Flight Instruments & Automatic Flight Control Systems PDF

GROUND STUDIES FOR PILOTS FLIGHT INSTRUMENTS & AUTOMATIC FLIGHT CONTROL SYSTEMS Related titles on the JAR syllabus Ground Studies for Pilots series Radio Aids R.B. Underdown and David Cockburn Sixth Edition 0-632-05573-1 Navigation R.B. Underdown and Tony Palmer Sixth Edition 0-632-05333-X Fligh...

GROUND STUDIES FOR PILOTS FLIGHT INSTRUMENTS & AUTOMATIC FLIGHT CONTROL SYSTEMS Related titles on the JAR syllabus Ground Studies for Pilots series Radio Aids R.B. Underdown and David Cockburn Sixth Edition 0-632-05573-1 Navigation R.B. Underdown and Tony Palmer Sixth Edition 0-632-05333-X Flight Planning P.J. Swatton Sixth Edition 0-632-05939-7 Flight Instruments and Automatic Flight Control Systems David Harris Sixth Edition 0-632-05951-6 Meteorology R.B. Underdown and John Standen 0-632-03751-2 Aviation Law for Pilots R.B. Underdown and Tony Palmer Tenth Edition 0-632-05335-6 Human Performance and Limitations in Aviation R.D. Campbell and M. Bagshaw Third Edition 0-632-04986-3 GROUND STUDIES FOR PILOTS FLIGHT INSTRUMENTS & AUTOMATIC FLIGHT CONTROL SYSTEMS Sixth Edition David Harris # 2004 by Blackwell Science Ltd First published 2004 a Blackwell Publishing company Library of Congress Cataloging-in-Publication Editorial offices: Data Blackwell Science Ltd, 9600 Garsington Road, is available Oxford OX4 2DQ, UK Tel: +44(0) 1865 776868 ISBN 0-632-05951-6 Iowa State Press, a Blackwell Publishing Company, 2121 State Avenue, Ames, Iowa 50014-8300, USA Set in 10/13pt Palatino Tel: +1 515 292 0140 by DP Photosetting, Aylesbury, Bucks Blackwell Science Asia Pty Ltd, 550 Swanston Printed and bound in Great Britain using acid-free Street, Carlton, Victoria 3053, Australia paper by T J International, Padstow, Cornwall Tel: +61 (0)3 8359 1011 For further information on Blackwell Publishing, The right of the Author to be identified as the visit our website: Author of this Work has been asserted in ww.blackwellpublishing.com accordance with the Copyright, Designs and Patents Act 1988. All right reserved. No part of this publication may be reproduced, stored in a retrieval system, or transmitted, in any form or by any means, electronic, mechanical, photocopying, recording or otherwise, except as permitted by the UK Copyright, Designs and Patents Act 1988, without the prior permission of the publisher. Contents Preface vii List of Abbreviations ix 1 Air Data Instruments 1 2 Gyroscopic Instruments and Compasses 32 3 Inertial Navigation Systems 74 4 Electronic Instrumentation 94 5 Automatic Flight Control 116 6 In-Flight Protection Systems 149 7 Powerplant and System Monitoring Instruments 181 Answers to Sample Questions 214 Index 218 Preface Since the publication of Roy Underdown's Fifth Edition of Volume 3 of Ground Studies for Pilots, the format of the JAR-FCL examinations has become established. Among other things, it has become clear that students need the materials assembled in such a way as to facilitate study for individual papers and that it is no longer practical to combine Instruments with Navigation General, for example. The scope of what was once called the Instruments paper (now paper number 022 in the ATPL syllabus) is now so broad that it justifies a volume in its own right. Whilst still covering the air data and gyroscopic instruments, inertial navigation systems and electronic naviga- tion instruments, it has now been extended to cover the additional syllabus requirements of engine and systems monitoring instruments and flight warning systems. As more and more students find the need to study subjects in the groupings as they appear in the examinations, it makes sense to group those subjects accordingly in the Ground Studies for Pilots series of books. This new Sixth Edition addresses all the subjects listed in the JAR Learning Objectives for the Instruments and Automatic Flight paper number 022. I have tried to make the book both readable and instructive, with the intention that it should be useful to those seeking general information as well as to the examination student. Whilst it is aimed principally at pilots studying for the JAR ATPL ground examina- tions, it will also be helpful to pilots at all professional and private levels. Just a few years ago, automatic flight systems and electronic instrument systems were almost exclusively the preserve of large passenger transport aircraft. In more recent times, however, it has become increasingly common for smaller short-haul and executive-type jet and turbo-prop aircraft to be equipped with such systems. Consequently, there is a need for those pilots intending to progress from light and general aviation into commercial flying to have knowledge of at least the basic principles of these systems and instruments, even though they may not be immediately intending to sit the professional examinations. The text and diagrams in this volume have been deliberately designed to be understandable without preknowledge of the subjects. viii Preface Acknowledgement The assistance of Roger Henshaw, Peter Swatton and David Webb, all of Ground Training Services Ltd at Bournemouth International Airport, is gratefully acknowledged for their advice and for scrutinising the manuscript of this book for errors. Their professional knowledge of the JAR-FCL examination requirements has helped to ensure that it will be extremely useful to pilots undertaking ground studies. David Harris Minehead List of Abbreviations a.c. alternating current ACARS aircraft communications addressing and reporting system ACAS airborne collision avoidance system ADC air data computer ADI attitude director indicator agl above ground level AIDS aircraft integrated data system amsl above mean sea level AOM Aircraft Operating Manual ASI airspeed indicator ASIR airspeed indicator reading BITE built-in test equipment CADC central air data computer CAS calibrated airspeed CDU control and display unit CEC compressibility error correction CG centre of gravity CHT cylinder head temperature COAT corrected outside air temperature CRT cathode ray tube CSDU constant speed drive unit CWS control wheel steering d.c. direct current DEC density error correction DG directional gyro DH decision height EADI electronic attitude and direction indicator EAS equivalent airspeed ECAM Engine Centralised Aircraft Monitoring EFIS Electronic Flight Instrument System EGT exhaust gas temperature EHSI electronic horizontal situation indicator EICAS Engine Indicating and Crew Alerting System EPR engine pressure ratio x List of Abbreviations FFRATS Full Flight Regime Autothrottle System FL Flight Levels FMC flight management computer FMS flight management system FOG fibre optic gyro GPWS Ground Proximity Warning System HDG heading HSI horizontal situation indicator hPa hectopascal IAS indicated airspeed ICAO International Civil Aviation Organisation IE instrument error IEC instrument error correction in Hg inches of mercury INS Inertial Navigation System IRS Inertial Reference System ISA International Standard Atmosphere IVSI instantaneous vertical speed indicator JAA Joint Aviation Authority JAR Joint Aviation Regulations K degrees kelvin kg kilogram(s) lb pound(s) LCD liquid crystal diode LED light emitting diode LNAV lateral navigation LSS local speed of sound LVDT linear voltage displacement transmitter M mach number MAP manifold air pressure MCDU multi-purpose control and display unit Mcrit critical mach number msl mean sea level nm nautical miles OAT outside air temperature P pitot pressure PE position error PEC pressure error correction Q dynamic pressure QFE height above a chosen ground datum QNH height above mean sea level RA resolution advisory RAS rectified airspeed  List of Abbreviations xi RCDI rate of climb/descent indicator RLG ring laser gyro RMI radio magnetic indicator rpm revolutions per minute S static pressure SAT static air temperature SBY standby SSR secondary surveillance radar TA traffic advisory TAS true airspeed TAT total air temperature TAWS Terrain Avoidance Warning System TCAS Traffic Collision Avoidance System TCS touch control steering TOD top-of-descent V airspeed VNAV vertical navigation VOR VHF omnidirectional ranging VSI vertical speed indicator Chapter 1 Air Data Instruments Two of the most important pieces of information for a safe flight are height and speed. Almost from the beginning of powered flight these have been provided to the pilot by instruments that utilise the ambient atmospheric pressure by means of a pitot/static system. Pitot and static systems Static pressure The ambient atmospheric pressure at any location is known as the static pressure. This pressure, in a standard atmosphere, decreases by 1 hecto- pascal (hPa) for each 27 feet (ft) increase in altitude at mean sea level. For simplicity this figure is usually approximated to 1 hPa per 30 ft gain in altitude. The rate of change of pressure with height is fundamental to the operation of the pressure altimeter, the vertical speed indicator and the mach meter. Each of these instruments uses static pressure to measure aircraft altitude, or rate of change of altitude. Static pressure, that is the pressure of the stationary air surrounding an aircraft, irrespective of its height or speed, is sensed through a set of small holes situated at a point on the aircraft unaffected by turbulence. This sen- sing point is known as the static source. It is typically on the side of the fuselage or on the side of a tube projecting into the airstream. Pitot pressure As an aircraft moves through the air it displaces the surrounding air. As it moves forward it compresses the air and there is a pressure increase on the forward-facing parts of the aircraft. This pressure is known as dynamic pressure. Suppose a cup were to be placed on the front of an aircraft, with its open end facing forward. When the aircraft is stationary the pressure inside the cup will be the same as the surrounding air pressure. In other words it will be static pressure. When the aircraft begins to move forward the air inside the cup will be compressed and dynamic pressure will be added to the static 2 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems pressure. The faster the aircraft moves, the greater the dynamic pressure will become, but static pressure will always also be present. The pressure measured on the forward-facing surfaces of an aircraft will be the sum of static pressure and dynamic pressure. This is known as pitot pressure, or total pressure, which is sensed by a forward-facing, open-ended tube called a pitot tube, or pitot head. Figure 1.1 is a simplified diagram of a pitot head. pitot connection electrical heater connection Figure 1.1 Pitot head. The pitot head comprises an aerodynamically shaped casing, usually mounted beneath one wing, or on the side of the forward fuselage, clear of any turbulent airflow. Within the casing is a tube, the rear of which is con- nected to the pitot system, which conveys pitot pressure to the pilot's instruments. An electrical heating element is fitted within the tube to pre- vent the formation of ice, which could otherwise block the tube and render it useless. Drain holes are provided in the bottom of the tube to allow water to escape. The dynamic element of pitot pressure is required to operate those air data instruments that display speed relative to the surrounding air, the airspeed indicator and the mach meter. Pitot and static pressure is supplied to the air data instruments through a system of tubes known as the pitot/static system. A schematic layout of the pitot/static system for a light aircraft is shown in Figure 1.2. The static source is duplicated on either side of the rear fuselage. This is to compensate for false readings that would occur if the aircraft were side-slipping or in a crosswind. The pitot and static pressure supplies are connected to a dupli- cate set of instruments in aircraft that have a pilot and co-pilot. In many aircraft the pitot and static sources are combined in the pitot head, as illustrated in Figure 1.3. The static source consists of a number of small holes in the side of the pitot head, connected to an annular chamber surrounding the pitot tube. This chamber is connected to the static system, which conveys static pressure to the pilot's instruments. A separate pipe connects the pitot tube to the pitot  Air Data Instruments 3 pitot system alternate static system ASI VSI static pitot ALT source head changeover cock static source Figure 1.2 Pitot/static system for a light aircraft. static source static pressure pitot pressure heater element drain holes direction of flight Figure 1.3 Pitot/static head. system. As with the pitot head shown in Figure 1.1, an electrical heating element is fitted to prevent blockage of the pitot and static sources due to icing and water drain holes are provided in the bottom of the casing. In some aircraft this type of pitot head is mounted on the fuselage near the nose and it may be duplicated, one each side, to compensate for crosswind effects. Static pressure error The static system of air pressure measurement will be incorrect if the airflow is turbulent, if there is a crosswind, or if the aircraft is side-slipping. The effect of turbulence is minimised by locating the static source clear of pro- tuberances and disturbed airflow. To eliminate the effect of crosswind or side-slip the static source is duplicated and this is known as static balancing. Despite all the measures taken by the aircraft designer there is often some small error in the sensed static pressure, but this can usually be measured and compensated for by a correction card or table. In many aircraft the correction values will differ according to the position of flaps and/or landing gear. 4 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems Malfunctions. Correct and reliable indications from the various air data instruments can only be achieved if the pitot and static sources are kept clear of any blockages and the pitot/static system within the aircraft remains unda- maged and pressure-tight.. Blockage of the pitot or static source may occur due to icing, insects, dirt or dust. It is also not unknown for aircraft painters to forget to remove masking tape from the perforated discs that form the static vents on the fuselage sides. Icing can be prevented by the use of heaters, but this may affect the sensed pressure to some small extent. The effect of blockages is to render the instruments dangerously inaccurate or useless.. Blockage of the static sources will cause the altimeter reading to remain constant regardless of changing aircraft altitude, the vertical speed indi- cator will not indicate rate of change of height and the airspeed indicator will be dangerously inaccurate.. Blockage of the pitot source will not affect the altimeter or vertical speed indicator, but it will render the airspeed indicator useless and the mach meter grossly inaccurate.. Leakage in the piping of the pitot/static system will also seriously affect the accuracy and usefulness of the air data instruments. Loss of pitot pressure due to leakage in the pitot pressure system will cause the air- speed indicator to underread.. Leakage in the static pressure system within the cabin of a pressurised aircraft is a serious problem, since the altimeter will register an altitude equivalent to cabin altitude, which will almost certainly be much lower than aircraft altitude. The vertical speed indicator will not function at all and the airspeed indicator will be inaccurate.. In unpressurised aircraft the effect of a static system leak is less serious, since internal pressure is much the same as external. However, it may change at a slightly slower rate when the aircraft is climbing or des- cending and this would clearly affect the accuracy of the pressure instruments during height changes. Alternate static source Blockage of the static source is a more probable hazard in flight and for this reason many aircraft are fitted with an alternate static source. This may take pressure from within the cabin in the case of unpressurised aircraft, or from a separate external source. In either case there is likely to be a slight differ- ence in pressure compared with that from the normal source. The aircraft flight manual usually contains correction values to be used when the alter-  Air Data Instruments 5 nate static source has been selected. Changeover is made by means of a selector cock easily accessible to the pilot. In light aircraft, not fitted with an alternate static source, if the static source is blocked an alternative source can be obtained by breaking the glass of the vertical speed indicator (VSI). The pressure altimeter The function of the pressure altimeter is to indicate the aircraft height above a given pressure datum. It operates on the principle of decreasing atmo- spheric pressure with increasing height and is, in fact, simply an aneroid barometer that is calibrated to read pressure in terms of height. To do this, the manufacturer assumes that air pressure changes at a given rate with change of height. The International Standard Atmosphere (ISA) values are the data used for this assumption. In the ISA the temperature at mean sea level is +158C and the air pressure is 1013.25 hectopascals (hPa). The temperature lapse rate (the rate at which the temperature will decrease with increase of height) is 1.988C per 1000 ft (6.58C per kilometre) up to a height of 36 090 ft. Above that height the tem- perature is assumed to remain constant at 756.58C up to a height of 65 600 ft. Air is a fluid and it has mass, and therefore density. If we consider a column of air, its mass exerts pressure at the base of the column; the taller the column the greater the pressure exerted at the base. At any given height in the column the pressure exerted is proportional to the mass of air above that point and is known as hydrostatic pressure. Atmospheric pressure is assumed to decrease at a rate of 1 hPa per 27 ft gain in height at sea level, this rate decreases as height increases, so that at a height of say 5500 metres (18 000 ft) the same 1 hPa change is equivalent to approximately 15 metres (50 ft) change in height. The pressure altimeter is calibrated to read height above a selected pres- sure datum for any specific atmospheric pressure. The element of the pressure altimeter that measures atmospheric pressure changes is a sealed capsule made from thin metal sheet. The capsule is partially evacuated, so that the surrounding atmospheric pressure tends to compress the capsule. However, a leaf spring attached to the capsule pre- vents this. The capsule may be of the diaphragm or the bellows type, as illustrated in Figure 1.4. The simple altimeter The capsule is mounted in a sealed casing, connected to the static source. Increased static pressure will cause the capsule to be compressed against the restraining force of the leaf spring, decreased static pressure will allow the 6 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems bellows type diaphragm type Figure 1.4 Aneroid capsule types. leaf spring to expand the capsule. The spring force ensures that the extent of compression or expansion is proportional to the static pressure being mea- sured. This compression or expansion of the capsule is converted into rotary motion of a pointer against a calibrated scale by a system of linkages and gears. A schematic diagram of a simple pressure altimeter is shown in Figure 1.5. Expansion of the aneroid capsule will cause a lever to pivot about its attachment to the instrument casing. This lever is connected to a drum by means of a chain and its pivoting motion causes the drum to rotate. The drum is attached to a pointer, which will consequently rotate against a calibrated card scale. from static source leaf pointer spring glass aneroid face capsule card scale pointer setting knob Figure 1.5 Simple pressure altimeter.  Air Data Instruments 7 The setting of the height-indicating pointer can be adjusted by means of the pointer setting knob. The purpose of this is to allow the pilot to set the altimeter so that it displays height above a chosen datum. For example, if the pointer is set to read zero with the aircraft on the airfield the altimeter will thereafter indicate height above the airfield. Alternatively if, when the air- craft is on the airfield, the pointer is set to read height above mean sea level it will thereafter show height above mean sea level. In both these examples the surface pressure is assumed to remain constant. The pressure altimeter is calibrated to read height in feet above the ISA mean sea level pressure of 1013.25 hPa. Some altimeters of US manufacture use inches of mercury (in Hg) as the unit of calibration and the equivalent of 1013.25 hPa is 29.92 in Hg. To convert inches Hg to hPa it is necessary to multiply by 33.8639. Conversely, to convert hPa to inches Hg it is necessary to multiply by 0.02953. Some simple altimeters are calibrated on the assumption that static pressure reduces with increasing height at a constant rate of 1 hPa per 30 ft. Whilst this is a reasonable approximation up to about 10 000 ft, the height increase represented by 1 hPa decrease in pressure becomes progressively greater above this altitude. This is indicated in Table 1.1. Table 1.1 Rate of pressure change with altitude. Height in ISA (ft) Static pressure (hPa) Height change (ft) represented by 1 hPa change in static pressure Mean sea level 1013.25 27 10 000 697 34 20 000 466 47 40 000 186 96 60 000 72 263 In order to have a linear scale on the altimeter display it is necessary to compensate for this non-linear distribution of atmospheric pressure. This is achieved during manufacture by careful design of the capsules and the mechanical linkages. The simple altimeter is rarely used in aircraft because it lacks sensitivity, furthermore the altitude scale is compressed and causes confusion. More accurate height measurement is achieved with the sensitive altimeter. The sensitive altimeter The principle of operation of the sensitive altimeter is essentially the same as 8 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems that of the simple altimeter. Sensitivity of response to static pressure changes is improved by incorporating two or three aneroid capsules connected in series with each other. In a typical instrument the movement of the capsules as altitude changes is transmitted to gearing via a rocking shaft. The gearing in turn operates the height-indicating pointers, of which there are two or three. Clearly, as an aircraft climbs the air temperature will fall. The altimeter is calibrated to be accurate when operating at a sea level air temperature of +158C (ISA mean sea level temperature) and without compensation it will become increasingly inaccurate as instrument temperature deviates from that datum value. In the sensitive altimeter, temperature compensation is effected by a bimetallic strip inserted between the aneroid capsules and the transmission shaft. Note that this temperature compensation has nothing to do with altitude temperature error, which is discussed in sub-paragraph (6) under Altimeter errors below. Figure 1.6 illustrates the display of a sensitive altimeter with three pointers. It is perhaps easiest to think of the display as being similar to a clock face, in which one revolution of the long hand takes one hour and one revolution of the shorter hand takes 12 hours. In the altimeter display one revolution of the longest pointer represents a height change of 1000 ft, thus each num- bered division on the instrument scale represents a height increment of 100 ft when read against this pointer. One revolution of the medium length pointer represents a height change of 10 000 ft, thus each numbered division on the instrument scale represents a height increment of 1000 ft when read against this pointer. The height scale is divided into ten equal parts, so for every complete revolution of the 1000-ft pointer the 10 000-ft pointer moves one-tenth of a revolution. 9 0 1 8 2 7 3 6 4 5 Figure 1.6 Sensitive altimeter display.  Air Data Instruments 9 The shortest pointer shows increments of 10 000 ft for each numbered division on the instrument scale. The sensitive altimeter display in Figure 1.6 is therefore indicating a pressure altitude of 12 500 ft. Whilst the sensitive altimeter is usable to greater altitudes than the simple altimeter it becomes inaccurate at high altitudes where the pressure change becomes small for a given increase in height. Its greatest disadvantage, however, is its multi-pointer display, which is very open to misinterpretation. Subscale settings Sensitive altimeters incorporate a pressure datum adjustment, so that height above any desired datum pressure will be indicated. The selected datum pressure is indicated on a subscale calibrated in hPa or in Hg, and this is usually referred to as the subscale setting. QFE If airfield datum pressure is set on the subscale the altimeter will indicate height above the airfield once the aircraft is airborne and zero when on the ground at that airfield. This setting is assigned the ICAO code of QFE for communication purposes. QNH If mean sea level pressure is set on the subscale the altimeter will indicate height above mean sea level. Thus, when the aircraft is on the airfield the altimeter will indicate airfield elevation above mean sea level. This setting is assigned the ICAO code of QNH for communication purposes. Alternatively, the subscale may be set to 1013 hPa (29.92 in Hg), in which case the altimeter will indicate pressure altitude, that is the altitude above this pressure datum. This setting is used for aircraft flying Flight Levels (FL). At this point it is appropriate to define the terms height and altitude with reference to altimeter subscale settings. Height Height is the vertical distance above a specified datum with known eleva- tion, such as an airfield. Hence, if QFE is set on an altimeter it will indicate height above that airfield. Altitude Altitude is the vertical distance above mean sea level and it is therefore altitude that is indicated by an altimeter with QNH set. Pressure altitude, as we have already seen, is the altitude indicated on an altimeter with 1013 hPa, or 29.92 in Hg, set on its subscale. 10 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems Density altitude is the height in ISA conditions at which a given air density will occur and it is temperature dependent. If the local air temperature is ISA temperature, then density altitude will be the same as pressure altitude. If local air temperature is higher than ISA, then density altitude will be higher than pressure altitude and vice versa. True altitude is the exact vertical distance above any point on the surface. A pressure altimeter cannot be relied upon to show true altitude, even with QFE set, since its calibration assumes a rate of pressure and temperature change with height that may not exist in the prevailing atmospheric conditions. The servo-assisted altimeter Whereas in the simple and sensitive altimeter the movement of the aneroid capsule directly moves the altimeter pointer, in the servo-assisted altimeter the capsule movement is transmitted to the indicator by an electro- mechanical system. When the capsules expand or contract their movement causes an electro-magnetic transducer to transmit an electrical signal to a motor, which drives the altimeter's single pointer and its height counters. The servo-assisted altimeter has several advantages. Since the only mechanical transmission is from the capsules to the transducer there is considerably less resistance to motion and therefore less lag (see Lag error below). It is much more sensitive to small pressure changes and therefore remains accurate to greater altitude, where the pressure change is less for a given change of height. Because its output is electrical it can readily be used in conjunction with remote displays, altitude alerting and height encoding systems. The servo-assisted altimeter display with its height counters, as shown in Figure 1.7, is less likely to be misinterpreted than that of the sen- sitive altimeter. 9 0 1 8 2 pointer completes 12 5 00 1 revolution per 1000 feet 7 3 6 4 5 Figure 1.7 Servo-assisted altimeter display.  Air Data Instruments 11 Altimeter errors Pressure altimeters are prone to a number of errors, as listed below. (1) Lag error. There is inevitably a delay between an atmospheric pressure change and the response of the aneroid capsules, with the result that the movement of the instrument pointer lags behind a change in altitude. The magnitude of the error depends upon the rate of change of altitude and is clearly unacceptable. In sensitive altimeters it is often reduced by means of a vibrator, or knocking, mechanism, which has the same effect as tapping a barometer to make the pointer move after a small pressure change. In servo-assisted altimeters lag error is virtually eliminated by the reduction of mechanical resistance. (2) Blockage of static source. If the static source becomes blocked the altimeter will cease to indicate changes of static pressure, and therefore of altitude. The effect will be that the altimeter will continue to indicate the reading at which the blockage occurred. (3) Instrument error (IE). As with any mechanical device, the pressure altimeter is manufactured with small tolerances in its moving parts and these give rise to small inaccuracies in its performance. They are usually insignificant, but in some cases a correction table may be supplied with the instrument. (4) Position error (PE). The static source is positioned at a point on the airframe where disturbance to the airflow is minimal, so that the static pressure measured is as close as possible to the undisturbed ambient static pressure. However, there is usually some small error due to the positioning of the static source. Use of the alternate static source may also cause pressure error, since its siting is different to that of the normal source and thus subject to different effects during manoeuv- ring. The PE for an aircraft type is determined during its initial flight tests and is supplied in tabular or graphical format so that the pilot may make the appropriate corrections. In servo-assisted altimeters the cor- rection is usually incorporated into the transducer system. Corrections to be applied to the altimeter reading for position error may be listed in the Aircraft Operating Manual (AOM). These are usually small and take account of the effects of aircraft speed, weight, attitude and configuration, since all of these factors have some effect on airflow over the static source. Corrections are also usually given for both the normal and alternate static source, since position error may not be the same for both. The correction is applied as indicated in the AOM tables. For example, suppose the tables state that the correction to be applied with flaps at the landing setting of 458 is +25 ft, then 25 ft must be added to the altimeter reading. 12 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems (5) Pressure error. Also known as barometric error and subscale-setting error, this occurs when the actual datum pressure differs from that selected on the subscale setting of the altimeter. Suppose, for example, the subscale has been set for a regional QNH of 1020 hPa, but the air- craft is now operating in an area where the mean sea level (msl) pres- sure is 1000 hPa. Let us assume that the aircraft is at an altitude where the ambient pressure is 920 hPa, so the altimeter will indicate 3000 feet: (1020 7 920 = 100 hPa 6 30 ft/hPa = 3000 ft). However, because of the lower msl pressure the actual height of the aircraft above msl is: 1000 7 920 = 80 hPa 6 30 ft/hPa = 2400 feet. (6) Temperature error. The altimeter is calibrated to assume a Standard Atmosphere temperature lapse rate of 1.988C per 1000 ft. Actual tem- peratures at any given altitude usually differ from this assumption and the altimeter will be in error. In cold air the density is greater than in warm air and a given pressure will occur at a lower altitude than in warmer air. Consequently, the altimeter will overread and, with QNH set, the aircraft will be lower than indicated, with the error increasing from zero at msl to a significant amount at altitude. The AOM contains a table of corrections for this situation. These corrections must be added to published or calculated heights/altitudes when temperature is less than ISA. A `rule of thumb' calculation of altitude temperature error is that the error will be approximately 4% of indicated altitude for every 108C temperature deviation from ISA. Altimeter tolerances For altimeters with a test range of 0 to 9000 metres (0 to 30 000 ft) the required tolerance is + 20 metres or + 60 ft. For altimeters with a test range of 0 to 15 000 metres (0 to 50 000 ft) the required tolerance is + 25 metres or + 80 ft. The test is required to be carried out by the flight crew prior to flight, with QNH or QFE set and the altimeter vibrated either manually or by the mechanical vibration mechanism. The airspeed indicator (ASI) It is essential for the pilot of an aircraft to know its airspeed, because many critical factors depend upon the speed of flight. For example, the pilot needs  Air Data Instruments 13 to know when the aircraft is moving fast enough for take-off, when it is flying close to the stalling speed, when it has accelerated to the speed at which landing gear and flaps must be raised and when it is approaching the maximum safe flying speed, to name but a few. This critical information is provided by the airspeed indicator (ASI). The aircraft's speed relative to the surrounding air is proportional to the dynamic pressure that results from the air being brought to rest on forward facing parts of the airframe. The airspeed indicator measures dynamic pressure and converts this to an indication of airspeed. We have already seen that pitot pressure (P), as measured in the pitot tube, is a combination of dynamic pressure (D) and static pressure (S), i.e. P = D + S. It therefore follows that dynamic pressure is pitot pressure less static pressure, i.e. D = P 7 S. Thus, the function of the ASI is to remove the static pressure element of pitot pressure and use the resulting dynamic pressure to move a pointer around a graduated scale. The instrument comprises a sealed case connected to the static source and containing a capsule supplied with pitot pressure. Hence, the static pressure element of pitot pressure inside the capsule is balanced by static pressure surrounding the capsule. Consequently, the capsule will respond only to changes in the dynamic pressure element of pitot pressure. The faster the aircraft flies through the atmosphere, the greater will be the resultant dynamic pressure and the capsule will expand. This expansion is trans- mitted to a pointer by means of gearing and linkages. The principle is illustrated schematically in Figure 1.8. A simple ASI typically uses a single pointer that moves around a scale calibrated in knots. More complex instruments may be used in high-speed aircraft, incorporating an angle-of-attack indicator or a mach meter. pointer static pinion pressure quadrant pitot pressure capsule Figure 1.8 Principle of operation of the airspeed indicator. 14 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems Clearly the airspeed indicator must be calibrated to some fixed datum, for the pilot must be confident that an indication of, say, 90 knots is aero- dynamically the same in all conditions. The datum used is the air density when ISA mean sea level conditions prevail, that is the density when the air temperature is +158C and the air pressure is 1013 hPa. Only when these conditions prevail will the ASI indicate the true airspeed. Any point where air is brought to rest, such as in the pitot tube, is known as a stagnation point. At this point the kinetic energy of the air is converted to pressure energy. The pressure resulting is, of course, dynamic pressure and is denoted by the symbol Q. It can be shown mathematically that Q = 2 2rV , where r is the air density and V is the airspeed. 1 As the aircraft climbs the air density decreases and the airspeed indicator reading will be lower than the true speed of the aircraft through the sur- rounding atmosphere. A simplistic way of thinking of this is to consider that as the aircraft moves through the air it is colliding with the air molecules; the faster it flies the more molecules it strikes in a given time period. The ASI indication of airspeed is based on the dynamic pressure measured, assuming an ISA mean sea level value of air density. Therefore, when the air density is lower, as with an increase of altitude, a given dynamic pressure (and therefore indicated airspeed) will only be achieved at a higher true airspeed. Let us assume that it is flying in ISA msl conditions and is colliding with air molecules at a rate that causes the ASI to indicate 90 knots flight speed. When the aircraft climbs to a greater altitude the air density is less and so the molecules are further apart. In order for the ASI to continue to read 90 knots the aircraft must fly faster relative to the surrounding air in order to strike the same number of molecules within a given time period. Thus the true air speed, which is the speed of the aircraft relative to the surrounding atmo- sphere, increases. Airspeed may be quoted in a number of ways and these are listed below: ASIR: Airspeed indicator reading. IAS: Indicated airspeed (IAS) is the speed indicated by the simple airspeed indicator reading (ASIR) corrected for errors due to manufacturing toler- ances in the instrument (instrument error), but not corrected for static pressure errors occurring at the static source (pressure error). CAS: Calibrated airspeed (CAS) is the airspeed obtained when the correc- tions for instrument error (IEC) and pressure error (PEC) have been applied to IAS. These correction values are usually found in the Aircraft Operating Manual and may be reproduced on a reference card kept in the cockpit. CAS = IAS + PEC Calibrated airspeed used to be referred to as rectified airspeed (RAS).  Air Data Instruments 15 EAS: Because air is a compressible fluid it tends to become compressed as it is brought to rest. At low to medium airspeeds the effect of compression is negligible, but above about 300 knots TAS the compression in the pitot tube is sufficient to cause the airspeed indicator to overread significantly. The greater the airspeed above this threshold, the greater the error due to com- pression. The effect is also increasingly noticeable at high altitude, where the lower density air is more easily compressed. The error produced by this effect is known as compressibility error. Compressibility error correction (CEC) can be calculated and when applied to the calibrated airspeed the result is known as equivalent airspeed (EAS). EAS = CAS + CEC. TAS: The airspeed indicator only indicates true airspeed (TAS) when ISA mean sea level conditions prevail; any change of air density from those conditions will cause the indicated airspeed to differ from true airspeed. The greater the altitude, the lower will be air density and therefore IAS (and consequently EAS) will be progressively lower than TAS. The error due to the difference in density from ISA msl density can be calculated. When this density error correction (DEC) is applied to EAS the result is the aircraft's true airspeed (TAS). TAS = EAS + DEC. The com- pressibility error and density error corrections are embodied in the circular slide rule (navigation computer), from which TAS can be found using CAS and the appropriate altitude, speed and temperature settings. Square law compensation Given that dynamic pressure increases as the square of airspeed it follows that expansion of the capsule in the ASI must do likewise. Thus, at low airspeeds the amount of expansion for a given increase in speed will be small, whereas at higher airspeeds the amount of expansion will be rela- tively large for the same speed increase. This is known as square law expansion and, transmitted to the instrument pointer, will require an expanding scale on the face of the ASI, as illustrated at Figure 1.9. Such a scale makes accurate interpretation of airspeed difficult at low speeds and limits the upper speed range that can be displayed. Most ASIs incorporate an internal compensation device that permits a linear scale on the dial, i.e. equal spacing of the speed divisions over the whole range, as depicted in Figure 1.10. The effects of temperature variations on the sensitivity of the capsules and their associated linkages are typically compensated by the inclusion of bimetallic strips in the lever system. The expansion or contraction of these strips varies the degree of lever movement in the system of linkages. 16 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems 150 200 100 50 250 AIRSPEED knots 0 Figure 1.9 Square law calibration of ASI dial. 200 40 AIRSPEED 60 180 KNOTS 80 160 140 100 120 Figure 1.10 Typical ASI presentation. The ASI scale is often marked with coloured arcs and radial lines. The arcs indicate operating speed ranges and the radial lines indicate limiting speeds. These speed ranges and limiting speeds are as follows: Vne This is the `never exceed' speed, beyond which structural damage may occur. It is indicated on the ASI presentation by a red radial line. Vno This is the maximum speed under normal operating conditions with the aircraft `clean' (i.e. flaps and landing gear retracted). Vs1 This is the stalling speed with the aircraft `clean'. A green arc extends around the ASI scale from Vs1 to Vno. Vso This is the stalling speed with flaps and landing gear fully extended. Vfe This is the maximum permitted speed with flaps extended. A white arc extends around the ASI scale from Vso to Vfe.  Air Data Instruments 17 Vmc This is the minimum control speed with one engine inoperative on a multi-engine aircraft. It is indicated by a red radial line on the ASI scale. Vyse This is the best rate-of-climb speed with one engine inoperative on a multi-engine aircraft. It is indicated by a blue radial line on the ASI scale. Vmo/Mmo Some ASIs incorporate a red and white striped pointer showing the maximum allowable airspeed (Vmo) which is often the mach-limiting airspeed, at which the airflow over parts of the airframe will be approaching the critical mach number (Mcrit) for the aircraft. The pointer is actuated by a static pressure capsule and specially calibrated mechanism. This allows the Vmo pointer reading to increase progressively up to about 25 000 ft, as the air becomes less dense. Above this altitude the Vmo pointer reading is pro- gressively decreased, since Mcrit will be reached at progressively lower values of indicated airspeed. Vlo This is the maximum speed at which the landing gear may be safely extended or retracted. Vle This is the maximum speed at which the aircraft may be flown with landing gear extended. Note: Mach number and critical mach number will be explained in detail in the section dealing with the mach meter. ASI errors (1) Instrument error. This is effectively the same as for the pressure alti- meter. It is the error between the airspeed that the ASI should indicate and that which it actually does, due to manufacturing tolerances and friction within the instrument. (2) Position or pressure error. This is the error caused by pressure fluc- tuations at the static source. These may be due to the position of the static source, hence the term position error. The error can be determined over the speed range of the aircraft and recorded on a correction card. Pressure error may also occur when the aircraft is at an unusual attitude or when flaps or landing gear are extended, in which case it is known as manoeuvring error. An example of ASI values with a normal static source in use, and the corrections to be applied when the alternate source is in use, is shown in Table 1.2 below. These data are normally found in the AOM. (3) Compressibility error. At true airspeeds above about 300 knots the air brought to rest in the pitot tube is compressed to a pressure greater than dynamic pressure, causing the ASI to overread. The effect increases with altitude, since less dense air is more readily compressed. As pre- 18 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems Table 1.2 Airspeed calibration card. Condition Indicated airspeed (knots) Flaps up Normal 40 50 60 70 80 90 100 110 120 130 140 Alternate 39 51 61 71 82 91 101 111 121 131 141 Flaps 108 Normal 40 50 60 70 80 90 100 110 Alternate 40 51 61 71 81 90 100 110 Flaps 408 Normal 40 50 60 70 80 85 Alternate 38 50 60 70 79 83 viously stated, the compressibility error correction can be found using the circular slide rule (navigation computer). (4) Blocked pitot tube. A blocked pitot tube will mean that the pressure in the pitot system will be unaffected by changes in airspeed. In level flight the ASI will indicate a false airspeed, dependent upon the pres- sure locked in the system, and will not indicate changes in airspeed. In the climb the decreasing static pressure will cause the ASI capsule to expand and the instrument will falsely indicate an increasing airspeed. In the descent the reverse will be the case. (5) Blocked static source. A blocked static source will cause the ASI to underread as the aircraft climbs above the altitude at which the blockage occurred, since the pressure trapped in the instrument case will be progressively greater than ambient static pressure. In the des- cent the reverse will happen and a hazardous condition arises, since the progressively overreading ASI may cause the pilot to reduce power and airspeed may fall below stalling speed. A blockage to the static source during level flight will not be readily apparent, since the ASI indication will not change. ASI tolerance Typical required accuracy for the airspeed indicator is +3 knots. The mach meter The speed at which sound travels through the air is known as the speed of sound, or sonic speed. This speed varies with air temperature and therefore with location. The speed of sound at any specific location is known as the local speed of sound (LSS). Aircraft that are not designed to fly at supersonic speeds usually suf-  Air Data Instruments 19 fer both control and structural problems when the airflow over the air- frame, particularly over the wings, approaches the LSS. Consequently it is essential that pilots of such aircraft be aware of the aircraft's speed rela- tive to the LSS. This is especially important at high altitude, since the speed of sound decreases with temperature, and air temperature decrea- ses with altitude in a normal atmosphere. Hence, the greater the alti- tude, the lower the LSS. The aircraft's speed relative to the LSS is measured against a scale in which the LSS is assigned a value of 1. The limiting speed above which control problems may be encountered is known as the critical mach speed and is assigned a value known as the critical mach number (Mcrit). Supposing that, for a particular aircraft, this speed happens to be 70% of the LSS, then the critical mach number would be 0.7. The mach number (M) is the ratio of the aircraft's true airspeed (TAS) to the LSS. This may be represented by the equation: TAS Mˆ LSS Thus, if an aircraft is flying at a TAS of 385 knots at an altitude where the LSS is 550 knots, the aircraft's mach number is 385 7 550 = 0.7. Clearly, if the aircraft were flying at a TAS of 550 knots in these conditions its mach number would be 1.0 and it would be flying at sonic speed. Since mach number is the ratio of TAS to LSS it follows that it is pro- portional to the ratio of EAS, CAS or RAS to LSS, but the requisite corrections must be applied. Once again, the navigation computer, or circular slide rule, is equipped to do this. The speed of sound in air is entirely dependent upon the air temperature. The lower the air temperature, the lower the LSS, therefore LSS decreases with increasing altitude in a normal atmosphere. The LSS can be calculated using the formula: p LSS ˆ 38:94 TK where TK is the local air temperature in degrees kelvin. The kelvin, or absolute temperature, scale is based upon absolute zero, which is equal to 72738C. Thus, 08C is equal to 273K and an ISA mean sea level temperature of +158C is equivalent to 273 + 15 = 288K. From this it can be shown that the local speed of sound at ISA mean sea level is: p 38:94 288 ˆ 661 knots However, at 36 090 ft in the standard atmosphere, where the ambient air temperature is 756.58C: 20 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems p LSS ˆ 38:94 273 ÿ 56:5 p ˆ 38:94 216:5 ˆ 573 knots Since mach number is the ratio of TAS to LSS it follows that mach number is also dependent upon local air temperature. At any given true airspeed the aircraft's mach number varies inversely with temperature. Thus, an aircraft climbing in standard atmospheric conditions at a true airspeed of 573 knots would reach mach 1 at an altitude of 36 090 feet, whereas at sea level its mach number would have been 573 7 661 = 0.87. The mach meter is essentially a pressure altimeter and an ASI combined in one instrument. Its purpose is to indicate the aircraft's airspeed relative to the LSS. We have seen that the LSS decreases with air temperature and, therefore, with altitude in a normal atmosphere. The mach meter contains an altitude capsule similar to that in the pressure altimeter. Static pressure is supplied to the sealed instrument casing and the aneroid altitude capsule will expand against a light spring as static pressure decreases with increasing altitude. Figure 1.11 is a schematic diagram showing the principle of operation of the mach meter. The instrument also contains an airspeed capsule, the inside of which is connected to pitot pressure. As airspeed, and therefore dynamic pressure, increases the airspeed capsule will expand, exactly as in the ASI. altitude capsule static pressure pitot ratio arm pressure ranging arm airspeed capsule Figure 1.11 Mach meter principle of operation. The movement of the altitude and airspeed capsules is transmitted to the mach meter pointer through a system of mechanical links and gears. The pointer rotates against a dial calibrated to show the aircraft speed in terms of mach number. As a pressure instrument the mach meter cannot measure the ratio of TAS to LSS, but it satisfies the requirement by measuring the ratio of dynamic pressure (pitot 7 static) to static pressure. This can be expressed as:  Air Data Instruments 21 p ÿ s† Mˆ s From the foregoing it follows that an increase in airspeed will raise the mach number, bringing the aircraft's speed closer to the LSS. An increase in altitude will reduce the LSS, also bringing the aircraft's speed closer to the LSS, raising the mach number. Let us examine the operation of the mach meter under these circumstances. Increased airspeed at constant altitude Movement of the airspeed capsule is transmitted to the instrument pointer through a ratio arm, a ranging arm and gearing. Let us assume there is an increase of airspeed in level flight. The consequent increase in dynamic pressure causes the airspeed capsule to expand and the ratio arm rotates to bear against the ranging arm. This in turn rotates a quadrant and pinion gear system, which is connected to the instrument pointer. Thus, the expansion of the airspeed capsule causes the pointer to move against a calibrated scale, indicating an increased mach number. A decrease in airspeed will cause the airspeed capsule to contract and the above sequence will be reversed, resulting in a decreased mach number indication. Increased altitude at constant airspeed An increase in altitude will cause the altitude capsule to expand and the linear movement of the ratio arm against the ranging arm will again rotate the ranging arm and its associated quadrant and pinion to rotate the mach meter pointer, indicating an increased mach number. A decrease in altitude will cause the altitude capsule to contract and the above sequence will be reversed, resulting in a decreased mach number indication. Figure 1.12 shows a typical mach meter display. Mach/TAS calculations Calculation of mach number given TAS and LSS has already been demon- strated. Clearly, by transposition of the formula, it is possible to calculate TAS given mach number (M) and LSS. TAS ˆ M  LSS LSS can be calculated if the ambient air temperature is known and converted to kelvin. 22 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems 1.3 0.5 MACH 0.6 1.2 0.7 1.1 1.0 0.8 0.9 critical mach number Figure 1.12 Typical mach meter display. Mach meter errors (1) Instrument error. In common with all pressure instruments the man- ufacturing tolerances inevitably lead to small errors of measurement due to friction and lost motion in the linkages and gearing. These are typically of the order of +0.1 M over a range of 0.5 M to 1.0 M. Instrument error increases slightly with increasing speed and altitude. (2) Pressure error. The mach meter is more sensitive than the ASI to errors in static pressure measurement, since it uses the ratio of dynamic pressure (p7s) to static pressure, whereas the ASI uses the ratio of pitot pressure to static pressure. (3) Blockages. Blockage of either pressure source will cause the measured ratio to be incorrect. If the static source is blocked, changes in altitude will not be sensed and the instrument will underread in the climb. If the pitot source becomes blocked the instrument will not respond to speed changes. The exact effect of a pitot blockage depends largely upon what the aircraft is doing at the time and is therefore difficult to predict. The vertical speed indicator (VSI) The vertical speed of an aircraft is otherwise known as its rate of climb or descent and the VSI is alternatively known as the rate of climb/descent indicator (RCDI). The purpose of the VSI is to indicate to the pilot the air- craft's rate of climb or descent, typically in feet per minute. Since it is rate of change of height being indicated it is necessary to create a pressure difference within the instrument whilst a height change is occur- ring, and to arrange that the magnitude of the pressure difference is pro- portional to the rate of change of height. This is achieved by a metering unit within the instrument, which is illustrated in schematic form in Figure 1.13.  Air Data Instruments 23 pinion quadrant pointer metering unit capsule static pressure Figure 1.13 Vertical speed indicator principle of operation. Static pressure is led directly to the inside of a capsule and also to the inside of the sealed instrument casing via a metering orifice. The capsule is connected to a pointer through linkages and a quadrant and pinion gear. Whilst the aircraft is in level flight the pressure in the capsule is the same as that in the casing and the pointer is in the horizontal, nine o'clock position indicating zero. When the aircraft enters a climb, static pressure begins to fall and this is sensed virtually immediately within the capsule. Pressure in the instrument casing is now greater than that in the capsule, since air can only escape at a controlled rate through the restricted orifice of the metering unit. The pressure difference causes the capsule to contract, driving the pointer in a clockwise direction to indicate a rate of climb. The faster the aircraft's rate of climb, the greater will be the pressure difference between capsule and casing and the greater the capsule compression, driving the pointer further around the scale. During a descent the increasing static pressure will be felt virtually instantly within the capsule, but the pressure in the instrument casing will rise at a slower rate due to the effect of the metering unit. Hence the capsule will expand, against a spring, driving the pointer in an anti-clockwise direction to indicate a rate of descent proportional to the pressure difference, which is in turn proportional to the rate of descent. When the aircraft levels out at a new altitude the pressure in the instru- ment casing will equalise with that in the capsule and the pointer will return to zero. A typical VSI presentation is shown in Figure 1.14. It should be noted that the scale graduation is logarithmic, having greater spacing at lower rates of climb. This is deliberate, to facilitate easy interpretation when small changes of height are being made. 24 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems 1 2.5 FT/MIN 4 0.5 X1000 4 1 2 Figure 1.14 VSI presentation. Instantaneous vertical speed indicator (IVSI) A disadvantage of the basic VSI described above is that there is an inherent time lag in displaying climb or descent when either is initiated. This is due to the hysteresis of the capsule, which needs a significant pressure differential before it begins to expand or contract. In aircraft of moderate performance this is unimportant, but in higher performance aircraft an instantaneous response to height change is necessary. This is achieved by introducing a small cylinder connected to the static pressure supply to the capsule, containing a lightly spring-loaded free pis- ton. This device is known as a dashpot accelerometer. When a climb is initiated, inertia causes the piston to move down in the cylinder, creating an instantaneous, but temporary, pressure drop in the capsule. The capsule immediately contracts, causing the instrument to indicate immediately the commencement of a climb. Thereafter, the decreasing static pressure due to the climb causes an indicated rate of climb as previously described. When a descent is initiated, inertia causes the dashpot piston to rise in its cylinder, creating a small instantaneous pressure rise in the capsule, suffi- cient to expand it and indicate the commencement of descent. A minor disadvantage of the IVSI is that, on entering a turn in level flight, the centrifugal acceleration force is liable to displace the dashpot piston and cause a temporary false indication of climb. Upon exiting the turn the reverse will be the case. The IVSI is also more sensitive to turbulence and is liable to give a fluctuating indication in such conditions. This is usually damped out by the inclusion of a restriction in the static connection to the capsule and the metering unit. An adjusting screw is provided below the face of the instrument, by which the pointer can be set to read zero. The range of adjustment available is, typically, between +200 ft per minute and +400 ft per minute, depending upon the scale range of the instrument.  Air Data Instruments 25 Errors (1) Instrument error. Manufacturing tolerances lead to small errors in the internal mechanism. When these cause displacement of the pointer from the zero position with the aircraft stationary on the ground it can be corrected with an adjusting screw on the front of the instrument. (2) Pressure error. Disturbances at the static source may cause the instru- ment to display an incorrect rate of change of height. (3) Lag. A delay of a few seconds before a rate of change of height is indicated is normal with the basic VSI. The IVSI virtually eliminates lag. (4) Transonic jump. A transonic shock wave passing over the static source will cause the VSI briefly to give a false indication. (5) Blockage. Blockage of the static source will render the VSI useless, since it will permanently give a zero indication. Failure of the instrument will usually result in a fixed indication of zero rate of change of height. Dynamic vane-type VSI Sailplanes and a few very light aircraft often use a different type of VSI known as a variometer. Figure 1.15 shows a schematic illustration. An enclosed casing shaped rather like a shallow tin can contains a pivoted vane, held in a central position by a light leaf spring. The casing on one side of the vane is connected to the static pressure source and on the other side of the vane it is connected to an enclosed air reservoir. In level flight air can leak past the vane sufficiently for the pressure in the reservoir to equalise with from static source vane pointer to reservoir Figure 1.15 Variometer. 26 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems static air pressure. Under these conditions the leaf spring will centralise the vane, which is connected to a pointer indicating zero rate of climb/descent. When a climb is initiated the static pressure will begin to fall below the pressure stored in the air reservoir. The differential pressure across the vane will deflect it and the attached pointer will indicate a rate of climb. The faster the rate of climb, the greater the differential pressure and the more the vane/ pointer will be deflected. When the aircraft levels off, air will leak past the vane to equalise the reservoir with static pressure once more and the leaf spring will centralise the vane/pointer to the zero position. In a descent the sequence of events will be the opposite of that described above. The vane-type variometer is incapable of displaying large rates of height change and is therefore unsuitable for most powered aircraft. It does, however, have the advantage that it suffers less from time lag (i.e. initial response) than the basic VSI. There is also a type of variometer used specifically in sailplanes that responds to height gain or loss not initiated by the pilot, as in a thermal. This type is known as the total energy variometer. The air data computer (ADC) In addition to the four instruments already covered there are numerous systems in modern aircraft that require air data inputs in terms of static pressure, pitot pressure and air temperature. Since most, if not all, of these systems are electronic in operation it is logical to supply such data in elec- tronic form. The air data computer receives pitot and static pressure from the normal and alternate sources and converts these into electrical signals and transmits them to the various indicators and systems. As far as indicators are concerned, this removes the need for bulky instruments that take up sig- nificant panel space and means that the information can be presented in digital form if necessary. The centralised air data computer can also be programmed to apply the necessary corrections for pressure error, barometric pressure changes and compressibility effects. With the addition of air temperature data inputs, true airspeed can be calculated by the computer. Analogue and digital ADCs Air data computers are usually of the digital type; that is, they transmit data in digital format which is compatible with other computer-based systems. Analogue air data computers, which transmit their output data to servo- operated devices, are less common, although a few are still in existence. Figure 1.16 is a block diagram showing the data inputs and outputs of a typical ADC. It will be seen that the data inputs are pitot and static pressure  Air Data Instruments 27 inputs computer outputs altitude altitude memory hold static pressure static altitude pressure transducer dynamic indicated pressure airspeed mach mach number number pitot pressure pitot pressure transducer true true airspeed airspeed total air temperature static static air air temperature temperature air density density systems using adc outputs flight director automatic thrust control automatic flight control cabin pressurisation altitude reporting flight management gpws flight recorder stall warning Figure 1.16 Data inputs and outputs of an air data computer. and total air temperature (TAT). From these, electrical signals are generated and transmitted as electronic data to operate the pilots' air data instrument displays, plus TAS, TAT and SAT (static air temperature) displays. Addi- tionally, these signals are transmitted to the various flight management control systems, listed in Figure 1.16. Loss of air data input activates a warning logic circuit within the ADC, which causes warning flags to appear on the associated indicators and annunciators to illuminate on the computer control panel. At this point it is perhaps appropriate to define the various forms of air temperature measurement. 28 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems (1) SAT (static air temperature). This is the temperature the air at the surface of the aircraft would be at if there were no compression effects due to the aircraft's movement. At very low airspeeds these effects are negligible, but for most transport aircraft normal flight speeds are such that the direct measurement of SAT is virtually impossible. SAT is also known as outside air temperature (OAT). (2) TAT (total air temperature). Total air temperature is the temperature of the air when it has been brought completely to rest, as in the pitot tube. The ram rise, that is the temperature increase due to compression, can then be subtracted from TAT to give corrected outside air temperature (COAT). The value of ram rise can be calculated for any given mach number, so clearly the air data computer can be programmed to make this correction. Temperature measurement probes Aircraft that operate at low airspeeds, such as some helicopters and light aircraft, usually employ a simple bimetallic thermometer, which operates a rotary pointer against a temperature scale to indicate what is essentially static air temperature. TAT sensors are more complex because they must, as far as possible, recover the temperature rise due to adiabatic compression of the air as it is brought to rest; what is known as the `ram rise'. To do this it is necessary to use sophisticated probes to capture and slow the air and then convert the air temperature to an electrical signal for transmission to the ADC. The sensitivity of the probe in terms of the extent to which it truly achieves, or recovers, the full ram rise effect is known as its recovery factor. For example, a probe that senses SAT plus 85% of the ram rise in temperature would be said to have a recovery factor of 0.85. TAT probes are typically contained within an aerodynamically shaped strut with an air intake mounted on the outer end, clear of any boundary layer air. Air is drawn into the hollow strut, through the air intake, where it is brought virtually to rest. Its temperature is sensed by a platinum resis- tance-type element mounted within the strut, which produces an electrical signal proportional to the temperature. Most modern TAT probes have a very high recovery factor, usually very close to unity (1.0). An alternative type of probe uses engine bleed air to create a reduction of pressure within the casing of the probe. This has the effect of drawing air into the hollow strut at a higher rate, so that the de-icing heating element within the strut cannot affect the sensed temperature of the indrawn air. Figure 1.17 is a diagram of an air temperature measurement probe.  Air Data Instruments 29 air flow air flow sensing engine element bleed air out heating element electrical engine output bleed air connection in Figure 1.17 TAT measuring probe. Pressure transducers Devices called transducers convert pitot and static pressure into suitable electrical signals for transmission to the air data computer. In the case of analogue ADCs the transducer is often of the electro-magnetic type, the amplified output of which drives a servomotor and operates a synchro system, which in turn operates the analogue instrument displays. Digital ADCs more commonly utilise piezoelectric transducers that form part of a solid-state circuit. Some crystalline materials, such as quartz, can be made to generate varying electrical signals when subjected to pressure. A diaphragm composed of thin quartz discs impregnated with metallic par- ticles is subjected to pitot or static pressure and the subsequent flexing of the diaphragm creates an electrical charge in the discs, the polarity of which is dependent upon the direction of flexing. Thus a signal proportional to increasing or decreasing pressure is generated. Sample questions 1. At lower altitudes, near to sea level, the change of atmospheric pressure with height is approximately: a. 1 hPa per 50 ft? b. 1 hPa per 27 ft? c. 27 hPa per 1 ft? d. 1 hPa per 10 ft? 30 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems 2. ISA mean sea level pressure is: a. 1013 hPa? b. 1025 hPa? c. 29.02 in Hg? d. 1030 hPa? 3. An aircraft is at 600 ft above an airfield, the elevation of which is 250 ft amsl. With QFE set, the altimeter will read: a. 850 ft? b. 350 ft? c. 600 ft? d. 250 ft? 4. An aircraft altimeter has been set for a regional QNH of 1010 hPa, but it is operating in an area where msl pressure is 1000 hPa. The altimeter is indicating 2000 ft; actual height amsl is: a. 2300 ft? b. 1730 ft? c. 1940 ft? d. 1970 ft? 5. Which of the following is correct? a. Static pressure = pitot pressure + dynamic pressure? b. Dynamic pressure = pitot pressure + static pressure? c. Pitot pressure = dynamic pressure 7 static pressure? d. Dynamic pressure = pitot pressure 7 static pressure? 6. An aircraft is flying at constant IAS: a. As altitude decreases EAS will increase? b. As altitude increases CAS will increase? c. As altitude increases TAS will increase? d. As altitude decreases TAS will increase? 7. The stalling speed with flaps and landing gear fully extended is assigned the notation: a. Vso? b. Vse? c. Vfe d. Vno?  Air Data Instruments 31 8. An aircraft is flying at a TAS of 400 knots at an altitude where the LSS is 550 knots. The aircraft's mach number is: a. 1.4? b. 0.65? c. 0.82? d. 0.73? 9. The white arc on an ASI scale: a. Extends over the safe speed range with flaps retracted? b. Extends over the safe speed range with flaps fully extended? c. Extends over the safe speed range with one engine inoperative? d. Extends over the speed range from Vno to Vne? 10. The error suffered by air data instruments that is caused by manu- facturing tolerances is known as: a. Pressure error? b. Lag? c. Position error? d. Instrument error? 11. The type of VSI commonly used in sailplanes is: a. IVSI? b. Simple VSI? c. Dynamic vane type? d. Variation meter? 12. The data inputs to an ADC are: a. TAT, pitot and static? b. SAT, pitot and static? c. COAT, dynamic and static? d. OAT, pitot and static? Chapter 2 Gyroscopic Instruments and Compasses Gyro fundamentals Gyroscopic instruments are of great importance in aircraft navigation because of their ability to maintain a constant spatial reference and thereby provide indication of the aircraft's attitude. The principal instruments that use the properties of the gyroscope are the directional gyro, the artificial horizon or attitude indicator and the turn and bank indicator. Gyroscopic properties The gyroscope used in these instruments comprises a rotor, or wheel, spinning at high speed about an axis passing through its centre of mass and known as the spin axis. A simple gyro rotor is illustrated in Figure 2.1. When a rotor such as that in Figure 2.1 is rotating at high speed it exhibits two basic properties, known as rigidity and precession. It is these properties that are utilised to give gyroscopic instruments their unique features. pla ne of r ota tion A xis na spi Figure 2.1 Gyro rotor.  Gyroscopic Instruments and Compasses 33 Rigidity The spinning rotor of the gyro has rotational velocity and therefore, if we consider any point on the rotor, that point has angular velocity as indicated by the arrow A in Figure 2.1. Since the rotor has mass, that angular velocity produces angular momentum, which is the product of angular velocity and mass. As stated in Newton's First Law of Motion, any moving body tends to continue its motion in a straight line and this is known as inertia. In the case of the spinning gyroscope there is a moment of inertia about the spin axis, which tends to maintain the plane of rotation of the gyro. Consequently, the spin axis of a gyroscope will maintain a fixed direction unless acted upon by an external force. This property is known as rigidity. Another way of putting this is that the spin axis of the gyro will remain pointing toward a fixed point in space unless it is physically forced to move. Since rigidity is the product of angular velocity and mass it follows that the rigidity of a gyroscope may be increased by increasing either its angular velocity, or its mass, or both. Increasing the speed of rotation of the rotor, or its diameter, will increase angular velocity and therefore angular momen- tum. The rotor diameter is constrained by the need to keep the instrument as compact as possible and so the gyro rotor is made to spin at very high speed. Similarly, the mass of the rotor is constrained by its size limitations, but angular momentum is improved if the mass is concentrated at the rim of the rotor, as seen in Figure 2.1. Precession Precession is defined as the angular change in direction of the spin axis when acted upon by an applied force. Let us suppose that the axis of the gyro rotor in Figure 2.2 has a force applied to it as shown in (a). Application of a force to the spin axis as shown is exactly the same as if the force had been applied at point X on the rotor. If the rotor were stationary then it, and its spin axis, would tilt as shown at Figure 2.2(b). However, when the rotor is rotating it has not only the applied force acting upon it, but also the angular momen- tum previously described. The combination of the two displaces the effect of the applied force through 908 in the direction of rotation, as shown in Figure 2.2(c). Thus, the spin axis of the gyroscope will precess as shown in Figure 2.2(d) in response to the force applied in Figure 2.2(a). The rate at which a gyro precesses is dependent upon the magnitude of the applied force and the rigidity of the rotor. The greater the applied force, the greater the rate of precession. However, the greater the rigidity of the rotor the slower the rate of precession for a given applied force. A gyro will continue to precess so long as the applied force is maintained, or until the applied force is in the same plane as the gyro plane of rotation, as 34 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems applied force X applied (a) (b) force gyro rotor movement applied applied force force precession precession of of applied applied force applied applied force force (c) force (d) applied force (e) Figure 2.2 Gyroscopic precession. shown in Figure 2.2(e). If the applied force is removed, precession will immediately cease. Free gyro Clearly the rotor of the gyroscope must be contained within a supporting structure. The rotor spindle is mounted within a ring known as a gimbal and this is in turn mounted within a framework, the design of which depends upon the gyro function. All gyroscopes must have freedom for the rotor to rotate and to precess. The gyroscope cannot precess about the axis of rotation, but precession may take place about either of the two axes at right angles to the plane of rotation. A gyroscope that has freedom to precess about both these axes is known as a free gyro, and is said two have two degrees of freedom of precession. Such a gyroscope is illustrated in Figure 2.3. The number of  Gyroscopic Instruments and Compasses 35 Z Y rotor inner frame gimbal X spin axis X outer spindle gimbal Y Z Figure 2.3 Free gyro. degrees of freedom of precession of any gyroscope is the same as its number of gimbals. It will be seen that the gyro rotor spindle is mounted in bearings within a ring, or gimbal, known as the inner gimbal. This is in turn mounted in bearings that are attached to a second gimbal ring, known as the outer gimbal. Thus, the gyro rotor is free to spin about spin axis XX and it also has freedom of movement about the inner gimbal axis YY. The outer gimbal is mounted in bearings attached to the frame of the assembly and therefore has freedom of movement about the third axis, ZZ. If the frame of the free gyro were to be fixed to the instrument panel of an aircraft, the aircraft could be pitched, rolled or even inverted and the spin axis of the spinning gyroscope would remain aligned with the same fixed point in space. In point of fact the free gyro has no practical application in aircraft, but gyroscopes having freedom of precession about two axes, known as tied gyros, are extremely useful. Tied gyro The aircraft instruments that employ gyros provide a fixed reference, about which aircraft movement is indicated to the pilot. The directional gyro provides the pilot with aircraft heading information and so its reference axis is the aircraft's vertical axis and the gyro rotor must be sensitive to movement about that axis and no other. The function of the attitude indi- cator is to provide the pilot with indications of aircraft attitude with refer- ence to the pitch and roll axes of the aircraft and so its gyro must be sensitive to aircraft movement about these axes. The turn indicator is 36 Ground Studies for Pilots: Flight Instruments & Automatic Flight Control Systems required to indicate rate of turn and so it must be sensitive to aircraft movement about the vertical axis. A gyroscope is not sensitive to movement about its spin axis, so its rotor must be maintained at right angles to the required axis for maximum sen- sitivity. Consider the situations depicted in the illustration in Figure 2.4. A B C C B A F E D Figure 2.4 Gyroscope spin axis alignment. Axis AA is the aircraft's vertical, or yaw, axis. Any movement about this axis involves a change in aircraft heading and so we require the directional gyro to be sensitive to movement about this axis. Since a gyro is not sensitive to movement about its spin axis it is clear that a gyro with its spin axis vertical (gyro D in Figure 2.4) would not be suitable, but that the spin axis of a directional gyro must be maintained horizontal. The attitude indicator is required to indicate aircraft attitude with refer- ence to the aircraft pitch an