input
axis ZZ 1 was aligned with the E-W component of we 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 , then there will be no apparent drift.
Real Drift
Real drift results from imperfections in a gyroscope such as bearing
friction and gimbal 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 set up
initially at the North Pole, with its input axis aligned with that of
the earth. As we have already seen, it will exhibit an apparent drift
equal to <.cJe. Assume now that the gyroscope is transported to a
lower latitude, and with its input axis aligned with the local vertical
component of we. During the period of trnnsport, it will have
appeared to an observer on the earth that the spin axis of the gyroscope has tilted in a vertical plane, until at the new latitude it
appears to be in the position shown in diagram (c) of Fig 5.6.
Apparent tilt, or transport wander as it is called, 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 we,
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 transport rate of this movement will be decreased or
increased depending on whether the E-W component of the
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 North or South component and also on
whether the gyros.cope is situated in the Northern or Southern
hemisphere.
The relationship between we, transport wander, and input axis
alignment are summarized in the table opposite.
If the gyroscope input axis were to be positioned such that its spin
axis is 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 as an attitude reference
in aircraft flight instruments and other associated navigational equip-
124
Input axis alignment
Local North
Earth Rate
we cos"
Transport Wander
nil
u
V
R
R
we= earth's angular velocity;
V
Local East
Local Vertical
we sin A
!!.. tan A
R
A= latitude; R = earth's radius;
= N-S component of transport velocity; U = E-W component
of transport velocity.
ment, drift and transport wander must be controlled so that the
gyroscope's plane of spin is maintained relative to the earth; in other
words, it requires conversion to what is tenned an earth gyroscope.
Control of drift which, as already pointed out, relates only to horizontal-axis gyroscopes, can be achieved either by (i) calculating
corrections using the earth-rate fomula 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 the earth rate
we (a practical example of this is given on page 179): (iii) applying
torques having a similar effect to that stated in (ii) but which can be
varied according to the latitude (see page 203).
The control of transport wander is normally achieved by using
gravity~sensing devices to automatically detect 'tilting of the gyroscope's spin axis, and to apply the appropriate corrective torques.
Examples of these devices are later described.
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 error.
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 the vertical-axis gyroscope
shown in Fig 5.5 becoming coincident with the ZZ 1 axis of the outer
gimbal ring. This means that the gyroscope would 'lose' its spin axis,
and since the rotor plane of spin would be at 90° to the ZZ 1 axis but
125
in the same plane a~ 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 turned, 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 gimb
rings not mutually at right angles to each other. Since the error is par
cularly relevant to horizontal-axis gyroscopes when used in directionindicating instruments, a more detailed description of how it is produced is given in Chapter 6 (page 180).
Methods of Operating
Gyroscopic Flight
Instruments
126
There are two principal methods used for driving the rot~rs of
gyroscopic flight instruments: pneumatic and electric. In the pneumatic method the case of an instrument is connected to either an
engine-driven vacuum pump, or a venturi located externally and in
the slipstream of a propeller. The pump, or venturi, creates a
vacuum which is regulated by a relief valve at between 3.5 and
4.5 in Hg. Certain types of turn-and-bank indicator operate at a
lower value, and this is obtained by an additional regulating valve
in the indicator supply line.
Each instrument has two connections: one is made to the pump
or venturi line, and the other'is made internally to a spinning jet
system and is open to the surrounding atmosphere. When vacuum
is applied to the instruments, the pressure within their cases is
reduced to allow the surrounding air to enter and emerge through
the spinning jets. The jets are adjacent to 'buckets' cut in the
periphery of each instrument rotor so that the jet stream turns the
rotor at high speed.
At high altitudes vacuum-driven gyroscopic instruments suffer
from the effects of a decrease in vacuum due to the lower atmospheric pressure; the resulting reduction in rotor speeds affecting
gyroscopic stability. Other disadvantages of vacuum operation
are weight due to pipelines, special arrangements to control the
vacuum in pressurized cabin aircraft, and, si nce air must pass through
bearings, the possibility of contamination by corrosion and dirt
particles.
To overcome these disadvantages and to meet instrumentation
demands for high-performance aircraft, gyroscopic instruments
were designed for operation from aircraft electrical systems. In
current applications this applies particularly to gyro horizons and
turn-and-bank indicators; electrically driven directional gyros form
part of remote-indicating compass systems, the principles of which
are dealt with 1n Chapter 7.
The power supplies generally used are 115 V, 400 Hz, 3-phase
current derived from an inverter or engine-driven alternator, and
28 V direct current, the latter being required for the operation of
some types of turn-and-bank indicator. The gyroscopes of alternatingcurrent instruments utilize the principle of the squirrel-cage
induction motor, and because the frequency of the power supply
is high, greater rotor speeds (of the order of 24,000 rev./min .)
are possible, thus providing greater rigidity and stability of
indications. The design of direct-current operated gyroscopes is
based on the principle of the conventional permanent-magnet type
of motor.
The Gyro Horizon
The gyro horizon , or artificial horizon as it is sometimes called,
indicates the pitch and bank attitude of an aircraft relative to the
vertical, and for this purpose employs a displacement gyroscope
whose spin axis is maintained vertical by a gravity-sensing device,
so that effectively it serves the same purpose as a pendulum but with
the advantage that aircraft attitude changes do not cause it to
oscillate.
Indications of pitch and bank attitude are presented by the
relative positions of two elements, one symbolizing the aircraft
Figure 5. 7 Gyro horizon
presentations. (a) Bottom
bank scale; (b) top bank
scale.
LEFT BANK INDICATION
(a)
(bl
127
itself, and the other in the form of a bar stabilized by the gyroscope
and symbolizing the natural horizon. Supplementary indications of
bank are presented· by the position of a pointer, also gyro-stabilized,
and a fixed bank angle scale. Two methods of presentation are
shown in Fig 5.7.
The operating principle may be understo0d by referring to Fig 5.8.
The gimbal system is arranged so that the inner ring forms the rotor
casing, and is pivoted parallel to the aircraft's lateral axis YY 1 ; and
the outer ring is pivoted parallel to the aircraft's longitudinal axis ZZ 1 •
The outer ring pivots are located at the front and rear ends of the
instrument case. The element symbolizing the aircraft may be either
Figure 5.8 Principle of
gyro horizon.
X
I
I
I
2
3
F'wo
I
4
5
8
X
I
X
I
X1
CLIMB Arnruoe
128
BAN K TO PORT
rigidly fixed to the case, or externally adjusted up and down for
pitch trim setting.
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 the aircraft's attitude, it goes into a climb say , the instrument case
and outer ring will tum 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. In a climb attitude the bar pivot
carries the rear end of the bar upwards causing it to pivot about the
stabilized actuating pin . The front end of the bar and the pointer
therefore move downwards through a greater angle than that of the
outer ring, and since movement is relative to the symbolic aircraft
element, a climbing attitude is indicated.
Changes in the lateral attitude of the aircraft, i.e. banking, displaces
the instrument case about the axis ZZ 1 and the whole stabilized
gimbal system. Hence, lateral attitude changes are indicated by
movement of the symbolic aircraft element relative to the horizon
bar, and also by relative movement between the bank angle scale and
the pointer.
Freedom of gimbal system movement about the roll and pitch
axes is 360° and 85° respectively, the latter being restricted by
means of a 'resilient stop'. The reason for restricting the pitch movement of a gyro horizon to 85° is to prevent 'gimbal lock' (see page
125).
Vacuum-Driven Gyro Horizon
A typical version of a vacuum-driven gyro horizon is shown in
Fig 5.9. The rotor is pivoted in ball bearings within a case forming
the inner ring, which in turn is pivoted in a rectangular-shaped
outer ring. The lower rotor bearing is fitted into a recess in the
bottom of the rotor casing, whereas the upper bearing is carried in a
housing which is spring-loaded within the top cap to compensate for
the effects of differential expansion between the rotor shaft and case
under varying temperature conditions.
A background plate which symbolizes the sky is fixed to the
front end of the outer ring and carries the bank pointer which
registers against the bank-angle scale.
The outer ring has complete freedom through 360° about the roll
axis. A resilient stop limiting the ±85° pitch movement is fitted on
the top of the rotor casing.
The horizon bar and pointer are an accurately balanced assembly
pivoted in plain bearings on the side of the outer ring and slotted to
engage the actuating pin projecting fro~ the rotor case. Pitch
129
Figure 5.9 Vacuum-driven
gyro horizon. 1 Sky plate,
2 inner gimbal ring, 3 resilient stop, 4 balance nut, S
temperature compensator,
6 rotor, 7 actuating pin, 8
outer gimbal ring, 9 actuator
arm, 10 pendulous vane unit,
11 buffer stops, 12 bank
pointer, 13 horizo11tal bar.
13
GIMBAL AND ROTOR ASSEMBLY
attitude changes are indicated by the pointer set at right angles to
the bar and positioned in front of the 'sky plate'.
In the rear end cover of the instrument case, a connection is
provided for the coupling of the vacuum supply. A filtered air inlet
is also proyided in the cover and is positioned over the outer-ring
rear-bearing support and pivot, 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 drilled in it.
With the vacuum system in operation, a depression is created so
that the surrounding atmosphere enters the filtered inlet and passes
through the channels to the 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 anticlockwise
direction as viewed from above. After spinning the rotor, the air
passes through a pendulous van~ unit attached to the underside of
the rotor casing, and is finally drawn of~ by the vacuum source.
The purpose and operation of the pendulous vane unit is
described under 'Erection Systems for Gyro .Horizons' on page 133
Electric Gyro Horizon
An example of an electric gyro horizon is shown in Fig 5 .10; as will
be noted, it is made up of the same basic elements as the vacuumdriven type, with the exception that the vertical gyroscope is a
3-phase squirrel-cage induction motor (consisting of a rotor and a
stator).
One of the essential requirements of any gyroscope is to have the
mass of the rotor concentrated as near to the periphery as possible,
130
Figure 5.10 Electric gyro
horizOJl.~.l,Miniature aircraft, 2 power-failure
indicator assembly, 3 gyro
assembly, 4 roll-torque
motor, 5 pitch-_torque motor, .
6 slip-ring assembly, 7 gyrogimbal contact assembly,
8 stator, 9 rotor, 10 pitch·
trim adjusting knob, 11 fast·
erection push switch.
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 inducation motor normally
has its rotor revolving inside the stator, but to make one small enough
to be accommodated within the space available would mean too small
a rotor mass and inertia. However, by designing the rotor and its
bearings so that it rotates on the outside of the stator, then for the
same required size of motor the mass of the rotor is concentrated
further from the centre, so that the radius of gyration and inertia are
increased. This is the method adopted not only in gyro horizons but
in all instruments and systems employing electric gyroscopes.
The motor assembly is carried in a hou sing which forms the inner
gimbal ting supported in bearings in the outer gimbal ring, which is
in turn supported on a bearing pivot in the front cover glass and in
the rear casting. The horizon bar assembly is in two halves pivoted at
the rear of the outer gimbal ring and is actuated in a manner similar
to that already described on page 129.
The 115 V 400 Hz 3-phase supply is fed to the gyro stator via slip
rings, brushes and finger contact assemblies. The instrument employs
a torque-motor erection system, the operation of which is described
on page 136.
When power is switched on a rotating magnetic field is set up in
the gyro stator which cuts the bars forming the squirrel-cage in the
rotor, and induces a current in them. The effect of this current is
131
to pr-:>duce magnetic fields around the bars which interact with the
stator's rotating field causing the rotor to turn at a speed of
approximately 20,000-23,000 rev./min. Failure of the power
supply is indicated by a flag marked OFF and actuated by a solenoid.
Standby Attitude hdicators
Many aircraft currently in service employ integrated flight systems or
flight director systems (see Chapter 15) which comprise indicators
that can display, not only pitch and roll attitude data from remotelylocated vertical gyroscope units, but also associated signal data from
radio navigation systems. Thus, it could be stated that there is no
longer a need for a gyro horizon. There is, however, an aitworthiness
requirement to meet the case of possible failure of the circuits
c·o ntrolling the display of aircraft attitude, and so the gyro horizon
still finds a place on the instrument panel, but in the role of a
secondary or standby attitude indicator.
An example of one such indicator which is, in fact, adopted as a
standby in the Douglas DCl 0, is illustrated in Fig 5.11 . The gyroscope is of the electrically-operated type and pcwered by 115 V,
3-phase a.c. supplied by a static inverter which, in tum, is powered
by 28 V d.c. from the battery busbar. Power from such a source is
always available, thereby further ens·uring that attitude indications
are displayed. In place of the more conventional stabilized horizon
bar method of displaying pitch and roll attitude, a stabilized spherical element is adopted as the reference against an aircraft symbol.
The upper half of the element is coloured blue to display climb
attitudes, while the lower half is black to display descending
attitudes. The dividing line between the two halves is engraved with
a circle at the centre of the line and represents the trne horizon.
Each half is graduated in ten degree increments up to 80° climb,
Figure 5. I I Standby attitude
indicator.
POWER 'OFF ' FLAG
PITCH ERECTION/TRIM KNOB
132
and 60° descent . Bank angle is indicated by a pointer and scale in
the normal manner.
The indicator has a pitch trim adjustment and a fast-erection facility, both being controlled by the knob in the lower righthand comer
of the indicator bezel. When the knob is rotated in its normally 'in'
position, the aircraft symbol may be positioned through ± 5° thereby
establishing a variable pitch trim reference. Pulling the knob out and
holding it energizes a fast-erection circuit (see also page 139). As with
any facility of this nature, time limitations are imposed on its operation.
Erection Systems For
Gyro Horizons
These systems are provided for the purpose of erecting the gyroscope
to its vertical position , and to maintain i_t in that 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. The
construction and operation of some systems which are typical of
those in current use are described on the following pages.
AIR FlOW FROM
ROTOR HOUSING
(a)
Figure 5.12 Pendulous vane
erection unit. (a) Construction; (b) precession due to
air reaction; ( c) gyro in
vertical position; (d) gyro
tilted.
(d)
133
Mechanical Systems
Pendulous Vane Unit
This unit, employed with the air-driven instrument described on page
120, is shown in detail in Fig 5.12. It is fastened to the underside of
the rotor housing and consists of four knife-edged pendulously
suspended vanes clamped in pairs on two intersecting shafts supported in the unit body. One shaft is parallel to the axis YY 1 and
the other parallel to the axis ZZ 1 of the gyroscope. In the sides of
the body there are four small elongated ports, one under each vane.
The air, after having spun the gyro rotor, is exhausted through the
ports, emerging as four streams; one forward, one rearward and two
lateral. The reaction of the air as it flows through the ports applies
a force to the unit body. The vanes, under the influence of gravity,
always hang in the vertical position, and it is this feature which is
utilized to govern the airflow from the ports and to control the
forces applied to the gyroscope by the air reaction.
When the gyroscope is in its normal vertical position as shown in
Fig 5.12 (b ), the knife-edges of the vanes bisect each of the ports
(A, B, C and D), making all four port openings equal . This means
that all four air reactions are equal and the resultant forces about
each axis are in balance.
If now the gyroscope is displaced from its normal vertical
position, 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) on the right-hand side of the body
and closing that (B) on the left. The increased reaction of the air
from the open port results in a torque being applied to the body in
the direction of the arrow, about axis XX 1 • •
This torque 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). As a gyroscope rotor always moves at a point 90° away from
the point of an applied torque, then in this case the rotor is precessed
at point P back to the vertical when the vanes again bisect the portsto equalize the air reactions.
Ball-type Erection Unit
This unit utilizes the precessional forces resulting from the effects
of gravity on a number of steel balls displaced within a rotating
holder suspended from the gyro housing.
The holder of the ball erector mechanism encloses from five to
eight balls, the number depending upon the particular design, which
are free to roll across a radiused erecting disc. A plate having a
number of specially profiled hooks is fixed around the inner edge of
the holder. The spacing of the hooks is chosen so as to regulate the
release of the balls when the gyroscope tilts, and to shift their mass
134
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 rotor shaft, the speed of the holder being 25
rev./min.
When the gyroscope is in its normal operating position as shown
in Fig 5.13 (a), the balls change position as the holder rotates but
their mass remains concentrated at the centre of the erecting 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.
SPIN AXIS
Figure 5. J3 Ball-type erection
,,X
unit. (a) Gyro vertical; (b) gyro
tilted away from front of .
instrument; (c) precession to
the vertical.
j
Z- - 1 - - - - T
z
---i-
0
-
-z 1
z,
A;iRO;
GRAVITY
CENTER
OF
OF MOVING
BALLS
Y1LEFT SIDE
(al
x,
(b)
I
y
I
p
4P
(cl
135
At (b) the gyro vertical axis is shown displaced about pitch axis
YY 1 away from the front of the instrument. The displacement of
the ball erector mechanism 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 baJls and the point at
which the force is acting will be carried round to the left-hand side
of the ball holder. In this position the balls remain hooked and their
mass remains.concentrated to allow a. torque to be exerted at the
left-hand side of the ball holder as indicated at (c). This torque may
also be considered as acting directly on the left-hand bearing of the
gyro housing and outer ring. Transferring this point of applied
torque to the rim of the rotor precession will then take place at a
point 90° ahead in the direction of rotation. As may be seen from
the diagram, the gyro housing will now start precessing about the
axis YY 1 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
more concentrated at this side allowing the torque and precession
to be maintained as they are C!1rried around to the left-hand side.
This action continues with diminishing movement of the balls as the
gyroscope erects to its normal operating 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 a gyroscope in other directions about its lateral
or longitudinal axis will result in actions similar to those described,
and it is left to the reader as a useful exercise to determine the
forces, torques and precession produced by both types of erection
system and for a chosen displacement.
Torque Motor and Levelling Switch System
This system is used in a number of electrically-operated gyro
horizons and consists of two torque control motors independently
operated by mercury levelling switches, which are mounted, one
parallel to the lateral axis, and the other parallel to the longitudinal
axis. The disposition of the torque motors and switches is illustrated
diagrammatically in Fig 5.14.
The laterally mounted switch detects displacement of the
gyroscope in roll and is connected to its torque motor so that a
corrective torque is applied around the pitch axis. Displacement of
the gyroscope in pitch is detected by the longitudinally mounted
levelling switch, which is connected to its torque motor so that
corrective torques are around the roll axis.
Each levelling switch is in the form of a sealed glass tube containing
136
Figure 5. 14 Auangement of a
torque-motor and levelling·
switch erectior. system. Pitch
torque motor: rotor fixed to
rear part of casing; stator fixed
to outer ring. Roll torque
motor: rotor fixed to outer
ring; stator fixed to gyrn
housing.
PITCH TORQUE MOTOR
SPIN AXIS
..,- z,
-
/
ROU MEf\CURV
(ON ~~C~ISI
:r·:
'
.......
,
,
ROLL TORQUE MOTOR
I
PITCH MERCURY I
(ONSWITCH
ROLL AXIS)
I
x,
three electrodes and a small quantity of mercury. They are mounted
in adjustable cradles set at right angles to each other on a switch
block positioned beneath the gyro housing. The tubes are filled with
an inert gas to prevent arcing at the electrodes as the mercury makes
contact and also to increase the rupturing capacity.
The torque motors comprise a squirrel-cage-type laminated-iron
rotor mounted concentrically about a stator, the iron core of which
has two windings, one providing a constant field and called the
'reference winding' , and the other in two parts so as to provide a
reversible field , and called the 'control winding'. Both windings are
powered from a step-down auto-transfonner connected between
phases A and B of the 115 V supply to the gyro horizon.
The electrical interconnection of all the components comprising
the system is indicated in Fig 5.15.
When the gyro is ·running and in its nonnal operating position , the
mercury in the levelling switches lies at the centre of the tubes and is
in contact with the centre electrode. The two outer electrodes, which
are connected across the control windings of the torque motor stators,
remain open. The auto-transformer reduces the voltage to a selected
value (typically 20 V) which is then fed to the centre electrode of the
137
&-1--- -- ~ - - o -
115V
I
~
i
fJ•
":lt__J~--j----~ 20V
'
'"---·'====~
·~
STEP· D
AUTO·TRAH
CAPACITOR
V AHO I
Figure 5. I 5 Circuit diagram
of torque-motor and
leveUing-switch erection
system.
138
ROTOR
...
¢::::I
t ll
TO OTHER lEVELUNG _
SWITCH AND TORQUE ":'
MOTOR
t==:>
NORMAL SUPPLY
. . FAST ERECTION SUPPLY
switches and the reference windings of the torque motors. Thus, in
the normal operating position of the gyroscope, current flows through
the reference windings only.
Let us consider what happens when the gyroscope is displaced
about one of its axes, to the front of the instrument , say, and about
the pitch axis YY 1 • The pitch-levelling switch will also be displaced
and the mercury will roll to the forward end of the tube to make
contact with the outer electrode. This -completes a circuit to one part
of the control winding of the pitch torque motor causing current to
flow through it in the direction indicated in Fig 5.15 . The stator of
the roll torque motor will also be displaced inside its fixed rotor, but
will receive no current at its control winding since the roll-levelling
switch is unaffected by displacement about the pitch axis.
The necessary corrective torque to the gimbal system must be
applied by the pitch torque motor, and in order to do this, the
magnetic field of its stator must be made to rotate. The voltage
applied to the reference winding is fed via a capacitor, and in any
alternating-current circuit containing capacitance, the phase of the
current is shifted S? as to lead the voltage by 90°. In the circuit to
the control winding there is no capacitance; therefore, the voltage
and current in this winding ue in phase, and since the reference and
control windings are both fed from the same source, then the reference winding current must also lead that in the control winding by
90°. This out-of-phase arrangement, or phase quadrature, applies
also to the magnetic field set up by each winding.
Thus, with current and flux flowing through the control winding
in the direction resulting from the gyro displacement considered, a
resultant magnetic field is produced which rotates in the stator in an
anticlockwi~e direction.
As the field rotates, it cuts the closed-circuit bar-type conductors
of the squirrel-cage rotor causing a current to be induced in them.
The effect of the induced current is to produce magnetic fields
around the bars which interact with the rotating field in the stator
creating a tendency for the rotor to follow the stator field.
This tendency is immediately opposed because the rotor is fixed
to the instrument case; consequently, a reactive torque is set up in
the torque motor which is exerted at the rear bearing of the outer
ring. We may consider this torque as being exerted at a point on the
gyro rotor itself so that precession will take place at a point 90°
ahead in the direction of rotation. This precession will continue
until the gyro and mercury switch are once again in the nonnal
operating position.
It will be clear from Fig 5.15 that displacement of the gyroscope
in the opposite direction will cause current to flow in the other part
of the levelling-switch control winding, thus reversing the direction
of the stator magnetic field and the resulting precession.
Fast-erection Systems
In some types of electrically-operated gyro horizon employing the
torque-motor method of erection, the arrangement of the levelling
switches is such that, if the gyro rotor axis is more than 10° from
the vertical, the circuits to the torque motors are interrupted so that
the gimbal system will never erect. For example, in one design a
commutator switch, known as a bank erection cut-ou t, is carried on
the outer gimbal ring about the roll axis, and serves to reduce erection
errors during turns involving bank angles greater than 10°, by opening
the circuits to both levelling switches. Thus, if on resuming level
flight the gimbal system has not remained accurately stabilized so as
to be within the 10° angle, the erection cut-ou t will maintain the
erection system in the inoperative condition.
Furthermore, it is possible for the gyroscope of a gyro horizon to
have 'toppled', or to be out of the vertical by too great an angle prior
to starting the instrument; then due to the low erection rate of the
system normally adopted, it would take too .long before the required
accuracy of indication was obt ained.
In order, therefore, to overcome these effects and to bring the
gyroscope to its normal operating position as quickly as possible, a
fast-erection system may be provided. Two typical systems in
current use are described in the following paragraphs.
Fast-erection Switch
This method is quite simple in operation. The switch (Fig 5. 15)
139
consists of several contacts connected in the power supply lines to
the erection-system torque motors and levelling switches.
Under normal operating conditions of the gyro horizon, the
switch remains spring-loaded to the 'off position and the lowvoltage supply from the auto-transformer passes over one closed
contact of the switch to the erection system, the other contacts
remaining open.
Whenever the gyroscope goes beyond the appropriate angular
limits, the erection system circuifmust be restored and the gyroscope's position brought back to normal as quickly as possible.
This is achieved by pushing in the switch so that the contact in the
low-voltage supply line opens to isolate the erection system from
the auto-transformer, and the upper contacts close. The closure of
these contacts completes the circuit to the torque motors and
levelling switches, but the power supply to them is now changed
over from the low-voltage value to the full line voltage of 115 V from
one of the phases. This results in an increase of current through the
stator windings of the torque motors, and the greater torque so
applied increases the erection rate from the normal value of 5° per
minute to between 120° and 180° per minute, depending on the
particular design.
There are two important precautions which have to be observed
when using this switch. Firstly, the switch must not be depressed
for longer than 15 seconds to prevent overheating of the stator coils
due to the higher current. The second precaution is one to be
observed under flight conditions; the switch must only be depressed
during straight and level steady flight and/or shallow angles of climb
or descent . If acceleration or deceleration forces are present, the
gyroscope will precess and produce false indications of pitch and
bank attitude.
Electromagnetic Method of Fast Erection
In this method, a circular-shaped electromagnet is secured to the
inside of the instrument case above an umbrella-shaped armature
mounted on the gyro rotor.housing. The armature is of approximately the same diameter as the magnet.
Control of the electromagnet and the erection time is achieved
by an auxiliary power control unit containing a three-phase transformer, bridge rectifier, thermally-operated time-delay relay and a
standard d.c. relay, all interconnected as shown in the circuit
diagram of Fig 5 .16.
When the normal 115 V alternating-current supply is initially
switched on, it is fed to contacts l and 2 of the standard relay, and
from one phase, through the time-delay relay, to the bridge rectifier.
The direct current obtained from the rectifier is then supplied to the
140
,--------------------0:}
--------1--------+-----~C
115V
3 PHASE
ELECTRO·
MAGNETIC
ERECTION
DEVICE
I
\GYRO
/ ROTOR
I
-
_.,.
/
I
BRIOGE .__......,.--.I......_--'
RECTIFIER
Figure 5. J6 Circuit diagram
of electromagnetic method
of fast erection.
coil of the electromagnet, which, on being energized, produces a
magnetic field radiating symmetrically from a small centre pole to
a circular outer pole.
If, at the moment of switching on the power supply the gyro
rotor housing and hence the armature are tilted away from the
centre of the magnet then the magnetic field is no longer symmetrical with respect to the centre of the armawre. Under these
conditions, therefore, a greater force is exerted on one side of the
armature than the other, and the applied torque is in such a
direction as to cause the gyro housing to erect to the vertical and
bring the top of the armature into line with the centre of the magnet
before the rotor is up to full speed.
In addition to passing through the electromagnet, the direct
current from the rectifier also passes through the coil of the standard
relay which is thus energized at the same time as the electromagnet.
The resulting changeover of the relay contacts causes the 115 V
supply to be fed to the tapping points 2 and 3 on the transformer
primary winding. This has the effect of reducing the number of
turns of the winding; in other words, the transformer is of the step-up
type, the voltage of the secondary winding in this particular application being increased to 185 V.
After approximately 20 seconds, the time-delay relay opens and
disconnects the direct current from the electromagnet. The standard
relay then de-energizes and switches the gyro rotor circuit from the
141
transformer to the normal 115 V supply, the rotor running up to full
speed some seconds later.
Erection Rate
This is the term used to define the time taken, in degrees per minute,
for the rotor axis of a vertical gyroscope to take up its vertical
position under the action of its gravity-sensing erection system.
For the ideal gyro horizon, the erection rate shou ld be as fast as
possible under all condit ions, but in practice such factors as speed ,
turning and acceleration of the aircraft, and earth's rotation all have
their effect and must be taken into account. During turns the erectior
system is acted upon by centrifugal forces and is displaced to make
the gyro follow it by precession. Therefore, the maximum erection
rate that can be used is limited by the maximum error that can be
tolerated during turns. The minimum rate is governed by the earth's
rotation, speed of the aircraft, and· random changes of precession due
to bearing friction, variations in rotor speed, and gjmbal system
unbalance.
Thus, erection systems must be designed so that, for small angular
displacements of the rotor axis 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
t o reduce the possibility of a slow cumulative error during manoeuvres
Normal erection rates provided by some typical erection systems
are 8° per minute for vacuum-driven gyro horizons and from 3° to 5°
per minute for electrically-driven gyro horizons.
Errors Due to Acceleration and Turning
As we have already learned, the erection devices employed in gyro
horizons are all of the pendulous gravity-controlled type. This being
so, 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 them the resulting torques will precess the gyro
axis to a false vertical position and so present a false indication of an
aircraft attitude. 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 (see Fig 5 .1 7).
The force set up by the acceleration will deflect the two athwartships-mounted vanes to the rear, thus opening the right-hand port.
The greater reaction of air flowing through the port applies a force
to the underside of the rotor and the torque causes it to precess
forward about the axis YY 1 • The horizon bar is thus displaced
downwards, presenting a false indication of an ascent.
142
Figure 5. 1 7 Acceleration
error. (a) Vane-type erection
system ; (b) levelling-switch
and torque-motor erection
system.
ACCELERATION
FALSE
VERTlCAI.
GREATER
REACTION
I
(at
x,
ION UNOERSIOE OF
ROTOR!
FALSE INDICATION
OF ASCENT
FALSE
VERTICAL
\
PlTCH TORQUE
MOTOR
q)
X
I
-=,.....:::~_-:_-3--·~
·r_______ -l
,,--t:,,,l,c-.,pffECESSIO;- • - - • -
• .......,
DISPLACEMENT
OF MERCURY
X
(b)
With the levelling-switch and torque-motor type of erection device,
the acceleration force will deflect the mercury in the pitch levelling
switch to the rear of the glass tube. A circuit is thus completed to
the pitch torque motor which also precesses the gyroscope forward
and displaces the horizon bar to indicate an ascent.
In both cases the precession is due to a natural response of the
gyroscope, and the pendulous vanes and the mercury always return
to their neutral positions, but for so long as the disturbing forces
remain, such positions apply only to a false vertical. When the forces
are removed the false indication of ascent will remain initially and
then gradually diminish under the influence of precession restoring
143
the gyro axis to its normally true vertical.
It should be apparent from the foregoing that, during periods of.
deceleration, a gyro horizon will present a false indication of a
descent.
When an aircraft turns, false indications about both the pitch and
bank axes can occur due to what are termed 'gimballing effects'
brought about by forces acting on both sets of pendulous vanes and
both levelling switches. There are, in fact, two errors due to turning:
erection errors and pendulosity errors .
Erection Errors
As an aircraft enters a correctly banked tum , the gyro axis will
initially remain in the vertical position and an accurate indication of
bank will be presented . In this position, however, the longitudinally
mounted pendulous vanes, or roll levelling switch, are acted upon by
centrifugal force. The gyroscope will therefore be subjected to a
torque applied in such a direction that it tends to precess the gyro
axis towards the aircraft perpendicular along which the resultant of
centrifugal and gravity forces is acting. Thus, the gyroscope is erected
to a false vertical and introduces an error in bank indication.
An analysis of the error can be made with the aid of Fig S.18,
which illustrates the case of an aircraft turning to starboard through
360° from a starting point A. The centrifugal force experienced by
the gyro axis in the false vertical position during the tum is constant
and at right angles to the instantaneous heading. This means that
when the aircraft changes its heading at a constant rate during a 360°
tum, the top of the gyro axis will trace out a circular path which is
90° in advance of the airc raft heading. The circle at the left of Fig
S.1 8 represents the path of the gyroscope axis, and any chord of this
Figure 5.18 Erection error.
090°
/
/ , __
PATH Of
TOPOf
GYRO AXIS
144
C'
.,
/
,/ D
circle will indicate the tilt of the axis in relation to the true vertical.
The chord AB', for example, represents the direction of tilt after the
aircraft has turned through 90°. In relating this tilt to the gyroscope
and the response of its gravity-controlled erection devices to the turn,
it can be resolved into two components, one forward and the other to
starboard. Thus, in addition to an error in bank indication an error in
pitch is presented when the aircraft is at point B of its tum. In a
similar manner, the chord AC' indicates the direction of tilt after
180° ; at this point the tilt is maximum and the bank error has been
reduced to zero, leaving maximum error in pitch indication. The
direction of tilt after 270° is indicated by chord AD', and resolving
this into its two components as at point D, we see that the pitch error
is the same as at B but the bank error is in the opposite direction. On
returning to point A the tilt of the gyro axis would be zero.
Compensation for Erection Errors
Erection errors may be compensated by one of the following three
methods: (i) inclination of the gyro spin axis, (ii) erection cut-out,
and (iii) pitch-bank compensation.
Inclined Spin Axis
The method of inclining the spin axis 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 forward (rearward in some
instruments) tilt to the gyro axis from the true vertical. The angle of
tilt varies but is usually either 1.6° or 2.5°. In vacuum-driven types
the athwartships-mounted pendulous vanes are balanced so that the
gyroscope is precessed to the tilted position ; in certain electric gyro
horizons the pitch mercury switch is fixed in a tilted position so that
090°
Figure 5.19 Compensated
erection error.
~
A'
I
\
\
!'',
Pit.lll Of
~
TOP Of
C-
,
G'fflO it.XIS
145
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.
The effect of the tilt is shown in Fig 5.19, where point A represents the end of the true vertical through the centre of the rotor, and
AA' represents the dirction of tilt (forward in this case). During
a tum to starboard the top of the gyro axis describes a circle about
point A at the same rate as the aircraft changes heading. The amount
of tilt and its direction in relation to the aircraft during the turn are
therefore constant.
Erection Cut-out
The erection cu t-out method is one applied to certain types of
electric gyro horizon and operates au tomatically whenever the ai rcraft banks more than 10° in either direction. It consists basically
of a commutator made up of a conducting segment and an insulated
segment , and two contacts or brushes connected in series with the
bank levelling switch. 1l1e commutator is located on the bank axis
at the rear of the outer gimbal ring, and in the straight and level flight
condition the two brushes bear against the conducting segment thus
completing the power supply circuit to the levelling switch.
During a turn there is relative movement between the commutator
and brushes due to banking, and when the angle of bank exceeds 10°
the insulated segment comes under o ne or other of the brushes and
so interrupts the supply to the bank levelling switch. Displacement
of the mercury by centrifugal force cannot therefore energize the
relevant torque motor and cause precession to a false vertical.
Owing to the function of the cut-out, no erection about the bank
axis is possible if the power supply is switched on to the instrument
when its gimbal system is tilted more than I 0° about this axis. However, the supply can be connected by means of a 'fast erection'
circuit which by-passes the cut-out in the manner described on page I
'Pitch-Bank Erection'
The third method , generally referred to as 'pitch-bank erection', is a
combined one in which the bank levelling switch is disconnected
during a turn and its erection system is controlled by the pitch
levelling switch. It is intended to correct the varying pitch and bank
errors and operates only when the rate of tum causes a centrifugal
acceleration exceeding 0 .18g, which is equivalent to a I 0° tilt of the
bank erection switch. The system is shown schematically in Fig 5.20,
and from this we note that two additional mercury switches,
connected as a double-pole changeover switch, are provided and are
interconnected with the normal pitch and bank erection systems.
Let us consider first a turn to the left and one creating a centri146
Figure 5.20 'Pitch-bank'
erection.
MERCURY DISPLACED
TO M-'l(E CIRCUIT TO
B-'NK TORQUE MOTOR
MERCURY DISPLAC:ED
TO INTERRUPT CIRCUIT
TO BANK TORQUE MOTOR
BANK MERCURY
SWITCH
....
PITCH
MERCURY SWITCH
---
SUPPLY WHEN CENTRIFUGAL
t::=c> ACCELERATION LESS THAN 0·18g
_ . SUPPLY WHEN CENTRIFUGAL
-'CCELERATION MORE THAN 0•18g
- - - SUPPLY IN CHANGEOVER FUNCTION
fugal accleration less than 0.18g. In such a turn, the mercury in the
hank levelling switch will be displaced to the right and will bridge the
gap between the supply electrode and the right-hand electrode, thus
completing a circuit to the bank torque motor. This is the same as if
the gyro axis had been tilted to the right at the commencement of
the turn; the bank torque motor will therefore precess the gyro to a
false vertical, left of the true one. At the same time, the gyro axis
tilts forward due to gimballing effect, and the mercury in the pitch
levelling switch, being unaffected by centrifugal acceleration, moves
forward and completes a circuit to the pitch torque motor, which
precesses the gyro rearwards. The two curved changeover switches,
which are also mounted about the bank axis, do not come into
operation since the mercury in each switch is not displaced sufficiently far to contact the right-hand electrodes. Thus, with centrifugal
acceleration less than 0.1 Bg there is no compensation.
Consider now a tum in which the centrifugal acceleration
exceeds 0.18g. The mercury in the bank levelling switch is
displaced to the end of the tube and so disconnects the normal
supply to the bank torque motor, i.e. it now acts as an erection cutout. However, the pitch levelling switch 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
changeover switches. Since the mercury in these switches is also
displaced by the centrifugal acceleration, a circuit is completed from
the lower switch to the bank torque motor, which precesses the gyro
axis to the right to reduce the bank error. At the same time, the
147
pitch levelling switch completes a circuit to the pitch torque motor,
which then precesses the gyro 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; hence the tenn
'pitch-bank erection'.
The changeover function of the curved mercury switches depends
on the direction of tilt of the gyro axis in pitch. This is indicated
by the broken arrows in Fig 5.20; the gyroscope and the pitch levelling switch now being tilted rearward, the latter connects a supply
to the upper of the two changeover switches and changes its
direction to the bank torque motor causing it to precess the gyroscope to the left.
The change in direction of the supply to the bank torque motor
is also dependent on the direction of the turn, as a study of Fig 5.20
will show.
As in the erection cut-out method of compensation, this system
requires a 'fast-erection' facility to bring the gyro axis to the true
vertical when it is tilted more than I 0° in either bank or pitch.
Since the forces and torques acting on the gyroscope depend on
the aircraft's speed and rate of tum, then obviously all erec;:tion errors
will vary accordingly, and this makes it rather difficult to provide
fore, pa~ticularly for instruments employing the inclined axis and
bank cut-out method of compensation, to base compensation on a
standard rate one tum of 180° per minute at an airspeed of 200
m.p.h. At other rates of turn and airspeeds the errors are small.
Pendulosity Errors
Pendulosity, or 'bottom heaviness' as it is sometimes called , is often
deliberately introduced in gyro horizons so that the gyroscope will
always be resting near its vertical position. This helps to reduce the
erection time when starting and also it prevents the gimbal system
from spinning about the bank and pitch axes during run-down of the
rotor. However, it can be acted upon by accelerating and decelerating
forces in straight and level flight, and centrifugal forces during turns;
consequently, it is an additional source of error; i.e. pendulosity error.
When acceleration takes place the base of the rotor assembly tends
to lag behind owing to inertia, i.e. it tends to swing directly rearwards.
In following the force through with the aid of the 90° precession rule,
it will be seen, however, that the rotor assembly will precess about the
bank axis to port or starboard depending on the direction of rotor
rotation. A deceleration has the opposite effect.
The pendulosity error resulting from a tum may be analysed in a
manner similar to that of erection errors. In Fig 5 .2 1 an aircraft is
again considered as turning to starboard through 360° from the point
148
Figure 5.21 Pendulosity error.
'
'
I
A2..__...;...;.,"'------""-----<-' C,
'
I
r
,
2
PATH OF TOP O
OF GYRO AXIS
WHEN INITIAL LEFT
TILT IS APPLIED
A. As the turn is entered the centrifugal acceleration tends to swing
the base of the rotor assembly to port, causing precession of the gyro
about the pitch axis, which again depends on the rotation of the
rotor. In this instance, the gyro axis tilts forward to a false vertical
and the instrument indicates an apparent climb. Throughout the
tum, the top of the gyro axis traces out a circular path which, unlike
that resulting from turning effects on erection systems, is synchronized with the aircraft's heading change. As before, any chord of the
circle from the point at which the tum commenced indicates the tilt
of the gyro axis in relation to the true vertical, and varying errors in
bank and pitch indications will be presented.
Compensation f or Pendulosity Errors
Compensation is usually effected by adopting the inclined-axis
method, the inclination in this case being about the bank axis, and
the direction being dependent on that of rotor rotation. The amount
of inclination is governed by the type of instrument, two typical
values \;>eing 0.5° and 1.75°.
The effect of the compensation, shown by the full circle in Fig
·s.2 1, is exactly the same as that produced by inclining the gy ro axis
in pitch, i.e. the top of the axis traces out a circular path about
itself to produce a singl e constant error.
Turn-and-Bank
Indicators
The tum-an d-bank indicator was the first of the aircraft flight
instrum ents to use a gyroscope as a detecting element, and in conjunction with a magnetic compass, it made a valuable contribution
to the art of flying without external references. It was thus
considered an essential primary 'blind flying' instrument for all
149
types of aircraft. However, with continued aircraft development,
changes in operational requirements, and the introduction of more
advanced flight instruments and systems, the place a turn-and-bank
indicator should occupy in the flight instrument group formed the
subject of much discussion. In the smaller types of aircraft, it still
functions as a primary instrument, but in many types of large and
more sophisticated aircraft, it may be used in a secondary role, or
even dispensed with altogether.
A tum-and-bank indicator contains two independent mechanisms:
a gyroscopically controlled pointer mechanism for the detection
and indication of the rate at which the aircraft turns, and a mechanism for the detection and indication of bank and/or slip. The dial·
presentations of two typical indicators are shown in Fig 5.22.
Figure 5. 22 Typical dial presentation of turn-and-bank
indicators.
Rate Gyroscope
For the detection of rates of turn, direct use is made of gyroscopic
precession, and in order to do this the gyroscope is arranged in the
manner shown in Fig 5.23. Such an arrangement is known as a
rate gyroscope.
It will be noted that the gyroscope differs in two respects from
those employed in directional gyros and gyro horizons; it has only
one gimbal ring and it has a spring connected between the gimbal
ring and casing to restrain movement about the longitudinal axis
YY 1 ; it is thus referred to as a single-axis gyroscope. Let us examine
Fig 5.23 a little more closely in order to understand how such an
arrangement can be made to indicate rates of turn.
When the instrument is in its normal operating position, due to the
spring restraint the rotor spin axis will always be horizontal and the
turn pointer will be at the zero datum mark. With the rotor spinning,
its rigidity will further ensure that the zero condition is maintained.
150
Figure 5.23 Rate gyroscope.
STRETCHING OF SPRING
FOR LEFT AND RIGHT TURNS
INPUT
AXIS
PRECESSION AXIS
y
.,....,--
" x,
Let us assume for a moment that the gyroscope has no spring
restraint and that the instrument is turned to the left abo ut a vertical
input axis. The gimbal ring will also tum , but as the rigidity of the
gyroscope resists this turning movement it will precess about axis
YY 1 • The direction of precession may be determined by the simple
rule already given. A tum to the left causes a force to be applied at
the front pivot of the gimbal ring, and this is the same as trying to
push the rotor round at the point F on its rim. In following this
through 90° in the direction of rotation, precession will take place
at point P, thus causing the gimbal ring and rotor to tilt about the
longitudinal axis. If a pointer were fixed to the gimbal ring, it t oo
would tilt through the same angle and would fndicate a turn and
also its direction. However, we are more interested in the rate at
which a turn is being executed, and to obtain an indication of this,
we control the angular deflection of the gimbal ring by connecting
it to the instrument case through the medium of a spring.
Considering once again the left tum indicated, the gyroscope
will now precess and will stretch the spring until the force it exerts
prevents further deflection of the gyro. Since precession of this
type of gyroscope is equal to the product of angular momentum of
the gyroscope and the rate of turn, then the spring force is a measure
of the rate of tum . If the spring is linear, i.e. its force is proportional
to the gimbal ring deflection, and the deflection is small, the actual
movement of the gimbal ring from the zero or rest position can be
taken as the required measure of turn 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-ofturn component not being at right angles to the spin axis during
gimbal ring deflection.
151
The rate-of-turn pointer is actuated by the gimbal ring and a
magnifying system the design of which varies between manufacturers. Scales are calibrated in what are termed 'standard' rates,
and although not always marked on a scale they are classified by the
numbers 1 to 4 and correspond to turn rates of 180, 360, 540 and
720° per minute respectively. The marks shown at either side of
zero of the scale in the right-hand upper corner of Fig 5.22
correspond to a Rate l tum.
A system for damping out oscillations of the gyrnscope 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
turn 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 system. For this
reason also, it is unnecessary for the rotor to turn at high speed, a
typical speed range being 4,000-4,500 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 the 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. Two principal mechanical methods may be
employed: one utilizing a gravity weight and pointer, and the other,
a ball in a curved liquid-filled glass t.ube (see Fig 5.22).
The gravity-weight method is illustrated schematically in Fig 5.24.
In normal flight, diagram (a), gravity holds the weight in such a
position that the pointer indicates zero. At (b) the aircraft is shown
turning to the left at a certain airspeed and bank angle. The indicator
case and scale move with the aircraft, of course, and because of the
turn, centrifugal force in addition to that of gravity acts upon the
weight and tends to displace it outwards from the centre of the turn.
However, when the turn is executed at the correct bank angle then
there is a balanced condition between the two forces and so the
weight and pointer .still remain at the zero position, but this time
along the resultant of the two forces. 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 the
weight and pointer will still remain at the zero position along the new
resultant of forces.
If the bank angle for a particular turn rate is not correct, say underbanked as in diagram (c), then the aircraft will tend to skid out of the
turn. Centrifugal force will predominate urider such conditions and
152
Figure 5.24 Gravity-weight
method of bank indication.
(a) Level flight; (b) correctly
banked; (c) underbanked
(skidding out of the tum);
(cf) overbanked (slipping into
the turn).
C.F.
(a)
W
R
C.F.
R
W
R
displace the weight, and the pointer from its zero 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 now have the predominant
effect on the weight. The pointer will thus be displaced from zero in
the opposite direction to that of an underbanked turn.
The effects of correctly and incorrectly banked turns on the balltype indicating element are similar to those described in the foregoing
paragraphs. The major differences are that the directions of ball
displacement are opposite to those of the pointer-type element
because the forces act directly on the ball. This is made clear by the
series of diagrams in Fig 5.25.
Typical Indicators
The mechanism of a typical air-driven indicator is shown in Fig 5 .26.
153
y
FiJ,'urC! 5. 25 Ball-type bank
indicatini: element . (a) Level
flight ; (b) correctly banked :
(c) undcrbanked /_skidding o u l
of turn) : (d) overbanked
(slipping into the turn).
, _j}_i,
~
l
W
A
w
(8)
(bl
1
C.F.
I
I
(cl
(d)
) ----':A
Figure 5.26 Mechanism of ::. n
air-0riven turn-and-bank
indicator. I Rotor, 2 instru·
ment frame, 3 damping cylin·
dcr, 4 buckets, 5 air bleed,
6 front plate, 7 rate-spring
adjusting screw, 8 dial, 9 rate
spring, 10 pointer, 11 agate
ball, 12 datum arrow, 13
gimbal front pivot, 14 slip
indicator, 15 expansion
chamber, 16 fluorescent card,
17 piston, 18 gimbal ring,
19 jet block, 20 jet.
9
10
14
13
Air enters the instrument through a filter situated at the rear of the
case and is led to a jet block via the inlet connecting tube. The jet is
set at an angle so that the air is directed on to the rotor buckets.
The direction of rotation is such that, with the indicator installed, a
point at the top of the rotor moves in the direction of flight. The
pointer moves over the scale, which has a centre-zero mark and a
mark at each end: Adjustment of the gyroscope sensitivity is provided by a screw attached to one end of the rate-control spring, the
screw protruding through a bracket mounted on the front plate of
the mechanism.
A stop is provided to limit the movement of the gimbal ring to an
154
angle which causes slightly more than full-scale deflection (left or
right) o f the 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
tube and its expansion chamber being concealed behind the dial and
clipped in position against a card treated with fluorescent paint.
The liquid used is white spirit.
The mechanism of a typical direct-current operated turn-and-bank
indicator is illustrated in Fig 5.27 . The gimbal system follows the
general pattern adopted for rate gyroscopes, varying only in construction attendant upon electrical operation.
Figurt 5.27 Mechanism of a
typic.tl direct current operated tum-and-bank indicator.
1 Case; 2 suppressor assembly ,
3 f~ spring, 4 rear end plate,
5 insulated connector, 6 magnetic dit!llping unit, 7 gimbal,
8 stirrup, 9 stirrup !llagnet,
10 flag spring, 11 ball, 12 bezel,
13 sli~ indicator, 14 pointer., 15
dial; 16 'off' flag, 17 rate scale,
18 .stirruf> arm, 19 dial frame,
20 froni 'frame plate, 21 gyroscope rotor, 22 brush feed
insulator.
2
4
Ill~- , - - -~
15..L!L._......-
-..;t,o\l
The rotor consists of a lap-wound armature and an outer rim
arranged concentrically, the purpose of the outer rim being to
increase the rotor mass and radius of gyration. The armature rotates
inside a cylindrical two-pole permanent-magnet stator secured to the
gimbal ring.
Direct current is fed -to the brushes and commutator via a radiointerference suppressor and flexible springs which permit movement
of the inner ring. The rotor speed is controlled by two identical
symmetrically opposed centrifugal cut-outs. Each cut-out consists of
155
a pair of platinum-tipped governor contacts, one fixed and one
movable, which are normally held closed by a governor adjusting
spring. Each cut-ou~ has a resistor across its contacts, which are in
series with half of the rotor winding. When the maximum rotor speed
is attained, centrifugal force acting on the contacts overcomes the
spring restraint causing the contacts to open. The armature current
therefore passes through the resistors, thus being reduced and
reducing the rotor speed. Both cut-outs operate at the same critical
speed.
Angular movement of the gimbal ring is transmitted to the pointer
through a gear train, and damping is accomplished by an eddy-current
drag system mounted at the rear of the gyro assembly. The system
consists of a drag cup, which is rotated by the gimbal ring, between a
field magnet and a field ring.
A power-failure warning flag is actuated by a stirrup arm pivoted
on the gimbal ring. When the rotor is stationary, the stirrup arm is
drawn forward by the attraction between a magnet mounted on it
and an extension (flux diverter) of the permanent-magnet stator. In
this condition the flag, which is spring-loaded in the retracted position
is depressed by the stirrup arm so that the OFF reading appears
through an aperture in the dial. As rotor speed increases, eddy
currents are induced in the rotor rim by the stirrup magnet, and at
a predetermined speed, reaction between the magnet and induced
current causes the stirrup arm to lift and the OFF reading to disappear from view.
Turn Co-ordinators
156
A turn co-ordinator (see Fig 5.28) is an interesting development of
the: turn and bank indicators just described, and is adopted in lieu of
such instruments in a number of small types of general aviation aircraft. The primary difference, other than the display presentation, is
in the setting of the precession axis of the rate gyroscope. The gyroscope is spring-restrained a·nd is mounted so that the axis is at about
30° with respect to the aircraft's longitudinal axis, thus making the
gyroscope sensitive to banking of the aircraft as well as to turning.
Since a turn is normally initiated by banking an aircraft, then the
gyroscope will precess, and thereby move the aircraft symbol to
indicate the direction of the bank and enable the pilot to anticipate
the resulting turn . The pilot then controls the turn to the required
rate as indicated by the alignment of the aircraft symbol with the
graduations on the: outer scale. In the example illustrated, the
graduations correspond to a rate 2 (2-minute) turn. Co-ordination
of the turn is indicated by the ball-type indicating element remaining
centred in the normal way (see page 152). In some turn co-ordinators,
a pendulous type of indicator may be adopted for this purpose.
The gyroscope is a d.c. motor operating at approximately 6,000
Figure 5.28 Turn coordinator.
rev ./min. In some types of tum co-ordinator the gyroscope may be
an a.c. brushless motor operating at constant frequency, and supplied
from a solid-state inverter housed within the instrument case. 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 a gyro horizon.
Damping of the gyroscope may be effected by using a silicone fluid
or, as in the instrument illustrated, by a graphite plunger sliding in a
glass tube. A small air line comes off the end of the tube to a valve
which can be adjusted to restrict the movement of the air inside the
tube. Its operation is similar to the damping device used in the turn
and bank indicator shown in Fig 5.26.
Questions
5.1
(a) Define the two fundamental p10perties of a gyroscope.
(b) On what factors do these properties depend?
5.2
5.3
5.4
5.5
How are the gyroscopic properties utilized in flight instruments?
What are the input and output axes of a gyroscope?
Why are displacement gyroscopes so called?
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'?
5.6
5.7
5.8
157
5.9
5.10
5.11
5.12
S .13
5.14
5.1 S
5.16
5.17
5.18
5.19
5.20
5.21
5 .22
5.23
158
With the aid of diagrams explain how a gyroscope precesses under the
influence of an applied torque.
What methods are adopted for driving the rotors of gyroscopic flight
instruments?
How is the gyroscopic principle applied to a gyro horizon?
Describe the construction and operation of an electrically-driven gyro
horizon including any special design features.
How are the gyroscopes of gyro horizons erected to and maintained in
their normal operating position?
Explain how the magnetic field set up in the stator of a torque motor
is made to rotate.
What are the functions of a 'fast-erection' system?
What precautions must be taken when using the levelling switch
method of fast erection?
(a) What effects does acceleration of an aircraft have on the indications of a gyro horizon?
(b) What do you understand by the terms 'erection error' and
'pendulosity error'?
What methods are adopted for the compensation of 'erection error'.?
Describe the operation of a method with which you are familiar.
How is compensation for 'pendulosity error' usually effected ?
Describe how the rate gyroscope principle is applied to a turn-andbank indicator.
(a) Why is it unnecessary to incorporate an additional erecting device
in a turn-and-bank indicator?
(b) Why is it important for a rate gyroscope to rotate· at a constant
speed?
(c) Describe how a constant speed is maintained in a d .c.-operated
turn-and-bank indicator.
With the aid of diagrams, describe how a ball type of bank indicator
indicates (a) a correctly banked turn, (b) a tum to starboard in which
the aircraft is overbanked.
Describe how a rate gyroscope may be utilized to sense both banking
and rate of tum.
6 Heading indicating
instruments
Direct-Reading
Magnetic Compasses
Direct-reading magnetic compasses 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 and navigational aids are concerned ,
however, such a directional reference is more accurately provided
by remote-reading compass systems, and flight director systems (see
Chapters 7 and 15) and so direct-reading compasses are relegated to
a standby role.
The operating principle of these compasses, and indeed of the
systems just referred to, is based on established fundamentals of
magnetism, and on the reaction between the magnetic field of a
suitably suspended magnetic element, and the field surrounding the
earth. It is useful at this stage therefore, to briefly study these fundamentals.
Magnetic Properties (Fig 6 .1)
First of all let us consider the three principal properties of a permanent magnet: (i) it will attract other pieces of iron and steel, (ii) its
power of attraction is concentrated at each end, and (iii) when
suspended so as to move horizontally, it always comes to rest in an
approximately North-South direc"tion. The second and third properties are related to what are tenned the poles of a magnet, the end of
the magnet which seeks North being called the North pole and the
end which seeks South the South pole.
When two such magnets are brought together so that both North
poles or both South poles face each other, a force is created which
keeps the magnets apart. When either of the magnets is turned round
so that a North pole faces a South pole again a force is created, but
this time to pull the magnets more closely to each other. Thus, like
poles repel and unlike poles attract ; this is· one of. the fundamental
laws of magnetism. The force of attraction or repulsion between
two poles varies inversely as the square of the distance between them.
The region in which the force exerted by a magnet can be detected
is known as a magnetic field . Such a field contains magnetic flux ,
which can be represented in direction and density by lines of flux .
159
Figure 6.1 Fundamental
magnetic properties.
H~~~:\!: :-
:'.:1!~f.d.{,
,,,,I~! "'~' ""~
'Wlt..,
• . ·.,
~
POLES
POLE
~ U N G POSITlON
APPROXIMATELY
NORTH-SOUTH
---------.. _
FIELO
6~~
~Y.§
SOfT-IRON
CORE
SOFT-IRON
CONCENTRATION OF FLUX
RING
MAGNETIC SCREENING
The conventional direction of the lines of flux outside a magnet is
from the North pole to the South pole. The lines are continuous and
unbroken, so that inside the magnet their direction is from South
pole to North pole. If two magnetic fields are brought close together
their lines of flux do not cross one another but form a distorted
pattern, still consisting of closed loops.
The symbol for magnetic flux is cl>, and its unit is the weber (Wb ).
The amount of flux through unit area, indicated by the spacing of
the lines of flux, is known as magnetic flux density (B); i.ts unit is
the weber per square metre, or tes/a (T).
Magnetic flux is established more easily in some materials than in
others: in particular it is established more easily in magnetic
160
materials than in air. All materials, whether magnetic or not, have a
property called reluctance which resists the establishment of magnetic
flux and is equivalent to the resistance of an electric circuit. It
follows that, if a material of low reluctance is placed in a magnetic
field , the flux density in the material wiil be greater than that in the
surrounding air.
Magnetic field strength, H, or the strength of a magnetic field at
any point is measured by the force, F , exerted on a magnetic pole at
that point. The force depends on the pole strength, i.e. the flux <I>
'emanating' from the pole* as well as on the field strength. In symbqls,
F
H =i
newtons per weber.
Thus the unit of His the newton per weber (N/ Wb) . A unit that is
more familiar to electrical engineers is the ampere per metre (Alm).
It can be shown that l N/Wb = l A/m.
Magnetic Moment
The magnetic moment of a magnet is the tendency for it to tum or
be turned by another magnet. It is a requirement in aircraft compass
design that the strength of this moment be such that the magnetic
detecting system will quickly respond to the directive force of a
magnetic field, and in calculating it the length and pole strength of a
magnet must be considered.
In Fig 6.2, suppose the pivoted magnet shown at (a) is of pole
strength <I> and the length of its magnetic axis isl, then its magnetic
moment m is equal to the product of the pole strength and magnetic
length, thus: m = lei>.
If now the magnet is positioned at right angles to a uniform
magnetic field H , the field will be distorted in order to 'pass through'
the magnet. In resisting this distortion, the field will try to pull the
magnet into alignment with it. Each pole will experience a force of
<PH newtons, and as the forces act in opposite directions they constitute a couple. Now, the torque, M, of a couple is the product of one
of the equal forces and the perpendicular distance between them, i.e.
M =l<l>H; but /<I> = m , so that M =mH.
From the foregoing it is thus evident that the greater the pole
strength and the longer the magnet, the greater will be its tendency
to tum into line with a surrounding magnetic field. Conversely, the
greater will be the force it exerts upon the surrounding field, or
indeed upon any magnetic material in its vicinity.
In Fig 6.2 (b ), the magnet needle is shown inclined at an angle(}
' 'Emanating' from the pole if it is a North pole; 'returning' to it if it is a South pole.
161
Fi}!lm! ti. 2 Mai:netic mom-·
enl. (a) Magnet at· right angles
to a unifo rm field; (b) magnet
at angle 8 to a uniform field.
FIELD H
(al
lb}
FIELD H
to the field H. The force on each pole is still cf>H, but the perpendicular distance between the forces is now SQ. Now SQ/ SN = SQ//
= sin 8 ; therefore SQ= / sin 8 . Thus the torque acting on the magnet
at an angle 8 is /4>H sin 8, or mH sin 8.
Magnet in a Deflecting Field
In Fig 6.3 , a magnet is situated in a uniform magnetic field H 1 and
a uniform deflecting field H1 is applied at right angles to H 1 . When
Figure 6.3 Magnet in a
deflecting field.
/
/
162
the magnet is at an angle B to field H 1 , as already shown, the torque
due to Hi is mHi sin O The torque due to H2 is mH2 cos O. Thus,
for the magnet to be in equilibrium, i.e. subjected to equal and
opposite torques, mH 1 sin B = mH2 cos O, so that the strength of the
deflecting field is H2 = Hi tan O.
Period of a Suspended Magnet
If a suspended magnet is deflected from its position of rest in the
magnetic field under whose influence it is acting, it at once experiences a couple urging it back into that position, and when the
deflecting influence is removed the magnet, if undamped, will
oscillate backwards and forwards about its equilibrium position
before finally coming to rest. The time taken for the magnet to swing
from one extremity to another and back again, i.e. the time for a
complete vibration, is known as the period of the magnet.
As t}:le magnet gradually comes to rest, the amplitude of the
vibration gradually gets less but the period remains the same and it
cannot be altered .by adjusting the amplitude. The period of a magnet depends upon its shape, size or mass (factors which affect the
moment of inertia), its magnetic moment, and the strength of the
field in which it vibrates. The period varies with these factors in the
following ways: (i) it grows longer as the mass increases;__(ii) it
becomes shorter as the field strength increases.
The vibrations of a magnet acting under the influence of a
magnetic field are very similar to those of an ordinary pendulum
swinging under the influence of gravity; the period T of a pendulum
is given by
- JIg
T- 211
where/ is the length of the pendulum and g the acceleration due to
gravity.
When a magnet of magnetic moment m is displaced through an
angle 8 then, as already shown, the torque, T, restoring it to the
equilibrium position is mH sin O. If I is the moment of inertia* of
the vibrating magnet about an axis through its centre of gravity
perpendicular to its length, then its angular acceleration is
If the displacement is small, sin B and O do not differ appreciably ,
•
Other symbols for moment of inertia are Kand J.
163
so that
111H sin O
b
.
mH
may e wntten - 1
1
and is constant. The motion is simple harmonic, having a period
given by
T,::; 2ir
/_j_
J"'n1H
Hard Iron and Soft Iron
'Hard' and 'soft' are terms used to qualify varieties of magnetic
materials according to the ease ,with which they can be magnetized .
Metals such as cobalt and tungsten steels are of the hard type since
they are difficult to magnetize but once in the magnetized state they
retain the property for a considerable length of time; hence the tenn
permanent magnetism . Metals which are easy to magnetize (siliconiron for example), and generally lose their magnetic state once the
magnetizing force is removed, are classified as soft.
These tenns are also used to classify the magnetic effects occ1,ming
in aircraft, a subject which is dealt with in detail in Chapter 8.
ferrestial Magnetism
164
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 magnet 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 on the earth's surface an imaginary
line called the magnetic meridian as shown in Fig 6.4.
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. However,. the origin of the field is still not exactly known,
but for purposes of explanation, the supposition of a bar magnet at
the earth's centre is useful in visualizing the general form of the
magnetic field as it is known to be.
The earth's magnetic field differs from that of an ordinary magnet
in several respects. Its points of maximum in tensity, or strength, are
not at the magnetic poles (theoretically they should be) but occur at
four other positions, two near each pole, known as magnetic foci.
Moreover, the poles themselves are continually changing their posi-
Figure 6.4 Terrestial mag-
ANGLE OF OIP
!INCREASING FROM EOUATO RI
netism. Lines AA , BB and CC
are isoclinals.
h
GLEOF
DIP ZERO
1
\
I
I
\
\
\
\
\
/
'
'',,,
'
HYPOTHETICAL MAGNETIC '
MERll>iAN
',
//
',
/ I
- - _ _,,__:'__.-'-- ACTUAL MAGNETIC
MERIOIAN
tions, and at any point on the earth's surface the field is not symmetrical and is subject to changes both periodic and irregular.
Magnetic Variation
As meridians and parallels are constructed with reference to the
geographic North and South poles, so can magnetic meridians and
parallels 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 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 at any place, the variation is said to be westerly.
When the inclination is to the right of the true meridian the variation
is said to be easterly. 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 natu·re of the rock disturbs the earth's 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
165
constant value in any one place, and the following changes, themselves not constant, may be experienced:
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 change (daily).
(i)
Infonnation regarding magnetic variation and its changes is given
on special charts of the world which are issued every few years.
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 needle 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. However, it will not
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 6.4. These lines emerge vertically from the North magnetic pole,
bend over and descend vertically into the South magnetic pole, and
it is only at what is known as the magnetic equator that they pass
horizontally along the earth's surface. If, therefore, a magnetic
needle 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
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.
Dip is conventionally considered positive when the red end of a freel)
suspended magnet needle dips below the horizontal, and negative
when the blue end dips below the horizontal. Hence all angles of
dip north of the .magnetic equator will be positive, and all angles of
dip south of the magnetic equator will be negative .
. The angle of dip at all pl~ces undergoes changes similar to those
described for variation and is also shown on charts of the world.
Places on these charts having the same magnetic dip 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 or Magnetic Intensity
When a magnet needle freely suspended in the earth's field comes to
166
rest, it does so under the influence of the total force of the earth's
magnetism. The value of this total force at a given place is difficult
to measure, but seldom needs to be known. It is usual, therefore,
to resolve this total force into its horizontal and vertical components,
termed H and Z respectively; if the angle of dip 8 is known, the
total force can be calculated.
A knowledge of the values of the horizontal and vertical
components is of great practical value, particularly in connection
with compass deviation and adjustment. Both components are
responsible for the magnetization of any magnetic parts of the aircraft which lie in their respective planes, and may therefore fluctuate
at any pl~ce for different aircraft or for different compass positions
in the same aircraft. The relationship between dip, horizontal force,
verticaHorce and total force is shown in Fig 6.5.
Figure 6.5 Relationship
betwe~n.dip, Z, Hand total
force.
a..- c = Vertical component Z
c - b = Horizontal component H
a - b = Total force T
Given angle of dip 8 and H,
l. = tan 8 and Z = H
H
tan 8
Ji= cos 8 and r = H cos 8
T
fl= H 2 + Z 2
As in the case of variation and dip, charts of the world are
published showing the values of Hand Z for all places on the earth's
surface, together with the mean annual change. Lines of equal
horizontal and vertical force 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 connection
with aircraft compasses this is the case; it is given relative to the
horizontal force at Greenwich.
Types of DirectReading Compass
Compasses have the following common principal features: a magnet
system housed in a bowl ; liquid damping~ and liquid expansion
compensation. The majority of compasses in use today are of the
card type as shown in. Fig 6.6.
Magnet System
An example of a magnet system is illustrated in Fig 6.7. It consists
167
Figure 6.6 Card-type compasses. (a) Suspended
mounting; (b) panel mounting.
Figure 6. 7 Typical
compass magnet system.
COMPASS
CARD
~~O:J~G__...--CUPHOLOER
BRACKET
SECURING
COMPLETE
SYSTEM TO
BOWL
solely of a single annular cobalt-steel magnet, to which is attached a
compass card. The suspension consists of an iridium-tipped pivot
secured to the centre of the magnet system and resting in a sapphire
cup supported in a holder or stem. The use of iridium and sapphire
in combination provides hard-wearing properties and reduces pivot
friction to a minimum . The card is referenced against a lubber line
fixed to the interior of the bowl and lying on or parallel to the longitudinal axis when the compass is installed in an aircraft.
Liquid Damping
The primary reason for filling compass bowls with a liquid is to
make the compass aperiodic. This is a term we apply to a compass
whose magnet system, after being deflected, will return to its
equilibrium position directly without oscillating or overshooting.
Other reasons for using a liquid are that it steadies the magnet
system and gives it a certain buoyancy, thereby reducing the weight
on the pivot and so diminishing the effects of friction and wear.
168
The liquids, which may be of the mineral or alcohol type, must
meet such requirements as low freezing point, low viscosity, high
resistance to corrosion, and freedom from discoloration.
Liquid Expansion Compensation
Compass liquids are subject to expansion and contraction with
changes in temperature, and the resulting changes in their volume
can have undesirable effects. For example, with reduction of
temperature the liquid would contract and so leave an air space in
the bowl thus reducing the damping effect. Conversely, expansion
would take place under high temperature conditions tending to
force the liquid out and resulting in leaks around bowl seals. It is
therefore necessary to incorporate a device within the bowl to take
up the volumetric changes and thus compensate for their effects.
The compensator takes the form of a flexible element such as a
bellows or a corrugated diaphragm which forms the rear part of the
bowl. When the bowl is filled the flexible element is compressed by
a specified amount by means of a special tool, the effect of this
compression being to increase the volume of the bowl. If now, the
compensated bowl is subjected to a low temperature the liquid will
contract, but at the same time the flexible element will respond to
the decrease in volume by expanding and filling up with liquid any
air space that may form. With an increase of temperature, the liquid
volume is further increased by expansion and so the flexible element
will be further compressed to take up the increase in volume.
The Effect of Dip on a Compass Magnet System
Dip, as we already know, is the angle that a suspended magnet needle
makes with the horizontal at any particular place owing to the
influence of the vertical component Z of the earth's field . Likewise
it is known that, for accurate indication of magnetic heading, we are
dependent only on the effect of the horizontal component H . This
being so, then maximum directional accuracy can only be obtained
at places where .dip is at, or approaching, zero, and since this is only
possible at the equator and its neighbouring latitudes, then a compass
utilizing a magnet system uncompensated for the effects of dip
would be very limited in its application as a heading indicator.
A compass must therefore be designed so as to neutralize the
effects of the vertical component Z over a much greater range of
latitudes, enabling its magnet system to remain horizontal or nearly
so. There are several ways of doing this, but the method which has
always proved to be the most effective is to make the magnet system
pendulous, i.e. to pivot it at a point above the centre of gravity, as
shown in Fig 6.8.
When the vertical component Z acts on the magnet system , the
169
PIVOT
Figure 6.8 Compensation
•• /
of magnetic dip .
POINT
z
latter is caused to tilt, drawing the centre of gravity of the system
away fr~m its position below the pivot point. A force is now acting
upward through the pivot and a second force is acting downward
through the centre of gravity, and as both forces are not acting along
the same line, a righting couple is introduced. The couple tends to
bring the magnet system once more into the horizontal position.
However, the vertical component Z is still being exerted on the
magnet system so that it .will not return to the horizontal position
exactly, except, of course, at the magnetic equator, where the value
of Z is zero.
Compasses are designed with a pendulosity such that the magnet
system is within approximately 2° of the true horizontal (when the
vertical component Zand gravity are the· only forces acting on it)
between latitudes 60° North and 40° South.
Compass Construction
170
The construction of one example of a card-type compass is shown in
Fig 6.9. The magnet system comprises an annular cobalt-steel magnet
and a light-alloy card mounted so as to be close to the inner face of
the bowl, thereby minimizing errors in observation due to parallax.
The card is graduated in increments of l 0°, intermediate indications
being estimated by interpolation. Observations are made against a
lubber line moulded on the inner face of the bowl.
Suspension of the system is by means of the usual iridium-tipped
pivot and sap 1 :.ire cup, the latter being supported in a holder
mounted on a stem and bracket assembly, secured to the rear of the
bowl. To prevent the magnet assembly from becoming detached
from the stem, should the compass be inverted , the clearance between
the top of the pivot nut and the bowl ceiling is less than the distance
between the sapphire cup and the top of the cupholder. When the
compass is returned to its normal attitude, the pivot is guided back
on to the sapphire cup by the cupholder, which has sloping and
polished sides.
·0· ANO ·c · CORRECTOR
INDICATORS
Figure 6.9 Sectional view of
a typical card compass.
HORIZONTAL
CORRECTORS
("B'
ANO
MOUNTING PLATE
'C')
BELLOWS
COMPASS CARO
The balance of the magnet system is such that its North-seeking
end is 2° down to compensate for the angle of dip (see page 169).
The bowl is moulded in Diakon and is painted on the exterior
with matt black enamel, except for a small area at the front through
which the card is observed. This part of the bowl is so moulded
that it has a magnifying effect on the card and its graduations. The
damping liquid is a silicone fluid (dimethylsiloxane-polymer), 1Y2oz
being required to fill the bowl. Changes in liquid volume due to
temperature changes are compensated by a bellows type of
expansion device, secured to the rear of the bowl.
The effects of deviation due to longitudinal and lateral
components of aircraft magnetism (see Chapter 8) are compensated
by permanent-magnet corrector assemblies secured to the compass
mounting-plate. An additional corrector assembly may also be
provided in some versions of this compass for use in aircraft requiring
compensation of a vertical magnetic component.
The card-type compass shown at (b) of Fig 6.6 is designed for
direct mounting on an instrument panel. The magnet system is similar
to that previously described except that needle-type magnets are
used . The brass case forms the bowl and is sealed by a front bezel
plate and a cover at the rear of the bowl. Changes in liquid volume
are compensated by a diaphragm type of expansion device sandwiched
171
between the rear part of the bowl and rear cover. A permanentmagnet deviation compensator is mounted on top of the bowl, the
adjusting spindles being accessible from the front of the compass. A
small lamp is provided for illuminating the card.
Acceleration And
Turning Errors
In the quest for accuracy of an indicating system, it is often found
that the methods adopted in counteracting undesirable errors under
one set of operating conditions are themselves potential sources of
error under other conditions. For example, when a compass magnet
system is made pendulous to counteract the effects of dip, the ·
compass can be used over a greater range of latitudes without
significant error; but unfortunately, any manoeuvre which introduces a component of aircraft acceleration, either easterly or
westerly of the earth's magnetic meridian, produces a torque about
the magnet system's vertical axis causing it to rotate in azimuth to a
false meridian.
There are two main errors resulting from these acceleration
components, namely acceferatto-n err_or and northerly turning error,
but before considering them in detail it is useful to consider·first
the effect which would be produced if a simple pendulum were to
be suspended in an aircraft.
So long as a constant course and speed are maintained, the
pendulum will remain for all practical purposes in the true vertical
with its centre of gravity directly below the point of suspension.
If, however, the aircraft turns, accelerates or decelerates, the
pendulum will cease to be vertical. This is because, owing to inertia,
the centre of gravity will lag relative to the pivot and move from
the normal position vertically below it. During a correctly banked
turn, the forces acting on the centre of gravity will cause the
pendulum to remain vertical to the plane of the aircraft and to take
up an angle to the true vertical equal to the angle of bank.
Since turns themselves are, in effect, acceleration towards their
centres, and whether correctly or incorrectly banked, always cause a
pendulum to take up a false vertical, it may be stated broadly that
any acceleration or deceleration of the aircraft will cause the centre
of gravity of a pendulum to be deflected from its normal position
vertically below the point of suspension.
From the foregoing, it is thus apparent that a magnet system
suspended to counteract the effects of dip will behave in a similar
manner to a pendulum; any acceleration or deceleration in flight
results in a displacement of the centre of gravity of the system from
its normal position.
Acceleration Error
Acceleration error may be broadly defined as the error, caused by
172
the effect of the vertical component of the earth's field, in the
directional properties of a suspended magnet system when the centre
of gravity of the system is displaced from its normal position, such
errors being governed by the heading on which acceleration or
deceleration takes place.
The force applied by an aircraft, when accelerating or decelerating
on any fixed heading, is applied to the magnet system at the point
of suspension, which is, of course, its only connection. The reaction
to this force will be equal and opposite and must act through the
centre of gravity below ·t he point of suspension and offset from it
d1:1e to the slight dip of the magnet system. The two forces constitute a couple which, dependent on the heading being flown by the
aircraft, causes the magnet system merely to change its angle of dip
or to rotate in azimuth.
Consider first an aircraft flying in the northern hemisphere and
accelerating on -a northerly heading. The forces bn:>Ught into play by
the acceleration will be as shown in Fig 6. l O (a). Since both the
point of suspension P and the centre of gravity are in the plane of
the magnetic meridian, the reaction R causes the N end of the
magnet system to nose-down, thus increasing. the dip angle without
any azimuth rotation. Conversely, when the aircraft decelerates, the
reaction at the centre of gravity tilts the needle down at the Send,
as shown at (b).
When an aircraft is flying in either the northern or southern
hemisphere and changing speed on headings other than the N-S
meridian, such changes will produce azimuth r0tation of the magnet
system and errors in indication.
Let us now consider the effects on the compass magnet system
Figure 6.10 Acceleration
errors. (a) Acceleration on
--N
northerly heading in
northern hemisphere;
(b) deceleration on northerly
(1)
heading in northern hemisphere;
(c) acceleration on easterly
heading in northern hemisphere;
(d) deceleration on easterly
heading in northern hemisphere.
DIP ANGLE
, , '
_j_~- -.. ~
---s
--,;,
................
\
.
C.G. \i---R
(b)
EASTERLY
DEVIATION
'\. N
N
-\
p
-E
-E
R
C.G.
R
(cl
WESTERLY
DEVIATION
I
s
(d)
s
173
when an aircraft flying in the northern hemisphere accelerates on an
easterly heading (Fig 6.10 (c )). The accelerating force will again act
through the point of suspension P and the reaction R through the
centre of gravity, but this time they are acting _away from each other
at right angles to the plane of the magnetic meridian. The couple
will now tend to rotate the magnet system in a clockwise direction,
thus indicating an apparent tum to the north, or what is termed
easterly deviation. When the aircraft decelerates as at (d) the
reverse effect will occur, the couple now tending to tum the magnet
system in an anticlockwise direction, indicating an apparent tum 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 result will be reversed in each case.
As northerly or southerly headings 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 errors
brought about by accelerations and decelerations, and that is the
effect of aircraft attitude changes. If an aircraft flying level is put
into a climb at the same speed, the effect on the compass magnet
system will be the same as if the aircraft had decelerated, because
the horizontal velocity has changed. If the change in attitude is also
accompanied by a change in speed, the apparent deviation may be
quite considerable.
Turning Errors
When an aircraft executes a tum, the point of suspension of the
magnet system is carried with it along the curved path of the tum,
whilst the centre of gravity, being offset, is subjected to the force
of centrifugal acceleration produced by the tum, causing the
system to swing outwards and to rotate so that apparent deviations,
or turning errors, are observed. In addition, during the tum the
magnet system tends to maintain 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 turn,
the angle of tilt would be maintained equal to the bank angle of the
aircraft, because the resultant of centrifugal force and gravity lies
normal to the transverse plane of the aircraft and also to the plane
through the magnet system's point of suspension and centre of
gravity. In this case, centrifugal force itself would have no effect
o_ther than to exert a pull on the centre of gravity and so decrease
the natural dip angle of the magnet system.
However, as soon as the magnet system is tilted, and regardless of
174
whether or not the aircraft is correctly banked, the system is free to
move under the influence of the earth's vertical component Z which
will 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 turning error is dependent upon
the aircraft heading, the angle of tilt of the magnet system, and the
dip. In order to form a clearer understanding of its effects on compass direction-indicating properties, we may consider a few examples
of aircraft heading changes from the magnetic meridian and in both
the northern and southern hemispheres.
Turning from a Northerly Heading towards East or West
Figure 6.11 (a) shows the magnet system of a compass in an aircraft
flying in the·northern hemisphere and on a northerly heading; the
north-seeking end of the sy~tem is coincident with the lubber line.
Let us assume now that the pilot wishes to make a change in heading
to the eastward. As 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 tum, and since the system is tilted, the
earth's vertical component Z exerts a pull on the N end causing
further rotation of the system. Now, the magnitude of magnet
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 by the compass: (i) a tum of
the correct sense, but smaller than that actually carried out when the
magnet system turns at a slower rate than the aircraft; (ii) no tum at
all when the magnet system and the aircraft are turning at the same
Figure 6. 11 Turning errors.
HEMISPHERE
NORTHERN
TURNING EAST
FROM
NORTHERLY
HEADING
TURNING EAST
FROM
SOIJTl,tERLY
HEADING
(cl
TURNING FROM
EAST TO SOVTH
OR
OR
WEST TO NORTH WEST TO SOUTH
c,@Jl}
(al
TURNING FROM
EAST TO NORTH
t
N
~Ar\
~E~E
(e
C.F.
(g)
S
t
s
N
SOUTHERN
C.F
N
-
C.F~C.G.
p
E
(d)
t -
®.·
~ - --
~C.G,
l p..-
''
(b)
t
E
C.F
(f)
s
~·----
(h)
I
~F.C.F.J
s
'
• •...........
175
rate; (iii) a turn in the opposite sense when the magnet 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 nothem hemisphere.
In the southern hemisphere (Fig 6.11 (b)) the effects are somewhat
different. The south magnetic pole is now the dominant pole and so
the natural dip angle changes to displace the centre of gravity to the
north of the point of suspension.
We may again consider the case of an aircraft turning from a
northerly heading to the eastward. Since the centre of gravity is now
north of the point of suspension, the centrifugal acceleration acting
on it causes the magnet system to rotate more rapidly in the opposite
direction to the tum , 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 6.11 (c))
then because the magnet system's centre of gravity is still south of
the point of suspension, the rotation of the system and the indications registered by the compass will be the same as when turning from
a northerly heading in the northern hemisphere.
In turning from a southerly heading in the southern hemisphere
(Fig 6.11 (d)) the magnet system's centre of gravity is north of the
point of suspension and produces the same effects as turning from a
northerly heading in the southern hemisphere.
In all the above cases, the greatest effect on the indicating
properties of the compass will be found when turns commence near
to northerly or southerly headings, being most pronounced when
turning through north . For this reason the term north erly 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 (Figs 6.1 1 (e)-(h)) no error in
indication results because the centrifugal acceleration acts in a
vertical plane th.rough the magnet system's point of suspension and
centre of gravity . As will be noted the centre of gravity is merely
deflected to the N or S of the point of suspension, 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 Wis 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 , and if the compass
has not been accurately corrected during a compass 'swing' any
uncorrected deviating force will become dominant and so cause
indications of apparent turns.
176
The Directional
Gyroscope
TI1e directional gyroscope was the first gyroscopic instrument to be
introduced as a heading indicator, and although for most aircraft
curren tly in service it has been replaced by remote-reading compass
systems and flight director systems, there are still some applications
of it in its vacuum-driven form. The instrument employs a horizontal-axis gyroscope, and , being non-magnetic, is used in conjunction with the magnetic compass; it defines the short-term heading
changes during turns, while the magnetic compass provides, a reliable
long-term headirtg reference as in sustained straight and level flight.
In addition, of course, the directional gyroscope overcomes the
effects of magnetic dip, and of turning and acceleration error
inherent in the magnetic compass.
In its basic form the instrument consists of an outer ring pivoted
about the vertical axis ZZ 1 and carrying a circular card graduated
in degrees. The card is referenced against a lubber line fixed to the
gyroscope frame . When the rotor is spinning, the gimbal system and
card are stabilized so that by turning the frame , the number of
degrees through which it is turning may be read on the card against
the lubber line.
The manner in which this simple principle is applied to practical
instruments is g<?verned by the manufacturer's design, but we may
consider the version illustrated in Fig 6. 12 and used in some light
aircraft not equipped with remote-indicating compasses.
The rotor is enclosed in a case, or shroud, and supported in an
inner ring which is free to tum about a horizontal axis YY, . The
inner ring is mounted in the vertical outer ring which carries the
compass card and is pivoted on a vertical axis ZZ 1• The bearings for
this ring are located in the top and bottom part of the instrument
case, which thus forms the gyroscope frame.
The front of the case contains a cut-ou t through which the card is
visible, and also the lubber line against which the card is referenced.
When the vacuum system is in operation a partial vacuum is
created inside the case so that the surrounding air can enter through
the filtered inlet and pass through channels in the gimbal rings to
emerge finally through jets. The air jets impinge on the rotor
'buckets' causing it to rotate at speeds between 12,000 and 18,000
rev./min.
To set the instrument so that it indicates the same heading as the
magnetic compass, a caging and setting knob is provided at the front
of the case. When this knob is pushed in, .an ann is lifted th ereby
locking the inner ring at right angles to the outer ring, and at the
same time meshing a bevel gear on the end of the caging knob
spindle with another bevel gear integral with the outer ring. Thus, a
heading can be set by rotating the caging knob and the complete
gimbal system . Once the correct setting is made the gyroscope is
freed by pulling the caging knob out. The reason for caging the inner
177
Figure 6.12 Air-<lriven
directional gyroscope.
(a) Airflow through instrument; (b) gimbal and gyro
rotor aisembly.
TO
VACUUM SOURCE
INLET
(a)
INNER GIMBAL
RING
CARD
ERE CTI ON
WEDGE PLATE
NOZZLE PILLAR
SYNCHRONISER RING
(b)
ring is to prevent it from precessing when the outer ring is rotated,
and to ensure that, on uncaging, its axis is at right angles to the outer
ring axis.
Control of Drift
Drift as we have already learned (see page 122 Chapter 5) is a funda178
mental characteristic of gyroscopes, particularly those of the
horizontal-axis type, and so for practical direction-indicating
purposes, earth-rate error, transport wander, and real drift (as
defined earlier) must be controlled . In the instrument shown in
Fig 6.12, control is effected by gimbal-ring balancing, and by erection
devices.
Gimbal Ring Balancing
The method of controlling earth-rate error is to deliberately
unbalance the inner ring so that a constant torque and precession are
applied to the gimbal system. The unbalance is effected by a nut
fastened to the rotor housing, and adjusted during initial calibration
to apply sufficient torque and precession of the outer ring to cancel
out the drift at the latitude in which it is calibrated. For all
practical purposes this adjustment is quite effective up to 60° of
latitude on the earth's surface. Above these latitudes the balancing
nut has to be readjusted .
Erection Devices
Erection devices 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 instrument already described, after
spinning the rotor the air exhausts through an outlet in the periphery
of the rotor case and is directed on to a wedge-shaped plate secured
to the outer ring.
In Fig 6.13 (a), the rotor axis is shown horizontal, and so the
exhaust air outlet is upright and directly over the high point of the
wedge located on the centre-line of the outer ring. The wedge therefore divides equally the air flowing from the outlet, and the air
reaction applies horizontal forces R I and R 2 to the faces of the
wedge. Since these forces are equal and opposite no torque is applied
to the outer ring and the rotor axis remains horizontal.
When the rotor axis is tilted from the horizontal position the air
outlet is no longer bisected by the wedge plate and a greater amount
of air strikes one face of the wedge. From Fig 6.13 (b ) , it can be
seen that the horizontal force R, is now greater than R 2 and a torque
will therefore be applied in the direction of R, about the vertical axis
ZZ,. This torque may be visualized as being applied to the rim of the
rotor at point F, causing the rotor to precess from this point about
axis YY I until its axis is horizontal and forces R, and R 2 are again equal.
If the rotor tilts to such an angle that the exhaust air is entirely off
the wedge plate, a secondary erection torque will be produced by the
air stream issuing from the rotor spinning jet. In the normal horizontal attitude of the rotor the air stream impinges on the ce~tre of
the buckets as shown in Fig 6.13 (c). In the exc~ssive tilt condition,
179
Figure 6. I 3 Erection device.
(a) Rotor axis horizontal ;
(b) rotor axis tilted ;
X---
(c ) secondary erection torque ;
(d) precession from excessive
tilt position.
11.,
~~
~· ·\..,i
··,:s
..... .
.
(
·~i>.~ . rt,
(a)
(b)
however, the air stream strikes nearer to one side of the buckets and
produces a force F which can be visualized as being a force Ft , acting
at the point shown at (d). This force produces corresponding
precession at point P 1 to return the rotor to its normal horizontal
attitude. The force F will diminish as the rotor returns, but by this
time the exhaust stream will again be in contact with the wedge plate
so that final erection will be achieved in the manner indicated at (b ) .
Gimbal Errors
A definition of gimbal error has already been given (see page l 26) and
in the practical case of a directional gyroscope, errors in indications
are dependent upon: (i) the angle of climb, descent or bank, and
(ii) the angle betw~en the rotor axis and longitudinal axis of the ai rcraft.
In Fig 6.14, a series of diagrams illustrates the gimbal system
geometry when an aircraft is in particular attitudes. At (a) the aircraft is represented as flying straight and level on an easterly heading,
and as the design of the gimbal system geometry is such that the rotor
axis lies along the N-S axis, the three axes of the gimbal system are
mutually at right angles, and the directional gyroscope will indicate
the aircraft's heading without gimbal error. The same would also be
180
Figure 6.14 Gimbal errors.
(a) Aircraft flying straight and
level on easterly heading; no
error; (b) aircraft banked to
port on easterly heading: no
error; (c) aircraft descending
on easterly heading: no error;
(cl) aircraft banking to port
and descending: error introduced; (e) aircraft flying on
intercardinal heading; errors
introduced.
FRAME
OUTER GIMBAL RING
(a)
(C}
&
(b)
(d)
(e}
true, of course, if the aircraft were flying on a westerly heading.
If the aircraft banks 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 with the aircraft about the axis of the stabilized
· inner gimbaJ ring (diagram (b)). In this condition also the instrument
would indicate, without gimbaJ error, the cardinaJ heading or change
of heading during a tum.
At (c) the aircraft is assumed to be descending so that, in addition
to the outer gimbal ring being tilted forward about the rotor axis, the
inner gimbaJ aJso rotates, both rings maintaining the same relationship
to each other. Again, there is no gimbal error; this would also apply in
the case of an aircraft in a climbing attitude.
When an aircraft carries out a manouevre which combines changes
in roll and pitch attitude, e.g. the banked descent shown at (d) , the
181
outer gimbal ring is made to rotate about its own axis, thus introducing a gimbal error causing the directional gyro to indicate 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
gimballing errors will occur during turns, banking in straight and level
flight, pitch attitude changes or combinations of these.
When the heading of an aircraft is such that its longitudinai axis-is
aligned with the gyroscope rotor axis, banking of the aircraft on a
constant heading will not produce a gimballing error because
rotation of the gimbal system takes place about the rotor axis. If,
however, banking is combined with a pitch attitude change, the effec1
is the same as the combined manoeuvre considered above and as sho'>'
at (d).
Calculation and plotting of the errors on all headings and for variot
bank, descent and climb angles produces the sine curves of Fig 6.15.
It will be observed that in the four quadrants there are alternate
positive and negative errors which, when applied, produce the characteristic acceleration and deceleration of the outer ring under the effec
of gimballing. It should also be noted that the errors plotted relate ta
directional gyroscopes, which on being set to indicate N or S, have th1
rotor axes aligned with the longitudinal axis of the aircraft. This
may not apply to all types of instrument; however,·with any other
setting of the rotor axis the error curves will merely be displaced .. .
relative to the aircraft heading scale, i.e. left or right by an amount".
equivalent to the angle between the rotor axis and the N-S axis. At
small angular displacements of the outer ring axis, the errors are
small and diminish to zero when an aircraft returns to straight and
level flight after executing a manoeuvre.
·
There is one final point which should be considered and that is
the effect of the erection device whenever the angular relationship
between the gimbal rings is disturbed. At all times this device will
Figure 6. 15 Gimbal errors
at variou.s bank angles.
182
be attempting to re-erect the rotor into a new plane of rotation and
will cause false erection, the magnitude of which depends on how
long the erecting force is allowed to operate i.e. the duration of the
manoeuvre. The magnitude of the erecting force itself will depend
on the angle of the rotor to the erection device . Thus, we see that on
completion of a manoeuvre, it is possible to have an error in indication due to false erection, and that during a manoeuvre, an error
can be caused which is a combination of both gimballing effect and
false erection.
Questions
6.1
6.2
6.3
6.4
6.5
6.6
6.7
6.8
6.9
6.10
6.11
6.12
6.13
6.14
6.15
Define the following: (i) flux density , (ii) inverse square law,
(iii) reluctance, (iv) magnetic moment, (v) period of a magnet.
The torque acting on a magnet at angle 8 to a magnetic field is given
by (a) mH sin 8, (b) mH, (c) FM cos 8. Which of these statements is
correct?
What do you understand by the terms 'hard-iron' and 'soft-iron'
magnetism?
Define the following: (i) magnetic meridian, (ii) magnetic variation,
(iii) isogonal lines, (iv) agonal lines.
Draw diagrams to illustrate the relationship between the earth's
magnetic components and magnetic dip at the equator and at the
magnetic poles.
Define the unit in which the values of the magnetic components are
given.
(a) What is an 'aperiodic' compass?
(b) How are the effects you have described obtained in an aircraft
compass?
Explain how the effects of temperature change in an aircraft compass are compensated.
(a) Describe the magnet system of a typical aircraft compass.
(b) How is the effect of dip overcome?
(a) Define acceleration error and northerly turning error.
(b) Assuming an aircraft is flying in the southern hemisphere, what
errors in compass readings will be introduced when (i) the aircraft accelerates on an easterly heading, (ii) the aircraft turns
from a southerly heading towards East.
Explain briefly the principle of operation of a directional gyro.
Why is it necessary for the gyroscope assembly of a directional gyro
to be caged when setting a heading?
Explain how the rotor and inner gimbal ri·ng of a directional gyro
are erected to the level position.
How is earth rate error. controlled in a directional gyro?
(a) How are gimballing errors caused?
(b) Under what flight conditions is the gimbal system unaffected?
183
7 Remote-indicating
compasses
In their basic form remote-indicating compasses currently in use are
systems in which a magnetic detecting element monitors a gyroscopic
indicating element. This virtual combination of the functions of
both magnetic compass and directional gyroscope was a logical step
in the development of instrumentation for heading indication and led
to the wide-scale use of such systems as ·the distant reading compass
and the Magnesym compass in Allied military aircraft of World War
II. Although successfully contributing to the navigation of such aircraft, these systems were not entirely free of certain of the errors
associated with magnetic compasses and directional gyroscopes, and
furthermore there were certain practical difficulties associated with
the synchronizing methods adopted. In order, therefore, to reduce
all possible sources of error and to provide subsequent designs of
compass systems with self-synchronous properties, new techniques
had to be adopted. The most notable of these were: the changeover
from a permanent-magnet type of detector element to one utilizing
electromagnetic induction as part of an alternating-current
synchronous transmission system; the application of electronics; and
the application of improved gyroscopic elements and precession
control methods.
The manner in which the foregoing techniques are applied to
systems currently in use depends on the particular manufacturer,
and for the same reason the number of components comprising a
system may vary. However, the fundamental operating principles of
Figure 7. 1 Essential
components of a remoteindicating compass system.
OEVIATION
COMPENSATOR
AMPUFIER
FLUX DETECTOR
ELEMENT
I
I
I
: ,---- - -,
I
I
I SERVO I
- , SYSTEM r
!_ ____ _ _!
184
1
--
the main components shown in the block diagram of Fig 7 .1 remain
the same and are dealt with in this chapter.
Flux Detector
Elements
Figure 7.2 General construction of a flux detector element.
Unlike the detector element of the simple magnetic compass, the
element used in all remote-indicating compasses is of the fixed type
which detects the effect of the earth's m agnetic field as an electromagnetically induced voltage and monitors a heading indicator by
means of a variable secondary output voltage signal. In other words,
the detector acts as an alternating-current type of synchro transmitter and is therefore another special application of the transformer
principle (see page 230).
In general, the construction of the element is as shown in Fig 7.2.
It takes the form of a three-spoked metal wheel, slit through the rim
between the spokes so that they and their section of rim act as three
individual flux collectors.
SECONDARY
PICl\·OFF COILS
LAMINATED
COLLECTOR HORNS
SPOKES
EXCITER (PRIMARY) COIL
Around the hub of the wheel is a coil corresponding t o the
primary winding of a transformer, while coils around the spokes
correspond to 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 tracing
through the development of only one of them.
·
We will first take the case of a single-tum coil placed in a magnetic
field. The magnetic flux passing through the coil is a m aximum 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. Figure 7.3 (a) shows that
for a coil placed at an angle O to a field of strength H, the field can be
resolved into two components, one along the coil equal to H cos 0
185
Figure 7. 3 Single coil in a
magnetic field. (a) Components; (b) total flux.
I
I
I
I
'
H COS 8
(a)
(b)
and the other at right angles to the coil equal to H sin 8. This latter
component of the field 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 may be represented as at (b ).
If the coil were positioned in an aircraft so that it Jay in the horizontal plane with its axis fixed on or parallel to 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 a compass
system able to detect the earth's magnetic field without the use of a
permanent magnet. Unfortunately, such a simple system cannot
serve any practical purpose 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. However,
if 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 in Fig 7.4.
Each spoke consists of a top and bottom leg suitably insulated
186
Figure 7.4 Vertical section
of a detector-element spoke.
LAMINATED SPIDERS
OF SPOKE
EXCITER COil SUPPLIED
AT 23·5V. 400 HZ
from each other and shaped so as to enclose the central hub core
around which the primary coil is wound . The secondary coil is
wound around both legs.
The material from which the legs of the spokes are made is
Permalloy, an alloy especially chosen for its characteristic property
of being easily magnetized but losing almost all of its magnetism once
the external magnetizing force is removed. 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 of the same frequency as the current and 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 Hand the
cosine of the angle between Hand the axis of the detector.
If we consider first that the axis of the detector lies at right angles
to H, the static flux linked with the coils will of course be zero.
Thus, with an 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 7 .5.
The transition from primary coil flux to flux in the legs of the
detector is governed by the magnetic characteristics of the material,
such characteristics being determined from the magnetization or
B/H curve. In Fig 7.6 the curve for Permalloy is compared with
that for iron to illustrate how easily it may be magnetized . There are
several other points about Fig 7 .6 which 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:
l. 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 mag187
Figure 7.5 Total flux when
detector spoke is at right
angles to earth's field.
EARTH'S COMPONENT H
rnrnrrrnn
:: : :: : :: Ir :
I I
I
I I I
I I I
I I I
+
: : : : : : : : : \':
,,
' ' '
I
\\ I
;
t-f"++-+if'f11'f+-rl I I
II
1-i,...-,.-,fl...;.,I~;:
I
-j,':
I
I
EXCITER COIL VOLTAGE: : ; : : I I I
1+i \
FLUX IN BOTTOM LEG
I
'
X
~ t-----r----+--::Tl;';'.M;;:E
//:
I
I
; ~1 :
rt!!!!!!!!!!!
I 1 I
I
I
I I
I I
I
I
I
I
FLUX IN TOP LEG
:1D®ci
5
Figure 7.6 B/H curve and
hysteresis loops.
TOTAL FLUX IS Z!RO
2nd HALF-CYCLE
+B
IRON
I
-H
I
+H
I
I
I
/
I
/
-B
netized. Note also _that 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
•
188
The lagging of an effect behind its cause (from Greek husteresis, 'coming after').
as remanence or remanent flux density and it will be noted that for
iron this is very high (distance OR 1 ) thus making for good permanent
magnetic properties. For detector elements, a material having the
lowest possible remanence is required, and as the distance ORp
indicates, Permalloy meets this requirement admirably.
4 . Coercivity: this refers to the amount of negative magnetizing
force (coercive force) necessary to completely demagnetize a material
and is represented by the distances OCi and OCp. Coercivity and not
remanence determines the power of retaining magnetism.
In order to show the character of the flux waves produced in the
legs of the detector, a graphical representation is adopted which is
similar to that used for indicating electron tube characteristics. As
may be seen from Fig 7. 7 the waveshapes of the alternating primary
fields are drawn across the flux density axis B of the B/H 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 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 flux
detector lies at any angle other than a right angle to the horizontal
component Has indicated in Fig 7 .8 (a). The alternating flux due to
the primary coil changes the reluctance of the material, thus allowing
Figure Z ~ Flux wave
characteristics.
B
FLUX IN BOTIOM LEG
:rf-
1
I
H
-
SATURATION
-POINT
-
.
\
01-----+, - - - - - 1-TI-M-E
I
\
I
__J __ _
\
I
I
EXCITER COIL
VOLTAGE AND
MAGNETIZING
FORCE IN TOP
LEG
-~
I
\
/
_.,._., - -
-
-
SATURATION
POINT
FLUX IN TOP LEG
B
0
- . . _ I EXCITER COIL
VOLTAGE AND
MAGNETIZING
FORCE IN
BOTIOM LEG
TOTAL FLUX AND SECONDARY
PICK -OFF COIL OUTPUT IS ZEAD
189
EARTH"S COMPONENT H
Figure 7.8 Effect of earth's
component H. (a) Detector at
an angle to componentH;
(b) displacement of axis due
to static flux.
I
\
I
\
+
~
,
EXCITER COIL VOLTAGE
{a)
DISPLACEMENT Of AXIS.
DUE TO EARTH'S COMPONENT
___ L
(b)
the static flux due to component H to flow into and out of the spoke
in proportion to the reluctance changes. During those parts of the
primary flux cycle when the reluctance 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 ).
If we now apply a graphical representation similar to Fig 7. 7 and
include the static flux due to component H, the result will be as
shown in Fig 7.9. 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 detector 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
190
Figure 7.9 Total flux and
e.m.f. due to earth's component H.
t
B
r
TOTAL FLUX
LINKED
WITH
~~~~,.__.~~....._~...__.'-'--;...,_............."--_......,._.'-'-------,1....L..'-'-..c;....,.SECONOARY
H-
\-j
PIC\~~
I I
I I
. I
MAGNETIZING
FORCE IN
TOP LEG
I I
E.M.F.
INDUCED IN
SECONDARY
PICK-OFF
COIL
driving the material into saturation at the other end of the B/H 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 indicated also in Fig
7 .9. When the flux detector 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 with the secondary 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 flux detector 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
half-cycle of the primary voltage, the reluctance of the flux detector
material goes from minimum to maximurµ and back to minimu·m,
and in flowing through the material the static flux cuts the pick-off
coil twice. Therefore,.in each half-cycle of {1rimary 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 pick-off coil.
191
The supply for primary excitation has a frequency of 400Hz;
therefore the resultant e.m.f. induced in the secondary .Pick-?ff coil
has a frequency of 800Hz, as shown in Fig 7.9, and an amplitude
directly proportional to the earth's magnetic component in line
with the particular spoke of the detector element.
Having studied the operation of a single flu~ detector the
reasons for having three may now be examined a little more closely.
If we again refer to Fig 7 .3, and also bear in mind the fact that the
flux density is proportional to the cosine of the magnetic heading,
it will be appa·r ent that for one flux detector 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 a.c.
voltmeter to the detector, the same voltage reading would be
obtained on the instrument for both maximum values because the
instrument 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, hy employing
three simple flux detectors positioned at angles of 120° to one
another, the paths taken by the earth's field through them, and for
360° rotation, will be as shown in Fig 7 .10. Thus, varying magnitudes of flux and induced voltage can be obtained and related to all
headings of the detector element without ambiguity of directiqnal
reference. The resultant voltage of the three detectors at any time
can be represented by a single vector which is parallel to the earth's
component H .
:
Figure 7.10 Path ofea.rth's
field through a detector and
signals induced in pick-off coils.
H
EARTH'S COMPONENT
' ',
I
'
I
I
I
I
\
'''
t H 't ;
\
000°
I
I
090°
I
I :
I
I
I
'
I
I\
.
I
\
I
1 1
I
ttH ' .. t
,so•
270°
Figure 7. I I is a sectional view of a typical practical detector
element. The spokes and coil assemblies are pendulously suspended
from a universal joint which allows a limited amount of freedom in
pitch and roll, to enable the element to s·e nse the maximum effect of
the earth's component H. It has no freedom in azimuth. The case in
which the element is mounted is hermetically sealed and partially fille
with fluid to damp out excessive oscillations of the element. The
complete unit is secured to the aircraft structure (in a wing or vertical
192
Figure 7.11 Typical flux
detector element. 1 Mounting
flange (ring seal assembly),
2 contact assembly, 3 terminal,
4 cover, 5 pivot, 6 bowl,
7 pendulous weight, 8 primary
(excitation) coil, 9 spider leg,
10 secondary coil, 11 collector
horns, 12 pivot.
3
stabilizer tip) by means of a mounting flange which has three slots for
mounting screws. One of the slots is calibrated a limited number of
degrees on each side of a zero position corresponding to an aircraft
installation datum for adjustment of coefficient A (see Chapter 8).
Provision is made at the top cover .of the casing for electrical connections and attachment of a deviation qompensating device.
Gyroscope And
Indicator Monitoring
Monitoring refers to the process of reproducing the directional
references establish ed by the flux detector element as quantitative
indications on the heading indicator. The principle of moni toring
is basically the same for all types of compass systems and may be
understood by reference to Fig 7 . 12.
When the fl.ux detector is positioned steady on· one h eading, say
000°, then a maximum voltage signal wiJI be induced in the pick-off
coil A, while coils Band C will have signals of half the amplitude
and opposing phase induced in them. These signals are fed to the
corresponding tappings of the synchro receiver stator, and the fluxes
produced combine to establish a resultant field across the centre of
the stator. This resultan t is in exact alignment with the resultant of
the earth 's field passing through the detector. If, as sh own in
Fig 7.12 (a), the synchro receive r rotor is at right angles to the resultan
flux then no vol tage can be induced in the windings. In this position,
the synchro is at 'null' and the directional gyro will also be aligned
with the earth's resultant field vector and so the heading indicating
element will indicate 000°. Now, let us consider what happens when
the flux det ector element turns through 90° say; the disposition of the
detector pick-off coils will be as shown at (b). No sign al voltage will
be induced in coil A, but coils B and C have increased voltage signals
induced in them, the signal in C being opposite in phase to what it
was in the ' null' position. The resultant flux across the receiver
193
-
Figure 7. 12 Directional gyroscope monitoring. (a) Heading
= 000°; (b) heading= 090°.
R
(a)
RESULTANT OF EARTH'S
FIELD COMPONENT
THROUGH DETECTOR
RESULTANT OF FIELD DUE
TO INDUCED VOLT AGE SIGNA~S
ICOIL A
V,;
PD
COIL C
-
- - -
EARTH'S FIELD
- - - INDUCED VOLTAGE SIGNALS
(b)
synchro stator will have rotated through 90°, and assuming for a
moment that the gyroscope and synchro are still at their original
position, the flux will now be in line with the synchro rotor and will
therefore induce maximum voltage in the rotor. This error voltage
signal is fed to a slaving amplifier in which its phase is detected, and
after amplification is fed to a slaving torque motor, which precesses
the gyroscope and synchro rotor in the appropriate direction until
the rotor reaches the ' null' position, once again at right angles to the
resultant field. In other words, the conditions will be the same as
shown in diagram (a) but on the new heading of 090°. With the aid
of Figs 7. IO and 7.12 the re:ider should have no difficulty in working
out the directions ofvoltage signals and resultant fluxes for other
positions of the detector element.
In practice of course, the rotation of the field in the receiver
synchro, and slaving of the gyroscope, occur simultaneously with
the turning of the flux detector so that synchronism between
detector and gyroscope is continuously maintained.
The synchronous transmission link between the three principal
components of a master gyroscope system is shown in Fig 7 .13.
194
-------,
INDICATOR
SLAVING AMPLIFIER
--------,
·~ _ _ _ _ _J
GNAL FLOW
:>+ DIRECTIONAL REFERENCE
-
HEADING ERROR
1
::>-I>
ANO
~,?e~lfs~~~
I
I
SERVO AMPLIFIE~
-+SERVO
--!>LOOP ERROR
...-!> SERVO DRIVE
*
FEEDBACK DAMPING
i_ _
-
Heading dot•
signals to other
syst1m1 e .g .
automatic fl ight
SET HEADING
KNOB
Figure 7. 13 Master gyroscope and indicator
monitoring.
SYNCHRONIZING
KNOB
The fundamental principle of monitoring already described also
applies to this system, but because the indicator is a separate unit, it
is also necessary to incorporate additional synchros in the wstem, to
fonn a servo loop. The rotor of the loop transmitter synchro (CX) is
rotated whenever the gyroscope is precessed, or slaved, to the
directional reference, and the signals thereby induced in the CX
stator are applied to that of the CT synchro in the indicator.
During slaving, the rotors of both synchros will be misaligned, and a
servo loop error voltage is therefore induced in the CT synchro rotor
and then applied to a servo amplifier. After amplification, the voltage
is applied to a servomotor which is mechanically coupled to the CT
rotor and to the heading dial of the indicator. Thus, both the rotor
and dial are rotated, the latter indicating the direction of the heading
change taking place. On cessation of the furn , the rotor reaches a
'null' position, and as there will no longer be an input to the servo
amplifier, the servomotor ceases to rotate and the indicator dial
indicates the new heading. The servomotor also drives a tachogenerator which supplies feedback signals to the amplifier to damp out
any oscillations of the servo system.
195
Directional Gyroscope
Elements
Depending on the type of compass system and application of
synchronized heading information, the directional gyroscope element
may be contained within a panel-mounted indicator, or it may form
an independent master unit located at a remote point and transmitting
information to a slave indicator. Systems adopting the master gyroscope reference technique are, however, the most commonly used,
because in serving as a centralized heading data source, they provide
for more efficient transmission of the data, particularly to flight
director-systems (see Chapter 15) and automatic flight control
systems with which they are now closely integrated .
An example of a master gyroscope unit is shown in Fig 7 .14. The
gyro rotor is spherical in shape and supported in the inner gimbal
ring which is fitted with hemispherical covers totally enclosing the
gyro motor. This assembly is filled with helium and hermetically
sealed. The rotors of the main slave heading synchro and an additional data synchro are mounted at the top of the outer·gimbal ring:_
The stator of a levelling torque motor is attached to the bottom of
the outer gimbal ring and its rotor is attached to the gyro casing.
The motor is controlled by a liquid level switch secured to the inner
gimbal ring shaft. A heading dial is also fixed to the outer gimbal
ring and is referenced against a fixed vernier scale. Both scales are
viewed through an inspection window in the gyro outer casing, their
purpose being to provide an arbitrary heading indication when
testing the system in an aircraft. The complete assembly is mounted
on anti-vibration mountings contained in a base whi'ch provides for
attachment at the required location, and also for the connection of
the necessary electrical circuits.
A 'spindown' brake system is also incorporated to prevent the
wobbling effect of the gyroscope (called nutation) by holding the
outer gimbal ring steady for a short period of time after power is
applied. The brake is also applied approximately 30 seconds after
the power supply has been switched off to prevent spinning of the
gyroscope about its vertical axis during rundown .
Heading Indicator
An example of the dial presentation of an indicator used with a
master directional gyroscope element is shown in Fig 7 .15. In
addition to displaying magnetic heading data, it also displays the
magnetic bearing of an aircraft with respect to the ground stations
of the radio navigation systems ADF (Automatic Direction Finding)
and VOR (Very high-frequency Omnidirectiqnal Range). F9r this
reason, the indicator is generally referred to as a Radio Magnetic
Indicator (RMI).
The bearing indications are provided by two concentrically
mounted pointers, one called a 'double-bar' pointer and the other
a 'single-bar' pointer. Both are referenced against the main compass
196
Figure 7.14 Example of master
directional gyro element.
Figure 7.15 Heading
HEADING "BUG"
indicator.
ANNUNCIATOR
DOUBLE -BAR
POINTER
SELECTOR KNOB
SINGLE·BAA
POINTER
SYNCHRONIZING
KNOB
COMPASS CARD
SET HE-"JP~ii
LUBBER LINE
~
EACH POINTER RESPONDS
TO SIGNALS FROM A VOR
STATION.
EACH POINTER RESPONDS
TO SIGNALS FROM AN ADF
STATION.
SINGLE·BAR POINTER RESPONDS
TO AN ADF STATION'S SIGNALS:
OOUBLE·BAR POINTER TO THOSE
OF A VOR STATION.
SINGLE -BAR POINTER RESPONDS
TD SIGNALS FROM A VOA STATION:
DOUBLE-BAR POINTER FROM
AN ADF ST ATION.
~
card , and are positioned by synchros connected to the ADF and VOR
receivers on board the aircraft. The display is controlled by manual
adjustment of a selector knob to the positions shown in the diagram.
The signals transmitted to the synchro systems are such that the
pointers always point to the stations from which the signals are
received. This may be seen from Fig 7 .16 which is a representation
of how the position of an aircraft may be determined from the three
sources of navigational information.
In order that a pilot may set a heading on which it is desired for
the aircraft to be flown , a 'set heading' knob is provided . It is mechanically coupled to a heading 'bug' so that when the knob is
rotated, the 'bug' is also rotated with respec t to the compass card.
The aircraft is then turned until the desired new heading is registered
against the 'bug'. For turning under automatically-controlled flight
conditions, rotation of the set heading knob also positions the rotor
of a CX synchro (see Fig 7.13) which then supplies turn command
signals to the autopilot system.
198
N
Figure 7.16 Example of
RMI indications.
ADf
Gyroscope Levelling
Figure 7.17 Levelling system.
In addition to the use of efficient synchronous transmission systems,
it is also essential to employ a system which will maintain the spin
axis of the gyroscope in a horizontal position at all times. This is a
requirement for all directional gyroscopes to overcome random drift
(see also page 178), of the gyro rotor housing due to bearing friction.
The levelling or erection systems utilized in the gyroscopes of
remote-indicating compasses are similar in design and operating prin-
A.C. SUPPlY TO FIXED
FIELD WINDING
zI
I
z,
A.C. SUPPi.Y TO FIXED FIELD
SECTION OF CONTROL WINDING
A.C. SUPPLY TO
LIQUID LEVEL SWITCH
199
Figure 7.18 Rotorace bearing
system.
ciple and consist of a switch and a torque motor. The switch is
generally of the liquid-level type mounted on the gyro rotor housing
so as to move with it. The torque motor is a two-phase induction
motor located so that its stator is attached to the outer gimbal ring,
and its rotor attached to the gyro casing. Fig 7 .17 schematically
illustrates the arrangement of the system. The switch has one of its
electrodes connected to a low alternating voltage source and the other
connected to one end of the torque motor control winding. It will be
noted from the diagram that the control winding is formed in two
parts; one is continuously energized, and the other is energized only
through the liquid-level switch. The fields set up by the currents in
the two windings produce torques, but since the fields are in opposition to each other precession to the level position is due to a resultant
torque.
Figure 7 .18 illustrates a method used in a current type of master
gyro unit for counteracting drift of the gimbal system due to bearingfriction torques. The inner gimbal ring is supported within .the outer
gimbal ring by two special bearings, known as Rotorace bearings.
Each bearing consists of an inner race, a middle race, an outer race,
and two sets of ball bearings; the middle race is constantly rotated by
a drive motor working through a gear train.
Rotation of the middle race uniformly distributes around the axis
of rotation any frictional torques present in the bearing. The gearing
to the middle races of the two bearings is such that the races are
rotated in opposite directions ; thus the frictional torques of the two
REVERSING SWITCH
GYRO ROTOR UNIT ASSEMBL V
!INNER GIMBALI
\
ROTORACE BEARING
200
bearings oppose each other. Drift that may be introduced by mismatching of the bearings is cancelled·by periodically changing the
direction of rotation of the middle races. This is achieved by reversing
the rotation of the drive motor through a switch which is operated by
a cam. The cam is, in turn, driven by a worm gear from the drive shaft
between the two bearings. Condensers connected across th e switch
suppress any tendency for contact arcing to take place.
Modes of Operating
Compass Systems
All compass systems provide for the selection of two categories or
modes of operation: slaved, in which the gyroscope is monitored by
the flux detector element, and f.&ee gyro , or D. G., in which the gyro
is isolated from the detector to function as a straightforward directional gyroscope. The latter operating mode. is selected whenever
malfunctioning of directional reference-signal circuits occur or when
an aircraft is being flown in.'-latijudes at which the horizontal component of the earth's magnetic field is an unreliable reference.
Synchronizing Indicators
The function of synchronizing indicators, or annunciators as they
are sometimes called, is to indicate whether or not the gyroscope of
a compass system is synchronized with the directional reference
sensed by the flux detector element. The indicators may form an
integral part of the main heading indicator (see Fig 7.15) or they may
be a separate unit mounted on the main instrument panel. They are
actuated by the monitoring signals from the flux detector and are
connected into the gyro slaving circuit.
..
A typical integral annunciator is shown in Fig 7 . 19 and serves to
illustrate the function of indicators generally. It consists of a small
flag, marked with a dot and a cross, which is visible through a small
Figure 7.19 Typical
integral annunciator
COILS -'ND POLE PIECES
201
window in one corner of the heading indicator bezel. The flag is
carried at one end of a pivoted shaft. A small permanent magnet is
mounted at the other end of the shaft and positioned adj acent to two
soft-iron cored coils connected in series with the precession circuit.
When the gyroscope is out of synchronism , the monitoring and
precession signals flow through the annunciator coils and induce a
magnetic field in them. The field reacts with the permanent magnet
causing it to swing the shaft to one side so that either the dot or the
cross on th e flag will appear in the annunciator window. The particular indication shown depends upon the direction in which the
gyroscope is precessing and will remain until synchronism is regained.
Under sy nchronized conditions the annunciator window should
ideally be clear of any image. During flight, however, the flag
oscillates slowly between a dot and a cross indication. This is due
to pendulosity effects on the flux detector element and does in fact
serve as a useful indication that monitoring is taking pla.ce.
Manual Synchronization
When electrical power is initially applied to a compass system operating in the 'slaved' mode, the gyroscope may be out of alignment by
a large amount. It will, of course, start to synchronize, but as the
gyroscopes normally have low rates of precession ( l O to 2° per
minute), some considerable time may elapse before synchronization
is effected. In order to speed up the process, systems always
incorporate a fast synchronizing facility which can be manually
selected.
The principle of a method applied to a system utilizing a master
gyro may be understood from Fig 7.13. The indicator employs a
synchronizing knob, marked with a dot and a cross and coupled to
the stator of the servo CT synchro. When the knob is pushed in and
rotated in the direction indicated by the annunciator, the rotation
of the stator of the CT synchro induces an error voltage signal in
the rotor. This is fed to the servo amplifier and motor which drives
the slave heading synchro rotor and gyroscope, via the slaving
amplifier, into synchronism with the flux detector. At the same
time, the CT synchro rotor is driven to the 'null' position and all
error signals are removed.
Indication of Magnetic
Heading in Polar
Regions
202
With the introduction of regular flight operations in the vicinity of
and over the polar regions by military aircraft and many civil aircraft
operators, the necessity for developing new navigational methods
and equipment became of paramount importance. The main difficulties of navigat ion in these regions are associated with the following problems: the convergence of the meridians and isogonals, the
long twilight periods limiting the use of astro navigation, magnetic
stonns and radio blackouts limiting the use of radio navigational aids,
and limitations in the directional references established by magnetic
compasses and some remote-indicating compass systems.
It will be recalled from · Chapter 6 that the magnetic North cannot
be defined as an exact point, but may be described as the area within
which the horizontal component Hof the earth's magnetic field is
zero, the entire field being vertical. It is therefore clear that the use
of a direct-reading compass or remote-indicating compass in this area
cannot provide the required accurate directional reference. Another
circumstance which makes magnetic steering impracticable in the
polar regions is the fact that the deflection of the magnet system
from magnetic North is very large and varies rapidly with distance.
The gyroscopes of present-day compass systems are designed to be
substantially free from drift due to gimbal-ring bearing friction and
unbalance, and so their use in the 'free gyro' mode of operation
offers an obvious solution to the problem. However, a gyroscope still
has an earth-rate error which varies with the latitude in which it
operates and is greatest in polar regions. To obtain accurate heading
inforr11ation, therefore, the errors must be eliminated either by
calculating the corrections required and applying them directly to
indicated headings, or by incorporating an automatic correcting
device in the gyroscope. In compass systems designed for polar
flying the latter method is used and takes the form of a controlled
precession circuit.
A typical circuit consists essentially of three main parts: (i) a
switch for selecting the hemisphere in which a flight is being made,
i.e . North or South, (ii) a selector knob, potentiometer and dial
calibrated in degrees of latitude, thus providing for the setting of the
latitude in which the aircraft is being flown, and (iii) two small d.c.
electromagnets built into the outer ring, one on either side of the
inner gimbal ring. The controlling circuit is illustrated in Fig 7 .20.
Wfiln the hemisphere selector switch is set to the appropriate
position it allows a high direct voltage from a special power unit to
flow to the coil of the electromagnet selected. The current also
passes through the potentiometer of the latitude compensator so
that it can be adjusted to give the control required at the latitude
selected. As the gyroscope drifts, the rotor turns through the magnetic field of the energized coil , and eddy currents are induced in
the rotor with the result that a torque is applied to the inner gimbal
ring, which precesses the complete gimbal system about its vertical
axis at the same rate, but in opposition to its drift. The precession
is counter-clockwise when the nothern-hemisphere magnet is
energized, and clockwise when the southern hemisphere magnet is
energized. A voltage-stabilizing circuit ensures that the current
through an energized coil is unaffected by :,ariations in the direct
voltage from the power supply unit.
203
Figure 7.20 Latitude
controlling circuit.
TO SLAVING
AMPLIFIER
INNER RING
l
r---
~LAVE IMA~ - - - ]
! ~
-~--.cro
~=------,,___-------;-.-,
N
s
q
D G.
- ' -. 115V
I
-v--v--,..~~i!
•I
;?. I
_ _ _ _J'
~
VOLTAGE
~sT_A_B1_uz_eR_
115V
Deviation
Compensation
Compensation for deviations resulting from the effects ofhard and
soft-iron magnetism is accomplished by applying either of the
methods described in Chapter 8 (see page 221 ).
Errors in Compass
Systems
In addition to the static errors resulting· from hard-iron and soft-iron
magnetic effects (see Chapter 8) remote-indicating compass systems
are also subject to two further principal errors; one is static and
referred to as transmission error, while the other is dynamic and
referred to as Coriolis error.
Transmission Error
This can be caused in several ways; for example, if the impedance
of the synchro windings connected to the flux detector element are
not exactly balanced, then the output signal will be affected in a
manner that will cause an error which varies through two positive
and two negative peaks in a 360° tum of the aircraft. Other sources
of transmission error are voltage unbalance in the flux detector
element itself, and synchro stators being elliptical. Transmission
error is normally compensated during manufacture by inserting
. impedances between the flux detector element and its associated
heading_ data synchro.
Coriolis Error
This error is caused by a deflecting force exerted by the rotation of
the earth upon any object in motion. In the case of an aircraft in
~ight therefore, it will be subjected to what is termed Coriolis
acceleration in following its curved path from one point on the
204
earth's surface to another. This acceleration causes the pendulous
element of the flux detector to be tilted from the horizontal such
that it will intercept a portion of the earth's field component Z.
This component, as detected in the flux detector, will rotate with
the aircraft during a turn, and will generate an error which is maximum on North and South headings. The extent of the error is
governed by the ground speed and track of the aircraft, latitude,
and magnetic dip (see page 397).
A usual method of correcting the error is to position a variable
resistor by means of a 'latitude set' knob in a controller, so as to
divide an input from a preset ground speed variable resistor. The
resultant d.c. signal is fed to the flux detector where it produces a
field that cancels the position of the Z component sensed as a
result of the tilt.
Questions
7.1
7.2
7.3
7.4
7.5
7.6
7.7
7.8
7.9
7.10
7 .11
7.12
7.13
Sketch and describe the construction of a heading detector unit
suitable for use in a remote-indicating compass system.
Sketch the hysteresis loops of soft-iron and Mumetal specimens and
explain the shape of the loops ..
By means of diagrams describe how fluxes and voltages are induced
in a detector element.
Explain how .unambiguous directional references are obtained.
What effect does the alternating current supplied to the exciter coil
of a detector element have. on the earth's field passing through the
element?
Explain the effects you consider the pendulous mounting of a
detector element will have on indicated headings during turns and
accelera.tions.
Describe the general construction of a remote-indicating compass
directional gyro element.
Explain how a directional gyro element is monitored by the flux
detector.
Describe how the principle of a servo synchro loop is applied to a
compass system .
(a) On which components of a directional gyro element are the rotor
and stator of a levelling torque-motor fitted?
(b) Explain the operation of a typical levelling system.
(a) Under what conditions is the 'D.G.' mode of operation selected?
(b) Describe how indications of synchronism between the directional
gyro and flux detector elements are obtained.
Describe how the 'earth rate error' is automatically corrected in a
polar compass system.
What do you understand by the terms 'transmission error' and
'Coriolis error'?
205
8 Aircraft magnetism and
its effects on compasses
A fact which has always been a challenge to the designers of aircraft
compasses is that all aircraft are themselves in possession of magnetism
in varying amounts. Such magnetism is, of course, a potential source
of error in the indications of compasses installed in any type of aircraft and is unavoidable. However, it can be analysed and, for any aircraft , can be 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 properly called, resulting
from such components.
The two types of aircraft magnetism can be divided in the same way
that magnetic materials are classified according to their ability to be
magnetized, namely hard-iron and soft-iron (see also page 164).
Hard-iron magnetism is of a permanent nature and is caused, for
example, by the presence of iron or steel parts used in the aircraft's
structure, in power plants and other equipment, the earth's field
'building' itself into ferrous parts of the structure during construction
and whilst lying for Jong periods on one heading.
Soft-iron magnetism is of a temporary nature and is caused by
metallic parts of the aircraft which are magnetically soft becoming
magnetized due to induction by the earth's magnetic field. The effect
of this type of magnetism is dependent on the heading and attitude
of the aircraft and its geographical position.
There is also a third type of magnetism due to the sub-permanent
magnetism of the 'intermediate' iron, which can be retained for varying periods. Such magnetism depends, not only on the heading,
attitude and geographical position of an aircraft, but also on the nature
of its previous motion, vibrations, lightning strikes and other external
agencies.
The various components which cause deviation are indicated by
letters, those for permanent, hard-iron, magnetism being capitals, and
those for induced, soft-iron, magnetism being small letters. It is also
important to note at this juncture that positive d eviations are termed
easterly and negative deviations westerly.
206
Effects of Magnetic
ComponP.nts o n
Compasses
Components of Hard-Iron Magnetism
The total effect of this type of magnetism at the compass position
may be considered as having originated from bar magnets lying
longitudinally, laterally and vertically. These are shown schematically in Fig 8.1. In order to analyse their effects they are respecttively denoted as components P, Q and R . The strength of these
components will not vary with heading or change of latitud e but
may vary with time due to a weakening of the magnetism in the aircraft. Referring again to Fig 8. I , we should note particularly when
the blue poles of the magnets are forward , to starboard, and
beneath the compass respectively, the components are positive,
and when the poles are in the opposite directions the components
are negative. The fields always pass from the blue pole to the red
pole.
...,
Figure 8. J Components of
hard-iron magnetism.
<
g,a:: :::1
Cl)
w·
>I
{ID RED
-R
POLES
0BLUE POLES
When an aircraft is heading North, the equivalent magnet due to
component P will, together with the compass, be in alignment with
the aircraft's longitudinal axis and the earth's component H, and so
will cause no deviation in compass heading. If we now assume that
the aircraft is to be turned through 360°, then as it commences the
tum, the compass magnet system will (ignoring pivot friction , liquid
swirl, etc.) remain attracted to the earth's component H. Component
P, however, will align itself in the resultant position and will cause a
deviatio_n in the compass reading of magnetic North , making it read
so many degre~s East or West of North, depending on the polarity of
the component. This deviation will increase during the turn and will
be a maximum at East, then it will decrease to zero at South,
increasing once again to a maximum but in the opposite direction at
West, and then decreasing to zero when turned on to North once
207
Figure 8.2 Deviation cuIVes
due to components P and cZ.
+P
/
+
-
.......
II
I
U)
z
Q
·
,·-
\
·-
·-.
i \ ~ -.:.?_ )3> -Ja···-2i?...~'.~ ..
0
\
-c1.
\
\
HEAOtNGIN
~DEGREES
;3:is ...-
so
I
' ,_,,,, I
-P
again. A plot of the deviations caused by a positive and a negative
component P on all headings results in the sine curves shown in Fig
8.2.
Component Q produces similar effects, but since it acts along the
lateral axis, deviation is a maximum on North and South headings
and zero on East and West headings. Plotting these deviations for
positive and negative components results in cosine curves shown in
Fig 8.3; thus deviations due to Q vary as the cosine of the aircraft's
heading.
Figure 8.3 Deviation cuIVes
due to components Q and /Z.
Component R acts in the vertical direction, and when the aircraft
is in its normal level flight attitude, it has no effect on the compass
magnet system. If, however, the aircraft flies with either its longitudinal axis or lateral axis in positions other than horizontal, then
component R will be out of the vertical and will have a horizontal
component affecting the compass magnet system.
Figure 8.4 (a) illustrates the horizontal components of
component R when an aircraft is in the nose-up and nose-down
208
Figure 8.4 Component R and
effects. (a) Effects in nose-up
and nose-down attitude;
(b) effects in banked attitudes.
'-
-R1
~
COMPONENT
DUE TO+R
I
•
,
'
I
(a)
PORT
COMPONENT
DUE TO -R
(b)
attitudes. When the aircraft is in the nose-up attitude, the horizontal
component of a positive component R acts in a forward direction,
i.e. in the same sense as a positive component P, and when the aircraft is in the nose-down attitude the horizontal component acts rearwards in the same sense as a negative component P. Thus, in either
of these attitudes R will cause maximum deviation on East and West
headings.
The actual amount of deviation depends on the value of R and the
angle between the aircraft's longitudinal axis and the horizontal. The
resultant curves of deviations will be sine curves similar to those
produced by component P.
If component R is negative, as may also be seen from Fig 8.4 (a),
the horizontal components will act in the opposite directions.
When an aircraft is in a banked attitude as shown at (b ), component R will again have a horizontal component with an effect
similar to the lateral component Q, causing maximum deviations on North and South. A positive component R produces an
effect similar to a negative component Q when the aircraft is banked
starboard wing down. When banked port wing down the effect is
similar to a positive component Q.
If component R is negative the horizontal components will have
the opposite effects. As before, the actual amounts of deviation
depend on the magnitude of R and the angle between the aircraft's
appropriate axis and the horizontal.
A question often raised in connection with component R is 'Are
the effects really serious enough to bother about?' As stated earlier,
they depend on the magnitude of the component and the angles
through which the aircraft's pitch and bank attitude is changed.
Normally, the angle of climb and descent of most aircraft is reason209
ably shallow, and so deviations are usually small. In some aircraft
with a large component R a large effect on deviation is experienced
in the nose-up attitude. During turns an aircraft is banked, but here
again angles and deviations are usually small. If direct-reading
compasses are used in an aircraft they suffer far more from errors
due to turns so that any additional errors resulting from component
R effects are of little or no practical interest. The arrangement of
detector elements of remote-indicating compasses is such that turn
errors are eliminated and component R is negligible.
Components of Soft-Iron Magnetism .
The soft-iron magnetism which is effective at the compass position
inay be considered as originating from a piece of soft-iron in which
magnetism has been induced by the earth's magnetic field. This field
has two main components Hand Z, but in order to analyse soft-iron
magnetism, it is necessary to resolve H into two additional horizontal
components X and Y and to relate them, and the vertical component
Z, to the three principal axes of an aircraft. This relationship and the
polarities are given in Table 8.1 ..
The polarities and strengths of components X and Y change with
changes in aircraft heading because the aircraft turns relative to the
Figure 8.5 Changes of X. Y,
and Z components with aircraft heading changes.
(X)()•
t
+X =H
Y= O
-Y
+Y = H
-Y = H
X =0
X=O
225°
- X=H
Y=O
180°
210
135°
Table 8.1
Earth's
magnetic
components
Aircraft
axes to
which related
Polarity
H{;
Longitudinal
Lateral
Vertical
+ Forward-Aft ·
+ Starboard - Port
+ Down-Up
z
fixed direction of component H. The changes occurring through the
cardinal and quadrantal points are indicated in Fig 8.5.
Component Z acts vertically through the compass and therefore
does not affect the directional properties of the compass magnet
system.
If an aircraft is moved to a different geographical position, then
because of the change in earth's field strength and direction, all three
components will change. A change in the sign of component Z will
only occur with a change in magnetic hemisphere when the vertical
direction of the earth's field is reversed.
Each of the three earth's field components produces three softiron components which are visualized as resulting from induction
in a number of soft-iron rods disposed along the three axes of the
aircraft, and around the compass position, in such a way that their
combined effect is the same as the actual soft-iron influencing the
directional properties of the compass. There are therefore nine rods
of soft-iron magnetism which are indicated conventionally by the
letters a, b, c, d, e, [, g, h and k, and when related to the field components X, Y and Z, their soft-iron components are designated aX,
b Y, cZ; d X, e Y, fZ ; and g X, h Y and kZ. Of the nine rods there are
six which will always affect the directional properties of a compass
since the components of their induced fields always lie horizontal:
three longitudinal due to rods a, b and c , and three lateral due to rods
d, e and f These are shown in pictorial and tabular form in Fig 8.6.
In passing we may note that slight deviations could be produced by
the vertically induced fields of rods g, h and k in a manner similar to
those resulting from component R of hard-iron magrretism.
There are two main points which should be noted with reference
to Fig 8.6. Firstly, each rod has alternative positions, designated
positive and negative. The reason for this is because each component
of the aircraft's magnetism may act in one· of two reciprocal
directions. The polarity designation is dependent only on the
position of the rods relative to the compass, and is not affected by
any polarity which the rods may acquire as a result of being
magnetized by the earth's field .
The second point to note is that with the exception of the - a and
211
Figure 8.6 Rods affecting
ROOS ANO l'OUIRl1Y
directional properties of a
compass.
+a
LONGITUOINAL
OMPONENT Of
INOUCEO FIELO
EARTH"S
MAGNETIC
COMPONENT
-a
ROO
COMPONENTS
-b
, rV
'
'
if ~ V1
t
y
X
+d
LATERAL
,_.OMPONENT OF
INDUCEO FIELO
+b
-d
+e
aX and dX
FACTORS C1N
WHICH
POlARITIES
OEPENO
AIRCRAFT HEADING
bY and eY
-e
+c
~
-c
~
~
z
t
+f
~
-f
~
cZ. and fZ
MAGNETx:
HEMISPHERE
- e rods, there are two positions relative to the compass in which
each rod is positive and two in which each is negative; thus the fields
pass from the poles nearer the compass. We often refer to these
fields as being due to asymmetric horizontal soft-iron.
The deviation curves due to rod components aX, ·e Y, b Y and dX
are illustrated in Fig 8.7. Rod components cZ and fZ have the same
effects as hard-iron components P and Q respectively. A further
point to note in connection with cZ and fZ is that their polarities
and direction depend on whether the aircraft is in the Northern or
Southern hemisphere.
Total Magnetic Effects
The total magnetic effects of aX and e Y are quadranta1 and may be
found by summating the individual curves of deviation shown in
Fig 8.7. The combining of the curves of deviation due to components b Y and dX produces the total effects shown in Fig 8.8, and
from this it will be apparent that, in addition ·to producing deviation
which varies as the cosine of twice the heading, components b Y and
dX may also produce constant deviation or a combination of both.
The total magnetic effect at a compass position is the sum of the
earth's field components X, Y and Z, hard-iron components P, Q and
R, and the respective soft-iron components acting along the three
axes of the aircraft. The effects are shown schematically in Fig 8.9,
together with the expressions for calculation of the combined forces
represented as X', Y' and Z' . The signs used in the expressions
212
Figure 8. 7 Deviation curves
due to rod components.
(a) ComponentsaX and eY;
(b) component b Y; (c) compo·
+
or
mentdX
-eY
I
,, __ , ~./ 90
-aX or +eY
(a)
+
-bY
, , ,. ... - ........
I
I
I
lo
I
,I
I
I
'
'
'
,
'
-
/
--r--
,
,,--,
- - - -
''
\- , - -
/
\
'
,I
I
HEADING IN
360 DEGREES
90
0
+ bY
(b)
+
z
~ 0 + --.----=~~--,----,.~
45; ·
-;,
- dX
90
'\35
180
--,---'.""'tc-,--,----,225/ 270
'
r
"·...._,/
'y~~
HEADING IN
DEGREES
\_
·,
213
Figure 8.8 Total effects of
components by·and dX.
(a) C6mponents of same sign
and equal value; (b) compon·
ents of same sign and unequal
value; (c) components of
opposite sign and equal value;
(d) components of opposite
sign and unequal value.
RESULTANT
DEVIATION
'
/f/
(8)
+
•
RESULTANT
/v--
I
/
fo
I
-, ~ - /
\
CONSTANT
DEVIATION
I
,I
I
(b)
+
+ dX
~
;o
"'
'
~ RESULTANT
~'
NSTANT DEVIATION
'/
I
I
/
,.
- bY
45
5
''
180
360H~:E~N
(c)
+
45
90
135
180
{d)
214
225
315
360
H~~t
Figure 8.9 Total magnetic
effect a~ CO!llpass position.
·1
II i-
TO,AL
x' = X + p
+
ax + bY +
TOTAL
cZ
z· = 2
TOTAL
v· = y +
Q + dX + eY + fZ
+ R + gX + hY + kZ
indicate that the sums of the quantities are algebraic.
In the foregoing explanation of aircraft magnetism, its components and effects, no direct reference has been made to the detector
elements of remote-indicating compasses. It should not be construed
from this that such elements are entirely immune from extraneous
magnetic fields. AB compasses, regardless of their method 0f
detecting the earth's component H, must suffer from the ultimate
effect which is deviation. The only difference is the mariner in
which the aircraft's magnetic components affect the detecting
elements and cause deviation.
TI1e detector element of a direct-reading compass, as we already
know, is of the freely suspended permanent-magnet type, and so
deviation is caused by direct deflection of the element from
magnetic North. This, of course, is not true of a remote-indicating
compass detector element, so that when an aircraft turns the
detector element turns with it and its associated hard-iron and softiron magnetic components. How then can deviation be produced?
The answer lies in the fact that the material of the detector
element has a high permeability and is therefore very receptive to
magnetic flux. It will be recalled that accurate heading indication
is dependent on the displacement of the H axis, about which the
magnetizing force alternates, by the earth's field component.
Since the material is easily magnetized by this component, then it is
just as easy for it to be magnetized by components of aircraft
magnetism. Consequently , the H axis is displaced to a false datum
in either the positive or negative direction. This, in tum, changes
215
the total flux linking the secondary pick-off coils of the detector,
thus causing an error in the induced voltage. The current resulting
from this voltage produces a corresponding error signal in the heading
or slaving synchro, and so the directional gyroscope and heading card
are deviated from the correct heading.
Deviation Coefficients
Before steps can be taken to minimize the deviations caused by hardiron and soft-iron components of aircraft magnetism, their values on
t:ach heading must be obtained and quantitatively analysed into
coefficients of deviation. There are five coefficients named A, B, C,
D and E, termed positive or negative as the case may be, and
expressed in degrees.
The relationship between aircraft magnetism and the coefficients
is shown in Fig 8 .10.
Figure 8.10 Relationship
between aircraft magnetism
and deviation coefficients.
AIRCRAFT MAGNETISM
IMSAUGNMEITT
Of COMPASS
SOFT IRON
VERTICAl.
HORIZONTAL
COMPONENT Q
HORIZONTAl.
COMPONENT p
_ _ _'_LA_TEt-R-A-LJ_ _ _ _'L_O_N_GI_'",'""'
<:OMPONENT R
'
l/
Ty
COEFFICIENT D COEFFICIENT f
r-.
- -- . -- -~ - . - - . - - - : COEFFICIENT C
COEFFICIENT 8
-,-
- -COEFFICIENT A I
- ·T· -- ·-- ·-·- · ·- ·-- + .
MAX.OEVIATlON
ON N-S
•
-
MAX DEVIATION
ON E-W
I
+
CORRECTOR MAGNETS
ELECTRO MAGN£TIC OR
MECHANICAL/MAGNETIC
CONSTANT
DEVIATION
t
RE-ALIGN
COMPASS
Coefficient A
This represents a constant deviation due to the combined effecfs of
components b Y and dX of unlike signs.
Referring again to Fig 8.7 we note that components -b Y and +dX
both cause easterly deviation, and a combination of these two will
cause a coefficient +A ; while components +b Y and - dX both cause
216
westerly deviation and in combination cause a coefficient - A. In
each case. the two components, as regards their maximum effects on
the compass, must be equal. Deviation coefficient A may be termed
either real or apparent.
Real A is caused by the induced magnetic components b Y and dX,
and is represented by the amount of displacement of the axis of the
total deviation curve (see Fig 8.8). Apparent A, on the other hand,
is a deviation produced by non-magnetic causes such as misalignment
of the compass or detector unit with respect to the aircraft's longitudinal axis, offsetting of the magnet system , etc. It is not
possible or practicable to separate real and apparent coefficients A.
Coefficient A is calculated by taking the algebraic sum of the
deviations on a number of equidistant compass headings and dividing
this sum by the number of observations made. Usually the average is
taken on the four cardinal headings and the four quadrantal headings.
Thus
Deviation on N + NE+ E + SE+ S + SW+ W + NW
A=-------------------8
Coefficient B
Coefficient B 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 (see Fig 8.2), but when of unlike signs
they tend to counteract each other.
A+ P or +cZ causes a +B, and a -Pora -cZ causes a -B.
Coefficient B is calculated by taking half the algebraic difference
between the deviations on compass heading East and West:
B = Deviation on E- Deviation on W
2
Since components P and cZ cause deviation which varies as the sine
of the aircraft heading, then deviation due to coefficient B may also
be expressed as
Deviation = B X sin (heading)
Coefficient C
Coefficient C represents the resultant deviation due to the presence,
either together or separately, of hard-iron component Q and soft-iron
component fZ (see Fig 8.3). These components when of like and
unlike signs cause deviations whose directions-are the same as those
caused by components P and cZ. Therefore, a +Q or a +/Z causes a
+C, and a - Q or a -fZ causes a -C.
Coefficient C is calculated by taking half the algebraic difference
between the deviations on compass headings N<;:>rth and South :
217
C = Deviation on N - Deviation on S
2
Since components Q and fZ cause deviation which varies as the
cosine of the aircraft heading then deviation due to coefficient C
may also be expressed as
Deviation= C X cos (heading)
Coefficient D
This coefficient represents the deviation due to the presence, either
together or separately, of components aX and e Y. It will be seen
from Fig 8.7 (a) that these components cause deviations of the same
type when they are of unlike signs and counteract each other when
of like signs. When a+ aX or - e Y predominates, or when they are
present together, coefficient D is said to be positive, whilst -aX or
+e Y predominating or together cause a negative coefficient D.
Coefficient D is one-quarter of the algebraic difference between
the sum of the deviations on compass headings North-East and SouthEast and the sum of the deviations on headings South-East and North·
West, i.e.
(Deviation on NE + Deviation SW) - (Deviation on SE
D = + Deviation on NW)
4
The deviations caused by components aX and e Y vary as the sine
of twice the aircraft heading; therefore deviation may also be
expressed as
Deviation= D X sin (twice heading)
Coefficient E
Coefficient E represents the deviation due to the presence of components b Y and dX of like signs. When a + b Y and a+ dX are combined,
coefficient E is·said to be positive, whilst a combination of a - b Y
and a - dX give a negative coefficient; the two components must in
each case be equal.
Coefficient Eis calculated by taking one-quarter of the algebraic
difference between the sum of the deviations on compass headings
North and South and the sum of the deviations on headings East and
West:
(Deviation on N + Deviation on S) - (Deviation on E
E = + Deviation on W)
4
218
The deviations caused by components h Y and dX vary as the
cosine of twice the aircraft heading: therefore deviation may also be
expressed as
Deviation =E X cos (twice heading)
The total deviation on an uncorrected compass for any given
direction of the aircraft's heading by compass may be expressed by
the equation
Total deviation= A + B sin 8 + C cos 8 + D sin 28 + E cos 28.
Adjustment and
Deviation Compensation of Compasses
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 adjustment can
be made and the deviation compensated.
Deviation
Compensation Devices
These devices fall into two distinct groups, mechanical and electromagnetic, the former being employed with simple direct-reading
compasses and detector elements of certain remote-indicating
compasses, and the latter being designed solely for use with detector
elements of remote-indicating compasses.
In both cases, the function is the same, i.e. to neutralize the eff~cts
of the components of an aircraft's hard- and soft-iron magnetism by
setting up opposing magnetic fields.
Mechanical Compensation Devices
One of the earliest mechanical devices is the micro-adjuster shown
in Figs 8.11 and 8.12. It consists of two pairs of magnets (a feature
common to all types of mechanical compensator), each pair being
fitted in bevel gears made of a non-magnetic material. The gears are
mounted one above the other so that, in the neutral condition, one
pair of magnets lies longitudinally for the correction of coefficient
C, and the other pair lies laterally for the correction of coefficient B.
Production of magnetic fields required for correction is obtained by
Figure 8. / / Micro-
adjuster type of deviation
compensator.
219
~;t
,-
Figure 8.12 Operation of
micro-adjuster.
·
~
LONGITUDINAL
OPERATING HEAD
(COEFFlCIENT
Bl
. - -
-
- - . •
g
.
.
-
• -- ,
LATERAL
't'~~~,OPERATINGHEAD
, ,\\ ,
a;i 1
(COEFFICIENT Cl
-~
~
L --
--
. - --- . -
LONGITUDINAL
FIELD
rotating small bevel pinions which mesh with the gears, causing
them to rotate in opposite directions. As can be seen from Fig 8.12,
the magnets are thus made to open up in the manner of a pair of
scissors, the fields being produced between the poles and in a direction dependent on that in which the operating head is rotated.
Let us now refer to Fig 8. 13 in order to see what effects the
fields produced by each pair of magnets have on a compass magnet
system. At (a) it is assumed that the aircraft is positioned on a
northerly heading, and due to a lateral component (Q and its allied
soft-iron) of aircraft magnetism, the detector system is deviated a
certain number of degrees from North, e.g. in an easterly direction.
To eliminate this deviation the lateral operating head must be
rotated in a direction which will open up the longitudinal magnets
and create-a lateral field sufficient to deflect the compass magnet
system westerly and back to North .
Similar compensation for deviation occurring. on a southerly
heading would also be effective as in (c). On easterly or westerly
220
Figure 8. J3 Effects of
fields produced by magnets.
LONGITUOINAl CORRECTOR MAGNETS
HEAOING NORTH
HEADING EASl
~+--
--=f
-:x:-
~
t
I
(a) COMPENSATION
(b) COMPENSATION
EFFECTIVE
NOT EFFECTIVE
HEADING SOUTH
-+--
HEAO<NG WEST
+
~
I
(c) COMPENSATlON
EFFECT!Vt
t Ji
~
' ' t
t
(d) COMPENSATION
NOT EFFEClM
LATERAL CORRECTOR MAGNETS
HEAOING NORTH
~
I
'
t
(e) COMPENSATION
NOT EFFECTM
-
HEAOING EAST
+
....-+--
~
-
(f)
cc
COMPENSATION
EFfECnvt
AIRCRAFT MAGNETIC COMPONENTS -
HEADING SOUTH
tl t
X
ttt
(g) COMPENSATION
NOT EFFECT!Vt
- - - - NEUTRAUSING AELDS
HEADING WEST
~+--
-:~
I
(h) COMPENSATION
EFFECTM
--!- COMPASS MAGNET
SYSTEM
headings, however, the longitudinal magnets are ineffective because
their fields are then aligned with the compass magnet system and so
cannot cause deflection (Fig 8.13 (b) and (d)).
From the series of diagrams relating to Component P and lateral
corrector magnets (Fig 8.13 (e) to (h) it will be noted that their
longitudinal fields are effective on easterly or westerly headings.
Thus for each pair of magnets compensation is effective only on
two cardinal headings ; namely longitudinal magnets, North and
South ; lateral magnets, East and West.
Two other versions of mechanical compensator are shown in Fig
8. 14, the one at (a) being employed with certain types of directreading compass (see page 171 ), and the one at (b) with certain fluxdetector units. Although differing in size and constru ction , they
both employ gear-operated corrector magnet arrangements, the
magnets being positioned side by side and not one above the other
as in the micro-adjuster.
The magnets are mounted on flat gears meshing with each other
and connected to operating heads , which in the compensator sh own
at (a) are operated by a key, and in that at (b) by means of a screwdriver.
As the magnets are rotated, their fields combine to set up
neutralizing components in exactly the same manner as those of a
micro-adjuster. Maximum compensation of deviation on either side
of a quadrantal heading is obtained when the magnets are in
complete alignment. Indication of the neutral position of the
221
Figure 8. 14 Mechanical
compensation de.vices.
(a) Direct-reading compass;
(b) flux detector element.
B
E/W ADJUSTER
(a)
B
·C
NJS ADJUSTER
DIAGRAM SHOWING N/S POLES OF
PERMANENT MAGNETS IN GEARS
(b)
magnets is given by aligning datum marks on the ends .of the magnet
operating spindles and on the casing.
Electromagnetic Compensation Devices
Figure 8./ 5 Electromagnetic
The design and construction of these devices vary depending on the
compensation.
COMPENSATOA
, ·- --
I
-- · - - i
!
_____Q
-~
N-S
FLUX
DETECTOR
I
9-=t~-=-.J=:=!__
-_=:}~~c5v
TO HEADING{---INDICATOR
----- ------ -----------------------'
*
DROPPING RESISTORS
222
type of remote-indicating compass, but in general they comprise two
variable potentiometers which are electrically connected to the coils
of the.flux detector as shown in Fig 8.15 . The potentiometers
correspond to the coefficient 'B' and 'C' magnets of a mechanical
compensator, and when they are rotated with respect to calibrated
dials, they inject very small d.c. signals into the flux detector coils.
The fields produced by the signals are sufficient to oppose those
causing deviations, and they accordingly modify the detec tor output,
which via the synchronous transmission link, will drive the compass
indicator to the corrected readings.
Questions
In connection with the effects of component Q which of the following statements is true?
(a) They correspond to those of a magnet lying longitudinally and
produce maximum deviation on North or South.
(b) They correspond to a magnet lying laterally and produce
maximum deviation on North and South.
(c) They correspond to a magnet lying laterally and produce
maximum deviation on East and West.
8.2 Under what attitude conditions of an aircraft will component R
produce the same effects as positive and negative components P and
Q?
8.3 (a ) ' How are the nine soft-iron components of magnetism designated,
knd on which axes do their induced fields lie?
(b) Which soft-iron components have the same effects as components P and Q?
8.4 Explain how the flux detector element of a remote-indicating compass system can be affected by components of aircraft magnetism.
8.5 (a) What are coefficients A, B, C, D and £?
(b) Name the components of magnetism associated with each
coefficient.
(c) What is the difference between coefficients real A and apparent A?
8.6 (a) Express the formulae used for the calculation of coefficients A. B
and C.
(b) Given the following information find values for each of the three
coefficients:
Magnetic
Magnetic
Compass
Compass
heading
deviation
heading
deviation
8.1
8.7
000°
+4°
180°
- 1°
045°
+2°
225°
-2°
090°
+4°
270°
-2°
135°
+3°
315°
0°
An aircraft has a component -P and a component +Q.
(a) Draw separate curves of deviation caused by them.
223
(b) What effect would the - P have on the compass when heading
8.8
8.9
224
360° and the +Q when heading 270°?
What do you understand by compass 'swinging'?
With the aid of diagrams explain how a deviation compensating device
neutralizes the fields due to aircraft magnetic components.
9 Synchronous
data-transmission
systems
With the introdu ction of large multi-engined aircraft the problem arose
of how to measure various quan tities such as pressure, temperature,
engine speed and fuel tank contents at points located at greater distances from the cockpit. Many of the instruments then available
operated on mechanical J?rinciples which could be adapted to suitably
transmit the required information. For example, on one very early
twin-engined aircraft, mechanically-operated engine speed indicators
were designed with large-diameter dials so that by mount ing the indicators in the engine nacelles they could be read from the cockpit .
However, as aircraft were further increased in size and complexity,
the adaptation of mechanically-operated instruments became severely
limited in application. A demand for improved methods of meas~rement at remote points therefore arose and was met by changing t'o
the use of electrical systems in which an element detects changes in
the measured quantity and transmits the information electrically to
an indicating element.
We can therefore consider most of the instruments used in a p resentday aircraft as being of the remote-indicating type, but many of them
are of a design in which the transmission of data is effected through
the medium of a special synch ro nous system .
·
Synchronous systems fall into two classes: direct-curren t and
alternating-current. The principles of some of those commonly used
form the subject of this chapter. Although varying in the method of
data transmission, all the systems have one common feature: they
consist of a transmitter located at the source of measurement and a
receiver which is used t o position the indicating element.
Direct-Current
Synchronous Systems
The Desynn System
This system, one of the earliest to be used in aircraft , may take one
of three forms, namely: rotary motion or toroidal resistance for
position and liquid-contents indications ; linear motion o r microDesynn for pressure iridication, and slab-Desynn also for pressure
indication. The principle of operation is the same in each case, but
the rotary-motion arrangement may be considered as the basic system
from which the oth ers h ave been developed.
225
The Basic System
The electrical element of the transmitter consists of a resistor wound
on a circular former (called the 'toroidal resistor') and tapped at three
points 120° apart. Two diametrically-opposed wiper contact arms,
one positive and the other negative, are insulated from each other by
a slotted arm which engages with a pin actuated by the appropriate
mechanical element of the transmitter.
The wiper contact arms are assembled in the form of a bar having
rotational freedom about a pivot which carries current to the positive
arm. Current to the negative arm is carried via a wiper boss the underside of which is in contact with a ring fitted on the inner side of the
terminal moulding. A circlip, fitted at the end of the pivot, holds the
complete assembly in place against a spring which gives the required
contact pressure on the toroidal resistor.
The receiver element consists of a cylindrical two-pole permanentmagnet rotor pivoted to rotate within the field of a laminated softiron stator, carrying a star-connected three-phase distributed winding
supplied from the toroidal resistor tappings. A tubular brass housing
is fitted inside the stator, and together with its end cover, provides a
jewelled bearing support for the rotor spindle. The front end of the
spindle projects through the end cover and a dial mounting plate, to
cany the pointer. Electrical connection between the transmitting and
recei":'ing elements may be either by terminal screws or plug-type
connectors.
The electrical elements of the receivers are common to all three
circuit arrangements of the Desynn system.
Operation
When direct current is applied to the transmitter contact arms, which
are in contact with the toroidal resistor, currents flow in the resistor
causing the three tapping points to be at different potentials. For
example, with the contact arms in the position shown in Fig 9. 1 the
potential at tapping No 2 is greater than that at No l because there is
less resistance in the circuit between the positive arm and the No 2
tapping. Thus, currents are caused to flow in the lines between transmitter and receiver, the magnitude and direction of which depend
upon the position of the contact arms on the toroidal resistor.
In turn, these currents flow through the coils of the receiver stator
and produce a magnetic field about each coil similar to that of a bar
magnet ; thus either end of a coil may be designated as a N-pole or a
S-pole, depending on the direction of the current through a coil.
The combined fields extend across the stator gap and cause the
permanent-magnet rotor to align itself with their resultant.
A pull-off ma-gnet is fitted to the end plate, its purpose being to
act as a power-failure device by exerting an attractive force on the
main magnet rotor so as to pull it and the pointer to an off-scale
226
Figure 9:J Circuit diagram of
TRANSMITTER
RECEMA
basic Desynn system.
+
28V
position when current to the stator is interrupted. The strength
of this magnet is such that, when the system is in operation, it does
not distort the main controlling field.
The Micro-Desynn System
In applications where the movement of a prime mover is small and
linear, the use of a basic-system transmitting element is strictly
limited. The micro-Desynn transmitter was therefore developed to
permit the magnification of such small movements and to produce,
by linear movement of contacts, the same electrical results as the
complete rotation of the contact arms of the basic transmitter.
In order to understand the development of this transmitting
element, Jet us imagine that a toroidal resistor has been cut in two,
laid out flat with its ends joined together, and three tappings made
as before together with positive and negative arms in contact with it.
Movement of the contact arms wiJJ produce varying potentials at tlie
tapping, but as will be clear from Fig 9.2 (a), the full range will not
be covered because one or othei: of the arms would run off the
resistor. We thus need a second resistor with tappings so arranged
that the contact arms can move through equal distances.
If we now take two toroidal resistors and join them in parallel
then, by cutting them both in two and laying them out flat, we
obtain the circuit arrangement shown at (b ) . By linking the contact
arms together and insulating them from each other, they can now be
moved over the whole length of each resistor to produce voltage and
current combinations which will rotate the receiver through 360°.
Since the contact arms .have to traverse a much shorter path, their
angular movement can be kept small (usually 45° ), a feature which
helps to reduce the energy required to operate the transmitter.
The resistors are wound on bobbins which may be of round or
227
Figure 9.2 Circuit diagram of
micro-Desynn system.
2
2
3
3
(a)
X
I
a
~"·-
-
i
,-
l
2b
Tb
I
b
3
l
·Y
+
2b
2
a
I
(b)
2
3
(C)
square section, the latter type being designed to help reduce cyclic
and friction errors.
Each resistance bobbin is secured in place against a set of miniature
spring contact fingers accurately positioned so as to provide the
necessary tapping points.
The contact arms are mounted on a rocker shaft supported between
the vertical parts of a U-shaped bracket, and movement of the transmitter's mechanical element is transmitted to the arms via a springloaded operating pin and crank arm connected to the rocker shaft.
Two beryllium-copper hair-springs conduct current to the contact
arms and also act together to return the rocker shaft and .con tact
arms to their starting position.
The Slab-Desynn System
In addition to the cyclic error present in the basic and micro-type
systems, small errors also arise due to friction set up by the contact
arms having to move over a considerable surface of resistance wire.
Although such errors can be reduced by providing a good contact
228
material and by burnishing the re$istance wire surface, the cyclic
error is still undesirable in certain measurements.
A solution to th.is problem was brought about by modifying the
basic system so as to change its three sawtooth waveforms into
sinusoidal waves, the instantaneous sum of which is always zero.
The transmitter so developed is shown schematically in Fig 9.3 ,
from which it will be noted that the resistor and contact arms have,
as far as electrical connections are concerned, virtually changed
places with each other. The resistor is now wound on a slab former,
hence the term 'slab-Desynn', and is connected to the direct-current
supply, while the contact arms themselves provide the three
potential tapping points for the indicator stator.
Figure 9.3 Slab-Desynn
transmitter.
SLAB RESISTOR
CONTACT SHAFT
PIVOT
CONTACT
ORM FROM MECHANICAL
ELEMENT
The three contact arms are insulated from each o.ther and pivoted
over the centre of the slab, and are each connected to a slip ring.
Spring-finger brushes bear against these slip rings and convey the
output currents to the stator coils. Movement of the mechanical
element is transmitted to the brushes via a gearing system.
·
Alternating-Current
Synchronous Systems
Systems requiring alternating current for their operation are usually
classified under the generic terms synchro and are manufactured
under various contracted trade names, e.g. Autosyn, Selsyn. All
these systems operate on the same principle and are normally divided
into four main groups accoi;-ding to their function; (i) torque synchros,
(ii) control synchros, (iii) differential synchros and (iv) resolver
synchros.
Before going into the operating principles of synchros, it is necessary to have a clear understanding firstly of the process of electromagnetic induction and secondly its application to a standard type
of transformer.
Electromagnetic Induction
Electromagnetic induction refers broadly to the production of an
229
e.m.f. within a conductor when there is relative movement between
it and a magnetic field .
If a conductor is moved in a magnetic field the lines of force will
be cut and an e.m.f. will be induced in the conductor of a magnitude
proportional to the rate at which the lines are cut. If the ends of the
conductor are connected together, thus forming a closed circuit, a
current resulting from the e.m.f. will flow so long as the conductor
is moving.
The effect would be the same if the conductor were stationary
and the field were moving. When current is passed through a
conductor and is increased or decreased , the field will proportionately increase or decrease, causing the flux lines around each turn of
the coil to move in an expanding and contracting manner thus
cutting adjacent turns and inducing an e.m.f. in the coil. If the
current is constant no e.m.f. will be induced.
The property described in the foregoing paragraphs is known as
self-inductance and is the fundamental principle on which motors
and generators operate.
Another inductive effect we may now consider is that when two
current-carrying coils are placed in close proximity to each .other
they become linked by the fluxes they produce, so that, when the
current through one coil is changed, an e.m.f. will be induced not
only in that coil but also in the adjacent coil. This property is
known as mutual inductance and is utilized in the transformer.
Principle of the Transformer
A transformer, shown in Fig 9.4, employs two electrically separate
coils on an iron core. One of the coils, the primary, is connected to
an alternating-voltage supply and the other, known as the secondary,
is provided with terminals from which an output voltage is taken.
The alternating-voltage applied to the primary causes an alternating
current to flow in this winding, the current being dependent on the
inductance of the winding. The effect of the alternating current is to
set up an alternating magnetic flux in the core, and since most of
this flux links with the secondary winding, an alternating e.m.f. is
produced in the secondary by mutual induction. If the secondary
terminals are connected to a closed circuit, a current will flow and all
the energy expended in the circuit is transferred magnetically through
the core from the supply source connected to the primary terminals.
As already stated, the primary and secondary windings are on a
common core, which means that the magnetic flux within the core
will, in addfrion to linking the secondary winding, also link the primary
winding. Thus an e.m.f. is set up in the primary winding which,
neglecting resistance, can be considered for practical purposes as
being equal and opposite to the voltage applied to the primary. If
230
Figure 9.4 Principle of
transformer.
LAMINATED IRON CORE
~~~~ ~~~~tRV
N, TURNS
v]\ll[v,
PRIMARY WINDING
N'p TURNS
the primary winding has Np turns the voltage per turn is Vp /Np.
Since the same flux cuts the secondary it will induce in it the same
voltage per tum and the secondary voltage will thus be
Vp
Vs =N- Ns
p
so that
Primary voltage _ Vp Np
Secondary voltage - Vs = Ns
The ratio of voltages in the two windings is therefore proportional to
the ratio of the turns, known as the turns ratio of the transformer.
There is one further effect concerning secondary current which
should be considered. When the current flows it too sets up a flux
which opposes the main core flux. A reduction in the core flux
reduces the primary e.m.f. opposing the applied voitage, and so allows
increased current to flow in the primary winding. This increased
current then restores the core flux to a value which is only very
slightly Jes!> than its no-load value.
Neglecting transformer losses, which are generally very small , ·the
input power may be equated to the output power:
Vpfp = Vsfs
so that
Ip
Vs
Ns
-ls - -VP Np
from which it is clear that the currents are inversely proportional to
the number of turns.
231
Torque Synchros
These are the simplest form of synchro and are used for the transmission of angular position information by means of induced signals,
and for the reproduction of this information by the position of a
shaft at an output or receiver element. A typical application of
torque synchros is in flight instrument systems.
A torque synchro system comprises two electrically similar
units interconnected-as shown in Fig 9.5, and by convention one is
designated the transmitter (TX) and the other the receiver (TR).
Figure 9.5 Torque synchro
ELECTRICAL
ELECTRICAL
ZERO
system.
ZERO
I
I
Sl
51
I
INPUT SH.AFT
OUTPUT SHAFT
A.C. SUPPLY
- - CURRENT
__,..
ROTOR FIELDS
==!>
STATOR FIELDS
CIRCUIT SYMBOL
Each unit consists of a rotor carrying a winding, and concentrically
mounted in a stator carrying three windings the axes of which are
l 20° apart. The principal physical differences between the TX and
the TR are that the rotor of the TX is mechanically coupled to an
input shaft, while the TR rotor is free to rotate. The rotor windings
are connected to a source of single-phase a.c. supply, and the
232
corresponding stator connections are joined together by transmission
lines. The similarity between these connection arrangements and a
conventional transformer may also be noted; the rotors correspond
to primary windings and the stators to secondary windings.
When the rotors are aligned with their respective stators in the
position indicated they are said to be at 'electrical zero' ; this refers
to the reference angle standardized for synchros at which a given set
of stator voltages will be produced, and by this convention enables
replacement synchros to be matched to each other.
With power applied to the rotors, due to transformer action a
certain voltage will be induced in the stator coils the value of which
will be governed, as in any transformer, by the ratio of the number
of turns of the rotor (primary) and stator (secondary) coils.
When the rotors of TX and TR occupy the same angular positions,
and power is applied, equal and opposite voltages will be produced
and hence no current can flow in the stator coils. The system (and
any other type of synchro) is then said to be at 'null'.
When the rotors occupy different angular positions, for example
when the TX rotor is at the 30° position and the TR rotor is at
electrical zero, an unbalance occurs between stator coil voltages
causing current to flow in the lines and stator coils. The currents
are greatest in the circui ts where voltage unbalance is greatest and
their effect is to produce magnetic fields which exert torques to tum
the TR rotor to the same position as that of the TX.
As the TR rotor continues to tum , the misalignment, voltage
unbalance and currents decrease until the 30° position is reached
and no further torque is exerted on the rotor.
In considering this synchronizing action one might assume that ,
since currents are also flowing in the stator coils of TX, its rotor would
be returned to 'null'. This is a reasonable assumption, because in fact a
torque is set up tending to turn the rotor in a clockwise direction.
However, it must be remembered that the rotor is being actuated by
some prime mover which exerts loads too great to be overcome by
the rotor torques.
Control Synchros
Control synchros differ from torque synchros, in that their function
is to produce an error voltage signal in the receiving element, as opposed to the production of a rotor torque. Typical uses of control
synchros are in servoed altimeters and airspeed indicators which
operate in conjunction with central air data computers (see page I 06).
The interconnection of the two elements of a control synchro
system is shown in Fig 9.6. By convention , the transmitter is
designated as CX, and the receiver designated as a con trol transformer
(CT). The CX is similar to a torque transmitter, and from the
233
Figure 9.6 Control synchro
system.
.ELECTRICAL
ZERO
CT
ELECTRICAL
ZERO----
-
AMPLIFIED ERROR
VOLT AGE TO CONTROL
PHASE
INPUT SHAFT
rv-vv-
APPLIED~/
VOLTAGE
VOOULTTAPGUTE O NO ERRO R SIGNAL
AC - W ERROR VOLTAGE
OUT ·OF ·PHASE
- - CURRENT
-
=
S1=©=
ROTOR FIELD
S2
STATOR FIELD
S3
Rl
R2
CIRCUIT SYMBOL
diagram it will be noted that the a.c. supply is connected to the CX
rotor only. The CT rotor is not energized since it acts merely as an
inductive winding for detecting the phase and magnitude of error
signal voltages which are supplied to an amplifier. The amplified
signals are then "fed to a two-phase motor which is mechanically
coupled to the CT rotor. Another difference to be noted is that a
control synchro system is at electrical zero when the rotor of CT is
at 90° with respect to the CX rotor.
If the rotor of CX is rotated through a certain angle, the resultant
flux in the CT stator will be displaced from its datum point by the
same angle, and relative to the CT rotor position at that instant.
An error voltage is therefore induced in the rotor, the phase and
magnitude of the voltage depending on the direction of CX rotor
rotation, and on the degree of misalignment between it and the CT
rotor. The error voltage is then amplified and fed to the control
234
phase of the motor, the other phase (reference phase) being
continuously supplied with alternating current. Since the control
phase voltage 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. As the motor rotates, it turns
the rotor of the CT in the appropriate direction, thereby reducing
its displacement relative to the ex rotor. Rotation continues until
both rotors are in alignment (bearing in mind, of course, that the
electrical zero points are at 90° from each other) at which position
no further error voltage is induced.
Differential Synchros
In some cases, 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 difference
or the sum of the two angles. This is achieved by introducing a third
synchro into either a torque or control system, and using it as a
differential transmitter. Unlike TX or ex synchros, a differential
transmitter (designated TDX or eDX) has an identically wound stator
and rotor which in the application to a torque synchro system are
interconnected as shown in Fig 9. 7.
At (a) the TX rotor is shown rotated clockwise through 60° while
the rotor of TDX remains at electrical zero; all the magnetic fields
rotate, and the rotor of TR takes up the same angular position as
the rotor of TX. If now the TX rotor remains at electrical zero, and
the TDX rotor is rotated clockwise through 15° say, the fields of
both synchros remain in the electrical zero position because their
position is determined by the orientation of the TX rotor (diagram
(b)). However, a 15° clockwise rotation of the TDX rotor without a
change in the position of its field is equivalent to moving the rotor
field 15° anticlockwise whilst leaving the rotor at electrical zero.
This relative angular change is duplicated in the stator of TR and so
its rotor will align itself with the field i.e. for a 15° clockwise rotation
of the TDX rotor, the TR rotor will rotate 15° anticlockwise.
Assume now that the TX rotor is rotated through 60° clockwise,
and the TDX rotor through 15° clockwise, then because the TR
rotor will rotate 15° anticlockwise, its final angular movement will
be equal to the difference between the two input angles i.e. it will
turn through 45° (diagram (c)). The differential effect is of course
reversed when the TDX rotor is rotated in the opposite direction to
the TX rotor, so that the TR rotor rotates through an angle equal to
~he sum of the two input angles. By reversing pairs of leads either
between TX and TDX, or between TDX and TR, any one of the
rotors can be made to assume a position equal to the sum or differ235
Figure 9. 7 Differential synchro
in a torque synchro system.
TX
TOX
TA
Al
INPUT SHAFT
INPUT SHAFT
OUTPUT SHAFT
SYNCHROS AT ELECTRICAL ZERO
A.C. SUPPLY
{a)
ELECTRICAL ZERO
{b)
{C)
Al~SI
A2
S2
R3
S3
CI RCUIT SYMBOL OF ·
DIFFERENTIAL SYNCHRO
236
-
ROTOR FIELDS
==I> STATOR FIELDS
Figure 9.8 Differential synchro
in a control synchro system.
ex
ence of the angular positions of the other rotors.
In the same way that differential transmitter synchros can be used
in torque synchro systems, so they can be used in systems utilizing
control synchros to transmit control signal information on the sum
or difference of two angles. The basic arrangement is shown in Fig
9.8.
cox
CT
Sl
Sl
Al
ERROR SIGNAL
PROPORTIONAL
:
- - - - - TO 0 1 -112
R2
S2
I
'
J'" "' ""
ANGLE OF FIELD
PROPORTIONAL TO 11 1
-
-
112
ROTOR FIELDS
~ STATOR FIELDS
Resolver Synchros
The function of resolver synchros (RS) is to convert alternating
voltages, which represent the cartesian coordinates of a point, into a
shaft position and a voltage, which together represent the polar
coordinates of that point. They may also be used in the reverse
manner for voltage conversion from polar to cartesian coordinates.
Typical applications of resolver synchros are to be found in flight
director and integrated instrument systems (see Chapter 15).
A typical arrangement of an RS for conversion from polar to
cartesian coordinates is shown in Fig 9.9 and from this it will be
noted that the stator and rotor each have two windings arranged in
phase quadrature, thus providing an eight-terminal synchro. An
alternating voltage is applied to the rotor winding R1-R2, and the
magnitude of this voltage, together with the angle through which the
rotor is turned, represent the polar coordinates. In this application,
237
Figure 9.9 Resolver synchro
ROTOR
STATOR
RI
S1
~ V01)0"'-7h..ll2lu4k .~
C,
"fl) [l.
INPUT SHAFT
S4 S2
S3
S1
CIRCUIT SYMBOL OF
RESOLVER SYNCHRO
- - - - - - -·
[
,
RI
FLUX
~onn~nnonnnonn~ru
]uuiJ1Ju111nn1qui::,
I
I
1
I
I
I
I
.
I
I
I
I
I
I
It C-·-~'"'"
L A.~. SUPPLY
R2
SI'
0
S2
ATI' fJ
,
t_Q_j
• S4
I
I
I
o•:
I
90"1
I
1so•1
I
I
270• 1 380° 1VALUE
~
OF/I
- -- - - O R SIN/I
53
I
I
I
the second winding is unused, and as is usual in such cases, it is shortcircuited to improve the accuracy of the RS and to limit spurious
response.
In the position shown, the alternating flux produced by the
current through rotor winding R 1-R2 links with both stator
windings, but since the rotor winding is aligned only with S 1-S 2
then maximum voltage will be induced in this winding. Winding
S3-S4 is in phase quadrature so no voltage is induced in it. When the
rotor is at a constant speed it will induce voltages in both stator
windings, the voltages varying sinusoidally. The voltage across that
stator winding which is aligned with the rotor at electrical zero will be
a maximum at that position and will fall to zero after rotor displacement of 90°; this voltage is therefore a measure of the cosine of the
displacement. The voltage is in phase with the voltage applied to
238
Figure 9.10 Conversion of
cartesian coordinates to polar
coordinates.
R 1-R2 during the first 90° of displacement, and in anti-phase from
90° to 270°, finally rising from zero at 270° to maximum in-phase at
360°. Any angular displacement can therefore be identified by the
amplitude and phase ·of the induced stator voltages. At electrical zero,
stator winding S3-S 4 will have zero voltage induced in it, but at 90°
displacement of rotor winding R 1-R2 , maximum in-phase voltage
will be induced and will vary sinusoidally throughout 360°; thus, the
S3-S 4 voltage is directly proportional to the sine of the rotor
displacement. The phase depends on the angle of displacement, any
angle being identified by the amplitude and phase of the voltages
induced in stator winding S 3-S4. The sum of the outputs from
both stators, i.e. r cos 8 plus r sin 8, therefore defines in cartesian
coordinates the input voltage and rotor rotation.
Figure 9 .10 illustrates an arrangement whereby cartesian coordinates may be converted to polar coordinates. An alternating voltage
Vx = r cos 8. is applied to the cosine stator winding S1-S2, while a
voltage Vy = r sin 8 is applied to the sine stator winding S3-S4. An
alternating flux representing cartesian coordinates is therefore
RI
, -,me~/
~
J
j,.
v, - rsinO
R2
a
0
R4
8
m
MECHANICAL ORIVE
/v,2 +v/
produced inside the complete stator. One of the rotor windings, in
this case R 1-R2 , is connected to an amplifier, and in the position
shown it will have maximum voltage induced in it, which will be
applied to the amplifier. The output from the amplifier is applied to
a servomotor which is mechanically coupled to a load and to the
rotor. When the rotor is turned through 90° the induced voltage in
winding R 1-R2 reduces to zero and the servomotor will stop. The
rotor winding R3-R 4 will now be aligned with the stator flux, and a
voltage will be induced in it which is proportional to the amplitude
of the alternating flux as represented by the vector r i.e. a voltage
proportional to..)( Vx 2 + Vy 2 ) . This voltage together with the angular position of the rotor therefore represents an output in terms of
the polar coordinates.
239
Synchrotcl
A synchrotel is generally used as a low-torque control transformer or
transmitter. It employs.a conventional 3-phase stator, but as will be
noted from Fig 9. l 1, unlike a conventional sy nchro the rotor section
is in three separate parts: a hollow aluminium cylindrical rotor of
oblique section, a fixed single-phase rotor winding, and a cylindrical
core about which the rotor rotates. The rotor shaft is supported in
jewelled bearings and is connected to the pressure-sensing element or
whatever element the application demands.
Figure 9./1 Synchrotel
,---~ ~ ~ ~ ~~
3-PHASE
SIGNALS FROM
~ ~ ~ - CONTROL
TRANSMITIER
STATOR
ROTOR COIL
3-PHASE STATOR
[>
SIGHA&.S TO
2·l'tlAS£
MOTOR
AMPUFIEA
In a typical pressure-measuring application, the synchrotel is electrically connected to a synchro control transmitter whose rotor is
made to follow the synchrotel rotor position ; in other words, it acts
as a servo loop system.
The synchro transmitter rotor is energized by a 26 V 400 Hz singlephase supply which induces voltages in the transmitter stator. As
this stator is connected to the 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
rotor winding is also fixed around the core, and therefore the core
flux will induce an alternating voltage in the winding, and the amplitude of this voltage will be a sinusoidal function of the relative posi240
tions of the rotor and stator flux. This voltage is fed, via an amplifier,
to the control phase of a two-phase servo motor which drives the
synchro transmitter rotor round in its stator thereby causing a change
in the synchrotel stator flux, to the point where no voltage is induced
in the rotor winding, i.e. the synchro transmitter is driven to the null
position. This position corresponds to the pressure measured by the
sensing unit at that instant.
Questions
9.1
9.2
9.3
9.4
9.5
9.6
9.7
9.8
9.9
9.10
Draw a circuit diagram of the basic Desynn system and explain its
operating principle.
Explain how the basic system is developed for such applications as
pressure measurement.
How is the pointer of a Desynn indicator returned to the 'off-scale'
position?
Explain the difference between the basic and slab-Desynn systems.
What is the difference between self-induction and mutual induction?
By means of a schematic diagram explain the operation of a torque
synchro system.
In what type of synchro system is a control transformer utilized?
Explain the operating principle.
In what manner does a Synchrotel differ in construction and operation
from. the more conventional synchro systems?
Explain how a synchro is applied to systems involving the measurement of the sine and cosine components of angles.
For what purposes are differential synchros used?
241
1O Measurement of engine
speed
The measurement of engine speed is of considerable importance,
since together with such parameters as manifold pressure, torque
pressure and exhaust gas temperature, it permits an accurate control
over the performance of the appropriate type of engine to be maintained.
With reciprocating engines the speed measured is that of the crankshaft, while with turboprop and turbojet engines the rotational speed
of the compressor shaft is measured, such measurement serving as a
useful indication of the thrust being produced. The indicating instruments are normally referred to as tachometers.
The method most commonly used for measuring these speeds is an
electrical one, although in several types of general aviation aircraft,
mechanically-operated tachometers are employed.
Mechanical
Tachometers
Tachometers of this type consist of a magnet whichjs continually
rotated by a flexible· shaft coupled to a drive outlet at the engine.
An alloy cup-shaped element (known as a drag cup) fits arouJ1d the
magnet such that a small gap is left between the two. The drag cup
is supported on a shaft to which is attached a pointer and a controllin!
spring. As the magnet rotates it induces eddy currents in the drag cup
which ttnd to rotate the cup at the same speed as the magnet. This,
however, is restrained by the controlling spring in such a manner that
for any one speed, the eddy current drag and spring tension are in
equilibrium and the pointer then indicates the corresponding speed
on the tachometer dial.
Electrical Tachometer
Systems
Systems of this type currently fall into two main categories: (i)
generator and indicator, and (ii) tacho probe and indicator.
Generator and Indicator Systems
A generator consists of a permanent-magnet rotor rotating within a
slotted stator which carries a star-connected three-phase winding.
The rotor may be of either two-pole or four-pole construction as
242
shown in Fig 10.1; in some applications a twelve-pole rotor may be
used. It will be noted that the poles of the four-pole rotor are
angled so that when one end of a pole leaves one stator tooth the
other end is entering the next tooth. This produces the best waveform and permits an even driving torque·. With the two-pole rotor
the same effect is achieved by skewing the stator teeth and the individual coils which make up a phase.
Figure 10.J Types of
generator rotor.
4·POLI:
The method most commonly used for driving a rotor is by means
of a splined shaft coupling, the generator as a whole being bolted
directly to a mounting pad at the appropriate accessories drive gear
outlet of the engine.
In order to limit the mechanical loads on generators, the operating
speed of rotors is. reduced by means of either four-to-one or two-toone ratio gears in the engine drive system. A sectioned view of a
typical spline-drive generator is shown in Fig 10.2.
Figure 10.2 Sectioned view of
spline-<irive generator. 1 Ball
bearings, 2 oil-seal retaining
ring, 3 oil seal, 4 two-pole
pennanent-magnet rotor,
5 grease retainer, 6 ball
bearings, 7 sealing cover, 8
connector, 9 driving spline.
243
Another example of a direct-drive a.c. generator is shown in Fig
10.3. It is of smaller construction, the rotor being either two-pole
or twelve-pole, and driven via a square-ended shaft. The two-pole
generator is utilized in conjunction with a three-phase synchronous
motor type of indicator, while the twelve;pole generator, which
produces a single-phase output at a much higher frequency and
sensitivity, is utilized in conjunction with servoed counter/ pointer
indicators, and also for supplying signals to engine control units.
Figure 10.3 Lightweight
generator.
A typical indicator, shown in .Fig 10.4, consists of two interconnected elements: a driving element and an eddy-current-drag
speed-indicating element.
Let us consider first the driving element. This is, in fact, a synchronous motor having a star-connected three-phase stator winding
and a rotor revolving on two ball bearings. The rotor is of composite
Figure /0.4 Sectioned view
of a typical synchronous
motor type tachometer
indicator. l Cantilever shaft,
2 terminal block assembly,
3 rear ball bearing, 4
magnetic cup assembly, 5
drag element assembly,
6 small pointer spindle and
gear, 7 outer spindle bearing,
8 bearing locking tag,
9 intermediate gear, 10 bearing
plate, 11 hairspring anchor tag, 9 -r-tttH--=- i
12 inner spindle bearing,
8
13 front ball bearing, 14 rotor,
15 stator.
7
6
244
14 15
2
construction, embodying in one part soft-iron laminations, and in the
other part a laminated two-pole permanent magnet. An aluminium
disc separates the two parts, and a series of longitudinal copper bars
pass through the rotor forming a squirrel-cage. The purpose of
constructing the rotor in this manner is to combine the self-starting
and high torque properties of a squirrel-cage motor with the selfsynchronous properties associated with a permanent-magnet type of
motor.
The speed-indicating element consists of a cylindrical permanentmagnet rotor inserted into a drum so that a small airgap is left
between the periphery of the magnet and drum. A metal cup, called
a drag cup, is mounted on a shaft aild is supported in jewelled
bearings so as to reduce frictional forces, in such a way that it fits
over the magnet rotor to reduce the airgap to a minimum. A calibrated hairspring is attached at one end of the drag-cup shaft, and at
the other end to the mechanism frame. At the front end of the dragcup shaft a gear train is coupled to two concentrically mounted
pointers; a small one indicating hundreds and a large one indicating
· thousands of rev./min.
System Operation
As the generator rotor is driven round inside its stator, the polt:i
sweep past each stator winding in succession so that three waves or
phases of alternating e.m.f. are generated, the waves being 120°
apart (see Fig 10.5). The magnitude of the e.m.f. induced by the
magnet depends on the strength of the magnet and the number of
turns on the phase coils. Furthermore, as each coil is passed by a
pair of rotor poles, the induced e.m.f. completes one cycle at a
frequency determined by the rotational speed of the rotor. Therefore, rotor speed and frequency are directly proportional, and since
the rotor is driven by the engine at some fixed ratio then the
frequency of induced e.m.f. is a measure of the engine speed.
The generator e.m.f. 's are supplied to the corresponding phase
coils of the indicator stator to produce currents of a magnitude and
direction dependent on the e.m.f.'s. 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 copper bars of the squirrel-cage rotor, inducing a current
in them which, in tum, sets up a magnetic field around each bar.
The reaction of these fields with the main rotating field produces a
torque on the rot.or causing it 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 the permanent245
Figure 10.5 Principle of a
generator and indicator system.
PERMANENT ·MAGNET
ROTOR
magnetic field, 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 of the drag-cup 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 shaffand 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 fitting a
thermo-magnetic shunt device adjacent to the magnet. It operates in
a similar manner to the compensator described on page 17.
Figure 10.6 shows another version of speed-indicating element
which is used in some types of indicator. In 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 a drag disc. Rotation of
246
Figure J0.6 Disc-type of drag
element.
DRAG DISC
SPINDLE TO
POtNlcA GEAR
MECH...NISM
TEMPERATURE
CDMPENS"TOR
SPACERS (3)
the disc is transmitted to pointers in a similar manner to the drag-cup
method.
Percentage Rev./Min. Tachometers
The measurement of engine speed in terms of a percentage is adopted
for turbojet engine operation, and was introduced so that various
types of engine could be operated on the same basis of comparison.
The dial presentations of three currently used percentage tachometers
are shown in Fig 10.7. The main scales are calibrated from Oto 100%
in l 0% increments, with l 00% corresponding to the optimum turbine
Figure JO. 7 Dial presentations
of percentage tachometers.
(a) Synchronous motor type;
(b) d.c. torques motor indicator;
(c) servoed counter/pointer
indicator.
speed. In order to achieve this the engine manufacturer chooses a
ratio between the actual turbine speed and the generator drive so
that the optimum speed produces 4,200 rev./min. at the generator
drive. A second pointer or qigital counter displays speed in I%
increments.
247
Servo-operated Tachometer Indicator
Figure 10.8 Servo-operated
tachometer indicator.
Indicators of this type are currently used in several types of public
transport aircraft in conjunction with a.c. generators . A schematic
diagram of the internal circuit arrangement and the construction of
an indicator are shown in Figs 10.8 and 10.9 respectively.
(·h··· .
I
.,
I
•
11
II
:
'
~ - - - - - - - ------,
~~
I
I
,ENERATOR
aGNAL
I
I
lS!_?NAL PROCES~NG MODULE
r--------,
~ ~t I
A.C. o-+j
I
POWER
SUPPLY
MODULE
_ _ _ _ ____jI
I
4
SERVC?_ ~D_MONITOR Mopu E
I
I
I
2s v D.c .L _ ___ ___ ...J
ll> SIGNAL SQUARING
~
BUFFER
IP SERVO DRIVE
B> FEEDBACK BUFFER
lCI: MECHANICAL DRIVE
The generator signals are firstly converted to a square waveform b)
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 signal shaping
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 d.c. 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 d.c. output from the wiper of a positiona
feedback potentiometer. Since the wiper is geared to the main
248
Fig11re JO. 9 Construction of
seJVo-operated tachometer
indicator.
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
motor; the indicator pointer and digital counter are then driven to the
demanded speed position. At the same time, the feedback potentiometer wiper is also re-positioned to provide a feedback voltage to
back-off the 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 solepoid-controlled warning flag which
appears across the digital counter display .
An overspeed pointer is also fitted concentrically wjth the main
pointer, and is initially positioned at the appropriate scale graduation.
If the main pointer exceeds this position, the limit pointer is carried
with it. When the speed has been reduced the main p0inter will move
correspondingly, but the limit pointer will remain at the maximum
speed reached since it is under the control of a ratchet mechanism.
249
It can be returned to its initial position by applying a separately
switched 28 V d.c. supply to a reset solenoid within the indicator.
Tacho Probe and Indicator System
This system is used in several types of large public transport aircraft,
and has the advantage of providing separate electrical outputs
additional to those required for speed indication, e.g. flight data
recording and engine control. Furthennore, there is the advantage
that a probe (see Fig 10.10) has no moving parts.
The stainless steel, hermetically-sealed probe comprises a pennanent 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 of probe) provide outputs
to the indicator and other equipment requiring engine speed data.
The probe is flange-mounted on the engine at a station in the highpressure compressor section of the engine 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 posi-
Figure IO.JO Tacho probe.
M &f;NFT
AXIS OF POLARIZATION
SENSING COILS
250
ELECTRICAL CONNECTOR
Figure JO.J l Simplified
schematic of a d.c. torquer
motor tachometer.
tion, 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
providP.d 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 the gear wheel teeth. As the flux
density changes, an e.m.f. is induced in the sensing coils, the amplitude of the e.m .f. varying with the rate of flux density change. Thus,
in taking the position shown in Fig 10.10 as the starting position,
maximum intensity would occur, but the rate of density change
would be zero, and so the induced e.m.f. would be at zero amplitude.
When the gear teeth move from this position, the flux density firstly
begins to decrease reaching a maximum rate of change and thereby
inducing an e.m.f. 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 e.m.f. 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 e.m.f. reaching a maximum coincident with the greatest rate
of flux density change. The probe and gear teeth may therefore be
considered as a magnetic flux switch that induces e.m.f.'s directly
proportional to the gear wheel and compressor or fan shaft speed.
The output signals for speed indication purposes are supplied to
an indicator of the d.c. torquer type, the dial presentation of which
is shown in Fig 10.7 (b ). The signals pass through a signal processing
module (see Fig 10.11) and are summed with an output from a
servo potentiometer and a buffer amplifier. After summation the
TACHO P R O B E - - - - PR8t1~~NG
SIGNAL
MODULE
115 V A.C.
400 Hz
POWER
SUPPLY
MODULE
>--- - - -
14 V O.C.
251
signal passes through a servo amplifier to the torquer which then
rotates the indicator pointers to indicate the changes in probe signals in terms of speed. The servo potentiometer is supplied with a
reference voltage, and since its wiper is also positioned by the torquer,
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 supply or signal failure , the
main pointer of the indicator is returned to an 'off-scale' position
under the action of a pre-loaded helical spring.
Synchroscopes
252
In aircraft powered by a multi-arrangement of either piston engines,
or turbopropeller engines, the problem arises of maintaining the
engine speeds in synchronism at 'on-speed' conditions and so
minimizing the effects of structural vibration and noise.
The simplest method of maintaining synchronism between
engines would be to manually adjust the throttle and speecl control
systems of the engines until the relevant tachometer indicators read
the same. This, however, is not very practical for the simple reason
that individual instruments can have different permissible indication
errors; therefore, when made to read the same operating speeds, the
engines would in fact be running at speeds differing by the indication
errors. In addition, the synchronizing of engines by a direct comparison of tachometer indicator readings is made somewhat difficult
by the sensitivity of the instruments causing a pilot or engineer to
overshoot or undershoot an on-speed condition by having to 'chase
the pointers'.
In order to facilitate manual adjustment of speed an additional
instrument known as a synchroscope was introduced. It provides a
qualitative indication of the differences in speeds between two or
more engines, and by using the technique of setting up the required
on-speed conditions on a selected master engine , the instrument also
provides a clear and unmistakeable indication of whether a slave
engine is running faster or slower than the master.
The instrument was designed at the outset for operation from the
alternating current g~nerated by the tachometer system, and it therefore forms an electrical part of this system. The dial presentations of
synchroscopes designed for use in twin and four-engined aircraft are
shown in Figs I 0 . 12 (a) and (b) respectively, while a combination
dual r.p.m. and synchroscope presentation is shown at (c).
The operation is based on the principle of the induction motor,
which, for this application, consists of a three-phase star-wound laminated stator and a three-phase star-wound laminated rotor pivoted in
jewelled bearings within the stator. The stator phases are connected
to the tachometer generator of the slave engine while the rotor phases
are connected to the master engine generator via slip rings and wire
Figure JO.I 2 Dial presentations of sy nchroscopes.
(a) Twin-engine; (b) four-
engine; (c) combined dual
tachometer and synchroscope.
brushes. A disc at the front end of the rotor shaft provides for
balancing of the rotor. The pointer, which is double-ended to symbolize a propeller, is attached to the front end of the rotor shaft and
can be rotated over a dial marked INCREASE at its left-hand side
and DECREASE at its right-hand side.
On some synchroscopes the left-hand and right-hand sides may be
marked SLOW and FAST respectively. Synchroscopes designed for
use in four-engined aircraft employ three separate induction motors,
the rotor of each being connected to the master engine tachometer
generator while each stator is connected to one of the three other
generators.
Operation
To understand the operation of a synchroscope let us consider the
installation of a typical twin-engined aircraft tachometer system, the
circuit of which is given in Fig l 0.13. Furthermore, let us assume
that the master engine, and this is usually the No I, has been adjusted
to the required 'on-speed' condition and that the slave engine has
been brought into synchronism with it.
Now, both generators are producing a three-phase alternating
current for the operation of their respective indicators, and this is
253
RESULTANT FIELDS
GENERATOR No.1
SYNCHROSCOPE
\
~
DIRECTlON OF ROTOR ROTATION
DUE TO REACll\/E TORQUE
~~
<-~~·
:::-t \ ,.,..,_" """' "'"'
.J-
RESULTANT OF
ROTOR FIELD
Figure I 0. 13 Operation of a
synchroscope.
254
DIRECTION OF ROTOR
ROTAT10N DUE TO STATOR
TORQUE
( f.Xl~. ')
LAGGING
I</
GENERATOR No.2
/
b
•
N
I ~X\,
\ji~)
11-b
RESULTANT OP
STATOR AELD
l.EAOING
cf
RESULTANT
OF ROTOR
FIELD
also being fed to the synchroscope, generator No I feeding the rotor
and No 2 the stator. Thus, a magnetic field is set up in the rotor and
stator, each field rotating at a frequency proportional to its corresponding generator frequency, and for the phase rotation of the system
rotating in the same direction. For the conditions assumed, and
because generator frequencies are proportional to speed, it is clear
that the frequency of the synchroscope stator field is the same as that
of the rotor field. This means that both fields reach their maximum
strength at the same instant; the torques due to these fields are in
balance, and the attraction between opposite poles keeps the rotor
'locked' in some stationary position, thus indicating synchronism
between engine speeds.
Consider now the effect of the slave engine running slower than
the master. The frequency of the slave engine generator will be lower
than the master engine generator, and consequently the stator field
will be lagging behind the rotor field; in other words, reaching its
maximum strength at a later instant at, say, point a in Fig I 0.13 .
The rotor, in being magnetized faster than the stator, tries to rotate
the stator and bring the stator field into alignment, but the stator is a
fixed unit; therefore, a reactive torque is set up by the interaction of
the greater rotor torque with the stator. This torque causes the rotor
to turn in a direction opposite to that of its field so that it is forced
to continually realign itself with the lagging stator field. The
continuous rotation of the rotor drives the propeller-shaped pointer
round to indicate that the slave engine is running SLOW and that an
INCREASE of speed is required to bring it into synchronism with
the master engine.
If the slave engine should run faster than the master then the
synchroscope stator field would lead the rotor field , reaching maximum strength at, say, point b. The stator field would then produce
the greater torque, which would drive the rotor to realign itself with
the leading stator field, the pointer indicating that the slave engine is
running FAST and that a DECREASE of speed is required to
synchronize it.
As the speed of the slave engine is brought into synchronism once
again, the generator frequency is changed so that a balance between
fields and torques is once more restored and the synchroscope rotor
and pointer take up a stationary position.
From the foregoing description we see that a synchroscope is, in
reality, a frequency meter, its action being due only to the relative
frequencies of two or more generators. The generator voltages play
no part in synchroscope action except to determine the operating
range above and below synchronism.
Rotation Indicators
In some aircraft using by-pass turbine engines, indicators are
provided to indicate that the shaft of the low-pressure compressors
commen.c es to rotate during the starting cycle, and that it is safe to
continue the cycle.
The basis of an indicator is a two-stage magnetic amplifier
operating from a 115 V 400 Hz a.c. supply and connected to one
phase of a normal tachometer generator. The signals from the generator are fed into the amplifier as a reference input in revolutions per
minute. An indicator lamp mounted on the main instrument panel,
or flight engineer's panel, is connected to the amplifier output stage.
When the speed voltage input signal reaches a critical level , usually
6 mV corresponding to a rotation speed of a fraction of I rev./min.,
sufficient output current is produced to light the indicator lamp.
The speed is reached in the first few degrees of rotation; therefore
the lamp provides an immediate indication that the low-pressure
shaft has rotated. Signals in excess of the critical cause the amplifier
to saturate and the lamp to remain alight but without being overloaded.
The power supply is fed to the amplifier via an engine-starting
255
circuit and is isolated once the starting cycle is satisfactorily concluded. In multi-engine installations, a single amplifier and lamp serve
to indicate rotation of each engine, being automatically selected
during each starting-cycle.
Questions
256
10.1
Describe a mechanical method of providing indication of engine
speed.
10.2 Explain how rotation of an alternating-current tachometer indicator
motor is transmitted to the indicating element.
10.3 With the aid of a circuit diagram explain the operation of a synchroscope system as fitted to a multi-engined aircraft.
10.4 (a) How is the necessary speed reduction between engine and tachometer generator obtained when direct-drive generators are
employed?
(b) Why are direct-drive generators essential for the measurement ot
turbine engine speeds?
10.5 Describe a method of compensating for the effects of temperature
on a synchronous motor type of tachometer indicator.
10.6 What is the purpose of a tacho probe?
10.7 Briefly de~cribe the construction and operation of a tacho probe.
10.8 With the aid of a diagram explain the operation of the indicator used
in conjunction with a tacho probe.
10.9 How is failure of tacho probe signals to an indicator presented?
10.10 Describe the operation of a servo-operated type of tachometer
indicator.
1 1 Measurement of
temperature
Methods and
Applications
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:
I. Most substances expand as their temperature rises; thus, a measure
of temperature is obtainable by taking equal amounts of expansion
to indicate equal increments of temperature.
2. Many liquids, when subjected to a temperature rise, 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 the pressure of the
vapour.
3. Substances change their electrical resistance when subjected to
varying temperatures, so that a measure of temperatures is obtainable by taking equal increments of resistance to indicate equal
increments of temperature.
4. Dissimilar metals when joined at their .e nds produce an electromotive force (thermo-e.m.f) dependent on the difference in temperature between the junctions. Since equal increments of temperature are only required at one junction, a measure of the electromotive force 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 what 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 .
The utilization of these various methods provides us with a very
convenient means o f classifying temperature-measuring instruments :
(a) expansion type (liquid or solid), (b) vapour-pressure type ,
(c) electrical type (resistance or thermo-electric) and (d) radia tion type.
The majority of instruments currently in use are, however, of the
resistance and thermoelectric type and are applied to the measurement of the temperature of such liquids and gases as fuel, engine
lubricating oil, outside air, carburettor air, and turbine exhaust gas.
In certain types of turbojet engine, the radiation pyrometry technique
is also applied to the measurement of actual turbine blade temperature.
257
Heat and Temperature
Heat
Heat is a form of energy possessed by a substance, and is associated
with the motion of the molecules of that substance. The hotter the
substance the more vigorous is the vibration and motion of its molecules. We may regard heat, therefore, as molecular energy.
The quantity of heat a substance contains is dependent upon its
temperature, mass and nature of the material from which the body is
made. A bucketful of warm water will melt more ice than a cupful _of
boiling water; the former must therefore contain a greater quantity of
heat even though it is at a lower temperature.
The transfer of heat from one substance to another may take place
by conduction, convection and radiation. Conduction requires a
material medium, which may be solid, liquid or gaseous. Hot and colc
substances in contact interchange heat by conduction. Convection is
the transmission of heat from one place to another by circulating
currents and can only occur in liquids and gases. Radiation is the
energy emitted by all substances, whether solid, liquid or gaseous.
Temperature
Temperature is a measure of the 'hotness' or 'coldness' of a substance
or the quality of heat. Therefore, in the strictest sense of the term,
temperature cannot be measured. The temperatures of substances
can only be compared with each other and the differences observed,
and so the practical measurement of temperature is really the
comparison of temperature differences. In order to make such a
comparison, the selection of a standard temperature difference, a
fundamental interval and an instrument to compare other temperature differences with this, are necessary.
Melting Point and Boiling Point
For pure substances, the change of state from solid to liquid and
from liquid to vapour takes place at temperatures which, under the
same pressure conditions, can always be reproduced. Thus, there
are two equilibrium temperatures known as (i) melting point, the
temperature at which solid and liquid can exist toge ther in equilibriun
and (ii) boiling point, the temperature at which liquid and vapour can
exist together in equilibrium.
Fundamental Interval and Fixed Points
The fundamental interval is the temperature interval or range between
two fixed points: the ice point at which equilibrium exists between
ice and vapour-saturated air at a pressure of 760 mm Hg, and the stean
point, at which equilibrium exists between liquid water and its
vapour; the water boiling also under a pressure of 7 60 mm Hg.
258
The term fundamental interval is used in resistance thermometry;
it refers to the increase in resistance of the temperature-sensing
element between the fixed points.
Scales of Temperature
The fundamental interval is divided into a number of equal parts or
degrees, the division being in accordance with two scale notations
Celsius (centigrade) and Fahrenheit.
On the Celsius scale, the fundamental interval is divided into I 00
degrees, the ice point being taken as 0° C and the steam point as
100°C. One Celsius degree is thus I/ 100 of the fundamental interval.
On the Fahrenheit scale, the fundamental interval is divided into
180 degrees, the ice point and steam point in this case being taken
as 32°F and 212°F respectively. One Fahrenheit degree is thus 1/180
of the fundamental interval.
If all the heat were removed from a body its temperature would be
as low as possible. This temperature is -273, J 5° C, or 273-15° C below freezing point, and is known as absolute zero or O on the Kelvin*
scale:
0 K = -273.l5°C
0°C = 273.15 K.
On the Fahrenheit scale, absolute zero is -459 .67° F, or 0° R on
the Rankine scale:
0°R = -459.67°F = 0 K = -273.15°C
32°F = 491.67° R = 273.15 K = 0°C
Conversion Factors
Since I 00 divisions on the Celsius scale correspond to 180 divisions
on the Fahrenheit scale,
I°C= 180/100= l.8°F and l°F = 1/ 1.8 = 0.55 °C
Expressed as fractions,
l°C=9/5°Fand 1°F=5/9°C
The fact that the zero points on the two scales do not coincide
makes conversion from on~ scale to another more difficult than just
dividing by the conversion factors, 1.8 and 0.55 . For example, if we
wish to convert I 0° C to° F and simply multiply it by 1.8 or its
equivalent 9/5 , we shall obtain J8°F, but this is only 18° above
freezing point and the value required must be with reference to the
zero on the Fahrenheit scale. We therefore add 32° to obtain +50° F.
Thus the formula for converting ° C to ° F is
or
•
° F={°C X 9/ 5)+32
The name 'degree Kelvin' (°K) was replaced by 'kelvin' (K) in 1967.
259
To convert° F to ° C is simply the reverse procedure. From 50° F
we subtract 32 to determine the number of Fahrenheit degrees above
freezing point (18°) and divide by 1.8 or 9/5. Thus, the formula for
converting ° F to ° C is
°C
= {°F
- 32)/1.8
or
° C = (°F -· 32) X 5/9
When converting values on the minus side of scales care must be
taken to observe the signs. For example, in converting - I 0° C to ° F
we obtain ( - 18) + (+32) giving us a difference of+l4, the equivalent
of-I0°Cin°F.
To convert °C and °F to degrees absolute we simply add 273 or
460 respectively.
Electrical Temperature
Indicating Systems
As we have already noted, these systems fall into two main categories:
variable-resistance and thermoelectric: the methods are termed
resistance thermometry and pyrometry respectively. In b.oth cases,
an understanding of the principles involved requires a knowledge of
certain fundamental electrical laws and their applications.
Ohm's Law
This law may be stated as follows: When a current flows in a conduc-
tor, the difference in potential between the ends of such conductor,
divided by the current flowing through it, is a constant provided
there is no change in the physical condition of the conductor.
The constant is called the resistance (R) of the conductor and is
measured in ohms (n ) . In symbols,
V/I = R
(I)
whe~e Vis the potential difference in volts (V), and I the current in
amperes (A).
Calcu lations involving most conductors are easily solved by this
law, for if any two of the three q uan ti ties ( V, I and R) are known,
the third can always be found. Thus, from eqn ( 1),
IR= V
(2)
V/R = I.
(3)
A characteristic of the majority of metallic conductors is that their
resistance increases when subjected to increases of temperature.
Resistances in Series
In all electrical instruments and associated circuits, we always find
certain sections in which conductors are joined end to end . For
example, in a thermo-electric temperature-indicating system, some
260
Figure I 1. 1 Series circuit.
ADJUSTING RESISTOR
(a) Practical circuit;
(b) equivalent circuit Rr =
Rrh + RJ + Rv + R;.
LEADS
__J
INDICATOR
(a)
'/,Rt
(b)
of the essential conductors are joined as shown in Fig 11.1. A circuit
formed in this manner is known as a series circuit and its combined or
total resistance (Rr) is equal to the sum of the individual resistances.
Thus
Rr =Rrh +R1+Rv +R;.
From a direct application of Ohm's law the current and the
potential difference may also be derived :
Ir::;; Vr/Rr
this being the current at any point in the circuit. The total voltage
(VT) is equal to the sum of the voltage drops across the individual
conductors and is always equal to the applied voltage, which in this
case is the thermo-e.m.f. Thus
Vr::;; !Rr1i + IR1 + IRv +IR;= IRr.
Example. In a typical thermoelectric temperature-indicating circuit,
the following resistance values apply: R,1, = 0.79 n , R 1 = 24.87 n ,
Rv = 7 n , and R; ::;; 23 n, so that
Rr = 0.79 + 24 .87 + 7 + 23 = 55.66 n.
The total voltage VT is governed by the temperature at the thermocouple.; therefore VT and the total current Ir flowing through the
indicator are variables. Assuming the temperature to be 500° C,
the voltage generated by a typical the,rmocouple is 20.64 m V. Thus
261
h = Vr = 20 ·64 = 0.37 mA very nearly.
Rr
55.66
Resistances in Parallel
When two or more conductors are connected so tha t the same voltage
is applied across each of them, they are said to form a parallel circuit.
The total resistance of such a circuit may be obtained by applying
the following rnle. The reciprocal of the total resistance (Rr) is
equal to the sum of the reciprocals of the individual conductor
resistances.
Figure 11.2 shows a parallel circuit made up of six thermocouples,
an arrangement which is employed for the measurement of turbineengine exhaust gases.
If the total resistance of the thermocouple section of the circuit is
represented by RTH, then
..L.
= _l_
RTH
R1111
+ _I_ +
R11i2
+ _j_
· · ·
Rch6 ·
The resistances of the leads from the thermocouple junction box
to the instrument terminals and the instrument resistance are
represented by R1 and R; respectively, and the circuit reduces to the
simple series circuit shown. The total resistance RT of the complete
circuit can therefore be calculated from
RT=Rrn+R1+R; .
Assuming that all the thermo-e .m.f.'s .are equal, the voltage acting
in the circuit will be equal to the thermo-e.m .f, V 111 , of a single
thermocouple, so that the current will be given by
Figure 11.2 Parallel circuit.
R;
-I
262
I= V,1z.
RT
Example. The resistance of each of the thermocouples and its leads,
shown in Fig 11.2, is 0.79 n; therefore, the total resistance RTH of
this part of the circuit is calculated as follows:
I
I
-=-X6
RTH R,1z
so that
R,h_0.79_01317
R TH=6D.
6 - .
The leads resistance R 1 is 24 .87 n and the instrument resistance R;
is 23 n, so that the total circuit resistance is
RT= RTH
+ R1 + R;
= 0.1317 + 24.87 + 23 = 48
n
very nearly.
Assuming that the thermocouples generate a voltage V,h of 20.64
m V ( at 500° C) the total current will be
! = ~'; =
20 4
4:
= 0.43 mA.
When only two conductors are connected in parallel, the total
resistance is given by
R1 _+ R2
- I= - I + - 1 = __,_
_....
RT R1
R2
R. R2
so that
_
R T-
R1R2
R1 + R2
Example. Two conductors having resistances of 14 n and 10
joined in parallel. Their combined resistance is therefore
RT
= 14 X l O = I 40 = S 833
14 + 10
24
.
n
are
n
The important point to note about parallel circuits is that the
total resistance is less than the resistance of any one of the individual
conductors .
Resistances in Series-parallel
A circuit made up of resistances in seri~s-parallel is sho wn in Fig l l .3
263
Figure 11.3 Series-parallel
circuit.
SHUNT RESISTOR
Rsh
+
',
SERIES RESISTOR
MOVING COIL
Rmc
THERMISTOR
R,
EQUIVALENT SERIES CIRCUIT
and is one based on an indicator used in conjunction with the thermo·
couple circuit of Fig 11. 2. The solving of such a series-parallel cir·
cuit is done by the methods already illustrated and by (i) reducing
each parallel group to an equivalent single resistance, (ii) solving the
resulting series circuit for total resistance (RT) and total current Ur),
(iii) obtaining the voltage drop across each parallel group, (iv) solving
for all other quantities.
Example. The resistances of the four conductors in a typical indi-
cator areRmc = 15.5 D.;Rse = 2 n;Rsh = 14 n andR 1 = 10
Rm c and Rse form the series part of the circuit, thus
R, =Rmc +Rse
n.
= 15.5 + 2 = 17.5 n.
The parallel part of the circuit is formed by Rsh and Rr , thus
R 2 = RshRr =14X 10
Rsh
+Rr
14+ 10
140=
24 5833
·
n
·
Reducing the circuit to an equivalent series circuit, the total resistance Rr is
Rr
= R,
+ R2
= 17.S + 5.833 = 23.333 n.
Factors Governing Resistance
The resistance of a single conductor is governed by three main factors:
(i) the material of which it is composed, (ii) its dimensions (length
and cross-:;P.ctional area), and (iii) its physical state, particularly its
temperature.
The resistivity p of a conductor is the resistance of a sample of
the material having unit length and unit cross-sectional area. At a
given temperature, it is therefore a constant for a given material and
is usually expressed in microhms per centimetre cube.
At a given temperature, the resistance R of a material is directly
proportional to the length / in centimetres, and inversely propor-·
tional to the cross-sectional area a in square centimetres. Thus
264
R
I
a
=p-.
Resistance and Temperature
The resistance of a conductor is dependent on temperature, the
effect of which is to change the dimensions by thermal expansion
and to change the resistivity. The first effect is comparatively
small, the main changes of resistance being due to changes of p.
In the case of pure metallic conductors, resistance increases with
increase in temperature, and this is the basis of t emperature
measurement in resistance thermometry.
The two metals most commonly used in aircraft resistance
thermometry are nickel and platinum, both of which are manufactured to a high degree of purity and reproducibility of resistance
characteristics.
The approximate resistance of a metal at a temperature t is given
by the linear equation
R, = R0 (1+ at)
(I)
where R 1 is the resistance at temperature t°C, R 0 the resistance at 0°C,
and a is the temperature coefficient of resistance.
In evolving temperature/resistance calibration laws, however, the
above simple linear relation cannot be applied because the temperature coefficients of nickel and platinum can only be regarded as
constant over the temperature range 0-100° C. Therefore, in order
to obtain greater accuracy and an extension of the temperature
range to be measured , it is necessary to introduce another constant
(13) into eqn (I) so that the resistance is represented by the quadratic
equation
R 1 = R 0 (I + 0tt + (3t 2 ).
(2)
The number of constants involved depends on the temperature
range, but in general, two are sufficient up to about 600°C.
The values of the constants which conform to typical standard
specification requirements for the calibration of resistance elements
and associated indicators, are as follows :
Nickel law:
Platinum law:
0t =
0t =
0.438 (3 = 0.006
0.0039583 (3 = 0.000000583.
Resistance values over appropriate temperature ranges are given in
the tables on page 402 .
Wheatstone Bridge
The most common method of measuring resistance is by means of
the well-known Wheatstone bridge network shown in Fig 11.4 (a ).
265
The circuit is made up of four resistances arms, R 1 , R2 , R 3 and
Rx . A moving-coil or moving-spot galvanometer is connected
across points B and D, and a source of low voltage is connected
across points A and C. Current flows in the directions indicated by
the arrows, dividing at point A and flowing through R 3 and Rx, at
strengths which we may designate respectively as / 1 and / 2 • At point
C the currents reunite and flow back to the voltage source.
Let us assume that the resistance of the four arms of the bridge
are so adjusted that B and D are at the same potential; then no .
current will flow through the galvanometer and so it will read zero.
Under these conditions the bridge is said to be 'balanced'. This
can be shown by applying Ohm's law; if the potential difference
between A and B or A and Dis, say, V 1 , and the potential difference
between B and C or D and C is V2 , then
(3)
and
V2 =I1R2 = I2R1.
(4)
If now eqn (3) is divided by eqn (4), we have
V1 R3 Rx
V2 =R2 =If;
(5)
from which
3 _
R X :R1R
R2
(6)
Hence, an unknown resistance can be calculated by adjusting the
values of the three others until no current flows through the galvanometer, as indicated by no movement of its pointer or spot.
It will be apparent that, if the resistor Rx is subjected to varying
temperatures and its corresponding resistances are determined, then
it is feasible for the network to serve as a simple electrical-resistance
thermometer system.
Obviously some rearrangement of the circuit is necessary in order
to obtain automatic response to the variations in the temperature/
resistance relation; the manner in which this can be effected is
illustrated in Fig 11 .4 (b ). The unknown resistance Rx forms the
temperature-sensing element and is contained within a metal
protective sheath, the assembly being called a bulb. The three other
resistances instead of being adjustable, are fixed and are contained
within the case of a moving-coil indicating element calibrated in
units of temperature. Both components are suitably interconnected
and supplied with direct current.
When the bulb is subjected to temperature variations, there will be
266
B
Figure J 1.4 Wheatstone
bridge network. (a) Basic
circuit; (b) application to
temperature measurement.
A
C
D
(a)
--+----'
J
INDICAT IN~ ELEMENT .
~ 1-- . - - . ---'----- - -- --f-+----"
(bl
TEMPERATURE - SENSING
ELEMENT
a corresponding variation in R x. This upsets the balance of the
indicator circuit, and the value of Rx at any particular temperature
will govern the amount of current flowing through the moving coil.
Thus, for a given value of Rx, the out-of-balance current is a measure
of the prevailing temperature.
267
It will be noted that there is .an important difference between the
two applications of this bridge network. In one the measurement of
a resistance is dependent upon the circuit current being in balance,
while in the other it is measured in terms of out-of-balance current.
There is in fact only one point in a bridge type of temperature
indicator at which the circuit is balanced and at which no current
will flow through the moving coil ; this is known as the null point.
It is usually indicated on the scale of the instrument by means of a
small datum mark and the pointer takes up this position when the
power supply is disconnected.
For accurate temperature measurement, however, this form of
indicating circuit has the disadvantage that the out-of-balance
current also depends on the voltage of the power supply. Hence,
errors in indicated readings will occur if the voltage differs from
that for which the instrument was initially calibrated. Bridge-type
thermometers have therefore been largely superseded by those operating on what is called the ratiometer principle described on page 276.
Temperature-Sensing
Elements
The general arrangement of a temperature-sensing element commonly
used for the measurement of liquid temperatures is shown in Fig 11.5.
The resistance coil is wound on an insulated former and the ends of
the coil are connected to a two-pin socket via contact strips. The
bulb, which serves to protect and seal the elemei;:it, may either be a
brass or stainless-steel tube closed at one end and soldered to a union
nut at the other. The union nut is used for securing the complete
element in the pipeline or component of the system whose liquid
temperature is required . The two-pin socket is made a tight fit inside
the male portion of the union nut, the receptacle of which ensures
correct location of the socket's mating plug.
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.
Figure 11.5 Schematic of a
UNION NUT
typical temperature-sensing
element.
PLUG
·· - ePTACLE
CALIBRATING OR
BALANCING COIL
268
A calibrating or balancing coil is normally provided so that a
standard constant temperature/ resistance characteristic can be
obtained, thus permitting interchangeability of sensing elements. In
addition the coil compensates for any slight change in the physical
characteristics of the element. The coil, which may be made from
Manganin or Eureka, is connected in series with the sensing element
and is adjusted by the manufacturer during initial calibration.
Air Temperature Sensors
Air temperature is one 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 overall would be the most ideal is that of air under pure
static conditions at the various flight levels compatible with the
operating range of the particular aircraft concerned. The measurement of static air temperature (SAT) by direct means is, however,
not possible for all types of aircraft or, in many instances, for one
type of aircraft, for the reason that measurements can be affected
by the adiabatic compression of air resulting from increases in airspeed. In general, the boundary layer at the outside surfaces of an
aircraft flying at speeds below 0.2 Mach number is very close to the
SAT. Howev~r, at higher Mach numbers the boundary layer can be
slowed down or stopped relative to the aircraft, and thereby produce
adiabatic compression which will raise the air temperature to a value
appreciably higher than SAT. Friction of high speed flow along the
aircraft surfaces will also raise the air temperature. This increase is
commonly referred to as ' ram rise', and the temperature indicated
under such conditions is known as ram air temperature (RAT) i.e.
SAT plus the ram rise.
The ram rise due to full adiabatic compression is always precalculated mathematically as a function of Mach number, and for
each type of aircraft values are presented in either tabular or graphical form in the operating manual or the flight manual for the type.
Thus, for air temperature sensors subjected to ram rise, the RAT
readings of the associated indicators can always be correct ed to
obtain SAT, either by direct subtraction of tabulated ram rise
values, conversion ch arts, or in the case of air data computers (see
page 106) by the automatic application of a correction signal. The
proportion of ram rise is dependent on the ability of the sensor to
sense or recover the temperature rise, the sensitivity in this case
being expressed as a percentage and termed the recovery factor. If,
for example, a sensor has a recovery factor of 0.80, it will measure
SAT plus 80% of the ram rise.
269
Various types of air temperature sensors may be adopted dependent on whether indications of SAT or RAT are required. The
simplest type, and one which is used in a few types of small private
aircraft for indicating SAT, is a direct-reading thermometer which
operates on the principle of expansion and contraction of a bimetallic
element when subjected to temperature changes. The element is in
the form of a helix anchored at one end of a metal sheath or probe;
the free end of the helix is attached to the spindle of a pointer. As
the helix expands or contracts, the helix winds or unwinds causing
the pointer to rotate against the scale of a dial fixed to the sheath
opposite to the fixed end of the helix . The thermometer is secured
through a fixing hole on one of the side windows of the cockpit, or
in the wrap-around portion of a windscreen, so that the probe protrudes into the airstream.
The majority of sensors are, however, of the platinum resistance
wire element type, the element being contained either in .a probe
similar to that adopted for the temperature measurement of liquids,
or mounted in what is termed a 'flush bulb' configuration. A probe
element is generally used for SAT measurement, the probe protruding
into the airstream through a fixing hole in the side of a fuselage. A
'flush bulb' (see Fig 11.6) is mounted flush with the aircraft skin and
senses RAT. The recovery factor generally varies from 0.75 to 0.90
depending on the geometry of the aircraft and the location of the
bulb.
Figure 11.6 RAT sensor
(' flush bulb').
For use at high Mach numbers, it is customary to sense and
measure the maximum temperature rise which is possible. This is
referred to as total air temperature (TAT) and is obtained when the
air is brought to rest (or nearly so) without addition or removal of
heat. For this purpose, probes of the type shown in Fig 11. 7 were
introduced, and are to be found on a number of present-day public
transport aircraft. They have several advantages over 'flush bulbs'
notably an almost negligible time lag, and a high recovery factor
270
Figt1re 11. 7 Total air
temp~rature probe.
AIRCRAFT FUSELAGE SKIN
SENSING ELEMENT
AIRCRAFT SKIN
_.. 5-POLE CONNECTOR
(approximately 1.00). The probe is normally connected to an
indicator on the flight deck instrument panel and to a Mach number
module of a central air data computer (see page 112).
The probe is in the form of a small strut and air intake made of
nickel-plated beryllium copper which gives good thermal conductivity
and strength. It is secured to the aircraft skin, at a pre-determined
location in the fuselage nose section, and outside of any boundary
layer which may exist. In flight, the air pressure within the probe
is higher than that outside, and the air flows in the manner indicated ;
separation of water particles from the air is effected by the air flow
being caused to turn through a right-angle before passing round the
sensing element. The bleed holes in the intake casing permit boundary layer air to be drawn off under the influence of the pressure
differential across the casing.
A pure platinum wire resistance element is used and is hermetically
271
sealed withi"n two concentric platinum tubes. The inner platinum
tube is used as the element fonner, thereby ensuring a close match
of thermal expansion and minimizing of thermal strain. An axial
wire heating element, supplied with 115 V a.c. at 400 Hz, is mounted
integral with the probe to prevent ice formation, and is of the selfcompensating type in that as the temperature rises so does the
element resistance rise, thereby reducing the heater current. The
heater dissipates a nominal 260 Wunder 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 being 0.9°C at 0.1 Mach decreasing to 0.15 at Mach
1.0.
Electrical Temperature
Indicators
The measuring elements employed in the indicators of resistance-type
and a majority of thermo-electric temperature-indicating systems,
depend for their operation on the fact that electric current flowing
through a conductor P!Oduces a magnetic field in and around the
conductor.
In order to utilize this effect as a method of measurement, it is
necessary to have a free-moving conduc_tor in a magnetic field which
is both permanent and of uniform strength. In this manner, as will
be shown, advantage can be taken of the interaction between the
two magnetic fields and the resultant forces to move the conductor
and its indicating element to definite positions.
Conductor Carrying Current in a Magnetic Field
When a current-carrying conductor is placed in a magnetic field , the
interaction of the field produced by the current and the field in
which the conductor is located exerts a force upon the conductor.
This force is in direct proportion to the flux density, the current and
the length of the conductor.
Figure 11.8 Magnetic field
around a straight conductor.
s
N
FORCE
F
272
In Fig I 1.8 the field set up around the conductor is shown as being
in the same direction as the main field above the conduc to r and in the
reverse direction to the main field below the conductor. Thus, the
interaction is similar to that of like and unlike poles of two bar magnets, i.e. lines of magnetic force flowing·in the same direction attract
each other while those flowing in opposite directions repel each other.
The lines of force of the main field thus become distorted so that
it is stronger on one side of the conductor than on the other, the net
effect being that the conductor is subjected to a force in the direction
of the weaker field. If the conductor is free to move in the main field,
as is the case with electric motors and moving-coil instruments, then
it is clear that the motion will be in the same direction as the force.
The force is also dependent upon the angle which the conductor
makes wi~h the main magnetic field, being maximum when it is at
right angles to the field.
The Moving-Coil Indicator
Consideration thus far has been given to a straight current-carrying
conductor fixed in a permanent magnetic field. If the conductor is
formed into a single coil and pivoted at a pGint Pas shown in Fig
11. 9, then, for the direction of current indicated, forces F,F will be
exerted on each side of the coil, each producing a torque Fr, so that ·
the total torque will be 2Fr and will cause the coil to rotate in a
clockwise direction. This is the basis of any moving-coil indicator.
F
Figure 11.9 Magnetic field
around a coiled conductor.
s
N
F
In the practical case, however, it is necessary -to intensify the
forces acth ,g on the coil in order to obtain reaso nably large
deflections of the coil and its indicating element for small values of
current. This is accomplished by placing a soft-iron core between
the pole pieces of the permanent magnet, and also by incre:ising the
number of turns of the coil.
273
Soft-iron has a lower reluctance to lines of force than air; therefore, when it is placed in a uniform magnetic field as indicated in
Fig I I . I 0, the lines of force from its surroundings are drawn into
the iron and it becomes magnetized by induction.
Figure 1J.10 Effect of soft·
iron in a magnetic field.
Referring to Fig 11 . 11 we see that by shaping the pole pieces and
soft-iron core cylindrically, a radial magnetic field of even greater
intensity and uniformity is obtained in the narrow air gap in which
the sides of the coil rotate.
Figure 11.J 1 Effect of a
CORE
cylindrical soft-iron core.
NON-MAGNETIC
MATERIAL
The constructional details of a typical moving-coil indicator are
shown in Fig 11.12. The permanent magnet is made from a special
alloy possessing high remanence and coercive force characteristics;
and may be either circular or rectangular in shape, and machined to
size within very close tolerances. An adjustable shunt is secured
across the pole pieces to vary the field in the air gap during calibratio,
of the indicator.
The cylindrical soft-iron core and the moving coil assembly are
usually built up into a single unit which can be positioned in the
magnet air gap. In the example shown, the unit is secured by screwing a bridge piece to the magnet pole pieces.
The coil consists of a number of turns of fine copper wire wound
on a rectangular aluminium former or frame , pivoted in jewelled
bearings so that when current flows through the coil the combination
of magnetic fields set up around each turn will increase the force
required to deflect the coil.
Current is supplied to the moving coil via two flat-coil hairsprings
the main function of which is to ensure that currents of varying
274
Fi!(ttre J J. J2 Construct ion of
a typical moving-co il indicato r.
POIN TER
HAIRSPRING
ADJUSTING
DEVICE
magnitudes shall produce proportionate deflections of the coil. In
other words, they form the controlling system, an essential part of
any moving-coil instrument without which the coil would move to
its maximum deflected position regardless of the magnitude of the
current, and moreover would not return to its zero position on
cessation o f th e current flow. The hairsprings are made of materials
having low resistance and low tempera ture coefficients, phosphorbronze and beryllium-bronze being most commonly used . The
effects of extreme temperature variations are minimized by coiling
the springs in opposite directions so that they act one against the
other.
The setting of the moving coil and pointer to zero is carried out by
means of an adjusting device similar to that described in Chapter 2
(page 14).
As in all indicating instruments, mechanical balance of the moving
system is necessary to ensu re uniform and symmetrical wear o n pivots
and bearings, so preventing out-of-balance forces from causing errors
in indication . This is usually effected by attaching balance arms on
the supporting spindle and providing them wi th either adjustable
balance weights or wire coiled round and soldered to the arms.
A further essen tial requirement of any moving-coil instrument is
that the coil should take up its deflected position without oscillation
and overshoot , i.e. it must be damped and rendered dead beat in its
indications. In this type of instrument damping is effected automatically by the aluminium former on which the coil is wound. As the
former moves in the air gap it cuts the magnetic flux, thus setting up
eddy currents within itself to produce a force opposing that causing
movement of the coil. Since the force is proportional to the velocity
275
of the moving coil, the latter will be retarded so as to take up its
deflected position without overshoot.
The complete instrument movement is enclosed in a soft-iron case
to screen it against the effects of external magnetic fields.
Ratiometer System
A ratiometer-type temperature-indicating system consists of a sensing
element and a moving-coil indicator, which unlike the conventional
type has two coils moving together in a pennanent-magnet field of
non-unifonn strength. The coil arrangements and methods of
obtaining the non-uniform field depends on the manufacturer's
design, but three methods at present in use are shown in Fig 11.13.
Figure 11.14 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 R 2 and the temperature- ·
sensing element Rx, Both arms are supplied with direct current from
the aircraft's main power source, but the coils are so wound that
current flows through them in opposite directions (see also Fig 11 .13)
As in any moving-coil indicator, rotation of the measuring element is
Figure 11.13 Methods of
produced by forces which are proportional to the product 9f t.he
obtaining a non-uniform
current
and field strength, and the direction of rotation depends on
magnetic field. {a) Parallel coll;
(b) cross coll; (c) twin former. the direction of current relative to the magnetic field. In a ratiometer,
FORMER
POLE
Pl~CE
MAGNETIC
FIELO
CORE PIECE
/
~ g f DMAGNETIC
lG
FIELO
MIN
·•
'
MAGNET
,
'
MOVING
COIL
le)
276
INOICATOR
POINTER
IRO N CORE PIECE
(FIXED)
(bl
MAGNETIC
POLE PI ECE
MAX
FIELO
] STR ENGTH
MIN
WINDING 'A'
\
WINDING 'B'
I
\
MAX[ ;
MAGNETIC
FIElO
STRENGTH
MIN
,
•
\
WI NDING 'B'
CORE PIECE
tel
Figure 11.14 Basic ratiometer
circuit.
+
A
-----~
B
INDICATOR
TEMPERATURE
SENSING EUMENT
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 furthermore, as the magnetic
field is of non-uniform 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 11 . 14 employs an instrument which utilizes the crossed-coil
winding method shown in Fig 11.13 (b ), and that winding B is in the
variable-resistance ann, and winding A is in the fixed-resistance arm.
The resistances of the arms are so chosen that at the zero position of
the instrument 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 ~trength of the field in
which the windings are positioned.
When the temperature at the sensing element Rx increases, then
in accordance with the temperature/resistance relationship of the
material ·used for the element, 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
277
altered and the force in winding A will rotate the measuring element
so that both windings are carried round the air gap; winding Bis
advanced further into the stronger part of the magneti~ field while
winding A is being advanced to a weaker part. When t~e temperature
at the sensing element stabilizes at its new Value the fotces pro_duced _.
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
made up of hairsprings which exert a controlling torque proportional
to the current flowing through the coil. Therefore, if the current
decreas_es 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 system, however, does not require hairsprings for
exerting a controlling torque, this being 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 coils remains
the same and tendencies for them to move to positiOIJS of differing
field strength are counterbalanced .
In practical applications of the ratiometer system, a spring is, in
fact, used and so at first sight this may appear to defeat the whole
object of the ratiometer principle. It is, however, essential that the
moving-coil former and pointer should take up an off~scale position
when the power supply is disconnected, and this is the sole function
of the spring. Since it exerts a very much lower torque than a
conventional moving-coil indicator control spring, its effects on the
ratiometer controlling system and indication accuracy, under power
supply changes normally encountered, are very slight.
The power of the output signals from resistance type temperaturesensing elements is very limited, so that the moving-coil mechanisms
of indicators need to be of delicate construction in order to provide
the necessary accuracy and response. For some applications, however, it may be required for indicators to operate in extreme environmental conditions which call for a more robust form of mechanism.
Indicators employing a powered moving coil were therefore
introduced and the dial presentation of one such indicator serving
a dual function is illustrated in Fig 11 .15. The low power output
signal from the temperature-sensing element is supplied to discrete
278
Figure /I. I 5 Powered
moving-coil indicator.
"'/
WARNING LIGHTS.
signal processing circuits within the indicator, which amplify the output to position the moving coil of the lower indicator in the
conventional manner.
Thermoelectric
Thermometers
Thermoelectric thermometers 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 rotors and
blading. The systems basically consist of a thermocouple sensing
element which, depending on _the application, is secured to an engine
cylinder head or exposed to turbine exhaust gases, and a movingcoil indicator connected to the sensing element by special leads.
Thermocouple Principle
Thermoelectric temperature-measuring instruments depend for their
operation on electrical energy which is produced by the direct
conversion of heat energy at the measuring source. Thus, unlike
resistance thermometers, they are independent of any external
electrical supply.
This form of energy conversion, known as the Seebeck effect, was
first demonstrated by Seebeck in 1871, when he discovered that by
taking two wires of dissimilar metals and joining them at their ends,
so as to form two separate junctions, A and B as in Fig 11.16, a
thermo-e.m.f. was produced when the junctions were maintained at
different temperatures, ca,using current to flow round the circuit.
The arrangement of two dissimilar metal wires joined together
in this manner is called a thermocouple, the junction at the higher
tc:mperature being conventionally termed the hot or measuring
junction, and .that at the lower temperature the cold or referen·ce
junction. (In practice , the hot junction is in the form of a separate
unit for sensing the temperature and this is regarded generally as the
thermocouple proper.)
279
Figure J 1.16 Thermocouple
principle.
~
A
-- --·
(
B
~01 JUNCTION
HOT JUNCTION
COLO JUNCTION
--COLD JUNCTION
Experiments which followed Seebeck's discovery proved the
existence of two other effects .
. When an electric current flows through the junction of two
different substances it causes heat to be either absorbed or liberated
at the junction, depending on the direction of the current This is
known as the Peltier effect. In a circuit in which the only generated
voltage is a thermo-e.m.f., current flows through a heated junction
in a certain direction. If, instead of heat being supplied to the system,
a battery is introduced into the circuit, of such polarity that a current
is driven in the same direction as the thermo-current, the junction
which was previously heated will be cooled, and the junction which
was previously held at a constant temperature wilt be heated.
Lord Kelvin (when he was Sir William Thomson) discovered that
effects similar to the Seebeck and Peltier effects occur in a single,
homogeneous conductor: if two parts of a conductor are at different
temperatures, an e.m.f. is generated; and whe~ current flows from a
part of a conductor to another which is at a different temperature,
heat may be either liberated or absorbed. These phenomena are
different aspects of a single Thomson effect.
Thermocouple Materials and Combinations
The materials selected for use as thermoelectric sensing elements fall
into two main groups, base metal and rare metal, and are listed in
Table 11.1. The choice of a particular thermocouple is dictated by
the maximum temperature to be encountered in service. Thermo·
couples required for use in aircraft are confined to those of the basemetal group.
In order to utilize the thermoelectric principle for temperature
measurement, it is obviously necessary to measure the e.m.f.'s
280
Table I I. I Thermocoupl e Combinations
Meta ls and composition
Gro111•
Maximu m
remperafll re
oc
Posiri1·c wire
Negative wire
(contim1011s)
Copper (Cu)
Constantan (Ni. 40%;
Cu, 60%)
400
Application
Cy linder-head temperature
measurement
Iro n (f'e)
Constantan (Ni, 40%;
Cu , 60%)
850
Chrome! (Ni, 90%; Cr. IO%)
Alumel (Ni, 90%; Al, 2%
+Si+ Mn)
1, 100
Base metal
Rare metal
Platinu m ( Pt)
Rhodium-platinum (Rh, 13%;
Pt, 87%)
1,400
Exhaust-gas temperatute
measurement
Not utilized in aircraft tempera·
lure-indicating systems
Cr, chromium ; Ni, nickel; Al, aluminium ; Si, silicon; Mn, manganese
generated at the various temperatures. This is done 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 instrument into the circuit involves the presence of additional
junctions which produce their own e.m.f.'s and so introduce errors
in measurement. However, the effects are taken into consideration
when designing practical thermocouple circuits, and any errors
resulting from 'parasitic e.m.f.'s', as they are called , are eliminated.
The dial presentations of typical indicators are illustrated in Fig
11.17. The SET marking on the exhaust-gas temperature indicator
dial indicates the temperature at which the pointer is positioned
during the setting-up procedure carried out after installatio n.
Types of Thermocouple
The thermocouples employed in aircraft thermoelectric indicating
systems are of two basic types: (i) surface contact and (ii) immersion.
Typical examples are shown in Figs 11.1 8 (a) and (b) .
The surface-contact type is designed to ."measure the temperature
of a solid component and is used as the temperature-sensing
element of air-cooled-engine cylinder-head temperature-indicating
systems .. The copper/constantan or iron/constantan element may be
in the form of a 'shoe' bolted in good thermal contact with a
cylinder head representative of the highest temperature condition
281
or in the form of a washer bolted between the cylinder head and a
sparking plug.
The immersion type of thermocouple is designed for the measureFigure 11. 17 Dial presentat·
ions of thermoelectric
temperature indicators (moving
coil). (a) Engine-cylinder-head
temperature indicator;
(b) tur1'ine-engine gas temperature indicator.
(a)
(bl
Figure 11.18 Types of
thermocouple probes.
(a) Surface contact;
(b) immersion; (c) stagnation;
(d) rapid response; {e) tripleelement probe.
(cl
(d)
282
ment of gases and is therefore used as the sensing element of turbineengine gas temperature-indicating systems. The chromel/alumel hot
junction and wires are usually encased in ceramic insulation within a
metal (typically, Inconel) protection sheath, the complete assembly
forming a probe which can be immersed in the gas stream at the
points selected for measurement.
The techniques for assembly of thermocouple elements include
vacuum brazing, induction brazing and argon are welding as well as
the technique of electron-beam welding.
Immersion-type thermocouples are further classified as stagnation
and as rapid response, their application depending upon the velocity
of the engine exhaust gases. In pure jet engines the gas velocities are
high, and so in these engines stagnation thermocouples are employed.
The reason for this will be clear from Fig 11.18 (c), which shows
that 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 of gas temperature.
·
Rapid-response thermocouples are employed in turboprop
engines since their exhaust-gas velocities are lower than those of pure
jet engines. As can be seen from Fig 11.18 (d) , the sampling holes
are diametrically opposite each other and of equal size; therefore the
gases can flow directly over the hot junction enabling it to respond
more rapidly. Typical response times for stagnation and rapidresponse thermocouples are l to 2 seconds and 0.5 to I second,
respectively.
Thermocouple probes are also designed to contain double, triple
and in some cases, up to eight thermocouple elements within a single
probe. A triple element arrangement is shown schematically in Fig
11.18 (e). The purpose of such multi-arrangements is to provide
additional temperature signals for engine systems which utilize
separate facilities, such as exhaust-gas temperature control (see page
350) and engine combustion analysing. Insulation of the thermocouple elements from each other is provided by compacted magnesium
oxide (MgO) which also serves to maintain the elements in position.
When the hot junctions.of immersion-type thermocouples are in
contact with the gas stream, it is obvious that not only will the stream
velocity b.e reduced, but also that 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 junc.tion . .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 of total gas temperature.
In Fig 11.19 the constructional details of a third type of immer283
Fig11re JJ.1 9 Nozzle guide
vane thermocouple probe.
1 Terminals, 2 nickel-alumi·
nium lead, 3 end-cap, 4 silicone
filter, 5 plug , 6 mica washers,
7 ceramic top clamp, 8 ceramic
bottom clamp, 9 glass seal ,
10 body, 11 mounting flange,
12 couple junction, 13 probe,
14 ceramic insulator, 15 washer,
16 ferrule, 17 nickel-chromium
lead.
14
ll'l't!---12
13
sion thermocouple designed to measure the gas temperatures between
turbine stages are shown. The hot junction is housed inside a sheath
which is shaped so as to form the leading edge of a stator guide vane,
and for this reason the assembly is usually called a nozzle-guide-vane
thermocouple.
Gases flow over the hot junction which is positioned between
sampling holes of equal diameter as in a rapid-response thermocouple.
Unlike the latter, however, nozzle-guide-vane thermocouples do not
exhibit the same characteristics, because the sampling holes are much
smaller in diameter, and furthermore the couple response is made
slower by the mass of the sheath and its proximity to the guide vane.
In some types of turbine engine, it is necessary to sense the·
t emperature o f the air which flows internally for the purpose cif
engine cooling. The temperature sensor in this case is also a chrome!/
alumel thermocouple element, but so arranged that it can be
positioned over a vent hole, and between a mounting boss on the
engine and an overheat detector switch. An example of a sensor is
shown in Fig I I .20.
Location of Exhaust Gas Thermocouple Probes
The points at which the gas temperature of an engine is to be measured are of great impo rtance, since they will determine the accuracy
with which measured temperature can be related to engine performance. The ideal location for measurement is either at the turbine
blades themselves or at the turbine entry, but certain practical difficulties are involved which preclude the application of thermocouples
at such locations (see also page 294). Consequently, thermocouple
284
Figure ll.20 Cooling air
temperature sensing probe.
-t- VE TERMINAL
probes are generally located at the exhaust, or jet pipe unit, and between the turbine stages at one of the stator positions. The temperatures at these locations are much lower, but they relate very closely
to those at the turbine entry.
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 and exhaust
unit , and so measurement at one point only would not be truly
representative of the conditions prevailing.
Therefore, the measuring system always consists of a group of
five or more thermocouple probes suitably disposed in the gas flow,
and connected in parallel so as to measure a good average temperature condition (see Fig 11.21 ). Nozzle guide-vane thermocouples
are arranged in pairs of long-reach and short-reach probes, named
according to the extent to which the hot junctions and gas sampling
holes reach into the gas stream.
Thermocouple Harness Assemblies
The thermocouple sensing elements and their leads are made up
into a harness assembly the design of which can vary dependent on
the type of engine and the number of probes required. Figure 11 .22
is intended to serve only as an example of the general 'make-up' of a
harness. The· five probes in this case each contain two thermocouple
elements; one for temperature indication, and the other for a
temperature control c_ircuit. Whereas for some engines, probes and
thermocouple lead junction boxes may be designed as separate items,
the probes in the example considered are welded to stainless steel
junction boxes thus forming single items. The parallel-connected
285
Figure J1.21 Disposition of
exhaust,.gas thermocouple
probes.
THERMOCOUPLE
HARNESS
/
Figure J J.22 Thermocouple harness.
thermocouple leads pass through lnconel conduits which are also
welded to ferrules at the junction boxes. The leads terminate at a
main junction, or 'take-off box, to which the leads of the remainder
of the circuits are connected. The forming of the conduits in the
286
manner shown provides sufficient flexibility to permit a high number
of harness removals and replacements.
Another type of harness designed for exhaust gas temperature
measurement and known a~ a hub array is shown in Fig 11.23. The
thermocouple probes radiate from the hub, and· since the bodies of
each probe are high-temperature brazed to the hub periphery, a very
robust assembly is provided. The thermocouple leads are connected
to rings of the same material as the thermocouples, i.e. chrome!
leads to a chrome! ring, and alumel leads to an alumel ring. The rings
are lo~ated within the hub and are insulated from each other, and
the hub, by ceramic mouldings. The hub is also packed with
magnesium oxide powder to provide support for the leads, and to
provide further electrical insulation. Connection of the harness to
the remainder of the circuit is done by means of mineral-insulated
lead wires and a separately mounted 'take-off box.
Figure 11. 23 Hub•array type of harness.
MINERAL-INSULATED
LEADS AND SLEEVE
Cold-Junction Temperature Compensation
As we have already seen, the indicator of a thermoelectric temperature measuring system forms the cold junction of the system, and
the e.m.f. produced depends upon the difference between the
temperature of this junction and the hot junction. It is thus
apparent that, if the ambient temperature of the indicator itself
changes while the hot junction temperatµre remains constant, then a
change in e.m.f. will result causing the indicator to read a different
temperature.
In applying this principle to the measurement of aircraft engine
temperatures, such temperature differences constitute indication
errors which cannot be tolerated, since it is essential for the
287
indicated readings to be representative (?f the temperature conditions
at the hot junction only . In order to achieve such readings it is
necessary to provide indicators with a device which will detect
ambient temperature changes and compensate for possible errors;
such a device is called a cold-junction compensator. Before going
into the mechanical and operating details of a compensator, it is useful to consider first how the changes in e.m.f. actually arise.
The various combinations of thermocouple materials specified for
use in aircraft conform to standard temperature/e.m.f. relationships,
and the indicators employed in conjunction with these combinations
are calibrated accordingly. The e.m.f.'s obtained correspond to a
cold-junction temperature which is usually maintained at either O~C
or 20°C. (See tables on page 403) .
Let us assume, for example, that the cold juntion is maintained at
0°C and that the hot junction temperature has reached 500°C. At
this temperature difference a standard value of e.m.f. generated by a
chrome/alumel combination is 20 .64 mV. If now the temperature at
the cold junction increases to 20° C while the hot junction remains
at 500°C, the temperature difference decreases to 480°C and the
e.m.f. equivalent to this difference is now 20.64 mV minus the e.m.f.
at 20°C, and as a standard value this corresponds to 0.79 mV. Thus,
the moving element of the indicator will respond to an e.m.f. of
19 .85 m V and move 'down scale' to a reading of 480° C.
A change, therefore, in ambient and cold-junction temperature
decreases or increases the e.m.f. generated by the thermocouple,
making the indicator read low or high by an amount equal to the
change of ambient temperature.
The method commonly adopted for the compensation of these
effects on moving coil indicators is quite simple and is, in fact, an
adaptation of the bimetallic strip principle described on page 15.
For this purpose, however, the strips of dissimilar metals are
fastened together and coiled in the shape of a flat spiral spring.
One end of the spring is anchored to a bracket which forms part of
the moving-element support, while the other end (free end) is
connected by an anchor tag to the outer end of one of the controllini
hairsprings, thus forming the fixed point for the hairspring. A
typical arrangement is shown in Fig 11 .24.
When the indicator is on open-circuit, i.e. disconnected from the
thermocouple system, the spring responds to ambient temperature
changes at the indicator, an increase in temperature causing the sprin;
to unwind so that its free end carries the hairspring and moving
element round to indicate the increase in temperature. Conversely, a
temperature decrease will wind up the compensator spring so that the
moving element will indicate the lower temperature. Therefore an
indicator disconnected from its e.m.f. source operates as a directreading bimetallic type of thermometer.
288
Figure J 1.24 Application of
bimetal type of cold junction
compensator.
HAIRSPRING
ANCHOR TAG
COMPENSATOR _
ANCHOR
'._Jb
__Jl..l-!L!---ltTir-...-~
BIMETAL
POINTER
SPIDER
HAIRSPRING
CONTACT TAG
JEWEL LOCKING NUT
With the thermocouple system connected to the indicator the circuit is completed, and if the two junctions are at the temperatures
earlier assumed, namely 0°C and 500°C, the e.m.f.' wi ll position the
moving element to read 500°C. If the temperature at the indicator
increases to 20°C, then, as already illustrated, the e.m.f. is reduced
but the tendency for the moving element to move do.wn 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.
Electrical Method of Compensation
The compensation for the effects of cold-junction temperature changes
can also be accomplished by applying the principle of an electrical
bridge circuit. The circuit arrangement based on that adopted in
exhaust gas temperature indicators of the servo-operated type (see
page 292) is shown in much simplified form in Fig 11. 25.
The thermocouple harness leads are connected to leads of the same
metals as the thermocouple and contained within the compensation
circuit module. The module leads are individually connected to
copper leads which are embedded in close proximity to each other
289
- - -- ·-----
FEEDBACK FROM
POSITIONAL - - - POTENTIOMETER
- - - - - - .. - - - -
TEST POINT
COPPER LEADS
THE RMOCOIJPLE HARNESS
I
ANO
~
EXTENSION LEADS
CHROMEL/ALUMEL LEADS
290
REFERENCE SUPPLY
OUTPUT TO POSITIONAL
FEEDBACK PQTENTIOMETER
+
REFERENCE SUPPLY
VOLTAGE
Figure 11.25 Cold-junction
compensation circuit.
.. ____ ---
-.l
OUTPUT TO
DIFFERENTIAL
ANO SERVO
AMPLIFIERS
within the former which supports the copper coil resistor R 4 ; thus,
together they form the effective cold junction of the system. The
bridge circuit is supplied with direct current at 7 V from a stabilized
reference supply module within the indicator, and the output from
the bridge is supplied to the indicating element of the indicator via a
servo amplifier.
As already noted, the standard values of e.m.f. produced by a
thermocouple are related to a selected value of cold-junction
temperature. In this case, the bridge circuit is adjusted by means of
a variable resistor (R V 1 ) so that an e.m.f. of the correct sense and
magnitude is injected in series with that of the thermocouple such
that in combination, the e.m.f. is equal to that which would be
obtained if the cold-junction temperature were 0°C. Since the ambie1
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 result in a reduction of the thermocouple output (see page 288). The resistor R I will, however, also be
subjected to the higher ambient temperature, but because under such
conditions the resistance of R I decreases, it will modify the bridge
circuit conditions so as to restore the combined e.m.f. output to the
standard value corresponding to a cold-junction temperature of 0° C.
The output is termed the demand exhaust gas temperature signal
and is supplied to the indicating element of the indicator in the
manner described on page 292.
Compensation of Moving-Coil Resistance Changes
The changes of thermocouple e.m.f. so far discussed are not,
unfortunately, the only effect brought about by changes of ambient
temperature at the indicator. A second effect requiring compensation is the change in -resistance of the moving coil itself. The manner
in which this is accomplished depends upon the particular design,
but the two methods most commonly adopted are the thermomagnetic shunt and thermo-resistor described in Chapter 2.
External Circuit and Resistance
The external circuit of a thermoelectric indicating system consists of
the thermocouple and its leads, and the leads from the junction box
at the engine bulkhead 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 appropriate copper leads or
cable normally used in aircraft, it is not acceptable to do so in a
thermoelectric system.
This may be explained by taking the case of a copper/constantan
thermocouple which is to be connected to a cylinder-head temperature indicator. If a length of normal copper twin-core cable is
connected to the thermocouple terminal box , then one copper lead
will be joined to its thermocouple partner, but the other one will be
joined to the constantan lead. It is thus apparent that the joining of
the two dissimilar metals introduces another effective hot junction
which will respond to temperature changes occurring at the junction
box, and in unbalancing the temperature/e .m.f. relationship will
cause the indicator readings to be in error. .Similarly, all terminal
connections which may be necessary for routing the leads through
the aircraft and connections at the indicator itself will create
additional hot junctions and so aggravate indicator errors.
In order to eliminate these hot junctions, it is the practice to use
leads made of the same materials as the thermocouple itself, such
leads being known as ex tension leads. It is sometimes the practice
also to use thermocouple materials having similar thermoelectric
characteristics in combination ; for example, a chromel/alumel thermocouple may be joined to its indicator by copper/constantan leads,
known as compensating leads. These not only compensate for
parasitic effects but also reduce the cost, since chromel-alumel is
expensive.
Additional hot junctions at the indicator itself are eHminated by
making the positive terminal of copper or brass, and the negative
terminal of constantan. This applies to indicators for use with the
various thermocouple combinations.
Another important factor in connection with the external circuit
291
is its resistance, which must be kept not only low but also constant
for a particular installation . Indicators are normally calibrated for
use with either an 8 n or a 25 n external circuit resistance and are
marked accordingly on their dials. The thermocouple, leads and
harness are made up in fixed low-resistance lengths, and for turbine engines they form component parts of the engine. Similarly,
the extension or compensating leads must also be made up in lengths
and of uniform resistance to suit the varying distances between
thermocouple ~ot junction and indicator installations.
Adjustment of the total external circuit resistance to the value for
which associated instruments have been calibrated is made possible
by connecting a trimming resistance spool in series with one of the
extension or compensating leads. The material used for the spool
may be either Eureka or Manganin, both of which have negative
temperature coefficients of resistance. The inclusion of the spool in
the circuit introduces anot~er junction, of course, and w~ether it
should be in the positive lead or the negative lead is governed by the
material of the spool and the leads. For example, if a Eureka spool
is to be used in a copper/constantan thermocouple system, it must be
connected in the constantan or negative lead, whereas a Manganin
spool must be connected in the copper or positive lead, in order to
have negligible thermoelectric effect. In a chromel/ alumel system,
Manganin is usually employed, .and for this circuit negligible thermo. electric effect is obtained by connecting it in the chrome! or positive
lead.
It is possible for the resistance of the extension leads to change
with changes of temperature, but as most of the total circuit resistance is within the indicator, the compensation methods described
earlier can also be effective in compensating for changes in external
circuit resistance.
Servo-Operated Indicating Systems
For some applications of the thermoelectric method of temperature
indication, the indicators can be of the type based on that already
described on page ·248, i.e. they can operate on servo control principlei
as opposed to the conventional moving coil principle. The circuit
arrangement is illustrated in Fig 11.26, and as will be noted the
thermocouple output is supplied to a signal processing module consisting primarily of a cold-junction reference circuit and an error
signal amplifier. The cold-junction reference circuit (see also Fig
11.25) compensates for changes in ambient temperature of the indicator, and automatically adjusts the thermocouple output to produce
a computed demand exhaust gas temperature signal. This i'> then
compared with a d.c. output from the wiper of a positional feedback
potentiometer, and since the wiper is geared to the main pointer and
292
01/ER· TEMPERATURE
WARNING LIGHT
I
~
QI/ER-TEMPERATURE LIMIT
II,. '\.~ /POINTER
,----
- - - -------,
:RMOCOUPLE
;NAL
II {
LAMP SUPPLY
u-----t---PtWARNING
LAMP
J
TEST SUPPLY
IL:.SER\/0
AMP. AND MONITOR MODULE
________
_
o---
II 400 Hz A.C.
'----- - - - - - - - - - - - 1
POSITIONAL
FEEDBACK
' - - - - - - - - - - - - -- - --.!pQTENTIOME ER
II D.C.
~ ERROR SIGNAL ~ FEEDBACK
V
VBuFFER
Figure 11.26 Servo-operated
indicator.
D>
SER\/0 DRIIIE
DI
M.ECHANICAL DRIIIE
digital counter of the indicator, then the d.c. output which is fed
back to the cold junction reference circuit represents the indicated
exhaust gas temperature. Any difference between computed and
indicate~ temperatures results in an error signal being produced by
the cold-junction reference circuit which then supplies the signal to a
servo amplifier. The servo amplifier output is fed to the armature
winding of the d.c. servomotor which then drives the indicator
pointer and digital counters causing them to display a coarse and fine
indication respectively of the exhaust gas temperature. The wiper
of the positional feedback potentiometer is also repositioned to provide a feedback voltage which backs-o~f t!1e demanded temperature
signal until the error signal is z~ro ; 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 servo loop to back-off the error signal voltage. Should
such failure occur, the monitor functions as an 'on-off' switch, and
in the 'off' 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 a.c. supply to the indicator falling below
JOOV.
An over-temperature warning light is incorporated in the indicator,
293
and is contro lled by a relay , a comparato r, and a solid-state switching
circuit. The function of the comparator is to compare the feedback
voltage from the positional potentiometer with pre-set volta~ the
level of which is equivalent to a pre-determined over-temperature
limit for the particula r type of engine . ln the event of this limit
being exceeded, the feedback voltage will exceed the reference voltage level, and the switch ing circuit will cause the relay to energize,
thereby closing its contac ts to illuminate the warning light. A
separate su pply voltage may be connected to the light by means of
an 'override' facility as a means of testing its fil ament at any point
over the t emperature range of the indicator.
An over-temperature pointer is also fitted concentrically with the
main pointer, and is initially positioned at th e appropriate scale
graduation. It operates in a similar manner to the over-speed pointer
of a servo-operated tachometer indicator (~ee page J_49).
a
Radiation Pyrometer
System
294
As we have already learned, th e measurement of exhaust gas
temperatures by means of thermocouples provides a fairly representative indication of the conditions prevailing at a turbine, but errors
can arise with varying efficiency of a compressor/ tu rb ine assembly,
and with ageing of thermocouples. The result is that the most
efficient engine could run at a turbine temperature lower than the
specified op timum. Moreover, the actual blade temperature of
engines with air-cooled blades, is a function of the cooling efficiency
as well as turbine inlet temperature, thereby giving a further derating
characteristic. Thus, in the course of turbine engine development,
the need arose for a system which could in fact, directly sense actual
turbine blade temperature.
The system adopted for this purpose is known as a radiatio n
py rometer system, and its principal components are illustrated in
Fig 11.27. It depends for its operation on the fac~ that the radiation
emitted by any body, at any wavelength, is a function of the
temperature of the body, and its emissivity. The temperature of
turbine blades can, therefore, be determined by measu ring their
raoiation at their ·known value of emissivity. This method does not
have the thermal inertia inherent in a thermocouple and therefore it
provides for more rapid response.
The pyrometer head and its sighting tube assembly is directly
mounted on the turbine casing, such that its sighting line is directly
on the turbine blades. Blade radiation is focussed by a synthetic
sapphire lens which is brazed into a titanium mount having high
fatigue strength, and ·a coefficient of expansion similar to that of the
lens material. The body of the pyrometer is manu factured from stainless steel, and has external cooling fins to minimize the temperature
at the junction with a fibre optic link onto which the lens focusses
Figure 11.27 Radiation pyrometer system.
DETECTOR
FIBRE OPTIC LINK
TO--+
AMPLIFIER
ENGINE CASING
blade radiation. The fibre optic link, which is contained with PTFE
tubing protected by flexible stainless steel braiding, transmits the
radiated energy from the pyrometer head as an optical signal to the
detector unit. The ends of the fibres adjacent to the pyrometer head
are bonded by a special process to withstand the high temperatures
encountered at the turbine casing.
·Conversion of the optical signal into an electrical output is effected
by a silicon photocell contained within the aluminium casing of the
detector unit. The temperature within the casing is thermostaticallycontrolled at a value slightly higher than maximum ambient so that
the cell characteristics are stabilized and accurate over a wide range
of ambient temperatures.
The output from the detector is a non-linear signal of small
magnitude, and is supplied to an amplifier which then produces ·a
linear and higher output necessary for the operation of the temperature indicator, or engine control unit as appropriate to the installa295
tion. The d.c. supplies and reference voltages required for system
operation are also incorporated within the amplifier.
Questio ns
11.1
11.2
11.3
11.4
11.5
11.6
11.7
Describe how temperature can cause variations in the properties of
substances.
Define the following: (i) convection, (ii) conduction, (iii) radiation,
(iv) temperature.
Define the fundamental interval of temperature and state how it is
divided in accordance with the Celsius and Fahrenheit scales.
(a) What is the relationship between the Celsius, Fahrenheit and
absolute [Kelvin] temperature scales?
(b) Convert 40° C into degrees absolute and degrees Fahrenheit.
Convert 77°F into degrees Celsius.
Define Ohm's law and include the expressions used for calculating
the three quantities, voltage, current and resistance.
Calculate the resistance (R r) of circuits (a) and (b ), and state
what current will flow in each circuit when 24 V d.c. is applied .
R1
FtgurelJ.28.
R1
R2
R3
~
200
160
400
(8)
7n
(b)
11.8
(a) Calculate the resistance (Rr) of a circuit containing two resistances in parallel , the values being R 1 = 20 n and R 2 = 45 n.
(b) Draw a diagram to illustrate a series-parallel circuit and explain
11.9
(a) What are the factors governing the resistance of a conductor?
(b) A round copper conductor 0.8 mm diameter has a length of
how you would solve such a -circuit from given values.
1,200 cm . Calculate its resistance at normal temperature. (The
resistivity of copper is approximately 1. 7 X 106 µ!'2-cm.)
11.10 Draw a diagram tb illustrate a Wheatstone bridge circuit, and from
this derive the expression for calculating the value of an unknown
resistance.
11.11 (a) Explain how the Wheatstone bridge circuit may be utilized for
the measurement of temperatures.
(b) Would the circuit be in balance at each temperature?
11.12 (a) What materials an~ most commonly used for resistance-type
temperature-sensing elements?
(b) Describe the construction of a typical element.
11.13 (a) Explain the operating principle of a moving-coil instrument.
(b) Does such an instrument have a linear or a non-linear scale?
296
11.14
Why is a soft-iron core placed between the magnet poles of a movingcoil instrument?
11.15 Explain how instruments are screened against the effects of external
magnetic fields.
11.16 (a) Describe the construction and operation of a ratiometer type of
temperature indicator.
(b) What principal advantage does the instrument have over a Wheatstone bridge type?
11.17 What would be the effect of an open-circuit between the sensing
element and indicator of a ratiometer system?
11.18 (a) Explain the thermocouple principle, and state to what temperature measurement it is applied.
(b) What metal combinations are used in thermocouple probes?
11.19 A thermocouple probe used for the measurement of engine cylinderhead temperature measurements is of: (a) the immersion type,
(b) the stagnation type, (c) the surface contact type. Which of these
statements is correct?
11.20 Describe the construction of a typical thermocouple probe assembly
used for turbine-engine exhaust-gas temperature measurement.
11.21 How is it ensured that a good average temperature condition of
exhaust gases is measured by an indicating system?
11.22 What effects can changes of cold-junction temperature have on the
indications of thermoelectric indicators? Describe the methods of
compensation.
11.23 What is the differen~e between extension leads and compensating
leads?
11.24 (a) What is the purpose of the resistor connected in series with one
of the leads in certain thermoelectric systems?
(b) Name the materials from which resistors are made and explain
why they are never connected in leads having the same polarity .
11.25 Briefly describe how the principle of servo control can be applied to
the measurement of turbine engine exhaust gas temperature.
11.26 On what fundamental principle does a radiation pyrometer system
operate? Briefly describe a practical system.
297
1 2 Measurement of
pressure
In many of the systems associated with the operation of aircraft and
engines, liquids and gases are used the pressures of which must be
measured and indicated. The gauges and indicating systems fall into
two main categories: (i) direct-reading, or those to which the source
of pressure is directly connected, and (ii) remote-indicating, or those
having a separate sensing element connected to a pressure source at
some remote point.
Methods of
Measuring Pressure
Pressure, which is defined as force per unit area, may be measured
directly either by balancing it against that produced by a ~olumn of
liquid of known density, or it may be permitted to act over a known
area and then measured in terms of the force produced. The former
method is the one utilized in simple U-tube manometers, while the
second enables us to measure the force by balancing it against a
known weight, or by the strain it produces in an elastic material.
In connection with pressure measurements, we are concerned with
the following terms:
Absolute Pressure
The absolute pressure of a fluid is the difference between the
pressure of the fluid and the absolute _zero of pressure, the latter
being the pressure in a complete vacuum. Thus, in using a gauge to
measure the fluid pressure, the absolute pressure o f the fluid would
be equal to the sum of the gauge pressure and the atmospheric
pressure.
Gauge Pressure
Most pressure gauges measure the difference between the absolute
pressure of a fluid and the atmospheric pressure. Such measurement
is called the gauge pressure, and is equal ·to the absolute pressure
minus the atmospheric pressure. Gauge pressure is either positive or
negative, depending on its level above or gelow the atmospheric
pressure reference.
In gauges which are spoken of as indicating vacuum or suction,
298
they are really indicating the amount the absolute pressure is less
than atmospheric pressure. Thus, gauge pressure is equal to
atmospheric pressure minus pressure of the fluid, and the absolute
pressure is equal to atmospheric pressure minus gauge pressure.
U-Tube Manometer
The simple U-tube manometer shown in Fig 12.1, consists of a glass
tube partially filled with a liquid, usually water or mercury, which
finds its own level at a point O within the open-ended limbs of the U.
If a low-pressure source is connected to the limb A, then a force
equal to the applied pressure multiplied by the area of the bore will
act on the surface of the liquid, forcing it down limb A. At the same
time the liquid is forced up the bore of limb B until a state of
equilibrium exists and the levels of the liquid stand at the same
distance agave and below the zero point. By taking into account
the area of the tube bore and the density of the liquid it is possible
to calculate the pressure from the difference in liquid levels, as the
following example shows.
Let us assume that the manometer is of the mercury type having a
bore area of 3 in2 , and that a pressure is applied to limb A such that
at equilibrium the mercury levels are 4 in below and 4 in above zero.
The difference in levels is Hand its value is obtained by subtracting
the lower level from the higher one; thus, H= hs- hA =4-(- 4) = 8 in.
Figure 12.1 U-tube manometer.
APPUED PRESSURE
p
= 3.92
L.B,IIN
2
6
5
LIMB A
4
3
2
1
T
0
-4 IN
2
_J__
3
4
hA=
5
6
LIMB B
t1
_l
H
=
8 IN
~
299
Now, we must know the weight of the mercury column being
supported, and this is calculated from volume multiplied by density.
The volume in this case is 3H and the density of mercury is usuallv
taken as 0.49 lbf/in 3 • Thus, the weight of the column is
3H X 0.49:::: 1.47 X 8:::: I 1.76 lb, and as the pressure balancing
this is weight divided by area, then I I .76/3:::: 3.92 lbf/in 2 is the
pressure being applied to limb A and corresponding to a difference
in mercury levels of 8 in. In the same manner, other pressures can
be calculated from the corresponding values of level difference H.
In practice, manometers are used for checking the calibration of
pressure gauges, and so it is usually more convenient to graduate the
manometer scale directly in pounds per square inch. If 3.92 lbf/in2
is represented by 8 in, then, for the mercury manometer we have
considered, 1 lbf/in2 is equal to 8/3.92, or 2.04 in, and so a scale
can be graduated with marks spaced this distance apart, each
representing an increment of 1 lbfiin 2 • The equivalent value 2. 04 in. I
to 1 lbf/in 2 is standard and results of calculations for differing bore
areas will show that they are independent of the areas.
If the water is used in the manometer the foregoing principles
also apply, but as water has a much lower density than mercury,
then for a given pressure the difference in level H for a water
manometer will be much greater than that of a mercury manometer
(2.04 in. Hg:::: 27.7 in. H2 0 very nearly).
Pressure/Weight Balancing
The measurement of pressure by balancing it against weights of
known value is based on the principle of the hydraulic press, and as
far as instruments are concerned, it finds a practical application in
a hydraulic device known as the dead-weight tester and used for the
calibration and testing of certain types of pressure gauge.
Let us suppose that we have a cylinder containing a liquid as shown
in Fig 12. 2(a), and that a tight-fitting piston is placed on the liquid's
surface. If now we try to push the piston down with a force F , we
shall find that the piston will only be displaced by a very sma11
amount, since the compressibility of liquids is very small. The pressur,
p produced in the liquid by pushing on the piston is equal to F/a and
is transmitted to every part of the liquid and acts on all surfaces in
contact with it.
In applying this principle to a hydraulic press we require essentially
two interconnected cylinders as shown at (b ), one of small crosssectional area a,, the other oflarge cross-sectional area a 2 • Each
cylinder is fitted with a piston and both are supplied with oil from a
common reservoir. If a force Fis exerted on the small piston then
the additional pressure produced is p :::: Fla, and is transmitted
throughout the liquid and therefore acts on the larger piston of
300
Figure 12. 2 Pressure/weight
balancing. (a) Pressure produced
in liquid; (bl hydraulic press;
(c) dead-weight tester.
F
(al
F
I
(bl
w
RESERVOIR - - - + t
CONNECTION FOR
GAUGE UNDER TEST
(cl
area a 2 • Thus, the force that can be exerted by this piston is equal
to pa 2 • If the press is designed to lift a weight W, then W will also
be equal to pa 2 • The weight that can be lifted by the application of
a force F is multiplied in the ratio of the areas of the two pistons.
Figure l 2.2(c) illustrates the hydraulic press principle applied to a
dead-weight tester. When the piston in the horizontal cylinder is
screwed in, a force is exerted and pressure is transmitted to the
301
weigh!ng piston in the vertical cylinder, so that it can be supported
in a balanced condition by the oil column. In this application we are
more interested in direct measurement of pressure and therefore need
to know what weights are necessary to balance against required
pressures. Now, the area constant A for a typical dead-weight tester
is 0.125 in 2 ; thus, assuming that we require to balance a pressure p
of 50 lbf/in 2 , then, from the relation W = pA a weight of 6.25 lb is
necessary. With this weight in position on the weighing piston the
piston in the horizontal cylinder is screwed in until the weight is
freely supported by the oil, which, at this point, is subjected to
50 lbf/in 2 • In practice, the weights are graded and are marked with
the actual pressures against which they will balance.
Elastic Pressure-Sensing Elements
For pressure measurements in aircraft, it is obviously impracticable
to equip the cockpit with U-tube manometers and dead-weight
testers. It is the practice, therefore, to use elastic pressure-sensing
elements, in which forces can be produced by applied pressures and
made to actuate mechanical and/or electrical indicating elements.
The sensing elements commonly used are Bourdon tubes, diaphragms,
capsules and bellows.
Bourdon Tube
The Bourdon tube is about the oldest of the pressure-sensing elements.
It was developed and patented in 1850 by a Parisian watchmaker
(whose name it bears) and has _been in general use ever since, particularly in applications where the measurement of high pressure is necessary. The element is essentially a length of metal tube, specially
extruded to give it an elliptical cross-section, and shaped into the
form of a letter C. The ratio between the major and minor axes
depends on the sensitivity required, a larger ratio providing greater
sensitivity 1 The material from which the tube is made may be either
phosphor-bronze, beryllium-bronze or beryllium-copper. One end of
the tube, the 'free-end', is sealed, while the other end is left open and
fixed into a boss so that it may be connected to a source of pressure
and form a closed system.
When pressure is _applied to the interior of the tube there is a
tendency-for the tube to change from an elliptical cross-section to a
circular one, and also to st raighten out as it becomes more circular.
In other words, it tends to assume its original shape. This is not such
a simple process as it might appear and many theories have been
advanced to explain it. However, a practical explanation sufficient
for our purpose is as follows. Firstly , a tube of elliptical crosssection has a smaller volume than a circular one of the same length
and perimeter. This being the case, an elliptical tube when connected
302
to a pressure source is made to accommodate more of the liquid, or
gas, than it can normally hold. In consequence, forces are set up which
change the shape and thereby increase the volume. The second point
concerns the straightening out of the tube as a result of its change in
cross-section. Since the tube is formed in a C-shape then it can be
considered as having an inner wall and an outer wall, and under 'no
pressure' conditions they are each at a definite radius from the centre
of the C. When pressure is applied and the tube starts changing shape,
the inner·wall is forced towards the centre, decreasing the radius, and
the outer wall is forced away from the centre thus increasing the
radius. Now, along any section of the curved tube the effects of the
changing radii are to compress the inner wall and to stretch the outer
wall , but as the walls are joined as a common tube, reactions are set
up opposite to the compressive and stretching forces so that a complete section is displaced from the centre of the C. Since this takes
place at all sections along the tube and increases towards the more
flexible portions, then the resultant of all the reactions will produce
maximum displacement at the free end. Within close limits the
change in angle subtended at the centre by a tube is proportional to
the change of internal pressure, and within the limit of proportionality of the material employed, the displacement of the free end is
proportional to the applied pressure.
The displacement of the free end is only small; therefore , in order
to transmit this in terms of pressure, a quadrant and magnifying
system is employed as the coupling element between tube and
pointer.
Diaphragms
Diaphragms in the form of corrugated circular metal discs, owing to
their sensitivity, are usually employed for the measurement of low
pressures. They are always arranged so that they are exposed at
one side to the pressure to be measured, their deflections being transmitted to pointer mechanisms, or to a warning-light contact
assembly. The materials used for their manufacture are generally
the same as those used for Bourdon tubes. The purpose of the
corrugations is to permit larger deflections, for given thicknesses,
than would be obtained with a flat disc. Furthermore, their number
and depth control the response and sensitivity characteristics; the
greater the number and depth the more nearly linear is its deflection
and the greater is its sensitivity.
Capsules
Capsules are made up of two diaphragms placed together and joined
at their edges to form a chamber which may be completely sealed or
open to a· source of pressure. Like single diaphragms they are also
employed for the measurement of low pressure, but they are more
303
sensitive to small pressure changes. The operation of capsules in
their various applications has already been described in the chapters
on height and airspeed measuring instruments.
Bellows
A bellows type of element can be considered as an extension of the
corrugated diaphragm principle, and in operation it bears sollle
resemblance to a helical compression spring. It may be·used for
high, low or differential pressure measurement, and in some applications a spririg may be employed (internally or externally) to
increase what is termed the 'spring-rate' and to assist a bellows to
return to its natural length when pressure is removed.
The element is made from a length of seamless metal tube with
suitable end fittings for connection to pressure sources or for
hermetic sealing. Typical applications of bellows are described on
pages 306 and 307.
Direct-Reading
Pressure Gauges
These are almost entirely based on the Bourdon tube principle
already described, and are used for such measurements as hydraulic
system pressure, and in a number of general aviation aircraft
powered by piston engines, fuel and oil pressures also. An example
of the general construction is given in Fig 12.3.
Figure 12.3 Direct-reading
pressure gauge.
Remote-Indicating
·Pressure Gauge
Systems
304
Systems of this type are available in a variety of forms but all have
one common feature ; they consist of two main components, a transmitter unit located at the pressure source, and an indicator mounted
on the appropriate panel. They have distinct advantages over directreading gauges; for example, the pressures of hazardous fluids such as
fuel, engine oil and certain hydraulic fluids can be measured at their
source and not brought up to the cockpit or flight deck; also long
pipelines are unnecessary thus saving weight.
The majority of systems in current use are of the electrical
transmission type, i.e. pressure is measur'ed in terms of the displacement of an elastic pressure-sensing element, and transmitted to an
indicator as a combination of varying voltage and current signals.
Transmitters are therefore made up of mechanical and electrical
sections, and according to the operating principle adopted for any
one system, indicators can either be synchronous receivers, d.c. or a.c.
ratiometers, and in some applications, servo-operated.
D.C. Synchronous System
Figure 12.4 shows the arrangement of a transmitter working on the
micro-Desynn principle (see Chapter 9). The pressure-sensing elemFigure 12.4 Micro-Desynn
transmitter. I Micro-Desynn
transmitting element, 2 eccentric pin, 3 push rod, 4 pressure·
sensing element, 5 bellows,
6 cup-shaped pressing, 7 spring,
8 rocking lever.
t
PRESSURE
IN
305
ent consists of a bellows which is open to the pressure source. A cupshaped pressing is fitted inside the bellbws and forms a connectio~
for a push-rod which bears against a rocking lever pivoted on a fixed
part of the mechanism. A spring is provided inside the bellows.
The electrical element is of the same type as that shown in Fig 9.4
and is positioned in the transmitter body in such a manner that the
eccentric pin is also in contact with the pivoted rocking lever. The
indicator is of the normal Desynn system type.
When pressure is admitted to the interior of the bellows it expands
and moves the push-rod up, thus rotating the rocking lever. This, in
turn, moves the eccentric pin and brushes coupled to it through a
small angle over the coils. The resistance changes produced set up
varying voltage and current combinations within the indicator, which
is calibrated for the appropriate pressure range.
D.C. Ratiometer System
An example of a pure d.c. ratiometer system which is still adopted
in one or two types of older generation aircraft, is that employing a
transmitter which is a special adaptation of the micro-Desynn pressure transmitter just described, the essential difference being in the
electrical circuit arrangement. The element still has two brushes and
resistance coils, but instead of the normal micro-Desynn method of
connection (see also ·Fig 9.5) they are connected as a simple twin
resistance parallel circuit. The two connections terminating at the
coils are joined to the appropriate terminals of a ratiometer similar to
that employed for temperature measurement (see page 276). The
operation is therefore quite simple; the movement of the bellows and
brushes results in a change of circuit resistance proportional to the
pressure change, which is measured as a coil current ratio.
A.C. Inductor and Ratiometer System
The operation of this system is dependent on the production of a
current ratio by a variable inductor transmitter, an example of which
is shown in Fig 12.5. It consists of a main body containing a bellows
and two single-phase two-pole stators each surrounding a laminated
salient-pole armature core. Both cores are on a common shaft and
are so arranged that, as pressure increases, the lower core (A) moves
further into its associated stator coil, while the upper core (B) moves
further out of its coil. The coils are supplied with alternating current
at 26 V, 400 Hz. The core poles are set 90° apart and the stators are
also positioned so that the poles produced in them are at 90° to each
other to prevent mutual magnetic interference. A spring provides a
controlled loading on the bellows and armature· cores, and is adjustable so as to set the starting position of the cores during calibration.
The essential parts of the indicator used with this particular trans306
Figure 12.5 Section view of
an inductor-type transmitter.
1 Overload stop screw, 2 centre
spindle bearing, 3 guide bush,
4 a!uminium cup, 5 armature
cores, 6 aluminium housing,
7 stators and windings, 8 centre
spindle assembly, 9 centre
spindle bearing, 10 guide bush,
11 bellows, 12 base plate,
13 radial ducts, 14 .body,
15 electrical connector, 16
main spring.
mitter are illustrated in Fig 12.6. The coils around the laminated
cores are connected to the transmitter stator coils, and as will be
noted, a gap is provided in one limb of each core. The purpose of
the gaps is to permit free rotation of two aluminium cam-shaped discs
which form the moving element. The positioning of the discs on
their common shaft is such that, when the element rotates in a clockwise direction (viewed from the front of the instrument in its normal
position), the effective radius ·of the front disc a decreases in its air
gap, while that of the rear disc b increases. The moving element is
damped by a circular disc at the rear end of the shaft, and free to
rotate between the poles of a permanent magnet. A hairspring is
provided to return the pointer to the off-scale position in the event of
a power failure.
When the bellows expand under an increasing pressure, the armature cores move in their respective stators, and since the latter are
supplied with alternating current, there is a change in the inductance
of the coils. Thus core A, in moving further into its stator, increases
the inductance and impedance, and core B, in moving out of its stator,
decreases the inductance and impedance. The difference between the
two may therefore be interpreted in terms of pressure.
As the stator coils are connected to the indicator coils in the form
of a bridge network, then the changes in impedance will produce a
change of current in the indicator coils at a predetermined ratio. The
current is alternating, and so produces alternating fluxes in the
laminated cores and across their gaps. It will be noted from Fig 12.6
that copper shading rings are provided at the air gaps. The effect of
307
.... - -::.::_
....,..
Figure 12.6 A.C. ratiometer
elements.
a
CAM·SHAPED
DISCS
r(),
I
:_
I
.
_) 0
DISC
a
,
G
I
J
DISC b
the alternating flux is to induce eddy currents in the rings, these
currents in tum setting up their own fluxes which react with the airgap fluxes to exert a torque on the cam-shaped discs. The resulting
movement of the cam discs is arranged to be in a direction determined
by the coil carrying the .greater current, and due to the disposition of
the discs, this means there will be a difference between their torques.
In the gap affected by the greater current the effective radius of its
disc (a) decreases, thereby increasing the impedance and decreasing
the torque, while in the gap affected by the weaker current, the
converse is true. We thus have two opposing torques controlling the
movement of the discs and pointer, the torques being dependent on
the ratio of currents in the coils.
The indicator, being a ratiometer, is independent of variations in
the supply voltage, but since this is alternating, it is necessary to
308
provide compensation for variations in frequency. For example. an
increase of frequency would cause the stator coils to oppose the
current changes produced by the transmitter, so that , in technical
terms, the coil reactance would increase. However, reactance changes
are overcome by the simple expedient of connecting a capacitor L,
parallel with each coil, the effects of frequency changes on a capacitor being exactly the opposite to those produced in a coil.
Changes in temperature can also have an effect on the impedance
of each coil: an increase in temperature reduces the ratio and so
makes the indicator under-read. Temperature effects are therefore
compensated by connecting a high-temperature-coefficient resistor
across the coils of the indicator.
Figure 12.7 illustrates another form of a.c. inductor type of
pressure transmitter. The sensing element construction differs from
the one already described in that it utilizes a capsule, and an :umature
that moves relative to air gaps in the stator core. With pressure
applied as indicated, the length of the air gap associated with stator
coil l is decreased, while that associated with coil 2 is increased. As
the 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 that of coil 2 will be decreased; the current
flowing in the coils will, respectively, be decreased and increased .
Another difference related to the use of this transmitter, is that its
associated indi::ator may be of the moving-coil type based on the
d.c. ratiometer principle (see also page 276).
Other types of indicator may be used with appropriate inductor
transmitters, and as will be seen from Fig 12.8 (c) and (d) the
RECTANGULAR AIR GAPS
Figure I 2. 7 Inductor pre11ure
transmitter.
SPRING
STATOR COIL 2
ARMATURE
STATOR COIL 1
309
fundamental principles already described on pages 248 and 278
can also be adopted.
Figure 12.8 Dial presentations
of pressure indicators.
(a) and (b) ratiometers,
(c) servo-operated,
(d) powered moving coil.
(a}
(b)
WARN"
LIGHTS
(c)
Pressure Switches
310
(d)
In many of the aircraft systems in which pressure measurement is
involved, it is necessary that pilots be given a warning of either low
or high pressures which might constitute hazardous operating
conditions. In some systems also, the frequency of operation may
be such that the use of a pressure-measuring instrument is not justified since it is only necessary for the pilot to know that an operating
pressure has been a~tained for the period during which the system is
in operation. To meet this requirement pressure switches are
installed in the relevant systems and are connected to warning or
indicator lights located on the cockpit or flight deck panels.
An example of a switch is illustrated in Fig 12.9. It consists of a
metal capsule open to ambient pressure, and housed in a chamber
open to the pressure source. On the other side of this cham her is an
electrical contact assembly arranged to 'make' ·on either a rising or a
falling pressure ; in the example shown, the contacts 'make' as a result
of a rise in pressure to the value pre-set by the micro-adjuster. The
capsule is constructed so that corrugations of each diagram half 'nest'
Figure 12.9 Pre~ure switch .
PRESSURE CONNECTION
ELECTRICAL CONNECTOR
FIXED CONTACT ARM
~>
VENTS
MO\/ING CONTACT ARM
CAPSULE
ACTUATING ARM
PRESSURE INLET
together when the capsule is fully contracted to fonn virtually a solid
disc which prevents damage to the capsule under an overload pressure
condition.
The pressure inlets of switch units are normally in the mounting
flange, and they may either be in the form of plain entry holes directly
over the pressure source, spigots with 'O' ring seals (as in Fig 12.9)
or threaded connectors for flexible pipe coupling.
Pressure switches may also be applied to systems requiring that
warning or indication be given of changes in pressure with respect to
a certain datum pressure; in other words, as a differential pressure
warning device. The construction and operation are basically the same
as the standard type, with the exception that the diaphragm is subjected
to a pressure on each side.
311
In some cases, a pressure switch may be incorporated with a pressure transmitter as shown in Fig 12.10.
Figure 12.10 Combined tran,-
mitter and pressure switch.
Questions
312
PRESSURE CONNECTION
J
Define the terms absolute pressure and gauge pressure.
Name some of the instruments which measure the pressures referred
to in 12.1.
12.3 Describe the operating principle of a U-tube manometer.
12..4 Briefly describe how pres~ures can be measured by balancing against
known weights.
12.S Explain the fundamental operating principle of the Bourdon tube.
12.6 Name three other types of elastic pressure-sensing elements and state
some specific applications.
12.7 Describe a method of measuring pressure based on a synchronous
transmission principle.
12.8 Explain briefly how an inductor type of pressure transmitter produces
the varying currents required for the operation of a ratiometer.
12.9 What effect does a change in frequency of the power supply have on
an inductor type of transmitter? Describe a method of compensation.
12.10 What types of indicator can be used with pressure transmitters operating on the variable inductance principle?
12.11 For what purposes are pressure switches required in aircraft?
12.12 With the aid of a diagram, explain how a pressure switch is made to
give a warning of a pressure in excess of a normal operating value.
12.1
12.2
13 Measurement of fuel
quantity and fuel flow
The measurement of the quantity of fuel in the tanks of an aircraft
fuel system is 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,
both measurements enable a pilot or. engineer to quickly assess the
remaining flight time and also to make comparisons between present
engine performance and past or calculated performance.
Fuel-quantity indicating systems vary in operating principle and
construction, the application of any one method being governed by
the type of aircraft and it5 fuel system. Two principal methods
currently applied utilize the principle of electrical signal transmission
from units located inside the fuel tanks. In one method, mainly
employed in the fuel systems of small and light aircraft, the tank
units consist of a mechanical float assembly which controls an
electrical resistance unit and varies the current flow to the indicating
element. The second method, employed in high-performance aircraft
fuel systems, measures fuel quantity in terms of electrical capacitance
and provides a more accurate system of fuel gauging.
Fuel-flow measuring systems also vary in operating principle and
construction but principally they consist of two units: a transmitter
or meter, and an indicator. The transmitter is connected at the
delivery side of the fuel system, and is an electromechanical device
which produces an electrical output signal proportional to the flow
rate which is indicated in either volumetric or mass units. In some
systems an intermediate amplifier/computer is included to calculate
a fuel-flow/time ratio and also to transmit signals to an indicator
which presents integrated flow rate and fuel consumed information.
Float-Type FuelQuantity Indicating
Systems
The components of a float-type system are shown schematically in
Fig 13.1, together with the methods of transmitting electrical signals.
The float may be of cork specially treated to prevent fuel
absorption, or it may be in the form of a lightweight metal cylinder
suitably. sealed . The float is attached to an arm pivoted to permit
angular movement which is transmitted to an electrical element
consisting of either a wiper arm and potentiometer, or a Desynn
313
Figure I 3. I Simple float
type of fuel quantity
indicator.
FLOAT
+
INDICATOR
FUEL TANK
type .of transmitter. As changes in fuel level take place the float arm
moves through certain angles and positions the wiper arm or brushes
to vary the resistance and flow of direct current to the indicator.
As a result of the variations in current flow a moving coil or rotor
within the indicator is deflected to position a pointer over the scale
calibrated in gallons.
Capacitance-Type
Fuel-Gauge System
In its basic form, a capacitance-type fuel-gauge system consists of a
variable capacitor located in the 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 quantity.
Before going into the operating details of such a system, however,
it is first necessary to discuss some 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
1
is known as capacitance.
The flow of a momentary current into a capacitor establishes a
potential difference across its plates. Since the dielectric contains
no free electrons the current cannot flow through it, but the
potential difference sets up a state of stress in the atoms comprising
it. For example, in the circuit shown in Fig 13.2, when the switch
is placed in position 1 a rush of electrons, known as the charging
current, takes place from plate A through the battery to plate B and
ceases when the potential difference between the plates is equal to
that of the battery.
314
2
Figure 13.2 Charging and
0
discharging of a capacitor.
/
I
~---
I
ll
r1
A
B
>->.-4-+-'~+;.;;+;..;;+~
I
I,
~I
I
I
1
t
I•
f
I
1
I
,',
I
ti
-------
-
f +I ____ ...
- - CHARGING CURRENT
- - - - DISCHARGING CURRENT
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 of the capacitor.
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 13. 2.
Units of Capacitance
The capacitance or 'electron-holding ability' of a capacitor is the
ratio between the charge and the potential difference 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 potential difference 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 micro farad (I µ F = l 0-6 F).
In the application of the capacitor principle to fuel gauge systems, an
even smaller unit, the picofarad (I pF = 10-12 F) is the standard unit
of measurement .
Factors on which Capacitance Depends
The capacitance of a parallel-plate capacitor depends on the area, a,
of the plates, the distanced, between the plates, and the capacitance,
ea, of a unit cube of the dielectric material between the plates:
C = Ea
da
farads
The unit of Ea is the farad per metre, so that a must be expressed in
315
square metres, and din metres; Ea is called the absolute permittivity
of the dielectric.
It is usual to quote pennittivities relative to that of a vacuum,
whose permittivity, e0 , is 1/ (4rr X 9 X 109 F)/m. Relative permittivity
e, is also called dielectric constant and is often denoted by K. In
terms of relative permittivity,
K
a
C = 4rr X 9 X 109 d
The relative permittivity of air at standard temperature and pressur,
is 1.00059, which for practical purposes may be taken as 1.0. Then,
for example,
·
C water = Kwater :;;; K water
Ca;,
Kair
very closely
i.e. K is the ratio of the capacitance of a capacitor with a given dielectric to its capacitance with air betwe1m its plates.
The relative permittivities of some pertinent substances are as
follows:
Air .
Water
Water vapour
Aviation gasolene
Aviation kerosene
l ·00059
81 ·07
1·007
1·95
2· 10
Capacitors in Series and Parallel
The total capacitance of capacitors connected in series or parallel is
obtained from formulae similar to those for calculating total resistance but applied in the opposite manner. Thus, for capacitors
connected in series, the total capacitance is given by
1 1
1
1
-=-+-+- +
Cr C1 C2
C3 . . .
and for capacitors connected in parallel,
Cr = C1 + C2 + C3 + . ..
This is because the addition of capacitors in a circuit increases the
plate area which, as already stated, is one of the factors on which
capacitance depends.
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
316
through the capacitor. In applying the capacitance principle to fuelgauge systems, however, a flow of current is necessary to make the
indicator respond to the changes in capacitance arising from changes
in fuel quantity. This is accomplished by supplying the capacitancetype tank units with an alternating voltag_e, 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.
Figure 13.3 shows a capacitor connected to an alternating-current
source. It will be observed from the graph that, as the voltage ( V)
rises rapidly at A a large current (!) will flow into the capacitor to
charge it up. As the voltage increases towards B, however, the current
decreases until at B, when the voltage is steady at some maximum
value for a brief instant, the current has decreased to zero. From B to
C the voltage decreases, the capacitor discharges and the current flows
in the opposite direction, being a maximum at C, where the voltage is
zero. From C to D the capacitor is charged in the opposite direction
and the current flows in the same direction as the voltage but reaches
zero at D, where the voltage is at some maximum value in the
opposite direction. From D to E the capacitor again discharges.
Thus, a charge and current flow in and out of the capacitor occurs
every half-cycle, the current leading the voltage by 90° .
Figure l J. J Capacitance in
VOLTAGE V
an a.c. circuit.
The ratio between the voltage V and the current/ is termed the
capacitive reactance, meaning the opposition or resistance a capacitor
offers to the flow of alternating current.
Basic Gauge System
For fuel quantity measurement, the capacitors to be installed in the
tanks must differ in construction from those normally employed in
electrical equipment. The plates therefore take the form of two
tubes mounted concentrically with a narrow air space 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,
317
while the third factor, dielectric constant, is variable since the mediun
between the tubes is made up of fuel and air. The mariner in which
changes· in capacitance due to fuel and air take place is illustrated in
Fig 13.4 and is described in the following paragraphs.
Figure 13.4 Changes in capaci:tance due to fuel and air.
I
- ·--- -u
IH
- ·- - - - -
r--t-
-
-- -
- ·--
- - --
(a)
c..,, =100pF
(b) K = 2 ·1
(cl
·---- -- - -.h
H
~
-
L
1.
2
At (a) a tank capacitor is fitted in an empty fuel tank and its capacitance in air is l 00 pF, represented by CA.
At (b) the tank is filled with a fuel having a K value of 2.1, so that
the capacitor is completely immersed. As stated on page 316 K is
equal to the ratio of capacitance using a given dielectric (in this case
Cr) to that using air; therefore,
(1)
From eqn (I) Cr = CA K, and it is thus clear that the capacitance of
the tank unit at (b) is equal to I 00 X 2.1, i.e. 210 pF. The increase of
110 pF is the added capacitance due to the fuel and may be representc
by Cp. The tank unit may therefore be represented electrically by tw
capacitors in parallel and of a total capacitance
(2)
In Fig 13.4(c), the tank is only half full and so the total capacitanct
is 100 + 55, or 155 pF. The added capacitance due to fuel is determined as follows. By transposing eqn (2), Cp = Cr - CA , and by sub·
stituting CA K for Cr we obtain Cp = CA K - CA, which may be
simplified as
(3)
the factor (K - l) being the increase in the K value over that of air.
Now, 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
eqn (3) the complete formula becomes
Cp
=nL (K -
l)CA.
The circuit of a basic gauge system is shown in Fig 13. 5. It is
divided into two sections or loops by a resistance R, both loops
318
(4)
TANK UNIT
Vs
1
R
/'
\.,..,.,,
' - - - - , I - - - - ' FUU.
11
BALANCE
Ii
POTENTIOMETER Ii
11
11
I
2-PHASE
ffi==JL=0
TOR
-
Figure 13.S Circuit of a basic
capacitance fuel-gauge system.
-
being, connected to the secondary winding of a power transformer.
Loop A contains the tank capacitor 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 the voltage VB 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 connected to the amplifier and is only energized when an
unbalanced condition exists in the bridge.
The amplifier, which is based on solid-state circuit techniques
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 Is is equal to the balancing current Is ;
the bridge is thus in balance and no signal voltage is produced across
R.
.
As the fuel level drops, the tank capacitor has Jess fuel around it;
319
therefore the added capacitance (Cp) has decreased. The tank unit
capacitance decreases and so does the sensing current Is, the latter
creating an unbalanced bridge condition with balancing current Is
predominating through R.
A signal voltage proportional to lsR is.developed across R and is
amplified and its ·phase detected before being applied to the control
phase of the indicator motor. The output signal is a half-wave pulse,
a feature of transistors in discriminator circuits, and in order to
convert it into a full-wave signal, a capacitor is connected in
parallel with the control winding. A capacitor is also connected in
series with the reference winding to form what is termed a seriesresonant circuit. This circuit ensures that the currents in both
phases are 90° out of phase, the current in the control phase either
leading or lagging the reference phase depending on which loop of
the bridge circuit is predominating.
In the condition we are considering, the balancing current is
predominating: therefore , the control-phase current lags behind that
of the reference phase causing the motor and balance potentiometer
wiper to be driven in such a direction as to decrease the balancing
current Is.
When the current Is equals the current Is, the bridge is once again
in balance, the motor stops rotating and the indicator ·pointer registers
the new, lower value.
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 13.6, which is a graph of the approximate changes
occurring in a given mass of fuel. From this it should be noted that
K - I is plotted, since for a gauge system measuring fuel quantity
by volume, the indicator pointer movement is directly dependent on
Figure 13. 6 Temperature
effects on fuel characteristics.
+5 ~~-,---.---r---;,
0
TEMPERATURE. 'C
320
this. It should also be noted that, although it varies in the same way
as density, the percentage change is greater.
Thus a volumetric gauge system will be subject to a small error due
to variations in fuel temperature. Furthennore, changes in K and
density also occur in different types of fuel having the same temperature. For example, a gauge 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
1.3
UX
GI.
100- 100,o = 18%
i.e. the gauge would over-read by 18%.
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 on the volume of fuel but on the energy it contains, i.e.
the number of molecules that can combine with oxygen in the .engine.
Since each fuel molecule has some weight and also because one·pound
of fuel has the same number of molecules regardless of temperature
and therefore voluine, 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 pevice
must therefore be sensitive to both changes in volume and density so
as to eliminate the undesirable effects due to temperature.
This will be apparent by considering the example of a tank helding
1,000 gal of fuel having a density of 6 lb/gal at normal temperature.
Measuring this volumetrically we should of course obtain a reading of
1,000 gal , and from a mass measurement, 6,000 lb. If a temperature
rise should increase the volume by 10%, then the volumetric measurement would go up to 1,100 gal, but the mass measurement would
remain at 6,000 lb because the density of the fuel (weight/volume)
would have decreased when the temperature increased.
For the calibration of gauges 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
A gauge system calibrated to this expression is still subject to indi321
cation 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.
Now, if-the same gauge 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 - l
0.85 = 1.529.
However, the percentage error is now
1.529
l .4 l 2 X l 00 - I 00% = 8% approx.
and this is the amount by which the gauge would over-read.
Compensated Gauge
Systems
Although the assumed constant permittivity and density relationship
results in a reduction of the indicator error, tests on properties of
fuels have shown that, while the capacitive varies from one fuel to
another, this variation tends to follow the permittivity. Thus, if a
gauge 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.
The gauge systems now 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 13.7.
A compensator is similar in construction to a standard tank unit
and is usually fittyd to the bottom of a unit to ensure that the
compensator is always immersed in fuel. Located in this manner, its
Figure 13. 7 Circuit connec-
TANK UNIT
tion of compensator capacitor.
I
o
itNSFORMER {
TAPPINGS
-
t...= .. =.J
-
-
---'INV>.MN----1
CR
TO BALANCE
POTENTIOMETER
-•- - - -..14
y--0-.....--- -'
C coMP
(BOTTOM Of TANK UNll)
322
]
BALANCE
LOOP
capacitance is determined solely by the permittivity, or K-value, of
the fuel, and not by its quantity as in the case of the tank unit. In
addition to variable voltage, the balancing loop will also be subjected
to variable capacitance, which means that balancing current ls will
be affected by variations in K as well as sensing-loop current ls .
Let us assume that the bridge circuit (see Fig 13.5) is in balance
and that a change in temperature of the fuel causes its K-value to
increase. The tank unit capacitance will increase and so current ls
will predominate to unbalance ,the bridge and to send a signal voltage
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 towards the 'tank full' condition, the indicator
will obviously register an increase in fuel quantity. The increase in K,
however, also increases the compensator capacitance so that balancingloop current ls is increased simultaneously with, but in opposition to,
the increase o f sensing-loop current ls. A balanced bridge condition
is therefore obtained which is independent of the balance potentiometer.
In practice, there is still 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 capacitance values in
conjunction with the reference capacitance, the greatest redu~tion in
overall gauge error is produced.
Location and
Connection of Tank
Units
In practical capacitance-type fuel-quantity indicating systems, a
number of tank units are disposed within the fuel tanks as illustrated
in Fig 13.8, and are connected in parallel to their respective '
indicators. The reasons for this are to ensure that indications remain
the same regardless of the attitude o f an aircraft and its tanks and also
of any surging of the fuel. This may be understood by considering a
two-unit system as shown in Fig 13.9. If the tank is half-full and in a
level at"titude, each tank unit will have a capacitance of half its maximum value, and since they are connected in parallel the total capacitance measured will produce a 'half-full' indicatio n. When the tank is
tilted , and because the fuel level remains the same, unit A is immersed
deeper in the fuel by the distance d and gains some capacitance,
tending to make the indicator over-read. Tank unit B, however, has
moved out of the fuel by the same distanced and loses an equal
amount of capacitance. Thus, the total capacitance remains the same
as for the level-tank attitude and the indication is unchanged.
In the majority of cases, tank units are designed for internal
mounting, their connections being brought out through terminals in
323
STANDARD TANK UNIT
LEVEL INDICATOR
COMPENSATOR UNIT
Figure 13.8 Location of tank
sensing units.
Figure 13.9 Attitude
compensation.
324
the walls of the tanks. For some aircraft tank systems flange-mounted
units may also be provided.
A typical standard tank unit and a compensator unit are shown at
(a) and (b) respectively of Fig 13 .I 0. The tubes are of aluminium
alloy held apart by pairs of insulating cross-pins. Electrical connections are made through coaxial connectors mounted on a bracket
attached to a nylon sleeve secured to the upper part of the outer tube.
Two further nylon sleeves, one at each end, are secured to the outer
tube, and to each sleeve a rubber ring is attached . The purpose of the
6
Figure 13. JO
Taruc units.
(a) Standard unit - 1 rubber
rubber ring is to hold the unit in position at supp'o rting fixtures within
the tank. The reference unit of the compensator consists of three
concentric tubes held apart by insulating buttons; the outer of the
three tubes acts as an earth screen.
ring, 2 n)tlon sleeve, 3 outer
tube, 4 inner tube, 5 rubber
ring, 6 nylon sleeve, 7 insulating crost-pin, 8 bracket,
9 miniature coaxial connector;
Characterized Tank Units
(b) compensator unit - 1 rubber
ring, 2 nylon sleeve, 3 outer tube, The fuel tanks of an aircraft may be separate units designed for
4 insulating cross-pin, 5 rubber
.
all a t·10n m
· wmgs
·
· cen t re sec t 10ns,
·
ring 6 inner tube, 7 nylon
mst
and m
or t h ey may form an
sl.ee;es, s reference unit, 9
integral sealed section of these parts of the structure. This means,
bracket, 10 miniature coaxial
therefore, that tanks must vary in contour to suit their chosen
connector.
locations, with the result that the fuel level is established from varying datum points.
Figure 13.11 represents the contour of a tank located in an aircraft
wing, and as will be noted the levels of fuel from points A, B and C
are not the same. When standard tank units are positioned at these
points the total capacitance will be the sum of three different values
due to fuel (CF), and as the units produce the same change of
capacitance for each inch of wetted length, the indicator scales will
be non-linear corresponding to the non-linear characteristic of the
tank contour.
325
Figure J3. 11 Non-linear fuel
B
A
C
tank.
The non-linear variations in fuel level are unavoidable, but the
effects on the graduation spacing of the indicator scale can be overcome by designing tank units which measure capacitance changes
proportional to tank contour. The non-linear tank units so designed
are called characterized tank units, the required effect being achieved
either by altering the diameter of the centre electrode or by varying
the area of its conducting surface at various points over its length, to
suit the tank contour.
Empty and Full Position Adjustments
In order that an indicator pointer shall operate throughout its range
between the two principal datums corresponding to empty- and fulltank conditions, it is necessary during calibration to balance the
current and voltage of the circuit sensin_g and balancing loops at these
datums. This is achieved by connecting two manually controlled
potentiometers into the circuit a!\ shown in Fig 13.12.
REFER1:NCE
CAPACITOR
BALANCE
POTENTIOMETER
"FUU.· ADJUSTMENT
- - ~ POTENTIOMETER
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 tank units, current will still flow through them. The balance
potentiometer wiper will also be at its empty position, but since it is
earthed at this point, no current will flow through the reference
capacitor. However, current does flow through the balance capacitor
326
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 and servomotor to the
empty position.
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 balance potentiometer.
Reference to Fig 13. 12 shows that, if the full potentiometer wiper is
set at the bottom, the full transformer secondary voltage will be
applied to the balance potentiometer. With the wiper at the top,
resistance is introduced into the circuit so that a smaller voltage is
applied to the balance potentiometer. Therefore, the distance
the balance potentiometer wiper needs to move to develop a given
balance voltage can be varied.
Fail-Safe and Test Circuits
In the event of failure of the output from an amplifier there can be
no response to error and balancing signals, and to ensure that a pilot
or engineer is not misled by an indicator remaining stationary at
some fuel quantity value, a fail-safe circuit or a test circuit is
incorporated in a gauge system.
A typical 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 the tank unit
signals flowing through the motor control winding. When these
signals cease due to a failure, the current in the parallel circuit
predominates and flows through the control winding to drive the
pointer slowly downwards to the empty position.
A test circuit 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 operable. When the switch is released
the pointer should return to its original indicatio n.
Fuel Quantity
Totalizer Indicators
In several types of public transport aircraft, the fuel quantity
indicating system incorporates an additional indicator known as a
'fuel totalizer'. A typical dial presentation is shown in Fig 13.13,
and from this it will be noted that digital counters display not only
the total quantity of fuel remaining, but also the aircraft's gross
weight.
327
Figure I 3. I 3 Totalizer
indicator. (By courtesy of
Smiths Industries.)
The indicator comprises basically a resistance network having 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 currents according to the amount of
fuel that each voltage represents, and then sums the currents to
represent total fuel quantity 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 preset to indicate the aircraft's
weight pr(or to flight, is also connected to·the primary fuel quantity
indicator system, but in such a manner that as fuel is consumed, the
counter continuously indicates a decreasing gross weight.
Volumetric Top-Off
Systems
328
In some aircraft, a system designated 'volumetric top-off' (VTO) is
provided which operates in conjunction with the tank sensing units
of the primary fuel quantity indicating system. The purpose of VTO
is to automatically terminate a tank refuelling operation at a preset
fuel level thereby permitting maximum volume fuel loading and to
allow proper expansion space in the tank.
A separate system is installed for each fuel tank, and comprises
a VTO unit, a compensator tank unit; and a set of tank units which
are shared with those of the primary indicating system. The output
from the VTO unit is supplied to the fuelling system shut-off valve.
During refuelling operations, the primary indicating system records
the increasing fuel quantity in the normal units of mass, and the
corresponding signals are applied to a bridge circuit in the VTO unit.
The signals from the VTO compensator relate to tank unit capacity
in volumetric units, and these are also applied 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 preset nominal full level, or volume, of the
fuel tank is reached . Calibration of the whole VTO system to the
desired nomlnal 'tank full' volume is accomplished by adjusting the
current through the VTO compensator so that it corresponds with
the total current from the tank units at th~ particular shut-off level.
Refuelling Control
Panels
On large aircraft, the location of the primary fuel quantity indictors
is far removed from the refuelling points on the aircraft, making it
difficult for the ground-handling crew to monitor the refuelling
operation. Furthermore, the quantity of fuel uplifted must be held
to within acceptable limits of accuracy to avoid carrying of unwanted
fuel. As a solution to these problems, control panels are located
adjacent to the refuelling points.
Figure 13.14 illustrates a panel which is used in conjunction with a
volumetric top-off system. 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 the refuelling operation. The switches below the indicators control the positions of their
respective tank refuelling valves.
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 on the appropriate valve
switch. When, however, a full load of fuel is required, the valves are
automatically closed by the shut-off action of the VTO and in the
manner described on page 328. On completion of a refuelling operation, the control panel is enclosed by a door which also actuates a
Figure 13. J4 Refuelllna
control panel.
TANK NO. 2
-
OFF~))
AUX. FUELING
POWER- CONTROL
,<::,...OPEN
«.u,1
CLOSED
CENTER TANK
_.,~OPEN
TANKN0. 1
6
:,.?PEN
1?r:)'1
1fr'";)'1
UcLoseo
UCLOSED
VALVE
CONTROL
SWITCHES
329
switch that isolates all electrical power to the fuelling system.
Another type of control panel in current use is shown in Fig 13.15.
In this case, the amount of fuel required is selected on load limit
controls prior to operating the refuelling valve switc~es. Each load
limit control has a dial calibrated similarly to those of the primary
quantity indicators, and a manually-adjusted knob and pointer.
The knob is coupled to a potentiometer which, together with the
re-balancing potentiometer of the respective primary indicator,
forms part of a. bridge circuit. When the required fuel load is selected
the load control potentiometer establishes a certain signal voltage
level in the bridge circuit and energizes a relay within the load control
The contacts of this relay are also in the control circuit to the
refuelling valve. As the appropriate tank is replenished, the
rebalancing potentiometer in the primary indicator will change the
signal voltage conditions of the bridge circuit until, at a value below
that set by the load control potentiometer, a polarity change causes
the relay to de-energize and thereby close the refuelling valve.
Figure 13. 15 Fuel load
limit control panel.
TANK
SELECTED
(g
TANK
SELECTED
0
REFUEL AUTO REFUEL AUTO
~
~
~
~
~OFF~
v&/oFF~
DEFUEL MANUAL DEFUEL MANUA
Fuel Flow Measure-
ment
TANK
SELECTED
( g + - - - - t - - - INDICATOR LIGHTS
REFUEL AUTO
~
~
VALVE CONTROL
~OFF~+----t---SWITCHES
DEFUEL MANUAL
In analysing currel'lt designs of fuel-flow measuring systems it will be
found that they come within two main groups: (i) independent fuel
flow, and (ii) integrated. There are a variety of types in use and it
is not possible to go into the operating details of them all. However,
a system which may be considered representative of each group has
been selected to illustrate the applications of fundamental requirements and prh:.:iples.
Independent Fuel Flow System
This system consists of a transmitter and indicator and requires 28 V
direct current for its operation.
330
The transmitter, shown in Fig I 3. I 6, 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 metering 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. Thus, its function
may be likened to an airspeed-indicator square-law compensator
which it will be recalled is a device for opening up an indicator scale.
The vane is mounted on a shaft carried in two bushed plain bearings,
one in each cover plate enclosing the metering chamber.
Figure J3.16 Section view of
ev-PIISS VALVE
a rotating-vane fuel-flow
transmitter.
CALIBRATED
SPRI NG
VANE
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. In this particular system the unit is a precision potentiometer; in some designs an
a.c . synchro may be used. The shaft of the potentiometer (or
synchro) 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 also protrudes through
its -bearing and carries the attachment for the inner end of a specially
calibrated control spring. The outer end of the spring is anchored to
a disc plate which can be rotated by a pinion meshing with teeth cut
331
in the periphery of the plate . This provides for adjustment of the
spring torque during transmitter calibration.
Any tendency for the metering assembly and transmission element
to oscillate under static flow rate conditions is overcome by a liquid
damping system, the liquid -being the fuel itself. The system comprise:
a damping chamber containing a counterweight and circular vane
· which are secured to the same end of the metering vane shaft as the
control spring. The damping·chamber is secured at one side of the
transmitter body, and except for a small bleed hole in a circular
blanking plate, is separated from the metering chamber. The purpose
of the bleed hole is, of course, to permit fuel to fill the damping
chamber and thus completely immerse the counterwight assembly.
The effectiveness of the damping system is uninfluenced by the fuel
flow. A threaded plug in the outer cover of the damping chamber
provides for draining of fuel from the chamber.
The indicator is of simple construction, being made up of a moving·
coil milliammeter which carries a single pointer operating· over a scale
calibrated in gallons per hour, pounds per hour or kilogrammes per
hour. The signals to the milliammeter are transmitted via a transistorized amplifier which is also contained within the indicator case.
In systems employing synchronous transmission, the indicator pointer
is operated by the rotor of a receiver synchro.
Operation
When fuel commences to flow through the main supply line it enters
the body of the transmitter and passes through the metering chamber.
In doing so it 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 only to take up an angular position at which the tension
of the spring 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 to be displaced, 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 the
amplifier, whose output current drives the milliammeter pointer to
indicate the fuel flow.
In a system employing synchros, the current flow due to differences in angular position of the rotors will drive the indicator
synchro rotor directly to the null position and thereby make the
indicator pointer read the fuel flow.
In the type of transmitter considered it is also necessary to provide
a by-pass 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.
It will be noted from Fig 13.16 that the valve is of the simple springloaded type incorporated in the metering chamber. The spring
332
tension is adjusted so that th e valve lift s from its seat and allows fuel
to by-pass the metering chamber when the difference of pressure
across the chamber exceeds 2.5 lbf/in 2 •
Integrated Flo1N111eter
Syste111
We may broadly define an integrated f[owmeter system as one in
which the element indicating fuel consumed is combined with that
required for fuel flow, thus permitting the display of both quantities
in a single instrument.
In order to accomplish this it is also necessary to include in the
system a device which will give directly the fuel consumed over a
period of time from the flow rate during that period . In other words,
a time integrator is needed to work out fuel consumed in the ratio
of fuel flow rate to time.
Such a device may be mechanical, forming an integral part of an
indicator mechanism, or as in electronic flowmeter systems it may be
a special dividing stage within the amplifier or even a completely
separate integrator unit. A typical system to which principles of
integration are applied will now be studied.
The system consists of three principal units : flow transmitter,
electronic relay or computer, and indicator. Its operation depends
on the principle that the torque required to accelerate a fluid to a
given angular velocity is a measure of the fluid's mass flow rate.
The angular velocity, which is imparted by means of a rotating
impeller and drum, sets up a reaction to establish relative angular
displacements between the impeller and drum. Inductive-type pickoffs sense the displacements in terms of signal pulses proportional
to the flow rate and supply them via the amplifier/computer, to
the indicator.
The transmitter, shown schematically sectioned in Fig 13.1 7,
consists of a light-alloy body containing a flow metering chamber,
a motor-driven impeller assembly, and an externally mounted
inductor coil assembly. 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 to impart angular velocity to fuel flowing through the
metering chamber. The drum and impeller are coupled to each
other by a calibrated linear spring. The motor is contained within a
fixed drum at the inlet end and rotates the impeller at a constant
speed. Straightening vanes are provided in the fixed drum 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 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 fuel by enclosing them in a chamber which is
333
Figure I 3.17 Fuel-flow
transmitter of a typical
integrated system. I Turbine,
2 fluid passage, 3 shaft,
4 light-alloy body, S fluid
passage, 6 impeller, 7 restraining spring, 8 pick-off assembly,
9 magnets, 10 pick-off assembly,
I 1 magnetic coupling, 12 rotor.
TO TRANSISTORIZED
BISTABLE SWITCH
Figure 13. 18 Integrated
fuel flowmeter system.
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 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 transistorized units
which amplify and switch the signals induced.
The computer performs the overall function of providing the
power for the various circuits of the system, detecting the number
of impulses produced at the transmitter, and computing and
SUPPLV TO TRANSMITTER SYNCHRONOUS MOTOR
,----------0
PICK· OFF SIGNALS
TRANSMITTER
REFEREN1
WINDING
SIGNALS
I
FLOW RATEI CONTROL
WINDING SIGNAL
-orv
AMP.
OMO"R.
I
3.
---+---- -
~LOW RATE SIGNALS
COMPLITER
FEEDBACK SIGNALS FOR COMPARISON
FUEL CONSUMED SIGNALS
334
integrating the fuel flow rate and amount of fuel consumed. It
consists of a number of stages interconnected in three distinct
sections as shown in the block diagram of Fig 13.18: The stages of
sections 2 and 3 consist of transistors and their associated coupling
capacitors and resistors. The power supply section (1) controls the
voltage and frequency of the supply to the transmitter synchronous
motor, and consists of a transformer, transistorized crystal oscillator,
output and power amplifier units.
From the diagram it will be ncted 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; to provide the time dividing factor ·,nd 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 alternating current
and produce a new output; to provide an operating signal to the
indicator servomotor control winding.
The indicator employs a flow indicating section consisting of a
400 Hz servomotor which drives a pointer and 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 computer servo amplifier. The
potentiometer is supplied with direct current and its wiper is
electrically connected to a transistorized time-base section, also
within the indicator. Transmitter output signals are also fed into the
time-base section via a pre-set potentiometer which forms part of the
computer signal 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 5-drum digital counter and a pulse amplifier. The amplifier
receives a pulse from the divider stage of the computer for each unit
mass of fuel consumed and feeds its output to the solenoid, which
advances the counter drums appropriately . A mechanical reset button
is provided for resetting the counter to zero.
Operation
When electrical power is switched on to the system, the synchronous
motor in the transmitter is operated to drive the impeller assembly at
a constant speed. Under zero fuel flow conditions the magnets of
the pick-off assemblies are effectively in line with one another,
335
although in practice there is a small angular difference established to
maintain a deflection representing a specific minimum flow rate.
This is indicated in Fig 13.,19 (a). As the fuel flows through the transmitter 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 a calibrated spring, a reaction torque is
created which alters the angular displacement between impeller and
drum, and their corresponding magnets. Thus angular displacement
is proportional to flow rate. Figs 13.19 (b) and (c) illustrate the
Figure 13. 19 Operation of
transmitter pick-offs.
(a) Zero fuel flow -A, Two
pick-off coils (one behind the
other), B, C, Ml!gnets, D, s·top
(gives 3° to 5° deflection),
8 Lag angle at which both
drums rotate together; (b)
cruising fuel flow; (c) maximum fuel flow ..
A
'
OUTER & INNER DRUMS
INTERCONNECTED BY
CALIBRATED SPRING
(al
'
'
(b)
displacement for typical cruising and maximum fuel flow rates.
The position of each magnet is sensed by its own pick-off coil, and
the primary pulses induced as each magnet moves past its coil are
fed to the dividing stage in the computer (see also Fig 13. 18). The
output from this stage is fed to the control winding of the indicator
servomotor via section 3 of the computer, and the indicator pointer
is driven to indicate the fuel flow. At the same time, the motor drives
the potentiometer wiper, producing a signal which is fed back to the
signal comparator stage and compared with the output produced by
the transmitter. Any resultant difference signal is amplified, modulated and power amplified to drive the indicator motor and pointer
to a position indicating the actual fuel flow rate.
336
The computer divider stage 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
for each kilogramme or pound of fuel consumed .
Questions
13.1
13.2
13.3
13.4
13.5
13.6
13.7
13.8
13.9 ·
13.10
13.11
13.12
13.1 3
13.14
13.15
13.16
13. 17
13.18
13.19
Describe the construction and operation of a float type of fuel-quantity gauge with which you are familiar.
Explain the operating principle of a capacitor and state the factors
on which its capacitance depends.
Define (a) the units in which capacitance is expressed, (b) permittivity.
(a) Explain how alternating current appli ed to a capacitor causes
current to flow across it.
(b) What is the term used to express the ratio of voltage to current?
Give the formula for calculating the total capacitance of a circuit
containing (a) capacitors in series, (b) capacitors in parallel.
(a) Which of the capacitance variables is utilized in a capacitancetype fuel-quantity rystem?
(b) Describe the construction of a typical tank unit.
Draw a circuit of a typical capacitance-type fuel-quantity indicating
system. Explain the operating principle.
(a) What effects do temperature changes have on the fuels used ?
(b) Explain how these are compensated in a fuel-quantity indicating
system.
.
Explain why it is necessary to install a number of tank units in a fuel
tank.
Why is it preferable for fuel-qu antity indicating systems to measure
fuel weight rather than fuel volume?
What adjustments are normally provided in a capacitance-type fuelquantity indicating system?
Explain the function of the test switch incorporated in some fuelquantity indicating systems.
Explain briefly how the total fuel remaining may be indicated.
What is the purpose of a 'volumetric top-off' system? Describe
briefly the operation of such a system.
Why are refuelling control panels provided on some types of aircraft?
Describe a method of refuelling control.
Describe the construction of a fuel fl owmeter indicator and ex plain
the basic principle of operation.
What is the function of the by-pass valve as fitted to certain types of
fuel flowmeter?
(a) What is an integrated flowmeter system?
(b) Describe a method of achieving integration.
337
14 Eng·ine power and
control instruments
Power of piston engines, turbojet and turboprop engines, refers to the
amount of thrust available for propulsion and is expressed as power
ratings in units of brake or shaft horsepower (bhp or shp) at a propeller shaft or in pounds of thrust at the jet pipe.
The power ratings of each of the various types of engine are determined during the test-bed calibration runs conducted by the manufacturer for various operating conditions such as take-off, climb and
nonnal cruising. For example, the rated thrust of a turbojet engine at
take-off may be limited to five minutes at a turbine temperature of
795 °C. This does not mean, of course, that the engine is suddenly
going to disintegrate should these limitations be exceeded, but
frequent excesses will obviously set up undesirable stresses leading to
deterioration of vital internal parts of the engine.
All ratings are therefore established with a view to conserving the
lives of engines and contributing to longer periods between overhauls.
It is thus necessary for engines to be instrumented for the indication
of power being developed. The instruments associated with power
indication for the various types of engine are listed in Table 14.1.
Table 14.1 Power indication instruments
Type of engine
Indicator
Piston
Turbojet
Unsuper- Supercharged charged
Tachometer
X
Centrifugal Axial flow
compressor compressor
X
X
X
Fuel flow
X
X
X
Manifold pressure
X
Torque pressure
X
Exhaust gas temp
Percentage thrust
338
X
X
X
X
or
Pressure ratio
Turboprop
X
X
Power Indicators For
Reciprocating Engines
The power of an unsupercharged reciprocating engine is dire~tly
related to its speed, and so with the throttle at a corresponding
setting, a tachometer system (described in Chapter 10) also serves as
a power indicator.
With the supercharged engine, however, additional parameters
come into the picture; fuel flow, increase of pressure ('boost'
pressure) produced by the supercharger at the induction manifold,
and on some engines either torque-meter pressure or brake mean
effective .pressure (bmep ). Thus, power is monitored on four instruments. Fuel-flow indicators having already been described in
Chapter 13, the following descriptions are con fined to the remaining
three instruments.
Manifold Pressure Gauges
Manifold pressure gauges or 'boost' gauges as they are colloquially
termed, are of the direct-reading type and are calibrated to measure
absolute pressure in inches of mercury. Before considering a
typical example, it will be useful to outline briefly the general
_principle involved in the supercharging of an aircraft engine.
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 k'nown 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
sea-level pressure of 14. 7 lbf/in 2 , and the cylinder pressure is reduced
to say, 2 lbf/in 2 , then the pressure difference is 12.7 lbf/in 2 , and it is
this pressure difference which '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 altitud e of l 0,000ft.
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/ in 2 surrounding
the engine and 1 \/2 lbf/in2 in each cylinder, leaving us with a little
more than 8.5 lbf/in 2 with which to 'push' in the useful charge.
This means then at 10,000ft 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 powe,; developed at sea-level.
339
This limitation on the high-altitude performance of a normallyaspirated engine can be overcome by artifically increasing the available
pressure so as to maintain as far as possible a sea-level value in the
induction system. The process·of in~reasing 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 between 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 vanes and a volute, the reduction in speed giving
the -required increase in pressure.
In order to measure the boost pressure delivered by the supercharger and so obtain an indication of engine power, it is necessary
to have a gauge which indicates absolute pressure. The mechanism
of a typical gauge working. on this principle .is illustrated in Fig 14.1.
The measuring element is made up of two bellows, one open to the
induction manifold and the other evacullted and sealed. A.controlling spring is fitted inside the sealed bellows and distension of both
bellows is transmitted to the pointer via the usual lever, quadrant
and pinion mechanism. A filter is located at the inlet to the open
bellows, where there is also a restriction to smooth out any pressure
surges.
When pressure is admitted to the open bellows the latter expands
Figure 14.1 Typical manifold
pressure gauge.
340
causing the pointer to move over the scale (calibrated in inches of
mercury) and so indicate a .change in pressure. 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 Pressure Indicators
These indicators supplement the power indications obtainable from
tachometers and manifold pressure gauges by measuring the pressures
created by a ·torquemeter system, such pressures being interpreted
as power available at the propeller shaft.
The torquemeter system forms part of the engine itself and is
usually built-in with the reduction-gear assembly between the crankshaft and propeller shaft. The construction of a system depends on
the type of engine, but in most cases the operation is based on the
same principles ; i.e. the tendency for some part of the reduction
gear to rotate is resisted by pistons working in hydraulic cylinders
secured to the gear casing. The principle is shown diagrammatically
in Fig 14.2.
Oil from the engine system is supplied to the cylinders via a
special torquemeter pump and absorbs the loads due to piston movement. The oil is thus subjected to pressures which are proportional
Figure 14. 2 Principle of torquemeter.
STATIONARY
RING GEAR
STATIONARY
RING GEAR
/,
'
"
PLANET GEARS
_, '-)==~== = ====-===-=
TORQUE PRESSURE
INDICATOR
----
DIRECTION IN WHICH RING
GEAR TENDS TO ROTATE
DIRECTION OF CRANKSHAFT
ROTATION
OlRECTION OF PROPEll ER
SHAFT ROTATION
341
to the applied loads or torques and are transmitted to the torque
pressure-indicating system, which is normally of the remoteindicating synchronous type.
The brake horsepower is calculated by the following formula:
BHP=pN/ K
where p is the oil pressure, N the speed (rev./min.) and Ka torquemeter constant derived from the reduction gear ratio, length of
torque arm, and number and area of pistons.
Power Indicators
For Turboprop
Engines
Turboprop 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 jet
thrust. They are therefore fitted with a torquemeter and pressure
gauge system of which the oil pressure readings are an indication of
the shaft horsepower. The torquemeter pressure gauge is used in
conjunction with the tachometer and turbine gas-temperature
indicators.
The indicating system used is governed by the particular type of
engine, but there are two main systems in current use and their
operating principles are based on the Desynn and alternatingcurrent synchro methods of transmission.
Desynn Torque Pressure Indicating System
This system operates on the slab-Desynn principle described on
page 228 and is used on Rolls-Royce Dart turboprop engines. In
common with the other Desynn systems the transmitter, shown in
Fig 14.3, comprises both mechanical and electrical elements.
The mechanical element consists of a Bourdon tube the open end
of which is connected by a flexible hose to the supply line from the
oil pump of the engine torquemeter. The free end of the Bourdon
tube is connected to the brushes of the electrical element via a
sector gear and pinion. A union mounted adjacent to the main
pressure connection is connected to a capillary tube accommodated
within the Bourdon tube and allows for the bleeding of the system.
The transmitter is mounted in a special anti-vibration mounting, the
whole assembly being secured to the engine itself.
With the engine running the pressure produced at the pistons of
the torquemeter system is sensed by the transmitter Bourdon tube,
the free end of which is distended so as to change the radius of the
tube. The movement of the free end is magnified by the sector and
pinion, which causes the brushes to be rotated over the slab-wound
resistor. The resulting currents, and magnetic field produced in the
indicator stator, position the rotor and pointer to indicate the
342
Figure J 4.3 Desynn torque
pressure transmitter.
OIL PRESSURE
CONNECTION
SlAB - DESYNN ELEMENT
BOURDON
TUBE
torque pressure on a dial calibrated from O to 600 lbf/in 2 • During
operation, and due to pulsations of torquemeter pressure, a certain
amount of pointer fluctuation is possible, but this is limited to
30 lbf/in1 on either side of a mean torquemeter pressure reading.
Synchro Torque Pr~ure Indicating System
This system is an application of the alternating-current control
synchro principle (see page 233).
As may be seen from Fig 14.4, the mechanical element of a
synchro torque pressure transmitter is very similar to that of the slabDesypn type. The Bourdon tube, sector gear and pinion, however,
are arranged to drive the rotor of a ex synchro. The transmitter is
designed for mounting directly on to an engine and is connecteJ by
flexible tubing to the torquemeter system.
The indicator consists of a· CT synchro connected to the transmitter
synchi:o ex, a two-stage transistor amplifier, a two-phase servomotor,
and two concentrically mounted pointers driven through a gearbox.
The smaller pointer indicates hundreds of pounds and rotates in step
with the synchro rotor, while the larger pointer rotates ten times as
fast.
When the Bourdon tube senses a change in torquemeter system
pressure it causes the ex rotor to rotate and to induce a signal
343
c::::,,,c. ~~~eiue TO PRESSURE
c::::::,+, INDUCED SIGNAL VOtTAGE
~
~~io ERROR SIGNAi.
2-SfAGE
AMPUFIER
,-,_::::::::.~
TOROUEMETER
PRESSURE .
GEARBOX
'f'D
2-PHASE MOTOR
Figure 14.4 Principle of control
syn~hr~type torque pressure
voltage in its stator which is then transmitted to the stator of the CT
synchro in the indicator. This signal results in a change in direction
of the resultant magnetic field with respect to the CT rotor position,
thus inducing an error voltage signal in the rotor. The error signal is
fed to the amplifier, which determines its direction, i.e. whether it
results from an increase or a decrease in pressure, as well as amplifying
it. The amplified signal is then applied to the control phase of the
servomotor, which, via the gearbox, drives the pointers in the appropriate direction and also drives the CT rotor until it reaches a new null
position at which no further error voltage signal is induced.
Power Indicators for
Turbojet Engines
With turbojet engines the number of instruments required for power
monitoring depends upon whether the engine employs a centrifugal
or an axial type of compressor. The thrust of a centrifugal
compressor engine is approximately proportional to the speed, so
that the tachometer, together with the turbine gas-temperature
indicator, may be used to indicate thrust at the specified ·throttle
setting.
The thrust produced by an axial compressor engine does not vary
in direct proportion to the speed, the thrust ratings being calculated
in such a way that they must be corrected for variations in temperature and pressure prevailing at the compressor intake. Since
1r1d1catmg system.
344
compressor intake pressure is related to the outlet or discharge
pressure at the turbine, then thrust is more accurately determined
by measuring, the ratio between these two pressures. This is done by
using an engine pressure ratio indicating system or, in some cases by
a percentage thrust indicator in conjunct"ion with the rev./min. , turbine gas-temperature and fuel-flow indicators.
Engine Pressure Ratio (EPR) Indicating System
In general, an EPR system consists of an engine inlet pressure sensing
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
system in current use, is schematically showr: in Fig 14.5.
The inlet pressure-sensing probe is similar to a pitot pressure
probe, and is mounted so that it faces into the airstream in the
engine intake or, as in some powerplant installations, on the pylqn
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. 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 pressure-sensing transducer, a linear
voltage differential transformer (LVDT) a two-phase servomotor,
amplifier and a potentiometer. The transducer bellows are arranged
in two pairs at right angles and supported in a frame which, in turn ,
is supported in a girnbal 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 EPR indicator in terms of changes in pressure ratio.
The indicator is of the servoed ty pe which is an adaptation of the
indicator described in Chapter IO (page 248). In some EPR systems,
the indicator may be of the vertical scale/moving tape type.
From Fig 14.5 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 sealed. Thus, the system together with its frame, gimbal
and yoke assembly, forms a pressure balancing and torsional
system. When a change in pressure 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 the
axis AA (Fig 14.6). The deflection displaces the core of the LVDT
to induce an a.c. signal which is amplified and applied to the
control winding of the servomotor. The motor, via the gear train,
345
Figure 14.5 EPR indicator
system.
l
INDl~~TOR
i--------'
,~~OMU~
---
~lCIClCICICICCIClc c = : J
LVOT
r----::::~~-~
4---1------4------REF. VOLTAGE
_J
POTENTIOMETER
,
1
I
t->j
M )1:1::::DxLl
r,
REF. PHASE
_ _ _ _ _ _ _ _ _ TRANSMIITER _ _ •
I
I
I
J
MECHANICAL LINKAGE
MANIFOLD
alters the potentiometer output signal to the indicator the pointer
and digital counter of which are servo-driven to indicate the new
pressure ratio. Simultaneously, the motor drives the transducer
gimbal and LVDT coils in the same direction as the initial yoke
movement so that the relative movement now produced between the
LVDT coils and core starts reducing the signal to the servomotor,
until it is finally cancelled and the system stabilized at the new
pressure ratio.
Percentage Thrust Indicator
In some cases, power monitoring is done by means of a direct-reading
type of pressure ratio indicator, which as will be noted from Fig 14. 7,
346
Figure 14. 6 Pressure
transducer.
,----A-.-----1
t
GIMBAL
I
I
INTAKE PRESSURE I
I
FRAME
EXHAUST
PRESSURE
I
I
~
I
I
I
A
I
I
I
I
t
I
_J
indicates power in terms of percentage thrust over the range 50 to
100%. In addition, it incorporates a manually controlled device
permitting the thrust indications to be compensated for variation in
ambient atmospheric conditions. The compensation is accomplished
by rotating a setting knob, which ac1: ·1sts a digital counter (in some
instruments a scale may be fitted) and at the same time rotates the
complete mechanism and positions the pointer to a new datum value
on the main dial.
With this compensation applied, the instrument normally indicates
100% thrust as a minimum take-off value under conditions least
favourable to engine performance. Under more favourable conditions the engine performance may indicate a take-off value greater
than 100% thrust.
The counter is of the three-digit display, and each number set on
it corresponds to an appropriate ambient atmospheric condition
obtained from performance curves plotted for specific aircraft/engine
combinations.
347
Figure 14. 7 Percentage-thrust
indicator. 1 Rocking shaft,
2 calibrating arm, 3 exhaust
unit pressure connection,
4 static pressure connection,
5 overload spring, 6 capsule,
7 overload spring, 8 sector
gear, 9 digital counter
(atmospheric datum) , 10
counter setting knc,b.
9
Turbine Temperature
Control
348
The power developed by a turbine engine is dependent on two main
factors: the air mass flow through it and the temperature drop. The
air mass flow varies with engine speed and also with air density,
which in turn is determined by altitude, atmospheric temperature
and forward speed. Temperature drop is the difference between the
temperatures immediately before and after the turbine and is therefore a measure of the energy extracted by the turbine.
It is thus apparent that the temperature drop will be a maximum
and so indicate the maximum energy extraction if the gas temperature at the entry to the turbine is maintained at the highest level.
There is,.however, a practical limitation to this temperature brought
about by the effects on the material of the turbine blades and
consequently on their life. For this reason, optimum temperatures
are established at which the maximum power may be obtained without impairing the structural integrity of the turbine blading, and
the operating conditions are carefully controlled to ensure that such
limitations are not exceeded.
Control of the conditions can be instituted at ground level, but in
flight a turbine must operate under changing conditions of atmospheric temperature, density and forward speed, and as already
mentioned, these variables determine the air mass flow through the
turbine. For given atmospheric conditions, the air mass flow is
controlled by the engine speed, and to maintain the maximum turbine entry temperature appropriate to these conditions, the flow of
fuel to the engine must be controlled so as to match the air mass
flow .
This process of fuel flow control is generally known as fuel
trimming; depending on particular engine installations it may be
effected by an electro-mechanical system under the direct control
of the pilot, or by a temperature control system which automatically
monitors the fuel flow in response to signals from the thermocouples
of the standard exhaust-gas temperature measuring system.
Fuel Trim Indicating System
An example of the electro-mechanical method is the one adopted for
the trimming of the fuel flow to the Rolls-Royce Dart turboprop
engine and serves as a useful illustration.
The air mass and fuel flows are matched to the designed ratio, in
the first instance, by mechanically interconnecting the fuel throttle
valve with the rev./min. controls. This results in an optimum gas
temperature for any selected engine speed.
To compensate for the changes in air mass flow, an electric
actuator operates a differential compound lever mechanism incorporated as part of the interconnection between throttle valve and
rev./min. controls. Electrical power to the actuator is controlled by
a switch which is accessible to the pilot and is placarded INCREASE
and DECREASE, indicating the trimming condition of fuel flow
required and consequently the change of turbine temperature. The
pilot must, of course, have some means of knowing by how much the
fuel flow should be trimmed, and so a fuel-trim position indicating
system is provided for his use in conjunction with datum position
tables or a datum computer supplied by the engine manufacturer.
Jhe indicating system consists of a position transmitter and
indicator operating on the basic Desynn principle described in
Chapter 9 (page 225. The transmitter is mechanically connected to
the trim actuator by a linkage suited to the engine installation and
electrically connected to the indicator, which is usually mounted
on the control pedestal in the cockpit. The scale of the indicator is
graduated from O (full decrease) to 10 (full increase), corresponding
to the range Oto I 00% fuel trim.
The percentage increase or decrease is related to the prevailing
air temperature and the pressure altitude, and is obtained from the
datum position tables or computer. When the required value has
been selected, the actuator is switched on so that its shaft will
retract or extend, depending on whether an increase or decrease of
the fuel flow is required. Movement of the actuator shaft positions
the throttle-valve lever, via control rods and th~ differential compound lever which permits trimming of the fuel without disturbing
the setting of the rev./min. controls. The actuator shaft also positions
the brushes over the toroidal resistance of the transmitter, the voltage
combinations so produced being supplied to the indicator stator
349
windings. Since the indicator rotor is a permanent magnet, it aligns
itself with the resultant magnetic field induced in the stator, and at
the same time moves the pointer to indicate the change in fuel trim.
Automatic Temperature Control System
A logical development of the fuel trimming process, particularly
since it is required to maintain turbine temperatures within
specified limits, is to utilize the signals generated by the thermocouples of the standard temperature indicating system and so let
these do the work of automatically changing the fuel flow.
Experiments along these lines resulted in a specially designt~r!
amplifier unit which, on being connected to the thermocouples,
amplifies the signals produced above a preselected datum temperature
and supplies them to a, solenoid-operated servo valve which then
progressively reduces the fuel flow to restore the datum temperature condition.
A block functional diagram of such an amplifier is shown in
Fig 14.8 and may serve as a basis for the explanation of temperature
control systems in current use.
The thermo-e.m.f. is fed as an input signal to a terminal block .on
the amplifier, the terminals forming the cold junction of the cor1Lrol
system. Compensation for temperature changes at this junction is
effected by a bridge circuit, one arm of which changes its resistance
Figure 14.8 Typical auto•
matic temperature-control
amplifier.
115V
AC. SUPPlY
SUPf'LY
TRANSFORMER
SIGNAL LINE FOR
TEMPERATURE
BELOW DATUM
RECTIFIER 2
STABlllZlNG
NETWORK
,11 STAGE
IF---=-J
2nd STAGE
~ -
MAGNETIC AMPUFIER
THERMOCOUPl.f
SIGNAL INPUT
REFERENCE
VOLTAGE UNrT
ERROR SIGNAL
WHEN TEMPERATURE
ABOVE DATUM
350
SELECTED
COLO
DATUM
TEMPERATURE
JUNCTION
COMPENSATION
FEEDBACK
STAGE
TO FUEL CONTllOL
SOLENOID OR
VALVE
with changes of temperature. In addition to the thermocouples, the
bridge circuit is also connected to a reference-voltage unit, the purpose of which is to inject a voltage in opposition to that of the thermocouples and corresponding with the desired operating temperature
range of the engine. This voltage is stabilized against changes in
supply voltage and frequency ( 115 V, 400 Hz) and ambient temperature, and is selected by a temperature selector.
The amplifying section is in three main stages: amplifier stages I
and 2, and an output stage connected to the solenoid valve.
When the engine is running at the selected temperature, the e.m.f.
from the thermocouple is opposed by an equal voltage from the
reference-voltage unit, so that there is no input signal to stage I of
the amplifier.
If the exhaust gas temperature is below the selected value, the
reference voltage is greater than that of the thermocouples and this
Figure 14.9 Operating
characteristics of a typical
temperature control system.
(a) Amplifier; (b) fuel control
system; (c) combined control
characteristic.
t
I
SELECTED DATUM
TEMPERATURE
~
I
i
a:
Ill
~
(a)
~--------l
~L..-- -THERMOCDUPLE
---TEMPERATURE
-
(b)
THROTilf SETilNG CONSTANT
SOLENOID CURRENT -
~--
tc1
THERMOCOUPLE TEMPERATURE -
351
Figure 14.10 (a) and (b)
Temperature control units.
provides a reverse-polarity input to stage I. This reverse signal is,
however, blocked by a network connected between stage 2 and the
output stage so that no output current is obtained .
When the exhaust gas temperature rises above the selected value,
and this is the more critical situation, the predominating voltage is
that from the thermocouples. As this voltage is of the correct polarity
the resulting signal is fed to stages 1 and 2 and, after amplification, to
the solenoid valve via the output stage. The valve operates so as to
restrict the fuel supply to the engine, thus reducing the exhaust gas
temperature and thermocouple e.m.f. until it balances the reference
voltage once again. The solenoid valve is then released and the
normal fixed fuel flow is restored.
The operating characteristics of a typical temperature control
system are shown graphically in Fig 14.9. It will be noted from
diagram (a) that, when the temperature reaches the limiting value, the
amplifier output continues to increas~. reaching a constant maximum
value after a small temperature increase of approximately 8°C. The
solenoid-current/fuel-flow characteristic depends on the type of fuel
system employed, but it is usually of the form shown at (b ), which
clearly indicates the decrease in the fuel flow with the increase in
solenoid current. A combination of (a) and (b) gives us the final
control characteristic shown at (c).
'Hunting' of the solenoid valve ·which would give rise to a fluctuation fuel flow and engine thrust, is prevented by feeding back
lrlERMOCOUPl.E CONNECTION
352
Figure 14. JO (b)
some of the output current into the amplifier stages, thus increasing
the time lag.
Examples of the physical construction of two typical control
units are shown in Fig 14.10.
Combined Temperature This system is a further development in the field of automatic
And Rev./Min. Control control of engine power, and is one in which fuel flow is regulated
System
by combinging the signals from thermocouples with those supplied
by_a tachometer generator. A block diagram of the system is shown
in Fig 14.11.
RP,M.
GENERATOR
SIGNAL INPUT
TO FUEL
CONTROL
SOLENOIO
OR VALVE
DISCRIMINATOR
AMPUFIER
SE1.EC'n0
DATUM RP.M.
Figure 14. 11 Typical temperature rev./min. control
system.
The temperature control channel operates on the same principle
as the system described in the preceding paragraphs, and so we need
only consider the r.p.m. control channel..
353
The three-phase output from the tachometer generator is supplied
to a rev./min. discriminator section which is made up of two
elements, the purpose of which is to provide two independent
voltages from the generator input. One element is of the resistance
type so that its voltage is largely independent of the generator
frequency , and the other element is of the reactance type, i.e. it
contains capacitors and so its voltage is proportional to frequency
and therefore to engine speed . Both these voltages are then rectified
and applied in opposition to the rev./min. control channel amplifier.,
which is of the two-stage type. The voltages may be equalized at
any desired engine speed by adjustment of the resistance-type
element with the aid of a selector. The output of the rev./min.
control channel amplifier is in turn fed to the output stage, which, as
may be seen from the diagram, is common to both rev./min. and
temperature control channels. At the selected datum speed, the
current from the output stage is the standing current in the fuel-valve
solenoid.
When the engine speed is below the selected datum, the rev./min.
control channel signal is suppressed by a discriminator network
similar to that used for the temperature channel. If, however, the ·
engine speed rises above datum, the voltage output from the reactive
element of the rev./min. discriminator section will predominate and
pass a signal on to the solenoid valve, which reduces the fuel flow to
.restore the datum speed condition.
The circuitry between both control ch,wnels is so arranged that
the solenoid valve remains under the control of the rev./min. channel
until such time as the exhaust gas temperature exceeds the datum limit
when the temperature control channel with its greater amplification
overrides the speed signal and assumes control of the fuel flow to
reduce turbine temperature.
The application of this system depends upon the requirements of
a particular engine ; for example, speed and temperature limiting may
be required at a fixed datum ('top limiting'), or it may be required to
Figure 14.1 2 Top temperature
limiting.
·
A1.mUD£ -
354
be variable ('range control'). In either case, appropriate datum
seiector units are used with the amplifiers.
An example of 'top limiting' is illustrated graphically in Fig 14.12
and relates to an aircraft climbing from ground level. Under these
conditions, engine speed is the limiting factor and maximum take-off
power is required. The solenoid valve modifies the fuel flow to
maintain constant rev./min. until the exhaust gas temperature rises to
its limiting value. When this limit is reached , the temperature channel
takes full control and overrid es the governing action of the rev./min.
channel, thus reducing the fuel flow to avoid exceeding the datum
temperature. The physical construction of a representative uni t with
its associated internal circuit. boards is shown in Fig 14. 13.
Figure 14.13 Limiter unit.
(By courtesy of Smiths
Industries.)
355
Engine Vibration
Monitoring And
Indicating Systems
Figure 14.14 Turbine vibration
monitoring system.
Engine vibration is, of course, something which is unwanted, but
unfortunately it cannot be entirely eliminated even with turbine
engines, which have no reciprocating parts. It can only be kept down
to the lowest possible level.
During operation, however, there is always the possibility of
vibration occurring in excess of acceptable levels, as a result of certain
mechanical troubles. 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 troubles will give rise to
an unbalanced condition of the main rotating assemblies. In order,
therefore, to monitor vibration and to indicate when the maximum
amplitude on any engine ex~eeds a preset level, systems have been
developed which come within the engine coritrol group of instrumentation.
A system consists essentially of a vibration pick-up unit mounted
on the engine at right angles to its axis, an amplifier monitoring unit
and a moving-coil microammeter calibrated to show vibration
amplitude in thousands of an inch (mils). A block diagram of the
system is shown in Fig 14.14.
The pick-up unit is a linear-velocity detector that converts the
mechanical energy of vibration into an electrical signal of proportional magnitudes. It does this by means of a spring-supported
permanent magnet suspended in a coil attached to the interior of the
case.
S USPENDED
MAGNET
DIFFUSER
PICK · UP
TURBINE PICK -UP
356
115V 400H Z
SINGLE· PHASE
SUPPLY
As the engine vibrates, the pick-up unit and coil move with it; the
magnet, however, tends to remain fixed in space because of inertia.
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 anq 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 direct current from the
amplifier rectifying section and it comes on when the maximum
amplitude of vi~ration exceeds the preset value. The test switch
permits functional checking of the system's electrical circuit.
In some engine installations, two pick-up units may be fitted to
an engine, one monitoring vibration levels around the turbine section
and the other around the diffuser section. In this case, a two-position
switch is included in the monitoring system so that the vibration level
at each pick-up may be selected as required and read on a common
indicator.
Questions
Briefly explain the principle of engine supercharging.
Describe the construction and explain the operation of the instrument
used for measuring manifold pressure.
14.3 What is the function of a torque pressure-indicating system? Describe
the construction and operation of a system with which you are
familiar.
14.4 List the instruments required for monitoring the power of turbojet
engines.
.
14.S By means of a schematic diagram explain the operation of an EPR
indicating system.
14.6 Describe a method of utilizing the voltage generated by exhaust-gas
thermocouples for controlling the fuel flow and gas temperature.
14.7 Describe the operation of an engine vibration indicator. Show with
the aid of a sketch the construction of the sensing device.
14.1
'14.2
357
15 Integrated instrument
and flight director
systems
The need for integrating the functions and indications of certain
flight and navigation instruments resulted in the main from the
increasing number of specialized radio aids linking aircraft with
ground stations. ·These were developed to meet the demands of safe
en-route nagivation and to cope with increasing traffic congestion
in the air space around the world's major airports.
The required information is processed by a multiplicity of 'black
boxes' which can be stowed in electrical compartments and radio
racks, but in order that the necessary precision flying may be
executed, information must still be presented to the pilot. This ,
requires more instruments and more instruments could mean more
panel space . The method of easing the problem was to combine
related instruments in the same case and to compound their indications so that a large proportion of intermediate mental processing on
the part of the pilot could be bypassed and the ind_ications more
easily assimilated.
In some respects, integration of instruments is not new ; for
example, a combined pressure and temperature indicator was in use
long before the present state of the art. Another early example and
one of those around which most of today's fully integrated flight
and navigation instrument systems have been built up is the Radio
Magnetic Indicator (see page 196).
During that phase of a flight involving the approach to an airport
runway, it is essential for a pilot to know, among other things, that
he is maintaining the correct approach attitude. Such information
can be obtained ·from the gyro horizon and from a special ILS indicator which responds to vertical and horizontal beam signals radiated
by the transmitters of an Instrument Landing System located at the
airport. It was therefore a logical step in the development of integration techniques in what are termed Flight Director Systems, to
include the information from both the gyro horizon and ILS
indicator. Furthermore, the RMI was developed to include the
presentation of ILS information. The methods adopted for the
integration of such information, and the manner in which it is presented vary between systems, but in basic form, they follow the pattern
illustrated in Fig 15. 1. A system normally comprises t wo
358
GLIDE
SLOPE
POINTER
BANK
POINTER
HORIZON
DISC
& BAR
LOCALIZER
BAR
GLIDE
SLOPE
POINTER
AIRCRAFT
REFERENCE
SYMBOL
HEADING
MARKER
GLIDE
~~m
PITCH
LOCALIZER
POINTER
BAA
(al
Figure 15. 1 Basic presentat·
ions of flight director system
indicators.
HEADING
SELECT
COURSE
SELECT
KNOB
KNOB
(b)
indicators: they are variously called (a) a flight director, an attitude
flight director or an approach horizon, and (b) a course deviation
indicator (CDI) or a horizontal situation indicator (HSI).
The flight director indicator has the appearance of a conventional
gyro horizon, but unlike this instrument the pitch and roll indicating
elements are electrically controlled. from a remotely located vertical
gyro unit. Furthermore, it employs a different method of referencing
the elements. These features are common to all integrated instrument
flight director systems. The horizon bar and roll scale are marked on
the background disc which is monitored by roll command signals from
the vertical gyroscope and rotates about the fore-and-aft axis. The
pitch, or command bar, corresponds to the miniature aircraft symbol
in a conventional gyro horizon, and is monitored by pitch command
signals and moves in a vertical plane above and below the centre-line
of the. instrument .
The approach attitude of an aircraft with respect to its verticalbeam (glide-path) signals and horizontal-beam (localizer) signals is
indicated by independen t pointers monitored by the relevant ILS
receiver channels. The glide slope pointer is referenced against a
vertical scale and the pitch bar, and shows the displacement of the
aircraft above or below the glide path. Displacement of the aircraft
to the left or right of the localizer beam is indicated by deflections
of the localizer pointer.
Electrical interconnection of the flight director indicator components primarily concerned with pitch and roll attitude information is
359
Figure 15. 2 Electrical interconnection of flight director
indicator elements.
shown in Fig l 5.2. Whenever a change of aircraft attitude occurs,
signals flow from pitch and roll synchros disposed about the relevant
axes of the vertical gyroscope to the corresponding synchros within
the indicator. Error signals are therefore induced in the rotors and
after amplification are fed to the servomotors, which rotate to
position the pitch bar and horizon disc to indicate the changing
attitude of the aircraft. At the same time, the servomotors drive the
synchro rotors to the 'null' position.
Figure 15.2 also shows the interconnection of the glide slope and
localizer pointer with the ILS. During an ILS approach the receiver
on board the aircraft detects the signals beamed from ground transmitters in vertical and horizontal planes. If the aircraft is above the
glide path, signals are fed·to the meter controlling the glide slope .
pointer causing it to be deflected downwards against-the scale, thus
directing the pilot to bring the aircraft down on to the glide path.
An upward deflection of the pointer indicates flight below the glide
path and therefore directs that the aircraft be brought up to the glide
path. The pointer is also referenced against the pitch bar to indicate any
pitch correction required to capture and hold the glide path. When
this has been accomplished, the glide slope pointer and pitch bar are
matched at the horizontal centre position.
BANK SIGNAL' ' ,~
,----A-IMPLIFIER
I'
~
,,
I~
GLIDE PATH
,:,,
'-s
TRANSMITTH!
~
,,
*~
l/~,,,/
II
II
II
Q
METER
MOVEMENT
11
:: BELOW
AIRCRAFT
ABOVE
GLIDE PATH
"FLY OOWN"
MOTOR
------- -----
==~
i
PITCH BAR
PITCH SIGNAL
AMPLIFIER
360
,,
AIRCR.AFT
ff;""
HORI ZON DISC
AIRCRAFT
RI GHT OF
LOCALIZER
"Fl Y LEFT'
I I AIRCRAFT
: : LEFT OF
•1 LOCALIZER
II "Fl V RIGHT"
--,--
Figure 15.] Interconnection
of course deviation Indicator
elements.
If, during the approach, the aircraft is to the left of the localizer
beam and runway centre-line , the localizer pointer is deflected to the
right directing that the aircraft be banked to the right. Flight to the
right of the localizer beam causes pointer deflection to the left,
directing that the aircraft be banked to the left. When either of these
directions has been satisfied, the pointer is positioned vertically
through the centre position of the horizon disc.
A course deviation indicator is similar in display presentation to
an RMI, except that it has the additional feature of displaying ILS
information. The display of magnetic heading, radio bearing and
localizer information is referenced against an aircraft reference
symbol fixed at the centre of the indicator to provide a plan view of
the aircraft's position; hence the alternative name of 'horizontal
situation' indicator. The interconnection of the various elements
comprising c!n indicator is shown schematically in Fig 15.3 ..
The compass card is monitored by signals from the directional
gyro unit of a remote-indicating compass system and indicates
magnetic heading against a lubber line. The localizer and glide slope
indicating elements are actuated in a similar manner to those in-a
flight director indicator, but as will be noted, the localizer or lateral
deviation bar as it is generally called, can also rotate with the
compass card as changes in aircraft heading take place. Indication
MAAKER
D.
..-,!.!..-.
II
II
II
11
II
II
$IGIWS
FROM
COMPASS
--+
GYRO
AllMU!l-i
RING
0
+
I@,
-~
HEAOINO
+
AIRCRAFT
REFERENCE
SYMBOL
I
DEi/iATION SCALE
-
srr
HEADING KNOB
'1
V.O.R.
TRANSMITTER.
+--~. /
~~-MO
I
I
- 1;-ccj:~~
I LOCALIZER BAR
;
COURSE
ARROW
.....---..
t
CTj~~~
-+-- -
METER
II
1,
6
COURSE KNOB
361
of flight either to or from a VOR station is indicated by an arrowtype element which is positioned by a meter. The course select and
heading knobs permit the selection of a desired localizer or VOR
radial, and desired magnetic heading respectively.
Description of a
Representative System
The indicators of a Flight Director system which may be considered
generally representative of those in current use are shown in Figs
15.4and 15.5.
Flight Director Indicator
Figure 15.4 Flight director
indicator.
362
In the indicator illustrated, ;iircraft attitude and ILS information are
presented in the form of a three-dimensional display. Attitude is
displayed by the relationship or" a stationary delta-shaped symbol
representing the aircraft, with respect to bank and pitch commands
displayed by two pointers, or command bars flanking the aircraft
Figure 15.5 Course deviation
indicator.
symbol and also by an horizon bar. The command bars form a
shallow inverted 'V', and are driven by separate servomotors within
the indicator such that they move up and down to command a change
in pitch, and rotate clockwise and anticlockwise to command a
change of bank attitude. The outputs of the two servos are
combined mechanically so as to provide an integrated pitch and
bank command . Sensing with respect to the aircraft symbol is such
that the pilot is always directed to 'fly into the 'Y". When a command
bas been satisfied, the command bars are aligned with the edges of the
aircraft symbol. The horizon bar is carried on a flexible tape which is
also driven by separate pitch and roll servomotors within the indicator.
Freedom of tape movement in pitch is ±90°, and 360° in roll. The
upper and lower sections of the tape are coloured to represent the sky
and ground respectively, and they also have index marks on them to
indicate pitch angles. In some types of flight director, the lower
section of the moving tape is also marked with lines converging on the
centre of the indicator display thereby enhancing its 'forward view'
363
effect. Roll angle is displayed by a pointer which rotates with the
flexible tape, and is referenced against a fixed scale. The servomotors
are supplied with signals from a vertical gyroscope unit located at a
remote point.
Deviations from the ILS glide path beam are shown by vertical
displacements of a pointer over a scale at the left-hand side of the
indicator display. Each of the inner dots on the scale represent a
Yi displacement from the beam centre line, while the outer dots
each represent a Yz° displacement. The pointer is driven by a d.c.
meter movement and, when not in use, is deflected out of view at
the top of the scale. A pointer at the lower part of the display
indicates deviations from the localizer beam, and is shaped to
symbolize a view of a runway during an approach. The reference
dots on the localizer or runway scale, indicate approximately 1Yi
displacement from the beam centre line. The pointer is also driven
by a d.c. meter movement, and , when not in use, is obscured by a
black warning flag ·as shown in Fig 15 .4. In some types of indicator,
the localizer pointer, or runway symbol, is also displaced in response
to signals from a radio altimeter so that during the last 200 ft of
descent, the pointer moves up to the fixed aircraft symbol thereby
presenting a 'live' display of the approach. This radio altitude display
concept is also adopted in the indicator shown but, in this case, it is
effected by a pointer moving over a fixed altitude scale.
Indications of slip or skid are provided by an inclinometer similar
to that adopted in conventional turn and bank indicators. In addition,
some flight directors have a rate of turn pointer incorporated in the
display, the pointer being actuated by signals from a rate gyroscope
sensor unit.
Another command function which may be displayed in some
flight director indicators, is that related to the speed of an aircrafi
when executing a go-around manoeuvre following a missed approach.
The display comprises a vertical scale and a pointer which is actuated
in response to signals corresponding to the difference between
indicated airspeed and a pre-determined go-around speed obtained
from a speed computing system external to the flight director system.
The scale has several graduations ranging from the computed speed
at the centre, to 'fast' and 'slow' at the top and bottom of the scale
respectively. In order to achieve the correct go-around speed, engine
power is adjusted so as to maintain the pointer at the centre of the
scale. The pilot selects the go-around mode by pressing a button
switch on the control wheel, the selection being indicated by the
illumination of an annunciator light marked 'GA', and by displacement of the flight director command bars to command a wings-level
climb attitude. In association with the go-around mode, a second
annunciator light marked 'MDA' (minimum decision altitude) is
provided . The light illuminates when the aircraft has descended to
0
0
364
the preset radio altitude at which the decision whether to land or
go-around must be made.
Course Deviation Indicator
A course deviation indicator presents a pictorial display of a navigation situation, and as will be noted from Fig 15.5, the situation is
shown as a plan view of the aircraft's position and heading with
respect to a selected reading and coµ rse. The aircraft reference
symbol is fixed at the centre of the display and it indicates the
position and heading of the aircraft in relation to the compass card,
and the localizer bar, (this bar is sometimes referred to as the lateral
deviation bar). The compass card is synchronous-linked with the aircraft's magnetic compass system;-and when changes in aircraft
heading take place, a position error signal is produced in a CT synchrq
within the indicator. After amplification by a servo amplifier in a
separate instrument amplifier unit, the signal is supplied to a servo
motor which, by means of a gear train system, drives the compass
card to indicate the new heading with reference to a lubber line at
the top centre of the indicator. Card movement is damped by means
of a rate feedback signal produced by a rate generator which is driven
by the motor. Feedback signals are mixed with position error signals,
and the composite s~gnal is amplified and supplied to the control
phase of the servomotor. At the same time, the servomotor drives
the CT synchro rotor to the null position corresponding to the new
heading. The localizer bar is a movable centre section of the course
arrow, and represents the centre line of the selected ILS localizer
course or VOR radial. The bar is deflected to the left or right by a
d.c. meter movement to indicate the appropriate commands necessary for beam interception and capture, and it also rotates with the
compass card as changes in aircraft heading take place. When
operating in the localizer mode, initial movement of the localizer
bar begins when the aircraft is approximately 4° from the localizer
beam centre, and the dots on the deviation scale represent
approximately 1¥.i and 2Y2° from beam centre. · In the VOR mode,
initial movement of the bar begins when the aircraft is approximately
16° from radial centre, and the deviation scale dots then represent
approximately 5° and I 0° from radial centre.
Selection of a desired ILS localizer course or VOR radial is carried
out by rotating the course selector knob until the course arrow coincides with the desired value on the compass card. The localizer bar
and deviation scale also rotate with the course arrow through the
gear train system driven by the selector knob. At the same time a
digital type of course counter is driven to the corresponding course
indication; in Fig 15.5, this is displayed as 075° . Once set , the course
arrow rotates with the compass card as aircraft heading changes. The
0
365
gear train system also positions the rotors of a course resolver synchro
associated with the V0R/LOC navigation receiver, and of course a
datum CT synchro. When the course resolver synchro rotor position
is changed, it shifts the phase of the reference 30 Hz signal in a phase
shift circuit of a VOR instrumentation unit. The signal is then
compared with the variable 30Hz signal in a phase comparator, the
output of which is supplied to the meter movement controlling the
localizer bar. When the output is such that it centres the bar, then
the aircraft is on the course selected. When the aircraft deviates from
the selected course, the phase-shifted refe1 vnce signal is maintained at
the angle determined by the resolver synchro rotor, but the variable
signal phase received by the VOR navigation receiver is changed. The
phase comparator will then produce an output which deflects the
localizer bar to the left or right of the selected VOR course. The
to-from arrow is positioned by a meter movement which is supplied
with the corresponding signals from the radio navigation receiver and
via a phase comparator in the instrumentation unit of the flight
director system. In Fig 15.5 a 'fly to' command is displayed. In the
LOC mode of operation, the localizer bar is similarly controlled by
changes in resolver synchro rotor position , except that the output to
the meter movement results fr0m amplitude comparison of the
signals either side of localizer beam centre. The to-from arrows
remain out of view since no to-from signals are transmitted in the
localizer mode.
Changes in the position of the course datum CT synchro rotor
produce a position error signal in the stator windings. The signal
is proportional to the difference between the selected course and the
actual heading of the aircraft and is transmitted to the roll control
channel of a steering computer or a flight director computer as a
turn command to capture the selected VOR/LOC course. The output from the appropriate computer is supplied to a roll command
servo amplifier cont ained in .a n instrument amplifier unit , and after
amplification it is fed to the roll servomotor coupled to the command
bars of the flight director indicator. Thus, the bars rotate to indicate
the direction of roll required to capture the VOR/LOC course. The
servomotor also drives a rate generator which produces a rate feedback signal for the purpose of damping display movements.
The selection of any desired magnetic heading is accomplished by
positioning a triangular-shaped heading marker over the compass
card, by means of the heading selector knob and its associated shaft
and gear train system. At the same time, the rotor of a heading error
CT synchro is rotated inside its stator, from its null position, and
this produces a position error signal proportional to the difference
between the selected heading and aircraft heading sensed by the
compass system. In Fig 15 .5 the headings displayed are respectively
110° and 085°. The signal is processed in the same manner as that
366
produced by the course ~atum CT synchro, and therefore results in
the flight director indicator command bars indicating the direction of
roll required to fly on the desired heading.
Warning Flags
In flight director systems it is necessary to make provision for the
warning of faulty display functions . In practice, warnings are
effected by monitoring the command signals produced such that
· when they are lost or are too weak to provide reliable information,
small red flags appear at appropriate parts of the flight director
indicator and course deviation indicator displays. The flags are
actuated by d.c. meter mechanisms which are connected to the
relevant signal sources. In the case of flight director indicators there
are, primarily, three warning flags labelled 'GS', 'GYRO' and
'COMPUTER' and respectively they indicate malfunctions of the
glide-slope receiver or signal, the vertical gyroscope unit and attitude
display systems, and the director or steering computer and command
display systems. The GS flag, when indicating a malfunction,
obscures the glide slope pointer and scale to prevent its use. If the
system is not being operated in the glide slope mode, the GS flag and
pointer are biased off-scale. Indication of localizer signal malfunction
and/or localizer mode not selected, is also provided and generally
takes the form of a black shutter which obscures the localizer pointer
and scale (see Fig 15 .4). Other warning flags may be provided
depending on any additional functions displayed; for example, a
flag is provided to give warning of malfunctions in the circuit of a
speed control display associated with a go-around manoeuvre.
In the case of a course deviation indicator, there are also three
primary warning flags and these are labelled 'GS', 'COMPASS' and
'VOR/LOC'. The GS flag operates in the same manner as that
provided in the flight director indicator, while the compass flag
indicates malfunctions of the magnetic heading signal circuit of the
compass system. The VOR/LOC flag serves the dual function of
warning of VOR radial signal and localizer signal malfunction.
Pitch Command Facility
In some types of flight director systems, a pitch command facility
is provided which permits the pilot to preselect a fixed climb or
descent command under certain modes of operation. Selection is
carried out by means of a selector knob which, in some cases, is
located in the bottom left-hand corner of a flight director indicator
and in others is located on a separate flight director mode selector
panel. The selector knob is mechanically coupled to the rotor of
a CT synchro, and after the knob is rotated a signal is induced in the
synchro. After amplification, this signal is transmitted to the pitch
367
servomotor/generator which drives the command bars to the selected
position. The aircraft attitude is then changed by 'flying the aircraft
symbol into the command bars'. In addition to pitch command, a
pitch trim adjustment is provided as a means of altering the position
of the flight director horizon bar with respect to the aircraft symbol.
The adjustment ls purely mechanical in operation and is used for
aligning the attitude display during installation of a flight director
indicator.
Modes of Operation
A number of flight director systems incorporate facilities for selecting
various modes of operation, and it is possible for mode selection to
be used on a common basis in cases where a flight director system is
employed in combination with an automatic flight control system.
The number of modes vary dependent on specific aircraft operating
requirements, and for a similar reason, the method by which modes
are selected can also vary; for example, selection may be effected by
the control knob of a rotary selector switch on a flight director
indicator or on a separate mode selector panel.
In some versions of a control panel, modes are selected by push
buttons which are push-on/push-off solenoid-held switches. The
push buttons illuminate when their corresponding modes are selected
and at the same time a mechanically-actuated flag with the word 'ON'
appears over a portion of each button engaged. The operating modes
which are fundamental to some typical flight director systems are
briefly described in Table 15. I. When each mode is selected, signal
circuits are completed through the appropriate computer and
amplifier sections, the outputs of which are supplied as command
signals at the flight director indicator. If a flight director system is
used in combination with an automatic flight control system, the
command signals are also utilized by this system for applying control
in the sense necessary to satisfy the relevant commands.
Table 15.1 Fundamental Operating Modes
368
OFF
Command bars deflected out of sight, and flight
director indicator used as an attitude reference
only.
HDG
Command bars provide lateral guidance to achieve
and maintain a compass heading, as selected on
the course indicator. Vertical guidance is from a
preselected pitch attitude.
VOR(NAV)/LOC
Command bars provide lateral guidance to
capture and track a VOR radial or localizer beam.
Vertical guidance is the same as in HDG mode.
GS
Command bars provide lateral and vertical
guidance to capture and track the localizer and
glide slope beams respectively. The GS and
LOC pointers monitor aircraft deviations of the
beam.
GS AUTO
As for GS except that interception and capture
of glide slope takes place automatically after the
localizer beam has been captured.
ALT
Command bars provide vertical guidance to hold
the aircraft at the desired altitude.
APPR I
Selected for capture and tracking of GS and LOC
beams on ILS approaches to Category I standards.
Command bars provide lateral and vertical
guidance.
APPR II
As for APPR I but produces tighter tracking of
beams to meet higher precision requirements of a
Category II ILS approach.
GA
Selected for a go-around manoeuvre after a missed
approach, and after selecting either one of the
approach modes. The command bars command a
wings-level, pitch-up attitude. HDG and IAS modes
may be selected after go-around power settings and
airspeed are established.
IAS
Selected to maintain a particular indicated airspeed
during climbout after take-off, and during letdown
over a VOR station. The command bars provide
vertical guidance.
V/S
Selected to maintain a particular vertical speed, i.e.
rate of climb or descent. The command bars provide vertical guidance.
MACH
As for IAS mode but selected at higher altitudes.
Computer and Amplifier Units
As already noted in the foregoing brief descriptions of flight director
indicator and course indicator operation, the appropriate command
signals are processed by computer and instrument amplifier units. The
primary function of a computer unit is to provide all the computation
necessary for determining any position or attitude errors, and to
develop the signals necessary to command position or attitude changes.
389
COURSE INDICATOR
FLIGHT DIRECTOR INDICATOR
/
~y
/
~
-- ___ \
,
~ )i\
...
mo~m
,,
~~ ~~o
~ •• _. , m•
L.,_
••
•
w
.
',lw
•
:·
-
!"""'
~~
..cl'
-
..
.."''!>'!, N J ~
'
,,.,..,.,'I:'. ,
~ I\
~L
..
'1,x.
<>
-
I
.
'
\ , -1!l..L ,
~--/~··
'
I
.
--
0
a:
:,:
u
z
>
,:
::e
<.>
~j
<.>
::Ea:
::e::e
!: w
5Za:
>a:
:l~
Ow
"'0
5~
,_ a:
a.w
-"'
00
PIT~H !:ERvn Mn, OR a:
a:
w
PITCH SYNCHRO ERROR
"'O
ROLL SYNCHRO ERROR
"'a:
<:i:
ROLL SERVO MOTOR
Q. <.>
00
:i:a: <.>o
xg u... uo
Ucc
Q."'
a:"'"'
DC TO
ACCO
-- -----
-.,
I
-- ·
CONVERTOR MODULE
1ii:
w
0w
0
t;
...
w
Iii
...<.>ww
,_<.>
~
X
ii:
~
g
::e
,_~
.,o
,_w
Ul/l
w
~~
Ow
PITCH
SERVO
AMPL
ROLL
SERVO
AMPL
0
u,::E
<o
<.>
~
e
COMPASS
SERVO
AMPL
~ ·9 ~
COMPUTER
FLAG
II GYRO
FLAG
M ONITOR
I+--
,_
::,
!~
......
·<
<
0
w
Cl)
a:
w
Cl)
... a:
Q.
WW
<!) "
INSTRUMENT
AMPLIFIER
NAVIGATION
RECEIVER
,_ww
"'...
...
0
a:
u
0
...
...
<
w
>
w
0
FLIGHT DIRECTOR COMPUTER
OME
RECEIVER
-·
:z:
..
<.>
:,:
<.>
,_
Ii:
1
STEERING
FLAG
OUTPUT
t
Figure I 5. 6 Signal flow
WW
Q.
"'Cl)
-'I~ I : : ) 0 0
~o oo
:i :;i
g> ,_ u
COMPASsr~
FLAG
MONITOR
,_
z
<!)
00
a: ......
gg ::.ow
a:
<.>
z
ii:
I; __-
370
g oi
:5... 0w
::,
0
<!)
w
diagram.
a:
~ ~
"5
Cl)
!
;
,
::i;r
...
...w~
,_~i
...
< .
...w
::,
g
... Q "s
"' ...
j
w
w
0
,_w
O>
<.>Cl>
ROLL
CMD
SERVO
MPL
----
I DC TO
I AC CONV
::.z
~
0::<I)
PITCH
CMD
SERVO
MPL
r- -
!!;
~
a:
0
,_
a
~
VERTICAL GYRO
PITCH
~.__
-+
ROL L
GYRO FLAG
OUTPUT
-+-
MAGNETIC
COMPASS
COMPASS
OUTPUT XVZ
FORM
COMPASS
FLAG
OUTPUT
X
<!)
'
...~
w
w :i:
• :z:
!:2
~ :5 :r
"... "...
o ..,
> ...
wO
...
.,;
i<.>
0
0
Cl)
0
Cl)
§
0
::e w
:!
ii!
">
When a flight director system is integrated with an automatic flight
control system, the computed signals are also utilized for the
application of control. A computer may in s~me cases be a single unit
containing solid-state signal circuits for both lateral and vertical
guidance information, while for some flight director systems separate
computer units are utilized. In addition to the signal circuits, a logic
network is incorporated, its purpose being to provide correct analogue
scaling of signals, to adjust computer gains, and logic to suit specific
types of aircraft. All signal and power supply circuits are on printed
circuit boards which are arranged as separate functional plug-in
modules.
The primary function of an instrument amplifier is to supply servoactuating power for the display mechanisms of the flight director and
course deviation indicators. The unit also contains separate plug-in
module circuit boards which, as shown in the overall signal flow
diagram of a representative system (Fig 15.6) correspond to five servo
channels, two signal converter channels, and three flag alarm circuits.
If additional warnings are required, then the number of alarm circuits
are increased accordingly . The converter channels accept d .c. input
signals and convert them to 400 Hz signals for use by the pitch and
roll command servo channels. The converters receive pitch and roll
steering signals from the computer, position error signal information
from the pitch and roll command control transformer synchros in the
flight director, and rate feedback signals from the pitch and roll
command servomotor/generators. These signals are mixed and after
filtering, the composite signals provide the input for the appropriate
command servo amplifier.
Questions
15.1
Draw .a diagram to illustrate the display of a typical flight director
indicator.
15.2 How are the localizer and glide slope elements of an indicator actuated?
15.3 What is the significance of the dots against which the localizer and
glide slope elements are registered?
15.4 What is the purpose of the annunciator lights provided in some types
of flight director indicators?
15.5 Briefly describe how the compass card of a course deviation indicator
is driven to indicate a change in heading.
When
a change in heading talces place, does the localizer bar rotate
15.6
with the compass card, or does it remain stationary?
15.7 What are the functions of the heading select and course select knobs?
15.8 In what operating mode does a 'to- from' arrow operate?
15.9 State the purpose and operation of a pitch trim facility .
15.10 What is the function of the warning flags provided in flight director
indicators and course deviation indicators and when do they come
into view?
371
15 .11 How is direction of roll required to fly onto a selected magnetic
heading indicated?
15.12 What are the primary functions of the computer and instrument
amplifier units?
372
16 Flight data recording
The recording of certain of the parameters measured by instruments,
and of any unusuaJ incidents connected with engines, airframes and
.performance generally, has always been a feature of in-flight management procedures. The earliest method of recording, and one which is
still mandatory, is that requiring entries to be made in the aircraft's
'tech log'. However, as aircraft and their systems became more
sophisticated in their operating procedures, the number of parameters
to be monitored increased, and this imposed limitations on the 'tech
log' entry method. Furthermore, the question of how to retrieve data
which would be of value in investigating the cause of a crash had also
to be taken into account. In this connection, therefore, it became
mandatory (in 1958 in the USA, and 1965 in the UK) to equip certain
categories of aircraft with an automatic flight recording system
comprising a flight data recorder and a cockpit voice recorder. The
number of parameters to be recc.rded progressively increases from a
common basic set established according to groups of aircraft weights,
and therefore according to the degree of complexity of aircraft and
systems. Since the majority of the parameters are measured by 'tapping off from the sensing elements or transducers of existing instrument systems then, consistent with the number of these employed in
an aircraft, there is virtuaJ!y no limit to the number of parameters that
can be recorded. Thus, the capacity of recorders can be increased to
cater not only for the data required by the regulations (mandatory
recording) but also for the acquisition of data which can be of value
to an aircraft operator in planning for economic operation, and for
the monitoring of aircraft, engines and systems reliability (nonmandatory recording). Recording systems of this nature utilize two
complementary recorders and are known variously as Aircraft Integrated Data Systems (AIDS) or Data Acquisition Systems (DAS).
Before outlining the operating fundamentaJs of flight data
recording systems, it is of interest to consider two instruments
which, it can be said, were the first to be used for recording purposes
under routine operating cqnditions, as an aid in collecting data on
gust loads. The instruments are the accelerometer and the fatigue
meter, and are still to be found in one or two current types of aircraft.
373
Accelerometers and
Fatigue Meters
The structure of an aircraft is designed to withstand certain stresses
which tnay be imposed on it during flight, the magnitude of such
stresses being dependent on the forces acting on the aircraft. All
these forces may be resolved into components acting in the directions of the three mutuaily perpendicular axes of the aircraft,
namely the longitudinal or roll axis, the lateral or pitch axis, and
the vertical or yaw axis.
Force is the product of mass and acceleration, and since the mass
of an aircraft may be considered constant, the forces acting on the
aircraft in flight may be expressed in terms of the acceleration
affecting it. During the manoeuvring of an aircraft, the largest
changes in the accelerations to which it is su_bjected take place in
the direction of the vertical axis. Consequently, the danger of
exceeding allowable stresses and the possibility of failure of some
part of the structure are greatest when excessive accelerations are
applied through the vertical axis. Vertical components of acceleration
are measured by accelerometers and fatigue meters, either of which
may be installed in an aircraft. They operate on the same basic
principle, but whereas the accelerometer provides instantaneous
indications of vertical acceleration, the fatigue meter is designed to
count and record the number of times predetermined threshold
values of vertical acceleration have been exceeded.
Basic Accelerometer
Principle
A basic mechanism is illustrated in Fig 16.1, from which it will be
noted that the principal components are a mass and two calibrated
springs tensioned to statically balance the weight of the mass. The
position of the mechanism shown in the diagram is the one corresponding to normal straight and level flight.
In this condition, the force along the vertical axis is due to the
aircraft's weight, and therefore the aircraft is subject to normal
gravitational force, which produces an acceleration of lg= 32 ft/sec 2 •
Similarly, the force on the mass is due to its weight and so it too is
subject to lg. The weight of the mass extends one spring and allows
the other to contract. The pointer, which is actuated by the lever
arm on which the weight is mounted, moves over a scale calibrated
directly in g units, and for the level flight condition it is positioned
at what may be termed the datum value of lg.
It will be noted from the diagram that the scale is calibrated in
positive and negative values of g; these corres;;,ond respectively to
upwards and downwards accelerations along the vertical axis. Thus,
the load supported by the wings of an aircraft in any manoeuvre is
the product of the aircraft's weight and the accelerometer indication.
Under vertical acceleration conditions brought about by manoeuvring of the aircraft, gusts or turbulent air, the mass will be displaced
thus changing the tension of the springs until it balances the force
374
Figure 16. J Basic accelero·
meter mechanism.
OISPt.ACl:MENT OUE TO
NEGAT1VE ACCELERATION
OISPlACEMENT DUE TO
POsmvE A CCU.ERAT10N
Figure 16.2 Typical accelero·
meter. 1 Instantaneous g
pointer, 2 pointer drive,
3 compensating gears, 4 control springs, 5 masses,
6 damping device.'
imposed and produces the corresponding change in indication. A
positive acceleration moves the mass downwards and a negative
acceleration moves it upwards.
The dial presentation and schematic arrangement of the mechanism
of a typical accelerometer are shown in Fig 16.2.
The mechanism consists of two spring-controlled masses mounted
on cantilever arms attached to two rocking shafts. A sector gear
375
attached to one of these shafts meshes with a pinion on the
instantaneous pointer spindle. At the rear end of the second rocking
shaft a sector gear meshes with a pinion coupled to a magnetic eddycurrent drag device. The purpose of this is to damp out violent
pointer fluctuations which could occur under short-period accelerations.
As the accelerometer scale is linear, there must be a linear relationship between acceleration and angular movement of the masses. This
is achieved by using linear springs and attaching them to links secµred
to the rocking shafts at right angles to the cantilever arms. The
rotation of the masses is therefore directly proportional to the
extension of the springs, and hence to the acceleration force imposed.
When the aircraft and instrument are subjected to a vertical
acceleration, each mass moves through an arc causing the rocking
shafts to move in opposite directions, thus driving the instantaneous
pointer to indicate the g force imposed.
In this particular design of accelerometer, it is also possible for
horizontal components of acceleration which may act in the plane
of the instrument dial to exert forces on the two masses causing them
to rotate the rocking-shafts to false positions. However, since the
rotation of each shaft would be in the same direction, any tendency
to rotate at all can be prevented by gearing the shafts together. This
is precisely the function of the two sector gears at the forward end
of each shaft.
In addition to knowing the instantaneous value of acceleration, it
is also very important to know the maximum acceleration experienced
during a manoeuvre or when flying in turbulent conditions. It is
customary therefore to provide accelerometers with two auxiliary
pointers, one to indicate maximum positive accelerations and the othe
to indicate maximum negative accelerations.
Both pointers are mounted concentrically with the 'instantaneous
g' pointer and are driven by a small plate attached to this pointer. Tot
plate has projections which move two ratchet wheels; engaging in
opposite directions, one wheel drives the maximum positive pointer,
and the other the maximum negative pointer, to the position of the
main pointer. A lightly-loaded pawl and spring hold the auxiliary
pointers at their respective maximum indications until that acceleration is exceeded, or until reset by the operation of the resetting knob
located at the front of the instrument. Both auxiliary pointers are
then returned to the position of the main pointer by the action of
hairsprings.
Since it is the purpose of an accelerometer to indicate acceleration
forces, it is obvious that such forces can be imposed on it in the
course of general handling. To avoid damage which might be sustainec
by careless handling, a locking device is provided to prevent movement
of the masses.
376
Fatigue Meter
In its basic form, a fatigue meter is the same as an accelerometer, i.e.
it comprises a suspended mass and controlling-spring type of verticalacceleration sensing element. However, since it is designed to count
and record the number of times predetermined acceleration threshold
values have been exceeded, its mechanism is of necessity a little more
complex. An external view and schematic of a typical meter are
illustrated in Figs 16.3 and 16.4 respectively.
The mass of the acceleration sensing element consists of a weight
mounted on a cantilever spring which is connected by secondary
springs and a fusee chain to a rotary inertia and damping unit. The
inertia unit is, in turn, conne'cted by a shaft to a wiper arm and brush
assembly which sweep around the face of a commutator. Each
segment of the commutator represents an equal change of applied
acceleration, e.g. O. lg. The centre portion of the commutator
corresponds to the normal gravitational force of lg.
Pairs of commutator segments are connected electrically to relays
which energize electromagnetic digital counter units corresponding
to the selected threshold values, in this case O.Sg, 0.4Sg, 0.75g,
1.25g, 1.55g and 1.95g. Each relay circuit is made up of two
sections referred to as 'lock' and 'release', and from the schematic
it will be noted that the 'lock' sections are connected to the segments
which also correspond to the selected threshold values. The 'release'
sections are connected to segments corresponding to lower values
Figure 16.J Fatigue meter external view.
377
L
R
L
OISPVCEMEHT
OUETONE~TM
ACCELERATION
FRAME
I
1 2 3 4
I~· 1·95 g
1 2 3 4
111· 1·55 g
3 4
11.. 1·25 g
l·Og
~·
-
II
DISP\.ACEMENT
DUE TO f'OSITM
0-45
3~
I
ACCELERAT10N
lo.75 9
SEU:CTEO THM~ VALUES
I
SECONOAII
SPRINGS
+
L
Figure 16.4 Fatigue meter schematic arrangement.
(}:
1 2 3 4
I•· 0-45 g
1 2 3 4
11•· (}06 g
which are a definite number of increments apart. For example, the
relay controlling the counter recording 0.75g has its 'lock' section
connected to segment 0. 75 and its 'ffilease' section connected to
segment 1.05, and incremental difference of 0.30g. It should also
be noted that, in the positive and negative directions, the connections
for the 'release' sections are taken from segments which precede those
connected to the 'lock' sections.
The power for operating the relays and counters is 24 V direct
current. The instrument is rigidly attached to the aircraft structure
in accessible underfloor compartments or special compartments
within a passenger cabin.
Operation
As in an indicating accelerometer, the mass can be moved from the
normal lg position by the forces acting on it. Assuming that the
aircraft and meter are subjected to a vertical acceleration in the
positive direction, then the mass will move downwards and will
cause the wiper arm to move in a clockwise direction through an
angular distance proportional to the applied acceleration.
When the wiper brushes reach the first commutator segment
connected to the 'lock' section of a relay and associated counter, a
circuit is completed through relay coil ( 1). The relay contacts are
thus closed and connect a positive supply to the counter solenoid
378
which operates and cocks the counter; no reading is recorded at
this moment. In addition to completing a circuit to the counter
solenoid, the relay contacts also feed a positive supply to a hold-in
coil (2) of the relay. If the acceleration is of such a value that the
mass and wiper arm cause the brushes to make contact with the
succeeding selected segments in the positive direction, then their
associated relays will be energized in a similar manner and the counters cocked. Thus the hold-in coils keep the relays energized even
through the brushes leave their particular commutator segment.
When the acceleration forces begin to decay to the Jg value, the
brushes return over the commutator, making contact with segments
connected to the 'release' sections of the relays. Reference to the
diagram shows that when a relay is still energized its contact also
feeds a positive supply to the segment connected to its 'release'
circuit segment. Therefore, when the brushes reach the segment a
circuit is completed which shorts out the relay hold-in coil, causing
the contact to open. The counter solenoid is thus de-energized and
thus releases a spring-loaded pawl of the cocking mechanism to
complete the counting operation.
The reason for arranging the circuit in this manner is to prevent a
second count being made when the brushes move to a high g level
and return over a selected segment. Furthermore, it also prevents
extra counts when accelerations of small amplitude and high frequency
are superimposed on the gust or manoeuvre acceleration.
Excessive overshooting of the mass is prevented by eddy-current
drag induced by the rotation of a bell-shaped damper in a permanentmagnetic field.
As fatigue meters are only required to count and record the vertical
accelerations experienced in flight, it is necessary to include in the
electrical circuit an automatic means of switching the power supply
on and off. This is accomplished by an airspeed switch unit (see page
96) whose contacts are set to make the circuit just after take-off and
to break it just before landing.
The acceleration counts recorded by a fatigue meter are essential
for relating the accelerations experienced by an aircraft to what is
termed its safe fatigue life. In consequence, readings have to be
noted at regular specified periods and entered on special tabulated
data sheets which on completion are used for analysis and calculation
of fatigue life.
Flight Data Recorders
As noted in the introduction to this chapter, the mandatory requirements for aircraft to be equipped with flight data recorders relate
primarily to the acquisition of data which will prove valuable during
investigations into the cause of a crash. Thus, the overall development of data recording systems has peen built up using such data
acquisition as the foundation.
379
The relative value of flight recording has, from investigators'
experience, been indicated in respect of two very general categories
of crash. The first category is that in which some form of aircraft, or
system, malfunction has been the primary cause. In attempting to
establish the primary failure , analysis of the wreckage would probably
play the principal role, and this would be supported by such 'precrash' recorded data as time, altitude, speed and acceleration. The
second category is that in which the crash has an operational cause,
i.e. one in which no engineering defect or deterioration in performan.ce
of the aircraft has occurred . Some examples of this are loss of control
resulting from improper handling by the flight crew, adverse weather
conditions, or navigational factors. In such a category of crash, the
role of wreckage analysis would be limited to establishing data concern
ing the aircraft's final configuration, and broad details of impact
attitude and speed . The role of the flight data recorder in this case
would be more prominent, since it would provide a time history of
the principal manoeuvres throughout the flight.
Parameters Selected
In the selection of the mandatory parameters for crash investigation
purposes, the objective is to obtain either directly or by deduction
from the recorded data, the following information:
1. the aircraft's flight path and attitude in achieving that path;
2. the basic forces acting on the aircraft, e.g. lift, drag, thrust and
control forces ;
3. the general origin of the basic forces and influences, e.g. navigational information, aircraft system status information.
The mandatory parameters are specified in national regulations
for civil aircraft operation and generally relate to the following :
(a) time (GMT or elapsed);
pressure altitude;
(c) airspeed;
(d) vertical (normal) acceleration;
(e ) magnetic heading;
(j) pitch attitude.
(b)
Methods of Recording
There are two principal methods of recording in use; (i) trace recording and (ii) electro-magnetic recording.
Trace Recording
As the name suggests, this is a method by which changes in the
380
required parameters are traced out in a graphical form. Fundamentally, it may be likened to that adopted in chart recording instruments used in industrial plants for the recording of temp~rature,
pressure, etc. However, in lieu of calibrated paper charts and recording pens, a flight recorder uses either an .aluminium or high nickel
content steel foil tape on which traces are engraved by metal-tipped
styli. An example of a trace recorded during the take-off of an aircraft is shown in Fig 16.5.
Figure 16.5 Example of a
trace recording.
VERTICAL ACCELERA'rlON
1 HOUR MARK
1 MINUTE MARK
15 MINUTE MARK
The foil tape which typically is 5 in wide, 200 ft long, and 0.001 in
thick, is wound on supply and take-up spools, sprockets and guide
rollers contained within a magazine assembly. The take-up spool is
driven by a high-ratio gear train and an electric motor, the spool in
tum transporting the foil tape from the supply spool. The motor
operates in conjunction with a timer to produce a controlled rate of
tape transport. Depending on the type of recorder, the motor may be
either of the permanent magnet d.c. type the speed of which is
controlled by a chronometric governor mechanism, or of the constantfrequency a.c. type. Typical speed control values and gear train ratios
are such that the tape transport rate is 6 in per hour ; thus, a 200 ft tape
provides for a recording time of 400h. The magazine and drive motor
assemblies are contained within the main housing of the flight recorder
together with the transducer and servomechanisms required for
actuating a stylus corresponding to each parameter to be recorded.
The recorded parameters are altitude, airspeed , magnetic heading,
verticaf acceleration, and time. Altitude and airspeed are measured by
independent transducers containing conventional type pressure-sensing
elements connected to an aircraft's pitot-static system. The sensing
elements are each mechanically coupled to a stylus-actuating arm in
such a way that displacement of an element moves its stylus to a
position on the tape equivalent to altitude or airspeed as appropriate.
Each stylus is pressed against the tape by a pressure bar in such a
381
manner that recording is done at one-second intervals.
Magnetic heading information is supplied from the aircraft's main
compass system, the signals being fed to a synchro control transformer
within the flight recorder. Error signals are amplified and fed to a
servomotor which drives the transformer rotor to the null position,
and at the same time, moves the heading stylus to a position on the
tape equivalent to magnetic heading. Recording is also done at onesecond intervals and with a pressure bar acting against the stylus.
Since the heading stylus is mechanically actuated through two sectors
each equal to 180°, it is necessary to differentiate between easterly
and westerly headings, and thereby resolve any ambiguity when
analysing the traces. This is effected by a separate solenoid-actuated
stylus which marks a line at"one position on the foil tape when the
heading stylus position corresponds to headings in the 0-180° sector,
and at another position when headings are in the 180- 360° sector.
The vertical acceleration stylus is also actuated by a servomotor
system which is supplied, via an amplifier, with signals from an
acceleration transducer suitably located near the aircraft's centre of
gravUy.
Elapsed time is recorded continuously from the moment power is
supplied to the recorder, in the manner shown in Fig 16.S;·the time
re~ording stylus is operated by a solenoid controlled by the timer
which operates in conjunction with the tape drive motor.
An indicator showing the amount of 'tape remaining' is also
provided on the recorder, and is operated by a linkage to a lever
which rides on the tape so that it senses the radius of the tape
remaining on the supply spool; the indicator is calibrated in hours.
A second indicator is also provided at a recorder control panel
location in the flight crew compartment, and is operated by signals
transmitted from a potentiometer connected to the lever linkage
.in the recorder.
Electro-magnetic Recording
In this method, the analogue signal data from the transducers
appropriate to each parameter are first encoded , i.e. transformed into
a digital coded form in accordance with the binary scale of
notation. The encoded signals then pass through such sub-channels
as conditioning, logic, multiplexing, and modulation, and are fed as
a series of pulse currents to an electromagnetic recording or
'writing' head, the currents being of either positive or negative
polarity and corresponding respectively to a '1 state' or a 'O state' of
the binary code. The magnetic core of the 'writing' head has a very
small air gap (typically 0.00 I in) between pole pieces, and is located
in very close proximity to the recording medium which, in some
recorders, may be a plastic-based tape coated with ferrite material,
382
PINCH AOLLEA ARMS
FLYWHEEL - - - --TAPE TENSION CONTROL ARM - - - - ~
TAPE HEADS
- WRITE'; '8' - 'READ'J - --
---.
- -- - --
Figure 16. 6 Example of
transporting magnetic tape.
-
RELAY
while in others, it may be a stainless steel wire. The air gap is kept
as small as possible to permit high packing density , and furthermore,
in conjunction with the small space between the g_ap and recording
medium surface, it minimizes any interference the generated flux
might have on adjacent sections of the medium already magnetically
coded.
The tape or wire is wound on spools and guide milers, and is
transported over the 'writing' head by a synchronous motor supplied
with a 115 V a.c. The transport rate is typically, 1.75 in per second
383
for magnetic tape and 2.5 in per second for wire. An example of one
method of transporting magnetic tape is illustrated in Fig 16.6.
Whenever a current pulse is applied to the coil of the 'writing' head,
.a leakage flux fringes out between the pole pieces of the core, and as
the tape or wire is transported past the pole pieces, magneticallypolarized 'spots' or sections are produced in a serial manner on the
tape or wire, the polarities being in accordance with the binary-coded
pulse currents applied to the coil. Thus, and as shown in Fig 16. 7, a
pulse current assigned the value equivalent to binary state 'I' will
;, __n___nsu
III II II
•
I
•2
(+)
SL
I
I
I
I
I
I
I
IIIII II
iiiiiiii
l
RECORDING MEOIUM
...
SIGNAL INPUT i 2
SIGNAL INPUT i1
Figure 16. 7 'Writing' head
operation.
384
BINARY
STATE
I~1.1-i
I, .i-i ·1-i
IO
I O • , l O I,
I I II I I
to,
I
I
magnetize with a positive polarity, while a pulse current equivalent
to binary state 'O' will magnetize with a negative po.larity . In binary
coding terms, each state is referred to as a 'bit' (derived from binary
digit) and a number of 'bits' corresponding to the parameters measured comprise a 'word'. Since the 'words' are polarized on the tape
or wire, then the latter provide what is termed a 'magnetic store' of
information. The number of 'bits' which can be stored in a unit lengtl
of tape or wire gives the packing density; some typical values are 900
bits per inch for tape, and 414 bits per inch for wire. In magnetic tap1
recording, coded data are stored on a number of tracks into which a
>/2 in wide, 850 ft long tape is divided. The 'writing' head has a corresponding number·of magnetizing coils each of which are automatically
selected to provide sequential track recording.
The technique of electro-magnetic recording is further illustrated in
Fig 16.8. To provide correct interpretation of the stored information
when 'reading' the tape or wire, the signal inputs to. the magnetizing
coils of the 'writing' head must first be assembled into an identifiable
format. This is done by multiplexing the pulses in a signal conditionir
unit, so that the format comprises a frame of 64 'words', each word
containing 16 'bits' ; a frame is of Is duration. The first word in each
frame is a frame synchronizing (or relative time) signal comprising a
repetitive series of 'I s' and 'Os', the remaining 63 words being avail-
Figure /6.8 Electro-magnetic
recording.
FRAME TIME • 1 SECONO
___
(64 WOROSi
I·
___j
I
'
'
:::~11m 11111111111 111mrrn=
f
WORO 1 WORO 2
(FRAME SYNC)
'
I
WORO ,· INFORMATION
SYNC
4 BITS
2 BITS
I
~
I
WORD 3
I
WORO 64
I
FRAME SYNC
I
I
WORO ,· INFORMATIO~ WORO i INFORMATIO~
SYNC
3 BITS
SYNC
3 BITS
2 BITS
2 BITS
I
I
TYPICAL WORO STRUCTURE • 16 BITS
MAGNETIZATION
0
BINARY COOE
I
0
0
1
0
0
0
1
0
0
FRAME SYNC. WORO I
TRANSOUCER 'A" - WORO 2
~
_.,,.._l•--
_
TRANSOUCER 'B' - WORD 3
-~
I-!+I -!+I -1 -1 -1•I • 1-1 + 1-1 • 1-1 •I• 1-1•1-1-1 -1- I+ I • 1-1 • I-I•1- I· I• I
2
·1 ·12· 2 2' 21 ·12· 2 2'j i2 2' 2°·1 ·12• 2' 2' 2•i ·1 2· 2• 2'il ·122 2' 20·1
60
8
BINARY
D
·
1 0
{
1
0
256
0 0
w
•
·
4
1
1 .0
i
40
DECIMAL
• 299
1
I I
1 1 0
i
3
O 1 O
I
H
1 0
0 0 0
I
O 1
1
0
•
43
I
40
1 0 1
•
0
j
1
1
3
1
·
I
able for pure information on the required parameters. Each of these
words include three interposed two-bit (0 I) word synchronizing
signals which divide the pure information into groups of 4 bits, 3 bits
and .3 bits respectively. Thus, each word contains 10 information bits,
and because in binary notation, the 'bits' are understood to be
multiplied by progressively higher powers of 2, then each word has
385
an equivalent full-scale decimal count of I 024 (2 1 0 ). In the example
shown in Fig 16. 7, the ten information bits of word 2 from a transducer 'A' have been magnetically encoded O I O 0 , I O I, and O I 1;
conversion to decimal requires only the 'ls' to be raised to higher
powers of 2 and always from right to left, so that the code therefore
converts to 256 + 40 + 3 and so is equal to 299. Similarly , the word
3 from transducer 'B' converts to 43 in decimal notation.
Information on all parameters is sampled within the signal
conditioning unit of the flight recorder system at periodic intervals
to construct the multiplex pattern, some parameters being sampled
more frequently than others. Some details of sampling rate and word
allocation based on one type of wire recorder in current use are
given in Table 16.1 .
Table 16.1
Parameter
Sampling rate
(per second)
Word number allocation
Time
Acceleration
8
6, 14, 22, 30, 38,46,54,62
Pitch
8
2, 10, It 26, 34,42, 50,58
Altitude
9
Airspeed
13
Heading
29
Flight number and date in
three groups of 10; l at and 2 a t -
;f
3,5, 35, 37
For identification purposes, it is necessary to record flight reference
information, i.e. flight number and date. This is done by operating
'thumbwheel' switches on a flight encoder panel (see Fig 16.9)
located in the flight crew compartment. The switches are numbered
and when selected they feed an output to the signal conditioning unit
for multiplexing and transmission to the recorder in groups of binary
coded information.
Data Recovery and Analysis
The handling of data derived from a flight recorder system falls into
two distinct phases: (i) the recovery of raw data from the recording
medium, and (ii) analysis of the information in relation to a particular accident or, in the case of data acquisition systems, in relation to
386
Figure 16. 9 Example of a
flight encoder panel.
data sampling for the purpose of routine monitoring. In addition to
the recovery of raw data, the first phase also includes the application
of cor:-ections for calibration and other system errors, e.g. pressure
errors of a pi tot-static system, and their presentation in a form
suitable for computation and analysis. The second phase requires
the introduction of other information such as design, performance or
operational data, relevant to the type of aircraft.
The nature of the equipment required for data recovery depends
on the recording method adopted. In the trace recording method,
and after a foil tape is removed from the recorder, the various
traces are translated into values of the recorded parameters, from a
transparent overlay having marked co-ordinates which can, with
precision, be registered against the traces. The analysis of electromagnetically recorded data requires the use of electronic processing
and presentation equipment to convert the digital-coded information
back into analogue, i.e. to decode.
Basically, the decoding process requires that the magneticallycoded tape or wire be transported past a 'reading' head of the
processing equipment. This. head is similar in its electro-mechanical
construction to a recorder 'writing' head . As each magnetized
section of the recording medium passes the head it induces a small
voltage in the coil of the head, the voltage polarity indicating whether
387
a binary 1 or O has been stored. The voltage signals are then
processed electronically to reproduce the recorded parameters in
analogue form and in accordance with the word code originally
assigned to them.
In some types of flight recorder a 'reading' head is also provided in
close proximity to the ' writing' head (see also Fig 16.6). This permits
'in situ' monitoring of the validity of recorded information.
Protection and Location of Flight Recorders
Flight recorders for the retrieval of data for accident investigation
purposes must be adequately protected and so located in an aircraft
that the recording media will survive in the event of a crash situation
(see Fig 16.10). The adverse factors against which protection is
necessary are mechanical damage due to impact and acceleration
forces, fire damage, and chemical attack by hydraulic, de-icing and fire
extinguishing fluids, fuels and acids. Protection must·also be afforded
against the effects of sea water, and to cater for cases where an
accident occurs resulting in wreckage entering deep water, the recording medium container must also be provided with separation and
flotation facilities.
The protection facilities are, of course, 'built-in' to the design of a
Figure I 6. IO Recorder
protection.
388
FIRE PROOFING
ARMOURED STEEL HOUSING
recorder and so may be considered as the primary means of ensuring
survival of the recording media. The location of a recorder in an aircraft is, however, also an important factor to be considered in this
respect, and from the evidence of aircraft accidents it is shown that
the rear fuselage and tail unit strncture is the section most likely to
survive, or be the least severely damaged . For this reason therefore,
a protected flight data recorder is installed in the rear of an aircraft.
In addition to the foregoing protection and location requirements,
it is also necessary to provide recorders with means to facilitate their
identification in conditions following a crash. This is usually done
by adopting a distinctive colour scheme; thus, contrary to the 'black
box' clicht frequently adopted in news media reports of a crash,
the box could well be coloured fluorescent orange.
Aircraft Integrated Data Systems
As stated in the opening paragraphs of this chapter, the capacity of
recorders can be increased to acquire data additional to that
associated with accident investigation. Except for the requirements
laid down for certain certification processes, the acquisition of such
data as airframe and engine systems operation, is non-mandatory.
It does, however, provide an operator with the means of monitoring
the performance of systems on a comparative basis, and from any
variations to obtain advance warnings of possible failures of a
component or a system overall. For example, from the monitoring
of parameters associated with engine performance, viz. pressures,
temperatures, vibration levels, an engine 'signature' can be detected
over a period of time, and this can be used as a performance datum
against which the recorded parameters of other engines can be
compared each time a 'read-out' is made.
Sy~tems of this nature which are known variously by the
acronyms AIDS and DAS (Data Acquisition System) may be used
separately, or in conjunction with an accident recorder. The
schematic arrangement of a system is shown in Fig 16.1 I. The
analogue inputs from the appropriate transducers are supplied to
one or more data acquisition units each containing analogue and
digital multiplexers, signal conditioning circuits, and an analogue-todigital converter. After processing and conversion, the signals are
'se rialized' and transmitted to the logic unit along what is termed
a digital high way; this provides a common data link for all the
acquisition units which may make up a system, and so keep interconnection cabling to a _m inimum. The logic unit converts the
information from the acquisition units into the particular formats
appropriate to the various output requirements. The control unit
enables the flight crew to insert flight information into the record,
and also provides for continuous system monitqring and simple test
389
Figure 16.11 Schematic
arrangement o.f an integrated
data system.
CONTROL
UNIT
ANALOGUE DATA INPUTS
TIME
AIRSPEED
ALTITUDE
HEADING
ACCELERATION
PITCH ATTITUDE
PRESSURES
TEMPERATURES
ENGINE POWER
LONG TERM
AND/ OR
QUICK-ACCESS
RECORDERS
t
CONTROL
SURFACE
POSITIONS
ILS SIGNALS
AFCS
ENGAGE
A ND MODES
DATA
ACQUISITION
UNIT/ 5
I
LOGIC UNIT
I
:
II
ACCIDENT
RECORDER
I
SERIAL DIGITAL HIGHWAY
facilities. The recorders used in these systems are normally of the
wire type, and are so installed in an aircraft as to be easily accessible
for random monitoring and 'read-out' of recorded data.
Questions
390
With the aid of a diagram explain the basic operating principle of an
accelerometer.
16.2 What are the essential differences between an accelerometer and a
fatigue meter?
16.3 How is it ensured that a fatigue meter does not record (a) second
'counts', (b) during taxiing and landing?
16.4 What parameters are mandatory for crash investigation purposes?
16.5 Explain briefly how changes in value of any of the parameters are
recorded by the trace method.
16.6 At what time intervals are parameters recorded by a trace recorder?
16.7 Describe how traces of easterly and westerly headings are differentiated.
16.8 Which of the stylus mechanisms of a trace recorder are solenoidactuated?
16.9 In connection with flight data recorders, what is meant by 'transport
rate'?
16.10 What is the recording medium used in recorders of the electro-magnet
type?
16.11 Briefly describe how the encoded signals appropriate to a parameter
are'written' on the recording medium.
16.12 What is meant by the term 'packing density' in relation to a recording
medium?
16.1
16.13
16. 14
16.15
16.16
16. 17
How many 'bits' and 'words' comprise a frame?
Convert the following information ' bits' of a recorded ' word' into
decimal notation: 1 0 0 I O O O O O I
How are recorded data recovered and analysed?
What are the protection and location requirements for an accident
data recorder?
Describe some of the advantages gained from the use of integrated
data systems.
391
Principal symbols and
abbreviations
Symbols for quantities are in italic type, and abbreviations for t'1e names of units (unit
symbols) are in ordinary type.
A
a
B
bhp
C
oc
c/s
F
ft/min
ft/h
2
ft/s
g
H
Hz
I
in Hg
K
lb/gal
lbf/in2
M
m
mA
392
ampere
speed of sound
magnetic flux density
brake horsepower
capacitance
degree Celsius
cycle per second
farad
foot per minute
foot per hour
foot per second per second
acceleration due to gravity
magnetic field strength
hertz (frequency)
electric current
moment of inertia
inch of mercury
Kelvin
pound per gallon
pound force per
square inch
Mach number
torque
magnetic moment
mass
milliampere
mbar
mV
m.p.h.
mmH,O
N
N
Nm
Pa
pF
R
rad
rev./ min.
T
V
V
w
Wb
Cl
µF
p
<I>
n
w
millibar
millivolt
mile per hour
millimetre of water
newton (force)
number of turns of a coil
Newton Metre (torque)
2
pascal (N/m )
picofarad
resistance
radian
revolution per minute
period, periodic time
volt
velocity
weight
weber (magnetic flux)
temperature coefficient of
resistance
microfarad
density
resistivity
magnetic flux
ohm
angular velocity
(radians per second)
Conversion factors
1. Pressure
To convert
Atmospheres
into
inches Hg (0° C)
inches H 20
kilogrammes per sq. cm
millibars
millimetres Hg (0° C)
pi,e zes
pounds per sq. in.
*pascals
Incnes Hg
atmospheres
inches H20
kilogrammes per sq. cm
millibars
millimetres Hg
pounds per sq. in.
Inches Ht)
atmospheres
inches Hg
kilogrammes per sq. cm
millibars
pounds per sq. in.
pascals
Kilogrammes per sq . cm atmospheres
inches Hg
millimetres Hg
piezes
multiply by
29.921
406.9
1.0333
1013.25
760.00
IO 1.331
14.696
101325.000
0.03342
13.60
0.03453
33.8639
25.40
0.4912
2.458 X 10-3
0.07355
2.540 X 10-3
2.490
0.03613
249.089
0.000987
28.96
735.54
1.0197
* The pascal (Pa) is an SI unit and is the pressure produced by a force of l newton applied,
uniformly distributed , over an area of 1 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 '!bf/sq. in.' is now adopted in lieu of 'lb/sq.in.'.
393
pounds per sq. in
14.223
Millibars
atmospheres
inches Hg
millimetres Hg
pounds per sq. in.
pascals
0.01450
0.02953
0.7450
0.0145
100.00
Millimetres Hg
inches Hg
kilogrammes per sq. cm
pascals
pounds per sq. in
0.03937
0.0013596
133.322
0.019337
Pounds per sq. in.
atmospheres
inches Hg
inches H2 0
kilogrammes per sq. cm
millibars
millimetres Hg
pascals
Pascals
pounds per sq. in.
0.06804
2.03596
27.68
0.0703
68.9476
51.713
6896.55
0.0001450
2. Velocity
multiply by
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
16.67
0.6214
Knots
feet per minute
feet per second
394
101.34
1.689
kilometres per hour
miles per hour
1.8532
1.1516
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
3. Volumetric
To convert
into
Cubic centimetres
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 ro-s
0.06102
2.1997 X 10_.
0.001
1.7598 X 10-3
8.7988 X 10_.
28316.85
1728.00
6.2288
28 .32
59.84
29.92
7.481
16.39
5.787 X 10-"
3.6047 X 10-3
0.01639
0.5688
0.2844
4546.087
0 .160544
277.42
4 .54596
1.201
395
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
1.13649
US gallons
cubic centimetres
cubic feet
cubic inches
imperial gallons
litres
3786.44
0.1337
231 .00
0.8327
3.785
4. Angular Measure
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
396
multiply by
60.00
0.0111
0.0175
0.0166
1.852 X 10-4
2.909 X I0-4
5. Angular Velocity
To convert
into
Degrees/second
radians/second
revolutions/second
revolutions/minute
Radians/second
degrees/ second
revolutions/second
revolutions/ minute
Revolutions/ second
degrees/second
radians/second
revolutions/minute
Revolutions/minute
degrees/second
radians/second
revolutions/second
multiply by
0.01745
2.778 X 10-3
0.1667
57.30
0.1592
9.5493
360.00
6.283
60.00
6.00
0.10472
0.01667
6. Coriolis Acceleration
Error
The value of coriolis acceleration is given by 2 nGsin X. ln the northern hemisphere it acts to
the left of track, tilting a flux detector element anticlockwise about an axis along magnetic
track. Therefore, the tilt o: is given by:
o:
horizontal acceleration
=- - - g- -- -- =
2 n Gsin X
g
d'
ra ians
where
n = earth rate in radians per second
G = ground speed in feet per second
g = gravity acceleration in feet per second per second
7. Gimbal Error
The equation for gimbal error is as follows:
tan hind = tan hact X cos 8 X sec 8 - tan 8 X sin 4>
where
hind
hact
= indicated heading
= actual heading
397
q,
8
= roll angle
= pitch angle
Since pitch angles are very much smaller than roll angles, the equation may be simplified as:
tan
398
hind
= tan hact X cos q,
Tables
1. Standard
Atmosphere
Temperature
Velocity of
sound
Altitude
Pressure
ft
-1,000
0
1,000
2,000
3,000
4,000
5,000
millibars
in. Hg
oc
ft/sec
1050·41
1013·25
977-17
942·13
908· 12
875· 10
843-07
31·019
29·921
28-856
27•821
26·817
25·842
24·896
15·234
14·696
14·172
13-664
13·170
12-691
12·226
+16·981
15.000
13·019
l 1·038
9·056
7-075
5-094
1119·9
1116·1
1112·3
1108·4
1104·5
1100·7
1096·7
6,000
7,000
8,000
9,000
10,000
811 ·99
781·85
752-62
724·28
696·81
23·978
23·088
22·225
21·388
20-577
l 1·775
11·338
10·913
10·502
10·104
3· 113
l· 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
619·43
595·24
571·82
19· 791
19·029
18·292
17·577
16·886
9·718
9.344
8·981
8·630
8·291
-6·793
- 8·774
-10·756
- 12·73 7
- 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
1044·7
1040·6
1036·5
21,000
22,000
23,000
24,000
25,000
446·45
427·91
410·00
392·7 l
376·0 1
13· 184
12-636
12·107
I 1·597
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
1019·8
10 15·6
26,000
27,000
359·89
344.33
10·628
10·168
5·216
4·991
- 36·511
-38·492
1011·4
1007-1
lbf/in 2
399
Altitude
Temperature
Premae
Velocity of
sound
fl
28,000
29,000
30,000
millibars
in Hg
9.725
9·298
8·885
oc
ft/sec
329·32
314·85
300·89
4.773
4·563
4·361
-40·474
-42-455
-44·436
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
36.000
37,000
38,000
39,000
40,000
227·29
216·63
206·46
196·77
187-54
6·712
6·397
6·097
5·811
5·538
3·293
3·139
2·991
2-851
2·717
-56·323
968-1
41,000
42,000
ft3,000
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
-56·500
967-7
46,000
47 ,000
48,000
49,000
50.000
140·56
133-96
127•67
121-68
115·97
4·150
3-956
3·770
2·036
1·940
1-849
1-16:l
1·679
lbf/in
3.593
3-425
2
2. Pressure/ Airspeed
Equivalents
Pressure in millimetres
of water
Pressure in millimetres
of water
Airspeed
Airspeed
When
speed
in
m.p.h .
400
When
speed
in
knots
When
speed
in
km/h
When
speed
in
m.p.h.
When
speed
in
knots
When
speed
in
km/h
640·3
676·0
706·9
752·9
793·0
10
20
30
40
50
60
70
80
90
100
1·25
5·00
11 ·25
20·00
31·29
45·08
61·40
80·24
101 ·6
125-6
1·66
6·63
14·93
26·55
41'51
59-81
81 ·47
106·5
134'9
166-7
0·5
1·9
4-3
7-7
12'1
17·4
23-7
30·9
39·2
48·4
360
370
380
390
400
410
420
430
440
450
1711
. 1813
1918
2026
2138
2254
2373
2496
2622
2752
2308
2447
2591
2740
2895
3054
3219
3389
3564
3745
110
120
152'1
181 ·2
202·0
240·7
58·5
69·7
460
470
2887
3024
3931
4124
Pressure in millimetres
of water
Pressure in millimetres
of water
Airspeed
Airspeed
When
speed
in
km/h
When
speed
in
m.p.h.
When
speed
in
knots
When
speed
in
m.p.h .
When
speed
in
knots
130
140
150
160
170
180
190
200
212-8
247 ·1
284·0
323·6
365-8
410·8
458-4
508·7
282-9
328·6
377-8
430·6
487·0
547·1
610·8
678·3
81 ·8
94-9
109·0
124·1
140·2
157'4
175-4
194-4
480
490
500
510
520
530
540
550
3166
3312
3462
3616
3774
3937
4103
4274
4322
4526
4736
4952
5174
5403
5638
5880
210
220
230
240
250
260
270
280
290
300
561·9
617-8
676·5
738·0
802-5
869 ·8
940·1
1013
1090
1169
749·6
824·6
903·5
986·3
1073
1164
1259
1358
1461
1569
214·5
235'6
257-7
280·8
304·9
330·1
356·2
383·5
411·7
441·0
560
570
580
590
600
610
620
630
640
650
4450
4630
4815
5004
5198
5397
5601
5810
6024
6243
6129
6384
6647
6917
7194
310
320
330
340
350
1251
1337
1425
1517
1612
1681
1797
1918
2043
2173
471 ·4
502'8
535·3
568·9
603·5
660
670
680
690
700
6467
6696
6931
7172
7418
When
speed
in
km/ h
3. Mach No./Airspeed
Relationship
Speed of
sound
Height
(ft)
Mach number
Abs
temp
oc
ft/sec.
m.p.h.
0.3
0.4
0.5
0.6
0.7
0.8
0.9
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
Airspeed in m.p.h.
0
5000
10000
15000
20000
25000
30000
35000
36090
288
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.J
198.l
304.6
299.4
294.0
288.5
282.9
277 .2
271.4
265.5
264.l
380.8
374.2
367.5
360.6
353.7
346.5
339.2
331.8
330.2
457.0
449.0
440.9
432.7
424.4
415.9
407 .1
398.2
396.2
533.l
523.9
514.4
504.8
495.1
485 .2
474 .9
464.6
462.2
401
4. Temperature/
Resistance Equivalents
Nickel sensing elements
oc
Minus
Plus
100
200
10
20
30
40
50
60
70
.8 0
90
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
0
100·0
152·2
218·8
Platinum sensing elements
oc
Minus
Plus
100
200
300
400
500
10
20
30
40 '
50
60
70
80
90
105-9
114· l
154·2
193-2
231 ·0
267-6
303·0
101·8
118· l
158·2
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 ·l
138·3
177'8
216·0
253 ·]
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
0
110·0
150·3
189·4
227·3
264·1
299·5
5. Temperature/Millivolt
Equivalent of Typical Iron
v Constantan Thermocouples
Cold junction at 0°C
oc
0
100
200
300
400
500
600
700
800
402
0
0
5.02
10.28
15.62
20.93
26 .28
31.75
37.77
44 .97
10
20
30
40
50
60
70
80
90
100
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
30.62
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
6. Temperature/Millivolt
Equivalent of Typical Copper
v Constantan Thermocouples
Cold junction at 0°C
oc
0
100
200
300
400
500
0
10
20
30
40
50
60
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
12.4 7
18.28
24.40
70
80
90
100
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
4.27
9.25
14.75
20.68
26.88
7. Temperature/Millivolt
Equivalent of Typical Chrome!
v Alumen Thermocouples
Cold junction at 0°C
oc
0
10
20
30
40
50
60
70
80
90
100
0
100
200
300
400
0
4·10
8·13
12-21
16·39
o..;o
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
1·6 1
5'73
9·74
13'87
18·08
2·02
6·13
10·15
14·29
18·50
2'43
6·53
10·56
14·7 1
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
20·2 1
4·10
8·13
12·21
16·39
20·64
500
600
700
800
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
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 ·3 )
1000
1100
1200
1300
1400
41 ·31
45· 14
48·85
52·41
55'8 1
41'70
45·52
49 ·21
52'75
42 ·08
45·89
49 ·57
53· 10
42'47
46·27
49·94
53-45
42'86
46·64
50·29
53'79
43'24
47·01
50·65
54· 13
43-6 2
47·38
5 1·00
54·47
44·00
47·75
5 1·36
54·8 1
44· 38
48· )2
5 1·71
55· 15
44·76
48·48
52·06
55·48
45 · 14
48·85
52'4 1
55·81
403
8. Nominal Dielectric
Constants and Densities
of Fuels
Density D
Obs/gall)
Dielectric constant
K
Fuel
type
91/98
100/130
115/145
JP-1
JP-3
JP-4
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
6·83S
6·464
6·520
6·414
6·379
6·308
7'177
6·827
6·880
9 . ARINC Standard
ATR Case Sizes
Approximate
vol. (in3 )
Short% ATR
Long% ATR
Short 31a ATR
Long 318 ATR
Short 'n ATR
Long 'Ii ATR
Short 3A ATR
Long '%ATR
1 ATR
l'h ATR
404
215
335
340
530
470
725
720
1,120
1,510
2,295
Width (in)
±0.031251n
2.250
2.250
3.5625
3.5625
4.875
4.875
7.50
7.50
10.125
15.375
Length (in)
12.5625
19.5625
12.5625
19.5625
12.5625
19.5625
12.5625
19.5625
19.5625
19.5625
Max
height '(in)
7.625
7.625
7.625
7.625
7.625
7.62S
7.625
7.625
7.625
7.625
Typical pressure (Pl
temperature (T) and
rotational speed (N)
notations.
0 Ambient, I intake, 2 L.P.
compressor delivery, 3 H.P.
compressor delivery , 4 turbine
entry, 5 H.P. turbine exit,
6 L.P. turbine exit, 7 exhaust,
8 propeUing nozzle.
P,
r,
405
Solutions to numerical
questions
.
B
Deviation on East - Deviation on West
Coe ffi1c1ent =
2
8.6
=
+4 - (-2)
2
+6
= - =+30
2
. tC
Deviation on North - Deviation on South
Coe ffi1c1en
=
2
= +4 -(-1) = +5 = +2.50
2
Coefficient A =
2
Deviation on N + NE + E + SE + S + SW + W + NW
8
_ (+4) + (+2) + (+4) + (+3) + (-1) + (-2) + (-2) + (0)
8
+8
=-=+lo
8
11.4(b)
40°C = 40 + 273.15 = 313.lSK
= 40 X
5
11.5
9
+ 32
~
77°F = (77 - 32) X
ll.7(a)
l 04 ° F
= 25°C
Rr=R 1 +R1+R3"'20+15+40:a:75n
l
l
I
I
l l.7(b) · Rr =
+R + R
R.
1
3
so that
I
Rr = - - "'2·847
0 ·351
406
=
n
1 I
"'sI + ii+
7= 0.125 + 0.083 + 0.143 = 0.351
The applied voltage is 24 V; therefore, from the expression I"" V/Rr,
the current flowing in the series circuit is 0.32 A, and in the parallel
circuit, 8.424 A.
11.8
For a parallel circuit containing only two resistances,
R
-
RI R1 - 20 X 45
+ R1 - 20 + 45 "" 13.84 n
T - R1
11 .9(b) Cross-sectional area of wire is
The resistance is R
=p
I/a
a=% X 0.08
= 1. 7 x
Io~
x
2
= 0.00503cm 1
°°
2
0.~
"" 0.406
0 5 03
n
407
Index
Absolute altitude, 74
Absolute permittivity, 316
Absolute pressure, 298
Absolute zero, 259
Acceleration errors
compasses, 1 72
gyro horizons, 142
Accelerometer, 374
Actinic line see Magnetic equator
A.C. ratiometer see Ratiometer pressure
gauge
A.D.F., 196
Agonlc lines, 166
AIDS, 373, 389
Air data computer, see Central air data
computer
Air mass flow, 348
Aircraft magnetic components
hard-Iron, 206
rod components, 211
soft-iron, 206
total effects, 212
Airspeed indicators, 86
Airspeed switch units, 96
Air temperature sensors, 269
Alpha-numeric display, 26
Alternate pressure sources, 59
Altimeter, 6 7
Altitude
absolute, 74
errors, 71
switches, 7 9
alerting system, 78
reporting system, 80
Ampere per metre, 161
Aneroid baro meter, 67
Angle of dip see Magnetic dip
Angular momentum, 117
Annual change, 166
Annunciator see Synchronizing indicators
Anti-backlash gear, 13
Aperiodic compass see Direct reading
-compasses
Apparent A, 217
Apparent drift, 122
Apparent tilt see Transport wander
Approach horizon, 359
ARINC specifications, 4
Artificial horizon see Gyro horizon
Atmospheric pressure, 64
Atmospheric temperature, 64
Attitude flight director, 359
ATR,5
Backlash, 13
Balancing coil, 269
'Banana-slot' compensator, 89
Bank indication
ball-in-tube method, 153
gravity-weight method, 152
Barometer conventions and tables, 67
Barometric pressure setting, 70, 74
Base metal thermocouple, 280
'Basic six' layout, 39
'Basic T' layout, 39
Bellows see Pressure measurement
B/H curve, 187
B!rnetal strip, see also Cold-junction
temperature compensation
Binary code, 82, 382
Binary notation, 82, 385
'Bits', 82, 384, 385
Boiling point, 258
Boost gauges see Manifold pressure
gauges
Blind flying panel, 39
'Bottom heacviness' see Pendulosity
errors
Bourdon tube see Pressure measurement
Bridge lights su Instrument illumination
British Standard, 4, 67
'Bug', 94, 198
Cabin altimeters, 77
Caging knob, 1 77
Capacitance
governing factors, 315
in a.c. circuits, 316
in series and parallel, 316
principle, 314
units of, 315
Capacitance-type fuel quantity indicators,
314
Capacitive index, 321
Capacitive reactance, 31 7
Capsules see Pressure measurement
Card compass, 167 see also Direct
reading compasses
Cartesian co-ordinates, 237
Central air data computer, 106
Ceramic type metering unit see Metering
urrits
Characterized tank units, 325
Circular scale, 19
Coefficients of deviation see Deviation
coefficients
Coercive force, 189
409
Coercivity, 189
Cold-junction temperature compensation,
287-291, 350
Command bars, 359, 362, 366
Compass operating modes, 201
Compass swinging, 219
Compass system monitoring, 193
Compensated gauge system, 322
Compensating leads, 291
Compensation cams, 110
Compensation for dip, 169
Compensator tank unit, 322, 324
Component P, 207
Component Q, 208
Component R, 208
Compressibility error, 86
Conduction, 258
Control of drift, 124, 178
Control synchro see Synchro systems
Control transformer see Synchro systems
Controlling system, 275 , 278
Convection, 258
Coloured displays, 28
Coriolis error, 204, 397
Coupling element, 6
Course deviation indicator, 359, 361 , 365
Critical Mach number, 92
Cross-coil ratiometer, 276
Cylinder-head temperature indicator, 281
Data acquisition systems, 373, 389
D.C. torquer indicators, 112, 251
Dead beat indication, 275
Dead weight tester see Pressure measurement
Declination se~ Magnetic variation
De-coding, 387
Density altitude, 74
Desynn system
basic, 225, 349
micro, 227, 305
slab, 228, 342
Detecting element (instrument), 6
Deviation, 206
Devia tion coefficients, 216
Deviation compensation devices, 204, 219
Diaphragms see Pressure measurement
Dielectric constant, 316, 318
Differential synchro see Synchro systems
Digital coding, 82, 384
Digital display, 26
Directional gyroscope, l 77 "-183
Directional gyroscope elements, 196
Directional gyroscope monitoring, 193
Director displays, 30
Direct-reading compasses
acceleration errors, 172-176
aperiodic, 168
construction, 167, 170
dip compensation, 169
functions, 159
liquid damping, 168
liquid expansion compensation, 169, 171
magnet systems, 167
turning errors, 174 - 176
Direct-reading pressure gauge, 298, 304
410
Displacement gyroscope, 122
Displays
director, 30, 359
high-range long-scale, 2 I
radio altitude, 364
qualitative, I 9
quantitative, 19
Diurnal change, 166
Double-bar pointer, 196
Double-tangent mechanism, 10
Drag elements, 245
Drains, 61
Drift, 123, see also Control of drift
Dual-indicator display, 27
Duplicate instruments, 3, 41
Dynamic counter display, 27
Dynamic-scattering liquid crystal display, 35
'E' and 'I' bar pick-off, 76, l 07
Earth's atmosphere, 63
Earth gyroscope, 125
Earth's magnetic components, 167, 203, 210
Earth rate, 122, 125
Elastic pressure-sensing elements, 302
Electric gyro horizon, 130
Electrical tachometers, 242
Electrical temperature indicators
resistance, 260, 265
thermoelectric, 279-285
Electrical zero, 23 3
Electromagnetic compensator, 222 see also
Deviation compensation devices
Electromagnetic induction, 229
Electromagnetic recording, 382
'Empty' adjustments, 3 26
Encoding altimeter, 84
Encoding disc, 84
Engine pressure ratio system, 345
Engine speed indicators see Tachometers
Engine vibration monitoring system, 356
Erection cu t·ou t see Erection error
compensation
Erection devices, directional gy ~oscopes, 179
Erection errors, 144
Erection error compensation
erection cut-out, 139, 146
inclined spin axis, 145
pitch-bank erection, 146
Erection rate, 142
Erection systems, gyro horizon
ball type, 134
fast-erection systems, 139
pendulous vane unit, 134
torque ·motor and levelling switch, 136
Eureka spool, 292
Exhaust gas temperature indicators,
248,292
Exosphere, 62
Ex tension leads, 291
Ex ternal circuit resistance, 291
Farad, 315
Fast-erection systems
electromagnetic method, 140
fast-erection switch, 139
purpose, 139
Fatigue meter, 374. 377
Fiducial point, 6 7
Field-effect liquid crystal display, 35
Fixed points, 258
Flight data recorders, 373, 379-389
Flight director indicator. 359, 362
Flight director systems, 35 8- 3 71
Flight encoder panel, 386
Flight instruments, 39
Float-type fuel,quantity indicators, 313
Flow lines, 44
'Flush' bulb, 270
Flux detector elements, 185
Force-balance transducer, J 07
Fortin barometer, 67
Free gyro mode see Compass operating
modes
Free gyroscope, 116
Fuel llowmcters, 330
Fuel quantity measurement, 313
Fuel quantity totalizer, 327
Fuel quantity by weight, 321
Fuel trim indicator, 349
Fuel trimming, 349
'Full' adjustments, 326
Fundamental interval, 258 see also
Temperature measurement
Gauge pressure, 298
Gears, 12
.
Geographic poles, 164
Gimbal errors, 126, 180, 397
Gimbal lock, 125
Gimbal ring balancing, 179
Gimbal system, 116
Gimballing effects, 144
Glide slope, 31, 359, 360
'Go-around', 364,367
Gyro horizon
electric, 130
erection systems, 133
presentations, 127
principle, 12 7
vacuum-driven, 129
Gyroscope, 116
Gyroscope levelling, 199
Gyroscopic rigidity , 117, 118
Hairsprings, 13
Hard iron, 164 see also Hard iron
magnetism
Hard iron magnetism, 206 see also Air·
craft magnetic components
Heading 'bug', 198
Heading indicator, 196
Head-up displays, 32
Heat, 258
High-range long-scale display, 21
Horizontal situatio n indicato r, 359, 361
Hub-array thermocouples, 287
Hysteresis curve, 188
ICAO Annex, 10, 84
!CAO standard atmosphere, 65
Ice point, 258
!LS indicator, 30 see also Integrated
instrument systems
Immersion thermocouple, 281
lnclinr.d spin axis see Erection error
compensation
Indicated altitude, 71
Indicated /computed airspeed indicator, 95
Indi~ating element. 6
Inductor-type pressure gauge see
Ratiometer pressure gauges
Inferred-density system, 322
Input axis, 118
Instantaneous vertical speed indicator, 105
Instrument displays, 19
Instrument elements, 6
Instrument grouping
flight instruments, 39
power plant instruments, 41
Instrument illumination
bridge lights, 46
pillar lights, 46
wedge-type lighting, 48
Instrument landing system, 30, 358, 360
Instrument layou ts see Instrument
grouping
Instrument mechanism, 6
Instrument panels, 2, 38
Integrated director display , 31 see also
Flight director systems
Integrated fuel llowmetcr, 333
Integrated instrument systems, 358
Interrogatio n modes, 81
Invar, 15
Ionosphere, 62
lsoclinals, 166
Isodynamic lines, 16 7
lsogonal lines, 166
Kelvin scale see Temperature measurement
Kew barometer, 67
Knot, 86
Lapse rate , 64
Lateral deviation bar, 361,365
Latitude control, 203
Levelling switches, 136
Levelling systems see Gyroscope levelli· .;;
Lever angle, 8
Lever length , 8
Lever mechanism, 8
Light-emitting diode (LED), 36
Light-emitting displays, 34
Lines of nux , 159
Linear scale, 20
Linear voltage differential transformer, 345
Liquid crystal display (LCD) , 34
Liquid damping, 168
Liquid expansion compensation, 169
Liquid-level switch see Levelling switches
Localizer, 31,361,364,365,366
Load-limit control panel, 330
Location of flight recorders, 388
Location of tank units, 323
Location of thermocouples, 284
Logarithmic scale, 20, l 00
411
Long-reach thermocouples, 285
Lubber line, 168, 198
Mach/ Airspeed indicator, 94
Mach number, 91
Mach warning system, 98
Machmeter, 92
Magnet systems, 167
Magnetic compass see Direct-reading
compasses
Magnetic dip
definition, 166
effect on a compass, 169
compensation, 170
Magnetic equator, 166, 170
Magnetic field, l59, 272
Magnetic field strength, 161
Magnetic flux, 159
Magnetic flux density, 160
Magnetic foci, 164
Magnetic inclination see Magnetic dip
Magnetic indicators, 44
Magnetic intensity see Total force
'Magnetic lock', 331
Magnetic meridian, 164
Magnetic moment, 161
Magnetic poles, 159, 164
Magnetic screening, 151
Magnetic tape, 382
Magnetic variation, 165
Magnetization curve see B/H curve
Magnetizing force, 187
Magnification of mechanisms, 8
Mandatory recording, 373
Manganin spool, 292
Manifold pressure gauges, 339
Manometric system, 50
Master directional gyro, 194
Maximum safe airspeed indicator see
Mach/ Airspeed Indicator
Measuring element, 6
Mechanical tachometer, 242
Mechanisms
double-tangent, 10
lever, 8
rod, 10
sine, 10
skew-tangent, 12
tangent, 10
Melting point, 258
'Mental focus ' lines, 41
Mercury barometer, 65
Mercury switches see Levelling switches
Meter movements, 364, 367
Metering units, 102
Micro-adjuster, 219
Micro-Desynn see Desynn system
Microfarad, 315
Minimum decision altitude, 364
Modes of operation, 368
Mounting of instruments, 43
Moving-coil indicator, 273
Moving-tape display , 26
Mutual inductance, 230
Nematic material, 35
412
Newton per weber, 161
Nickel law, 265
Non-linear scale, 20 see also Square
law scale
Non-uniform magnetic fields, 280-283
Northerly turning error, 172, 176
Nozzle guide vane thermocouple, 284
Null point, 268
Null position, 233
Nutation,
Ohm's law
definition, 260
parallel circuit, 262
series circuit, 260
series-parallel circuit, 263
Operating range, 24
Output axis, 118
Oz.onosphe1e, 62
Parallax errors, 23
Parallel circuit see Ohm's Law
Parasitic e.m.f. 281
Peltier effect, 280
Pendulosity errors, 148
Pendulosity error compensation, 149
Percentage tachometer, 24 7
Percentage thrust indicator, 346
Period of a magnet, 163
Permalloy, 187
Permanent magnetism, 164
Permeability, 187
.Permittivity, 316
Phase quadrature, 138
Phonic wheel, 251
Picofarad , 315
Pie~o-electric elements, 108
Pillar lights, 46
Pitch-bank erection, 146
Pitch command facility, 367
Pitot probe heating circuits, 55
Pitot pressure, 52
Pi tot-static pipelines, 6 3
Pitot-static system, 2, 50
Pitot-static probe, 53
Pitot probe, 57
Platform scale, 23
Platinum law, 265
Polar co-ordinates, 23 7
Polar navigation, 202
Pole strength, 161
'Porous pot' see Metering units
Position error see Pressure error
Power indication instruments, 338
Powered moving-coil indicator, 278, 310
Power plant instruments, 41
Precession, 117, 118
Pressure altitude, 71, 7 4
Pressure error, 55 , 111
Pressure error correction, 58, 111
Pressure head see Pilot-static probe
Pressure/height relationship, 64, 69, 71, 74
Pressure measurement
bellows, 304, 306
bourdon tube, 302
capsules, 303, 309
Pressure measurement (cont'd)
dead-weight tester, 300
diaphragms, 303
methods, 298
remote-indicating gauges, 304
U-tube manometer, 298, 299
Pressure switches, 310
Pressure transducer, 107, 345
Pressure/weight balancing, 300
Pyrometry, 260
'Q' CODE, 15
Qualitative display, 19, 29
Quantitative display , 19
Radiation, 257
Radiation pyrometer system, 294
Radio magnetic indicator, 196, 358
Radius of gyration, 118
Ram air temperature, 269
Range markings, 28
Rapid-response thennocoup!e, 283
Rare-metal thermocouples, 280
Rate of climb indicator see Vertical speed
indicators
Rate generator, 365, 366
Rate gyroscope, 150
Rates of turn, 152
. Ratio meter pressure gauges, 306 see also
Desynn system and Ratiometer system
Ratiometer systems, 276
'Reading' head, 387
Real A, 217
Real drift, 124
Recovery factor, 269, 283
Refuelling control panels, 329
Reluctance, 163, 189
Remanance, 189
Remote-indicating compasses, 184, 361
Reply pulse coding, 83
Requirements for instruments, 1
Resistance bulb see Temperature sensing
elements
Resistance thermometry, 260, 265
Resistivity, 264
Resolver synchro see Synchro systems
Rod components see Aircraft magnetic
components
Rod mechanisms, 10
Rotation indicators, 265
Rotorace bearings, 200
Saturation point, 187
Scale base, 19
Scale length, 21
Scale marks, 20
Scale range, 24
Scissor magnets see Deviation compensation
devices
Sealing of instruments, 17
Sktor gear, 12
Secular change, 166
Seebeck effect, 279
Self-inductance, 230
Series circuit see Ohm's law
Series-parallel circuit see Ohm's law
Servo altimeter, 75
Servo-operated indicating systems
tachometer, 248
temperature, 289 , 292
Short-reach thermocouple, 285
Sine mechanism, 10
Single-axis gyroscope, 150
Single-bar pointer, 196
Skew-tangent mechanism, 12
Sl.ab-Desynn, 228, 342
Slaved gyro mode see Compass operating
modes
Slaving amplifier, 194
Slaving torque motor, 194
Soft iron, 164 see also Soft iron magnetism
Soft iron magnetism, 206 see also Aircraft
magnetic components
Space gyroscope see Free gyroscope
Spinning freedom, 116
'Spin-0own' brake, 196
Spring-rate, 304
Square-law compensation, 87, I 08
Square-law scale, 20, 87, see also Non·
linear scale
Squirrel-cage, 130, 137, 245
Stagnation point, 52
Stagnation thermocouple, 283
Standard atmosphere, 64, 72
Standard tum rates, 152
Standards, 4
Standby attitude indicator, 132
Static air temperature, 269
Static counter di~play, 27
Static tube, 52
Static vents, 57
Steam point, 258
Straight scale, 25
Stratopause, 62
Stratosphere, 62
Subsonic speed, 91
Supercharging, 339
Supersonic speed, 92
Surface-contact thermocouple, 281
Synchro systems
control, 79 , 108, 233, 343, 365, 366
control transformer, 195, 233
definition, 229
differential, 110, 235
resolver, 287, 366
synchrotel, 96, 240
torque, 110, 232
transolver, 79
Synchronizing indicators, 201
Synchroscope, 252
Synchrotel see Synchro systems
Tachometers
drag elements, 245, 246
electrical, 242
functions, 242
generators, 24 3
indicators, 244
mechanical, 242
percentage, 24 7
probe, 250
servo-operated, 248
413
Tangent mechanism, 10
Tank units
characterized, 325
compensator, 322, 324
float-type, 313
standard (capacitance), 324
'Tape remaining', 382
Temperature , 258
Temperature bulb see Temperature
sensing elements
Temperature coefficient, 265
Temperature compensation methods
bimetal strip, 15, 87
bi-metal
bracket, 69
thermo-magnetic shunt, 17, 246, 291
thermo-resistance, 16, 291
Temperature effects on fuel, 320
Temperature scales and conversions, 259
Temperature measurement, 257
Temperature/ resistance laws, 265, 402
Temperature and r.p.m. control system, 353
Temperature sensing elements, 268
Terrestial magnetism , 164-.167
Tesla, 160
ll1ermistor see Thermo-resistor
Thermocouplr.
combinations, 280
harness assemblies, 285
location, 284
materials, 280
principle, 279
recovery factor, 283
types, 281- 285
Thermomagnetic shunt, 17, 246, 291
Thermometer display see Moving-tape
display
Thermo-resistor , 16,291
Thomson effect, 280
Three-dimensional display, 362
'Tied' gyrsocope see Earth gyroscope
Tilting freedom, 116
'Top limiting', 354
Toppled gyro, 139
Toroidal resistance, 225, 226
Torquemeter, 341
Torque motors, 136, 194
Torque motor erection system, 136
Torque pressu re indicators, 341, 343
Torque synchro see Syncluo systems
Torricellian vacuum, 66
Total air temperature, 270
Total force, 166
Total pressure see Pilot pressure
·u·
414
'Trace' disc, 69
Trace recording, 380
Transfo rmer principle, 230
Transmission error, 204
Transolver, 79
Transonic range, 92
Transponder, 81
Transponder codes, 82
Transport wander, 124
Trimming resistance, 292
Tropopause, 62
Troposphere, 62
True airspeed computation, 111
True alt.itude, 74
True poles see Geographic poles
Tuning spring, 90
Turbine temperature control, 348-355
Turn-and-bank indicators, 149- 156
Turn co-ordinator, 156
Turning errors
compasses, 172
gyro horizons, 142, 144
Turns ratio, 231
Two-axis gyroscope, 122
Units of capacitance, 315
U-tube manometer, 298
Vacuum-driven gyro horizon, 129
Veeder-counter display see Digital display
Veering freedom, 116
Vertical-axis gyroscope, 124
Vertical speed indicators, 99
Vibration monitoring, 356
Vibration pick-up, 356
Viscosity compensator valve, 103
Visual approach monitor, 33
Vmo, 94
Volumetric gauge system, 32 1
Volumetric top-off, 328
V.O.R., 196, 362, 365
Warning flags, 367
Water traps see Drains
Weber, 160
Wedge-type ligh ting, 48 see also Instrument
illuminatio n
Wheatstone bridge network, 265
Wire recorders, 383
'Words' , 384
'Writing head'. 382, 384
Zero adjustments, 13, 105
Was this manual useful for you? yes no
Thank you for your participation!

* Your assessment is very important for improving the work of artificial intelligence, which forms the content of this project

Download PDF

advertising