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Aircraft Instruments and Integrated Systems book. Read 6 reviews from the world's who have pdf copy for this book? like · 3 years ago · Add your answer. E.H.J. Pallett is the author of Aircraft Instruments and Integrated Systems ( avg rating, 70 ratings, 6 reviews, published ), Aircraft Instrument. Brief Review of Aircraft Instruments & Integrated Systems. IJIRAE::International Journal of Innovative Research in Advanced Engineering, Volume V,

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Integrated Systems, First Edition, by Pallett, published by Pearson Education Limited,. Cop)Tight '. . The title Aircraft instruments and integrated systems is. 14 Engine power and control instruments 1 S Integrated instrument and flight director systems 16 Flight data recording Tables. Principal symbols. Aircraft Instruments & Integrated System by e.h.j Pallett - - Ebook download as PDF File .pdf), Text File .txt) or read book online. A/c IS.

An indicator mechanism is shown in schematic form in Fig. We have -come a long way from the time when the engineer had only to undo four bolts and two unions, and out came the airspeed indicator. At a in Fig. These indicators consist of the same basic elements as conventional VSis. The magnification ratio between the two levers is therefore altered as the altitude mechanism divides p. In order to indicate whether temperatures are either positive or negative. A point which may be noted in connection with turns from E or W is that when the N or S end of the magnet system is tilted up.

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Readers Also Enjoyed. About E. Books by E. This balances out any pressure differences which might be caused by the location of the static holes along the fore-and-aft axis of the probes. Valves are of the self-closing type so that they cannot be left in the open position after drainage of accumulated water. Drain holes provided in probes are of such a diameter that they do not introduce errors in instrument indications.

Drain traps are designed to have a capacity sufficient to allow for the accumulation of the maximum amount of water that could enter a system between servicing periods. In order for an air data system to operate effectively under all flight conditions. The diameter of pipelines is related to the distance from the pressure sources to the instruments in order to eliminate pressure drop and time-lag factors.

The method of draining the pipelines varies between aircraft types. Such provision takes the form of drain holes in probes. They are commonly used in the more basic air data systems installed in many types o small aircraft. Servo-operated instruments are. The fundamental principles of these instruments will be described-in a later chapter. I3 Air data system drains. When the fluid flows at a certain velocity V over the probe it will be brought to rest at the nose.

It is also equal to the product of the ratio of the mass m to density p and pressure p. If the fluid is an ideal one. In connection with this probe. This means that work must be done by the mass of fluid and this raises an equal volume of fluid above the level of the fluid stream. TheRinetic energy ofa mass m before being brought to rest is t equ"afto ml72".. The work done in raising the fluid is equal to the product of its mass.: In coming to rest at the stagnation point.

Let us consider a pitot probe placed in a fluid with its open end facing upstream as shown in Fig. The term 'computed' appiies specifically to air data computer systems in which PE corrections are automatically applied to an airspeed sensing module via an electrical correction network. This coefficient is. CAS is automatically compensated for compressibility of air at a pitot probe to obtain EAS at varying speeds and altitudes..

In order therefore to minimize 'compressibility errors' in indication. In air data computer systems. Jndicated airspeed IAS fj: Errors and appropriate corrections to be applied are determined by comparison against calibration equipment having high standards of accuracy.

Pitot pressure p. This is also done automatically in air data computer systems. Mmo Maximum operating speed in terms of Mach number.

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The foregoing airspeeds are summarized pictorially in Fig Je which damps out pressure surges. Displacements of the capsule in accordance with what is called the 'square-law' are transmitted via a magnifying lever system.. The pressure-sensing element is a metal capsule. Except for this connector the case is sealed I 7 Square-law characteristics. If also the capsule were coupled to the pointer mechanism so that its deflections were directly magnified. The non-linearity of such a scale makes it difficult to read accurately.

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.

In some types of servo-operated indicators. As speed and differential pressure increase. Since the speed of sound The principle of a commonly used version of the foregoing method is one in which the length of a lever is altered as progressive deflections of the capsule take place.

Another type of square-law compensating device is shown in Fig. In other words. It consists of a special ranging or 'tuning' spring which bears against the capsule and applies a controlled retarding force to capsule expansion. Of the two methods the latter is the more practical because means of adjustment can be incorporated to overcome the effects of capsule 'drift' plus other mechanical irregularities as determined during calibration.

I Airspeed capsule. This ratio. Let us assume that the aircraft is flying under standard sea-level conditions at a speed V of mph.. Figure 2.! The speed of sound at sea-level This obviously is not a practical solution. C and D will be set to angular positions determined by this difference. At sea-level and as based on our earlier assumption. We may now consider how the altitude mechanism of the Machmeter fiinctions in order to achieve this. The speed of sound cannot be measured by the instrument.

The airspeed mechanism therefore tends to make the pointer indicate a lower Mach number. What happens at altitudes above sea-level? As already pointed out.. It is for this reason that critical Mach numbers Mer. The critical Mach number for a particular type of aircraft is indicated by a pre-adjusted lubber mark located over the dial of the Machmetc. The magnification ratio between the two levers is therefore altered as the altitude mechanism divides p.

Ps by p. Thus a Machmeter indicates the Mach number Via in terms of the pressure ratio p. It affects the pressure difference p.

Two external index pointers around the bezel may be manually set to any desired reference speed. 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..

A second pointer. The mechanism consists of two measuring elements which drive their own indicating elements. The pointer is striped red and white and can be pre-adjusted to the desired limiting speed value. In operation. When the limiting speed is reached.

The adjustment is made on the ground against charted information appropriate to the operational requirements of the particular type of aircraft. 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.

In addition to their basic indicating function. It has. The necessary computation is effected by calibrating the scales to logarithmic functions of pitot and static pressures. The indicating element for this purpose is a servomotor-driven digital counter. In the Figure 2. It will be noted from Fig. In aircraft having an autothrottie system. In the example illustrated. A readout of the command speed is given on a digital counter which is also mechanically set by the command speed knob.

A check on the operation of the failure monitoring and flag circuits. The dial presentation and mechanical features of a typical pneumatic type of altimeter are shown in Fig. The resultant of both curves produces the linear scale as at curve 4.

This conversion is represented by the graphical example shown in Fig. Since the ISA also assumes certain temperature values at all altitudes. In standard conditions. The bi-metal compensator is simultaneously affected by the decrease in ambient temperature. In a similar manner. At higher altitudes the same effects on elasticity will take place. We may consider these errors by taking the case of a simple altimeter situated at various levels.

Systems and aircraft pdf integrated instruments

In practice. As far as altimeters are concerned. Assuming that at the sea-level airfield the pressure falls to The relationship between the various altitudes associated with flight operations is presented graphically in Fig.

The altimeter will thus read a greater pressure Mop and will indicate an altitude greater than ft. The pressure of A 1B1 is. At point H. I 8 1- C'4: This may be seen from the three columns shown in Fig.. Variations in temperature cause differences of air density and therefore differences in weight and pressure of the air. It will be apparent from the foregoing that. In order.. If the temperature of the air in part AB increases.

Thus the altimeter. At point A the altimeter measures the pressure of the column AC. At a the altimeter is assumed to be subjected to standard conditions. If now the altimeter is raised through ft as at The underlying principle of this may be understood by considering the setting device to be a millibar scale having a simple geared connection to the altitude pointer as shown in Fig. If the setting is then changed to. The deflected position of the capsules appropriate to whatever pressure is acting on them at the time will not be disturbed by rotation of the mechanism.

When the knob is rotated then. Likewise it will be noted that the setting knob is also geared to the sensing element mechanism body.

Hg and the other in mb interconnected through gearing to a setting knob. In the altimeter shown in Fig. There are two code letter groups commonly used in connection with altimeter setting procedures. For this purpose. The requests and transmissions are adopted universally and form part of the ICAO 'Q' code of communication. In order to make the settings flight crew are dependent on observed meteorological data which are requested and transmitted from air traffic control. The pressure set is a value reduced to mean sea-level in accordance with ISA.

The zero reading is regardless of the airport's elevation above sea-level. When used for landing and take-off.

Aircraft Instruments and Integrated Systems

Any value is only valid in the immediate vicinity of the airport concerned. Since an altimeter with a QNH setting reads altitude above sea- level. QFE Setting the barometric pressure prevailing at an airport to make the altimeter read zero on landing at. Height is the vertical distance of a level. SAS Transition altitude Height.

Elevation is the vertical distance of a fixed point above or below mean sea-level. The following definitions. Where a runway is below the airport elevation. QFE Altitude. It is used for flights above a prescribed transition altitude and has the advantage that with all aircraft using the same airspace and flying on the same altimeter setting.

For altimeter settings the QFE datum used is the airport elevation which is the highest usable point on the landing area. The transition altitude within UK airspace is usually ft to ft. Flight levels.

Instruments systems aircraft pdf integrated and

Altitude is the vertical distance of a level. The other end of the metering unit is open to the interior of the case to apply static pressure to the exterior of the capsule. Let us now see how the instrument operates under the three flight conditions shown in the diagram. An indicator mechanism is shown in schematic form in Fig. This is accomplished by incorporating a special air metering unit in the sensing system. The dial presentation is such that zero is at the 9 o'clock position.

Since the rate at which the static pressure changes is involved in determining vertical speed. A typical example of this presentation is shown in Fig.

The reason for this is that a logarithmic scale is more open near the zero graduation. This tube serves the same purpose as the one employed in a pneumatic type of airspeed indicator. It is. Certain types of indicator employ a linear scale. A pneumatic type of indicator consists basically of three principal components: At the instant of commencing a descent.

The pressure inside the case. Metering units are designed to compensate for the effects of the variables over the ranges normally encountered. The construction of a typical indicator is shown in Fig.

Apart from the changes of static pressure with changes of altitude. In addition. It consists of a cast aluminium-alloy body which forms the support for I Rocking shaft assembly. During a climb. An adjustment device is provided at the front of the indicator for settiog the pointer to zero. The basic principle is illustrated in Fig. The accelerometer comprises a small cylinder. When a change in vertical speed is initiated.

The capsule displacement in turn produces instantaneous deflection of the indicator pointer over the descent portion of the scaie. Instantaneous vertical speed indicators IVS! These indicators consist of the same basic elements as conventional VSis. The flange of the metering unit connects with the static pressure connection of the indicator case.

At initiation of an ascent. Displacements of the capsule in response to differential pressure changes are transmitted to the pointer via a balanced link and rocking-shaft assembly. The purpose of the restrictor in the bypass line is to prevent any loss of pressure change effects created by displacements at the acceleration pump. The cylinder is connected in a capillary tube leading to the capsule. The upper spring and its adjusting screws control the rate of descent calibration.

The accelerometer response decays in each case after a few seconds. The range of adjustment around zero depends on tne scale range of any one' type of indicator. At speeds below The temperature which would overall be the most ideal is that of air under pure static conditions at the flight levels compatible with the operating range of any particular type of aircraft concerned.

The measurement of static air temperature SAT by direct means is. As the helix expands or contracts. An example of this thermometer and its installation in one type of helicopter is shown in Fig.

Details of the method by which this is normally accomplished will be given in Chapter 7. If the corresponding SAT value is to be determined and indicated. Various types of sensor may be adopted for the sensing of air temperature. The simplest type. The measurement of TAT requires a more sophisticated measuring technique. This parameter is referred to as total air temperature TAT and is derived when the air is brought to rest or nearly so without further addition or removal of heat.

The thermometer is secured through a fixing hole in the side window of a cockpit. For use in aircraft capable of high Mach speeds. The element is arranged in the form of a helix anchored at one end of a metal sheath or probe.

In this context.. The bled holes in the intake casing In flight. TAT sensors are of the probe type. The probe is in the form of a small" strut and air intake made of nickel-plated berylfa: It is secured at a pre-determined location in the front fuselage section of an aircraft typically at the side.

The heater dissipates a nominal W under in-flight icing conditions. A pure platinum wire resistance-type sensing element is used and is hermetically sealed within two concentric platinum tubes. The errors involved. A secord type of TAT probe is shown in Fig. An axial wire heating element. The principal differences between it and the one just described relate to the air intake configuration and the manner in which airflow is directed through it and the probe casing.

The probe has an almost negligible time- lag. The element is wound on the inner tube. The purpose of the engine bleed air injector fitting and tube is to create a negative differential pressure within the casing so that outside air is drawn through it at such a rate AIR fl.

In addition to TAT. The internal arrangement of an LCD see page 15 type of indicator is schematically shown in Fig. This controls an 'OFF' flag which under normal conditions is held out of view by an energized solenoid.

TAT indicators can. The power supply to the computer is connected via supply. Air temperature indicators As in the case of other instruments. The purpose of this element is to transmit a signal to other systems requiring air temperature information. The motor then drives the counter drums. Detection of failure of the 26 V ac power to the indicator.

The system is supplied with V ac which is then stepped down and rectified by a power supply module within the indicator. In order to indicate whether temperatures are either positive or negative. The generation of the appropriate temperature signals is also accomplished by means of a de bridge circuit. The probe element forms one part of a resistance bridge circuit. The circuit of a probe and a basic conventional pointer and scale type of indicator is shown in Fig.

An example of this would be the airspeed measuring circuit of an ADC for the computing of true airspeed see Chapter 7. The temperature data signals are transmitted from a digital type of ADC see Chapter 7 via a data bus and receiver to a microcomputer. In some cases. Some of the variations are illustrated in Fig. I Mechanical drive. TAT -,-,,--, ,-, ,-,. Function selector 0 push-button. TAS, each of which can be selected in sequence by a push-button function select switch.

When power is first applied, the indicator displays TAT, as in Fig. Pushing the switch in for a third time returns the display to TAT.

Aircraft Instruments & Integrated System by e.h.j Pallett -

A test input facility is provided, and when activated it causes the display to alternate between all seven segments of each of the three digits , 'ON' for two seconds, and blank for one second. Since this is normally done b: Details of the coloured display shown at d of Fig. Air data alerting and In connection with the in-flight operation of aircraft, it is necessary to warning systems impose limitations in respect of certain operating parameters compatible with the airworthiness standards to which each type of aircraft is certificated.

It is also necessary for systems to be provided which will, both visually and aurally, alert and warn a flight crew whenever the imposed operational limitations are being exceeded.

The number of parameters to be monitored in this way varies in relation to the type of aircraft and the number of systems required for its operation overall. As far as air data measuring systems are concerned, the principal parameters are airspeed and altitude, so let us now consider the operating principles of associated alerting and warning systems typical of those used in some of the larger types of public transport aircraft.

Mach warning system This system provides an aural warning when an aircraft's speed reaches the maximum operating value in terms of Mach number, i. Mmo a typical value is 0. The system consists of a switch unit which, as can be seen from Fig.

It will also be noted that in lieu of a pointer actuating system, the sensing units actuate the contacts of a switch which is connected to a 28 V de power source.. At speeds below the limiting value, the switch contacts remain closed and the de passing through them energizes a control relay.

The contacts of this relay interrupt the ground connection to an aural warning device generally referred to as a 'clacker' because of the sound it emits when in operation. When the limiting Mach speed at any given altitude is reached, the airspeed sensing unit causes the switch contacts to open, thereby de-energizing the control relay so that its contacts now complete a connection from the 'clacker' to ground.

Since the 'clacker' i:. A toggle switch that is spring-loaded to 'OFF' is provided for the When placed in the 'TEST' position, it allows de to flow to the ground side of the switch unit control relay, thereby providing a bias sufficient to de-energize the relay and so cause the 'clacker' to be activated.

In the exam: The other indicator, which is in the first officer's group of flight instruments, is also of the servo-operated The 'ciacker' units associated with the indicators are respectively designated as 'aural warning I' and 'aural warning 2'. The captain's indicator contains an overspeed circuit module that is supplied by the ADC with prevailing speed data and also the limiting V,,w and Mmo values appropriate to the type of aircraft. The contacts of the switch unit in the first officer's indicator are connected to a relay, and since these contacts remain closed at speeds below maximum values, the relay is de-energized.

When the maximum speed is reached, the relay coil circuit is interrupted and its contacts then change over to provide a ground connection for the de supply which activates 'aural warning 2' clacker unit. Test switches are provided for checking the operation of each clacker by simulation of overspeed conditions. When switch 1 is operated de is applied to the overs peed circuit module in the captain's inr!

The operation of switch 2 applies de to the relay coil such that it is shorted out against the standing supply from the closed airspeed switch; the relay 'is therefore de-energized to provide a ground connection for 'aural warning 2' clacker unit. The indicators themselves provide visual indications of overspeed and these are discernible when the airspeed pointers become positioned coincident with pre-set maximum limit pointers see Figs 2.

The selected altitude is set by means of a knob on the controller, and is indicated by a digital counter which is geared to the rotors of control and resolver synchros, so that they produce a corresponding signal. The signal is compared with the pressure altitude signal. At predetermined values of rotor voltages of both synchros, two signals are produced and n. The sequence of alerting is shown at b of Fig. As an aircraft descends or climbs to the preselected altitude the difference signal is reduced, and the logic circuit so processes the input signals that, at a pre-set outer limit H 1 typically ft above or below preselected altitude, one signal activates the aural alerting device which remains on for two seconds; the annunciator light is also illuminated.

The light remains on until at a further pre-set inner limit H2 typically ft above or below preselected altitude, the second. As an aircraft approaches the preselected altitude, the synchro system approaches the 'null' position, and no further alerting takes place. If an aircraft should subsequently depart from the preselected altitude, the controller logic circuit changes the alerting sequence such that the indications correspond to those given during the approach through outer limit Hi, i.

Angle of attack The angle of attack AoA , or alpha a angle, is the angle between sensing the chord line of the wing of an aircraft and the direction of the relative airflow, and is a major factor in determining the magnitude of lift generated by a wing. Lift increases as a increases up to some critical value at which it begins to decrease due to separation of the slow-moving air the boundary layer from the upper surface of the wing, which, in turn, results in separation and turbulence of the main airflow.

The wing, therefore, assumes a stalled condition, and since it occurs at a particular angle rather than a particular speed, the critical AoA is also referred to as the stalling angle. The manner in which an aircraft responds as it approaches and reaches a stalled condition depends on many other factors, such as wing configuration, i. Other factors relate to the prevailing speed of an aircraft, which largely depends on engine power settings, flap angles, bank angles and rates of change of pitch.

The appropriate responses are pre-determined for each type of aircraft in order to derive specificaliy relevant procedures for recovering from what is, after all, an undesirable situation. An aircraft will, in its own characteristic manner, provide warning of a stalled condition, e. It is, therefore, necessary to provide a means whereby a can be sensed directly, and at some value just below that at which a stalled condition can occur it can provide an early warning of its onset.

Stall warning systems The simplest form of system, and one which is adopted in several types of small aircraft, consists of a hinged-vane-type senso. The vane is protected against ice formation by an internal heater element.

When a reaches that at which the warning unit has been preset. Control switches for normal operation and for testing are also provided in this unit. In larger tyges of aircraft.

Since the pitch attitude of an aircraft is also changed by the extension of its flaps. If the aircraft's attitude changes such that a increases. The circuit of a typical system is shown in basic form in Fig. When the aircraft is on the ground and electrical power is on..

Sensing relays and shock strut microswitches on the nose landing gear are included in the circuit of a system to permit operational change-over from ground to air. Stick-shaking is accomplished by a motor which is secured to a control column and drives a weighted ring that is deliberately unbalanced to set up vibrations of the column. In normal level flight conditions. The complete unit is accurately aligned by means of index pins at the side of the front fuselage section of an aircraft.

Sensor signals. It consists of a precision counter-balanced aerod namic.. The only signal now supplied to the amplifier and demodulator is the modified a signal.

The output is then supplied to a demodulator whose circuit is designed to 'bias off the ac voltage from the contacts of K 1. The demodulator then produces a resultant voltage which triggers the switch SS 1 to connect a 28 V de supply direct to the stick-shaker motor. V to the circuit module amplifier.

Bias of! In normal flight. During take-off. When such aircraft first get into a stalled condition then. The positions are: The comparator is also supplied with signals from a central processor unit also within the module which processes a programme to determine maximum a angles based on the relationship between flap position and three positions of the leading edge slats.

If the latter is higher than a computed maximum. The manner in In certain types of aircraft the sensor signals are transmitted to an air data computer. A confidence check on system operation may be carried out by placing the circuit module control switch in the 'TEST' position. This energizes a relay which switches the sensor signal to the motor of an indicator. Since the switch isolates the sensor circuit from the amplifier.

In order to prevent the development of a deep stall situation. The aircraft then sinks rapidly in the deep Slalled attitude. In aircraft having computerized flight control systems. When selected for installation. Whenever stick-push is activated. Another type of indicator currently in use has a pointer which is referenced against horizontal yellow.

In some cases a conventional pointer and scale type of display is used.

Aircraft Instruments And Integrated Systems Ehj Pallett Free Similar PDF's

Indicators are connected to the alpha sensors of a stall warning system. Indicators There is no standard requirement for angle of attack indicators to be installed in aircraft.

In the more sophisticated types of aircraft. The field differs from that of an ordinary magnet in several This partly explains the fact that the magnetic poles are relatively large areas. That this is so is obvious from the fact that a magne. A plane passing through the magnet and the centre of the earth would trace out on the earth's surface an imaginary line called the magnetic meridian as shown in Fig.

Terrestrial magnetism 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. The origin of the earth's field is still not precisely known. The operating principle of a direct-reading compass is based on established fundamentals of magnetism. As far as present-day aircraft are concerned.

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. Its points of maximum intensity. If a map were prepared to show both true and magnetic meridians.

The horizontal angle contained between the true and the magnetic meridian at any place is known as the magnetic variation or declination. Figure 3. Magnet'ic variation As meridians and parallels are constructed with reference to the true or geographic North and South pcles.. BB and CC are isoclinals. I Terrestrial magnetism. While the variation differs all over the world. At some places on the earth. The angle the lines of force make with the earth's surface at any Figure 3.

Roll attitude, or turn commands, are established in a similar manner, the command bars in this case being rotated in the required direction; diagram e of Fig. As the aircraft's attitude changes the aircraft symbolic element moves with the aircraft, while the horizon symbolic element and bank pointer are driven in the opposite direction.

When the command has been satisfied, the display will then be as shown in diagram c. The scales and pointers shown to the left and bottom of the indicator also form a director display that is utilized during the approach and landing sequence under the guidance of an Instrument Landing System.

Details of the operation of this display and of the second indicator involved in a Flight Director System will be given in Chapter 9. Electronic displays With the introduction of digital signal-processing technology into the field colloquially known as 'avionics', and its application of microelectronic circuit techniques, it became possible to make drastic changes to both quantitative and qualitative data display methods.

In fact, the stage has already been reached whereby many of the conventional 'clock' type instruments which, for so long, have performed a primary role in data display, can be replaced entirely by a microprocessing method of 'painting' equivalent data displays on the screens of cathode ray tube CRT display units. Table I. Liquid crystal Passive monitoring indicators; radio frequency selector indicators; distance measuring indicators; control display units of "menial navigation systems.

Electron CRT beam Active Weather radar indicators; display of navigational data; engine performance data: Display configurations Displays of the light-emitting diode and liquid crystal type are usually limited to applications in which a single register of alphanumeric values is required, and are based on what is termed a seven-segment matrix configuration or, in some cases, a dot matrix configuration. Examples of these alphanumeric displays are illustrated at b of Fig. In a dot matrix display the patterns generated for each individual character are made up of a specific number of illuminated dots arranged in columns and rows.

In the example shown at c of Fig. When current flows through the chip it emits light which is in direct proportion to the current flow. Light emission in different colours of the spectrum can, where required, be obtained by varying the proportions of the elements comprising the chip, and also by a technique of 'doping' with other elements, e. In a typical seven-segment display format it is usual to employ one LED per segment and mount it within a reflective cavity with a Figure 1.

Depending on the application and the number of digits comprising the appropriate quantitative display, independent digit packs may be used, or combined in a multiple digit display unit. LEDs can also be used in a dot-matrix configuration, and an example of this as applied to a type of engine speed indicator is shown in Fig.

Each dot making up the decimal numbers is an individual LED and they are arranged in a 9 x 5 matrix. The counter is of unique design in that its signal drive circuit causes an apparent 'rolling' of the digits which simulates the action of a mechanical drum-type counter as it responds to changes in engine speed. Liquid crystal display LCD. The basic structure of a seven-segment LCD is shown in Fig.

It consists of two glass plates coated on their inner surfaces with a thin film of transparent conducting material referred to as polarizing film such as indium oxide.

The material on the front plate is etched to form the seven segments, each of which forms an electrode. A mirror image is also etched into the oxide coating of the back glass plate, but this is not segmented since it constitutes a common return for all segments. The space between the plates is filled with a liquid crystal Figure 1.

Courtesy of Smith's Industries Ltd. Figure 1,16 Application of LCD. Magnitude of the optical change is basically a measure of the light reflected from, or transmitted through, the segment area to the light reflected from the background area.

Thus, unlike an LED, it does not emit light, but merely acts on light passing through it. Depending on polarizing film orientation, and also on whether the display is reflective or transmissive, the segments may appear dark on a light background as in the case of digital watches and pocket calculators or light on a dark background. An example of LCD application is shown in.

A head-up display HUD is one in which vital in-flight data are presented at the same level as a pilot's line of sight when he is viewing external references ahead of the aircraft, i. This display technique is one that has been in use for many years in military aviation, and in particular it has been essential for those aircraft designed for carrying out very high-speed low-level sorties over all kinds of terrain.

As far as civil aviation is concerned, HUD systems have been designed specifically for use in public transport category aircraft during the approach and landing phase of flight, but thus far it has been a matter of choice on the operators' part whether or not to install systems in their aircraft. This has resulted principally from the differing views held by operators, pilot representative groups, and aviation authorities on the benefits to be gained, notably in respect of a system's contribution to the landing of an aircraft, either automatically or manually, in low-visibility conditions.

The principle adopted in a HUD system is to display the required data on the face of a CRT and to project them through a collimating lens as a symbolic image on to a transparent reflector plate, such that the image is superimposed on a pilot's normal view, through the windscreen, of the terrain ahead.

The display is a combined alphanumeric and symbolic one, and since it is focused at infinity it permits simultaneous scanning of the 'outside world' and the display without refocusing the eyes.

The components of a typical system are shown in Fig. The first real attempt at establishing a standard method of grouping was the 'blind flying panel' or 'basic six' layout shown in Fig. As control of airspeed and altitude are directly related to attitude, the airspeed indicator, altimeter and vertical speed indicator flank the gyro horizon and support the interpretation of pitch attitude.

Changes in direction are initiated by banking an aircraft, and the degree of heading change is obtained from the direction indicator; this instrument therefore supports the interpretation of roll attitude Figure1. Boeing series aircraft. The turn-and-bank indicator serves as a secondary reference instrument for heading changes, so it too supports the interpretation of roll attitude.

With the development and introduction of new types of aircraft, and of more comprehensive display presentations afforded by the indicators of flight director systems, a review of the functions of certain of the instruments and their relative positions within the group resulted in the adoption of the 'basic T arrangement as the current standard As will be noted from diagram b of Fig.