avionics systems

BASIC

COMPLEMENTARY COURSE FOR AIRFRAME AND POWER PLANT

ENGINEERS

ELECTRICAL SYSTEMS

EGYPTAIR TRAINING CENTER

BASIC TECHNICAL TRAINING

© EgyptAir Training Center - 2015

BASIC

COMPLEMENTARY COURSE FOR AIRFRAME AND POWER PLANT

ENGINEERS

AVIONICS SYSTEMS

EGYPTAIR TRAINING CENTER

BASIC TECHNICAL TRAINING

© EgyptAir Training Center - 2015

ELECTRONIC FUNDAMENTALS DIODES

PAGE 1 of 22

BASIC COMPLEMENTARY COURSE FOR AF & PP

ENGINEERS

TECHNICAL TRAINING DEPARTMENT

1 DIODES Semi-conductor diodes embrace a very wide field of devices using varied modes of operation. Before discussing these, it is necessary to briefly describe semi-conductors themselves.

1.1 SEMI-CONDUCTORS Germanium and silicon are the most common semi-conductor elements. Figure 1 shows an element in pure crystalline form. The circles represent atoms and the dots valence electrons, electrons able to combine with those of another atom.

Silicon Structure

Figure 1

1.1.1 INTRINSIC SEMI-CONDUCTOR Note that one of the atoms has lost an electron, leaving a 'hole' but the free electron is still present inside the crystal lattice, so the crystal as a whole remains. A crystal of pure semi-conductor material with no other atoms, such as in Figure 1, is called an intrinsic semi-conductor.

4

4

4

4

4

4

4

4

4

4

4

4

HOLE

ELECTRON


PAGE 2 of 22


ENGINEERS


Figure 2 shows current flow in an intrinsic semi-conductor. The electrons (negative charge) are attracted to the positive terminal of the battery, while the holes (positive charge) are attracted to the negative.

Intrinsic Semiconductor Figure 2

1.1.2 EXTRINSIC SEMI-CONDUCTOR Intrinsic semi-conductors are poor conductors. By adding an impurity to the crystal, conductivity can be improved. Figure 3a shows an impurity having five electrons added. The 'extra electron' is not needed for crystal bonding and so is free to move about the lattice as a conduction electron. Since it is not a part of the lattice, it does not leave a 'hole' when it moves; but a 'positive ion'. The more impurity atoms added, the more conductive the material. The semi-conductor is now 'extrinsic' and of the 'N type'. Electrons are the majority carriers, they are negative, and hence 'N' type. Figure 3b shows a lattice with an element having only three valence electrons added. This time there is a shortage of electrons and this produces 'holes' in the material and negative ions. With fewer negative electrons, the majority carriers are positive 'holes'. Now the material is described as 'P' type. The impurity added to give more electrons to make N type material is known as a ‘donor impurity’. The impurity added to give more holes to make P type material is known as an ‘acceptor impurity’. The process of adding either type of impurity is known as doping.

SEMICONDUCTORMATERIAL

HOLESELECTRONS


PAGE 3 of 22


ENGINEERS


Extrinsic Semiconductor Figure 3

4

4

4

3

4

4

4

3

3

4

4

4

4

4

4

5

4

4

4

5

5

4

4

4

(a)

(b)

EXTRAELECTRON

DONORIMPURITY

ATOM

ACCEPTORIMPURITY

ATOM

HOLE


PAGE 4 of 22


ENGINEERS


1.2 THE HALL EFFECT When experimenting in 1879 with current flowing in a strip of metal, E M Hall discovered that some of the charge carriers were deflected to one of the faces of the conductor when a strong magnetic field was applied. This gave rise to an emf (the Hall voltage) between opposite faces of the conductor. The emf is only a few microvolts in the case of a metal conductor, but is much larger when the current flows in a semiconductor. An experiment, making use of what is known as the “Hall Effect”, can be conducted to demonstrate that the majority carriers in a bar of semiconductor material are electrons in “N” type and holes in “P” type. Figure 4 shows the Hall Effect

The Hall Effect

Figure 4

20V

+10V

+10V

SEMICONDUCTORMATERIAL

CURRENTFLOW

+20V

0V

0VP.D.

20V

+11V

+9V

+11V

+9V

+9V

+11V

POSITIVE CHARGE CARRIERS (HOLES)

NEGATIVE CHARGE CARRIERS (ELECTRONS)


PAGE 5 of 22


ENGINEERS


Consider the arrangement illustarted in figure 4a, this shows a bar of semiconductor material, with a D.C. voltage of 20V applied. Conventional current will flow as indicated by the arrow. A further two connections “A” & “B” are taken from opposite faces of the bar at the mid-point along the axis. Thus under static conditions, the voltgae at connect A and B will be +10V relative to the negative terminal, and there is no voltage difference between them, i.e. no potential difference. No consider what happens when we place this bar in a transverse magnetic field as in figure 4b. the charge carriers moving in the semiconductor are deflected by the magnetic field in the direction given by “Fleming’s Left-Hand rule”. Thus, whether the charge carriers are holes or electrons, they are deflected upwards in figure 4b, towards connection A. This will result in a redistribution of charge carriers between A & B, with the consentration towards A. If the charge carriers are positive (holes), connection A becomes positive with respect to connection B as shown in figure 4c. Conversely, if the charge carriers are negative (electrons), connection A becomes negative with respect to B as shown in figure 4c. The voltage difference between connection A & B is called the “Hall Voltage” and has many pratical applications such as “Contactless switches (proximity detectors). It can also be used in a dc starter/generator system as a means of measuring generator output current and providing an input signal to a Generator Control Unit (GCU) which controls generator field current (voltage regulation)m and protection. Figure 5 shows Hall Effect Sensors in a DC starter/generator system as fitted to the ATR 42/72 aircraft.


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ENGINEERS


Hall Effect Sensors Figure 5

GENERATORCONTROL UNIT

HALL EFFECTSENSOR

STARTERGENERATOR

HALL EFFECTSENSOR

CURRENTMEASURING

TODISTRIBUTION


PAGE 7 of 22


ENGINEERS


1.3 THE JUNCTION DIODE So far “N” type and P-Type materials have been considered separately. However, most semiconductor devices contain regions where P-type material is joined to N-type material at one or more places. These places are called P-N junctions and the behaviour of the devices depends upon the electrical behaviour of the region around the junctions. By doping a semi-conductor so that there is N type material at one end and P type at the other, a Junction Diode is made. Refer Figure 6. In this arrangement, the electrons in the N type are repelled by the like polarity of the negative ions in the P type. Similarly the positive holes in the P type are repelled by the positive ions in the N Type. This leaves an area at the junction without any majority carriers and it is called the depletion layer.

Junction Diode Figure 6

DEPLETIONLAYER

POSITIVE IONS NEGATIVE IONS

N-TYPE P-TYPEHOLESELECTRONS


PAGE 8 of 22


ENGINEERS


By connecting a battery across a junction diode, positive to N type, negative to P type, (reverse biased), majority carriers cannot flow, hence there is no current flow in the circuit. If the battery is connected positive to P type, negative to N type, (forward biased) majority carriers are allowed to flow and there is current flow in the circuit. This is the characteristic of the diode. It will allow current flow in one direction only, when forward biased, but not in the other direction when reverse biased. Figure 7 shows a junction diode reversed and forward biased.

Junction Diode Reversed/Forward Biased Figure 7


PAGE 9 of 22


ENGINEERS


1.4 DIODE SYMBOL Figure 8 demonstrates, using the circuit symbol for a diode, how the device is placed in a circuit to allow or block current flow. Note that (conventional) current flows in the direction of the arrow in the symbol.

Diode Symbol

Figure 8

1.5 DIODE CHARACTERISTICS With all diodes there are four parameters to be considered, these are:

1. Maximum permissible forward current (mA). 2. Maximum voltage drop (V) at nominal operating current (mA). 3. Typical reverse current (µA). 4. Maximum permissible reverse voltage (V).

ANODE CATHODE

+ _

NO CURRENT CURRENT FLOW

REVERSEDBIASED

FORWARDBIASED


PAGE 10 of 22


ENGINEERS


Figure 9 shows the static characteristics of a silicon diode and figure 10 show s the characteristics for a germanium diode. Note: That the reverse current axes on both graphs are different.

Silicon Diode Characteristics Figure 9

mA

VOLTS0.25V 0.5V 0.75V 1V-50V-100-150-200V

µA

50

200

150

100

-0.08

-0.02

-0.04

-0.06REVERSEDBIAS

FORWARDBIAS


PAGE 11 of 22


ENGINEERS


Germanium Diode Characteristics

Figure 10

1.6 DIODES IN SERIES AND PARALLEL Diodes may be connected in series or parallel. For carrying high voltage, a series configuration would be used. If a high current carrying capability were required, the diodes would be connected in parallel.

mA

VOLTS0.25V 0.5V 0.75V 1V-50V-100-150-200V

µA

50

200

150

100

200

50

100

150REVERSEDBIAS

FORWARDBIAS


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ENGINEERS


1.7 RECTIFIER DIODES

Rectifier diodes are designed to convert A.C. to D.C. and to be able to achieve this effectively and efficiently, they must have:

1. Low resistance to current flow in the forward direction.

2. High resistance to current flow in the

opposite (reverse) direction.


PAGE 13 of 22


ENGINEERS


Because of the need for a very low reverse current and a high breakdown voltage, almost all semiconductors rectifier diodes are silicon junction types; they usually have a junction area that is large relative to their size to assist in the dissipation of heat. An elementary rectifier circuit is where the diode is inserted in series between the input and output, this is shown in figure 11.

Basic Rectifier Circuit Figure 11

The diode effectively passes current only in the forward bias condition. As can be seen from figure 10, when A.C. input is applied, pulses of unidirectional D.C. voltages are developed across the output load resistance. Note; The polarity of the output D.C. can be reversed by reversing the diode connections.

+

0

-

A.C. INPUT D.C. OUTPUT

+

0

-

INPUT OUTPUT


PAGE 14 of 22


ENGINEERS


1.8 EXAMPLES OF RECTIFIER DIODES Silicon rectifier diodes are available that are capable of supplying currents from about 200mA to about 2000A at voltages up to 3000 or 4000 volts. A sample cross-section of such diodes is illustrated in Figure 12. Compared with other rectifying devices, silicon junction rectifiers are small and lightweight. They are also impervious to shock and are capable of working at temperatures up to about 200°C.

Silicon Rectifier Diodes

Figure 12

250mA @ 200V

1A @ 1000V

1A @ 1500V10A @ 400V

1000A @ 2500V


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ENGINEERS


1.9 RECTIFIER DIODES 1.9.1 SELENIUM RECTIFIERS The aluminium base serves as a surface for the dissipation of heat. The rectifying junction covers one side of the base apart from a narrow strip at the edges and an area around the fixing hole, which is sprayed with insulating varnish. Figure 13 shows the construction of a selenium rectifier element.

Selenium Rectifier

Figure 13 The counter electrode is a thin layer of a low melting point alloy, sprayed over the selenium coating and insulating varnish. The counter electrode is the cathode, while the base is the anode. These rectifiers may be stacked in series, suitable for high voltages, or in parallel, suitable for high current. When stacking, pressure applied during assembly tends to reduce the reverse resistance. This is overcome by application of varnish at the mounting studs. Reverse resistance is a limiting factor in rectifiers, as is temperature. The maximum operating temperature of these rectifiers is in the order of 70°C.


PAGE 16 of 22


ENGINEERS


1.9.2 SILICON RECTIFIERS The silicon rectifier is a far smaller unit than the selenium rectifier. This type of rectifier is used in the brushless ac generator. The silicon slice is extremely small. On one face it has a fused aluminium alloy contact to which the anode and lead are soldered. The other face is soldered to a base, usually copper. This is the cathode and acts as a heat sink. The aluminium - silicon junction forms the barrier layer. The whole is enclosed in a hermetically sealed case to protect it from environmental conditions. These rectifiers operate at temperatures up to 150°C. Figure 14 shows a Silicon Rectifier.

Silicon Rectifier Figure 14


PAGE 17 of 22


ENGINEERS


Figure 15 shows the circuit for a “Full-Wave bridge” rectifier.

Full-Wave Bridge rectifier

Figure 15


PAGE 18 of 22


ENGINEERS


1.10 THE LIGHT-EMITTING DIODE (LED) LEDs are made from a semiconductor material, which emits light when current flows through the junction. The most common colour emitted is red but green and yellow are available at a lower intensity. Figure 24 shows the circuit symbol for an LED and its operation.

Light Emitting Diode (LED) Figure 24

The voltage drop across a LED is around 2 volts. Above this voltage, the current passing through it increases rapidly. For this reason a series resistor is used to limit the current to around 10 ma to prevent burnout of the junction. 1.10.1 USE OF LEDS LEDs can be used to replace filament lamps, with the advantage of less current consumption, less heat and no filament to burn out. They are often found on aircraft fault panels.

+5VEARTH

OFF

DIODE IS REVERSED BIASED

+5V EARTH

DIODE IS FORWARD BIASED

EMITS LIGHT

ON

CIRCUIT SYMBOL


PAGE 19 of 22


ENGINEERS


1.11 THE PHOTO CONDUCTIVE DIODE This device is a normal PN junction with a transparent case or window. All semi-conductor diodes are subject to some movement of hole/electron pairs when the junction is at room temperature and this gives rise to a small leakage current, even with the diode reversed biased but the current is measured in microamperes. When light falls on the junction, its energy produces a much larger number of hole/electron pairs and the leakage current is greatly increased. These devices have a rapid response to light and are used in the encoding altimeter to encode the grey code into binary code. Figure 25 shows the circuit symbol and construction of a Photo Conductive Diode.

Photo Conductive Diode Figure 25


PAGE 20 of 22


ENGINEERS


1.12 VARISTORS The varistor is a semi-conductor device used for clipping 'noise spikes' off ac voltage. Noise spikes are of very short duration and large amplitude. They may pass through a power supply and appear on a dc regulated output voltage. Low pass filters are often ineffective against noise spikes so the spikes are attenuated before rectification of ac to dc.


PAGE 21 of 22


ENGINEERS


1.13 TESTING DIODES Before testing a diode, the cathode must be identified and then an ohmmeter is applied as in Figure 27. In one direction the ohmmeter reading should be low but a very high resistance should be detected in the other direction.

Testing Diodes Figure 27

FLUKE 23 SERIES MUL TIMETER

0 10 20 30

0 0 0 . 2 3O HM S

OFF VV

300 mV

Ω

AA

COM

VΩ10A

300mA

FUSED

!10 00V 75 0V

PRESSRANGE

AUT ORANGE

SYMBOL

PN

STRUCTURE

LOW RESISTANCE

CATHODE

ELECTRONIC FUNDAMENTALS TRANSISTORS PAGE 1 of 8



ENGINEERS

1 TRANSISTORS The transistor can be a high or low resistance device, hence the name, which is derived from TRANSfer resISTOR. It is used in many switching and amplifier circuits where its resistive properties are controlled by small currents. 1.1 TRANSISTOR CONSTRUCTION The properties of semi-conductor materials, P and N type, were discussed in Module 4.1.1. A transistor is made up of these materials in the configurations shown in Figure 1. The circuit symbols for these transistors are also shown.

PNP & NPN Transistors Figure 1

N

N

PBASE

COLLECTOR

EMITTER

B

C

E

CIRCUITSYMBOL

NOT

POINTING

IN

THE NPN TRANSISTOR P

P

NBASE

COLLECTOR

EMITTER

B

C

E

CIRCUITSYMBOL

THE PNP TRANSISTOR




ENGINEERS

As can be seen from figure 1, there are two possible types of physical arrangement: 1. The N-P-N transistor, which consists of a thin region of P-type material,

sandwiched between two N-type regions. 2. The P-N-P transistor, which consists of a thin region of N-type material,

sandwiched between two P-type regions. The centre region of the device is called the “Base”; one outer region is called the “Emitter”, and the other the “Collector”. Although the emitter and collector regions are the same type of extrinsic semiconductor (N-type in N-P-N and P-type in P-N-P), they are constructed and doped differently and are not interchangeable on a practical device. The circuit symbol for both P-N-P and N-P-N are shows in figure 1. The only difference between them is the direction of the arrowhead on the emitter lead. For either type, the arrowhead indicates the direction of “Conventional” current flow when the base/emitter junction is forward biased (i.e. base +ve with respect to emitter for an N-P-N device, and base –ve relative to emitter for a P-N-P device).




ENGINEERS

1.2 TRANSISTOR OPERATION Figure 2 shows a NPN transistor and the corresponding diode circuit. It can be seen from the diode circuit that to operate, the base/emitter must be forward biased, whereas the base/collector is reversed biased.

NPN Transistor & Diode Circuit

Figure 2

N - TYPE

N - TYPE

P - TYPE

DIODE MODEL

B

C

E




ENGINEERS

Figure 3 shows a simple transistor circuit using electron flow to explain the operation.

NPN Transistor Operation Figure 3

C

B

E

IE HIGH(100%)

IB LOW(1%)

IC HIGH(99%)




ENGINEERS

1.3 SWITCHING TRANSISTORS When a transistor is to be used as a switching device, it operates either as an open circuit (i.e. in the cut-off region) or as a short circuit (i.e. in the saturation region). Figure 3 shows the solenoid switch and an alternative transistor switch.

Switching Transistors

Figure 3

SOLENOID ANALOGY

LAMP

E C

B




ENGINEERS

For a common base circuit, such as in figure 3, the output voltage taken from the collector is either equal to the supply voltage (saturated region), or zero volts. (cut-off).




ENGINEERS

1.4 TRANSISTOR CONFIGURATIONS Before a transistor can be used, it must be connected into an input circuit (by two wires) and an output circuit (two wires). However, because the transistor has only three terminals, one of the terminals must be in both the input and output circuits; this is then called “The Common terminal”. Three circuit configurations are possible and are illustrated in figure 9.

Transistor Configurations Figure 9

Note that the word ‘common’ refers to the transistor component connected to both the INPUT and OUTPUT. In the common emitter configuration for example, the emitter is connected to both the input and output.

INPUT

OUTPUT

COMMON EMITTER

B

C

E INPUT OUTPUTB

CE

COMMON BASE

INPUT

OUTPUTB

C

E

COMMON COLLECTOR




ENGINEERS

Table 1 shows the comparisons of the three transistor configurations

Common Emitter

Common

Base

Common Collector

Current Gain

20 -200

(0.95 – 0.995)

20 - 200

Voltage Gain

100 – 600

500 – 800

<1

Power Gain

Highest

Medium

Lowest

Input Impedance

500 - 2000Ω

50 - 200Ω

20kΩ - 100kΩ

Output Impedance

10 – 50 KΩ

100 kΩ - 1MΩ

20 – 500Ω

I/P – O/P Phase

180°

In Phase

In Phase

Typical Use

Normal Amp

Impedance matching (low to high)

Impedance matching (high to low)

Table 1

ELECTRONIC FUNDAMENTALS INTEGRATED CIRCUITS PAGE 1 of 19


ENGINEERS


1 INTEGRATED CIRCUITS

1.1 GENERAL Integrated circuits, or IC’s, have changed the entire electronics industry. Before IC’s were developed, all electronic circuits consisted of individual (discrete), components that were wired together, often requiring a large amount of physical space. Printed circuit Board (PCB) technology made it possible to reduce the amount of space required. Electronic circuits can be quite complex, requiring a large number of components, since discrete components have a fixed size, there is a practical limitation on the amount of size reduction that can be achieved. The development of integrated circuit technology has made it possible to fabricate large numbers of electronic components onto a single silicon chip. As a result, the physical size of a circuit can be significantly reduced, making it possible to design circuits and devices that would otherwise be impractical. IC’s are complete circuits containing many transistors, diodes, resistors and capacitors as may be necessary for the circuit operation. They are encapsulated in packages that are often no larger than a single discrete transistor. The technology and materials used in the manufacture of IC’s are basically the same as those used in the manufacture of transistors and other solid-state devices. In addition, IC’s are manufactured for a wide variety of applications and, as a result, are used throughout the electronics industry.

1.1.1 ADVANTAGES The small size of the IC is its most apparent advantage. A typical IC can be constructed on a piece of semiconductor material that is less than 4mm2. Even when the IC is suitably packaged, it still occupies only a small amount of space. The small size of the IC also produces other benefits such as they consume less power than the equivalent conventional circuit. They generate less heat and therefore generally do not require elaborate cooling or ventilation systems. IC’s are also more reliable than conventional circuits. This greater reliability result because every component within the IC is a solid-state device and is permanently connected together with a thin layer of metal. They are not soldered together like the components in a conventional circuit and a circuit failure due to faulty connections is less likely to occur.

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ENGINEERS


1.1.2 DISADVANTAGES It might appear that the IC has only advantages to offer and no real disadvantages. Unfortunately, this is not the case, since IC’s are an extremely small device it cannot handle large currents or voltages. High currents generate heat within the device and small components can be easily damaged if the heat becomes excessive. High voltages can break down the insulation between the components in the IC because the components are very close together. This can result in shorts between the adjacent components, which would make the IC completely useless. Therefore, most IC’s are low power devices, which have a low operating current (milliamps) and low voltages (5 – 20V). Also, most IC’s have a power dissipation range of less than 1 watt. At the present only four types of component are commonly constructed within an IC. This makes only a narrow selection of components available, these are:

1. Diode. 2. Transistor. 3. Resistor. 4. Capacitor.

Diodes and transistors are the easiest components to construct and are used extensively to perform as many functions as possible within each IC. Resistors and capacitors may also be formed, but it is much more difficult and expensive to construct these components. The amount of space occupied by a resistor increases with its value and in order to conserve space, it is necessary to use resistors with values as low as possible. Capacitors occupy even more space than resistors and the amount of space required increases with the value of the capacitor. Ic’s cannot be repaired because their internal components cannot be seperated. When one internal component becomes defective, the whole IC becomes defective and musty be replaced. This means that good components are often thrown away with the defective ones. This disadvantage is not as bad as it sounds, as the task of fault finding is simplified because it is only necessary to trace the problem to a specific circuit instead of an individual component. This greatly simplifies the task of maintaining highly complex systems and reduces the demands on maintenance personnel.

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ENGINEERS


1.2 IC CONSTRUCTION There are basically four methods of construction used for IC’s. These are:

1. Monolithic. 2. Thin-Film. 3. Thick Film. 4. Hybrid.

1.2.1 MONOLITHIC IC’S The monolithic IC is constructed in basically the same manner as a “Bipolar Transistor”, although the overall process requires a few additional steps because of the greater complexity of the IC. Its fabrication begins with a circular semiconductor wafer (usually silicon). This wafer is usually very thin (0.015mm – 0.3mm) and either 2.5cm or 5cm in diameter. The semiconductor serves as a base on which the tiny integrated circuits are formed and is commonly referred to as a “Substrate”. Figure 1 shows the IC construction.

IC Construction

Figure 1

2.5 - 5 CM DIAMETER

0.015 - 0.30mm

SILICON WAFER

IC’S ARE FORMEDON THE WAFER

NUMBER OF IC’S FORMEDDEPENDS ON THE SIZE

OF THE WAFER

ONCE THE IC’S HAVEBEEN FORMED, THE

WAFER IS SLICED INTOINDIVIDUALCHIPS



ENGINEERS


When all of the IC’s have been simultaneously formed, the wafer is sliced into many sections, which are commonly referred to as “Chips” or “Dice”. Each chip represents one complete integrated circuit and contains all the components and wiring associated with that circuit. Once the IC’s have been separated into individual chips, each IC must be mounted in a suitable package and tested.

1.2.2 BIPOLAR IC CONSTRUCTION As mentioned earlier, the components that are commonly used in IC’s are diodes, Transistors, resistors and capacitors. Diffusing impurities into selected regions of a semiconductor wafer (substrate) can form these components. This process produces PN junctions at specific locations and the basic manner in which these four components are formed and the manner in which they are interconnected are shown at Figure 2.

Basic Construction of Bipolar IC Figure 2



ENGINEERS


The circuit shown in figure 2 is a simple circuit consisting of a capacitor, a PN junction diode, an NPN transistor and a resistor. Operating voltages and currents can be applied to the circuit through terminals 1,2 and 3 as shown. This circuit could be easily constructed using four discrete components, however, it can also be produced as a monolithic IC.

1.2.3 MOS IC’S Not all IC’s are constructed using bipolar components, IC’s are often designed to utilize either bipolar transistors or “Field-Effect transistors” (FETS). The Field effect transistor is one in which the emitter-collector current is controlled by voltage rather than by a current. Figure 3 shows the construction and operation of a MOSFET.

MOSFET Figure 3

The FET may be constructed of a channel of either N-type or P-type silicon with a controlling gate sitting on top. One end of the channel is called the source, and the other end is called the drain. An N-channel FET has a P-type gate, so that when a positive voltage ios applied to the gate, the FET is forward biased. There will be current flow between the source and the drain. When a negative voltage is applied to the gate, the FET will be reversed biased, and the flow between the source and the drain will be pinched off.



ENGINEERS


The source and drain regions are diffused into the substrate. A thin layer of silicon oxide is formed over the substarte and the appropriate windows are cut into it so that metal electrodes ) terminals) can be formed at the proper locations. Note that the gate terminal is separated from the substrate by an extremely thin oxide layer, which is only 1 X 10-10 metres thick, but it completely isolates the gate from the substrate.

1.2.4 THIN-FILM IC Unlike the monolithic IC’s, which are formed within a semiconductor material (substrate), the thin-film circuit is formed on the surface of an insulating substrate. In the thin-film circuit, components such as resistors and capacitors are formed from extremely thin layers of metals and oxides, which are deposited onto a glass or ceramic substrate. Interconnecting wires are also deposited on the substrate as thin strips of metal. Components such as diodes and transistors are formed as separate semiconductor devices and then permanently attached to the substrate at the appropriate locations. The substrate on which the thin-film circuit is formed is usually less than 2.5cm2. Depositing tantalum or nichrome as thin films or strips on the surface of the substrate forms the resistors. These films are usually less than 0.00254cm thick. The thickness, length and width of each strip that is formed on the substrate determine the value of each resistor. The interconnecting conductors are extremely thin metal strips, which have been deposited on the substrate. Low resistance metals, such as gold. platinum, or aluminium, are generally used as conductors. The substrate is made from an insulating material that will provide a rigid support for the components. Glass or ceramic materials are often used as substrates. Figure 4 shows a portion of a thin-film circuit.

Thin-Film IC

Figure 4

INSULATINGSUBSTRATE

THIN-FILMCONDUCTORS

THIN-FILMRESISTORS



ENGINEERS


1.2.5 THICK-FILM IC’S Thick-film IC’s components are formed on an insulating substrate by using a “Silk-screen” process. In this process, a very fine wire screen is placed over the substrate and a metalized-ink is forced through the screen using a squeegee. Only certain portions of the wire screen are open (the remaining portions are filled with a special emulsion), thus allowing the ink to penetrate and coat the specific portions of the substrate. A pattern of interconnecting conductors is formed on the substrate, which is then heated to over 6000°C to harden the painted surface and become low resistance conductors. Resistors and capacitors are also silk-screened on top of the substrate by forcing the appropriate materials (in paste form) through the appropriate screen and then heating the substrate to a high temperature. This process is repeated using various pastes until the circuit is formed. Components such as diodes and transistors are formed as separate semiconductor devices and then permanently attached to the substrate at the appropriate locations.

1.2.6 HYBRID IC’S Hybrid IC’s are formed by utilizing various combinations of monolithic, thin-film and thick film techniques and may in certain circumstances contain discrete semiconductor components in chip form. Therefore many types of hybrid circuit arrangements can be produced. A typical hybrid circuit might consist of a thin-film circuit on which various monolithic IC’s have been attached or it could utilize monolithic IC’s thick-film components and discrete diodes and transistors that are all mounted on a single insulating substrate. A portion of a hybrid IC is shown at figure 5. An insulated substrate is used to support the circuit components as shown. The monolithic IC is mounted on the substrate along with thich-film resistors and a small discrete capacitor. All the components are interconnected with conductors that are formed on the substrate using film techniques. The monolithic IC is connected to the conductors with fine wires that are bonded in place. Thick-film resistors will usually have notches cut into them to trim their values. The capacitor used in these circuits can be formed either by using film techniques or miniature devices can be installed between conductors as shown.



ENGINEERS


Hybrid IC Construction

Figure 5

1.2.7 IC PACKAGES Like transistors and other types of solid state components, IC’s are mounted in packages, which protect them from moisture, dust and other types of contaminations. Many different types of IC packages are available and each type has its own advantges and disadvantages. The most popular IC package is the “Dual In-Line (DIL) package. The packages also make it easier to install the IC’s in various types of equipment, since each package contains leads which can be either plugged into matching sockets or plugged into DIL mounting frames.



ENGINEERS


Figure 6 shows typical DIL packages.

DIL Packages

Figure 6

The IC package shown in figure 6 contains three monolithic IC’s, also a network of conductors have been formed on the same base that supports the chip. Various conductor pads on the chips are connected to these conductors with fine gold wires that have been bonded in place. The conductors in turn are connected to two rows of connecting pins along the edge of the package. A lid or cover (not shown) is placed over the opening in the package and soldered into place to provide an air tight (hermetically sealed) unit. Integrated circuits may also be mounted in “Metal cans” that are similar to the types used to house transistors. The metal can have 8 or more connecting leads and can used to house either monolithic or hybrid type IC’s. The advantage of these packages is that they may be installed in a variety of ways. Metal cans can be used over a wide temperature range (-55° - +125°C) and are therefore suitable for military and space applications. Figure 7 shows the DIL and metal can type of packages.

MONOLITHICIC’s

CONNECTING PINS

INTERCONNECTINGCONDUCTORS



ENGINEERS


DIL and Metal Can Packages

Figure 7

1.3 TYPES OF INTEGRATED CIRCUIT Integrated circuits are placed into two general groups, these are:

1. Digital IC’s.

2. Linear IC’s.

1.4 DIGITAL IC’S Digital circuits use discrete values (0 or 1) to perform 3 general functions. These are:

1. AND Function.

2. OR Function.

TYPICAL MINIATUREDUAL IN-LINE (DIL)

PACKAGES

TYPICAL METAL CANIC PACKAGES



ENGINEERS


3. NOT Function.

Thgese three function are performed by logic circuits that are called the AND, OR and NOT logic gates. These gates or circuit configurations can be combined to make decision based on digital input information. In a digital logic gate it is only possible to have an output of either a 0 or 1.

1.4.1 AND GATE Figure 8 shows the AND gate truth table and logic circuit and a corresponding circuit to carry out this function.

AND Gate Figure 8

The AND gate has an output of 1 only when all of its inputs are equal to 1. This is similar to a multiplier function since the only possibilities in a digital circuit are 0 X 1 = 0 and 1 X 1 = 1. The schematic circuit in figure 8 shows two switches connected in series. Unless both switches are closed, there is no current flow to the output.

A

BA.B

SYMBOL

A B A.B

TRUTH TABLESCHEMATIC DIAGRAM



ENGINEERS


1.4.2 OR GATE Figure 9 shows the OR gate truth table and logic circuit and a corresponding circuit to carry out this function.

OR Gate Figure 9

1.4.3 NOT GATE The NOT gate provides an output that is always the opposite the input. This is called inversion or 180° phase shift. Thus, the NOT gate is commonly referred to as an inverter. In the bipolar transistor, the common emitter amplifier configuration was the only one capable of inverting the input so is used to carry out the NOT function.

A

BA+B

SYMBOL

A B A+B

TRUTH TABLE SCHEMATIC DIAGRAM



ENGINEERS


Figure 10 shows the NOT gate truth table and logic circuit and a corresponding circuit to carry out this function.

NOT Gate Figure 10

1.4.4 COMBINATION LOGIC CIRCUITS The three basic logic circuits can be combined into a single decision making circuit with more than 1 distinct outputs. Consider a circuit that compares two inputs and calculates three outputs as shown below.

Output X1

Input A < Input B

Output X2

Input A > Input B

Output X3

Input A = Input B

A AA A1 00 1

INPUT

OUTPUT

+VE

SYMBOLTRUTH TABLE

SCHEMATIC DIAGRAM



ENGINEERS


A combined logic circuit that would carry out the function is shown at Figure

11.

Combination Logic Circuit

Figure 11

A

BX1 (A<B)

X2 (A>B)

X3 (A=B)

AA BB X1X1 X2X2 X3X3

TRUTH TABLE



ENGINEERS


1.5 LINEAR (OR ANALOGUE) IC Figure 12 shows the type of analogue signal handled by the Linear Integrated Circuit.

Analogue Signal

Figure 1 1.6 THE OPERATIONAL AMPLIFIER (OP AMP) The integrated circuit operational amplifier is one of the most useful and versatile electronic devices available today. The name ‘operational amplifier’ is not new; it refers to a type of amplifier originally used in analogue computing to perform mathematical operations – e.g. multiplication or division by a constant. The modern integrated circuit device can be adapted (by feedback) to perform most general-purpose amplifier duties, as well as its use in mathematical operations.

0 TIME



ENGINEERS


The Op Amp can consist of many stages of amplification to ensure high gain, and will be arranged to have two input terminals, two power supply terminals and an output terminal. In addition it will normally have terminals for setting the output to zero when the input is zero. The Op Amp consists of a transistor circuit of considerable complexity, which has been found so useful that the whole circuit is manufactured on a single piece of silicon, fitted with input and output leads, and covered in plastic. It is the first “Integrated Circuit”, and can be treated just as if it were a new component. Figure 2 shows a type 741 Op Amp and circuit.

Op Amp and Circuit Figure 2

1

2

35

4

78

6

INVERTINGINPUT

NON-INVERTINGINPUT

VOLTAGEOUTPUT

POWERSUPPLY

(+)

POWERSUPPLY

(–)

14 3 2

5 6 7 8

V–

V+

NON-INVERTINGINPUT

INVERTINGINPUT

VOLTAGEOUTPUT

GROUND



ENGINEERS


In the Op Amp, two pins are marked supply + and supply - and are connected to the amplifiers power supply. The device also has two inputs, the “Inverting input” (VΙ) identified by a negative symbol. A “Non inverting input” (VN) identified by a positive sign and a single output (VO). Note: The negative/Positive signs on the inputs does not mean that negative/positive signals are applied, but identify the inverting and non-inverting terminals. The VΙ, VN and VO are the values of the voltages applied to the inputs and obtained form the output. These voltages are joined by the equation:

VO = AO (VN – VΙ) Here we have a slight problem. Voltages are measured between one point in a circuit and another. Usually one point is the negative or zero line. When calculating VN & VΙ it does not matter were the reference is as long as it is the same for both voltages. When we obtain the output VO we need to know the reference point used by the Op Amp. This is not the zero line but a voltage halfway between the positive supply and the zero line. The other unknown quantity in the equation is AO, the “Open Loop Gain”. This gain is constant for each particular Op Amp and is the ratio between two voltages. Open Loop gain in Op Amps is normally 105.



ENGINEERS


The following example will make use of the equation. Figure 3 shows an Op Amp with an open loop voltage gain of 400, connected between a 12V supply.

Op Amp Figure 3

VΙ = 5.88V VN = 5.87 AO = 400

Using the equation:

VO = AO(VN - VΙ) VO = 400(5.87 – 5.88)

= 400(-0.01)

= -4V

The voltage is relative to a point halfway between +12v and zero, that is 6V. The output voltage is therefore 4V below 6V, i.e. 2V. What would the output be if the input values were reversed? Ans:……………………….

VOUT

5.87V

5.88V

+12V

GAIN = 400

ZERO LINE



ENGINEERS


1.7 THE IDEAL OPERATIONAL AMPLIFIER Although the characteristics of an ideal operational amplifier are unattainable, modern integrated circuit types can provide an extremely close approximation. The ideal characteristics are: * A very large open loop gain, near infinite, * Output unaffected by signal frequency, no signal phase shift with

change in frequency, * A very large (infinite) input impedance so that the amplifier takes

negligible current, * A very small output impedance so that the output of the amplifier is

unaffected by loading, * Output voltage is zero for zero input voltage (offset zero applied). Naturally, no practical operational amplifier will be this perfect, which means of course that there will be small operational errors with such devices. Therefore, the closer to the ideal properties the amplifier is made, the smaller will be these errors.

ELECTRONIC FUNDAMENTALS PRINTED CIRCUITS BOARD PAGE 1 of 3


ENGINEERS


1 PRINTED CIRCUIT BOARDS Aircraft electronic systems necessitate the interconnection of many components; in the past this was done by soldered or crimped terminations. With the development of circuit technology and micro miniaturisation, weight saving and simplification of installation and maintenance became needful and these needs were met by the development of the printed circuit board. 1.1 CONSTRUCTION Printed circuit board is a laminated paper or fibreglass board coated on one side with a thin layer of copper. The areas of copper, called 'lands', required to connect the components are marked out by painting over the copper, and the remaining copper is etched away by a solution of ferric chloride. Holes are then drilled in the board for the component leads. The advantage is that the copper strips can be any shape and few additional wires are required. Industry can produce printed circuit boards in large numbers very cheaply so they have become the standard circuit construction method. Figure 1 shows the front face of a PCB, with Figure 2 showing the rear face.

Printed Circuit Board Figure 1

BASEBOARD

FRONT



ENGINEERS


Printed Circuit Board Figure 2

IC3 IC4

IC6

C2

IC1 IC2 IC5

REAR

CIRCUIT MODULEDESIGNATION(E.G. SIGNAL SELECTOR)

CIRCUITREFERENCE

FINGER OREDGE CONNECTOR

INTEGRATEDCIRCUIT CHIPS



ENGINEERS


1.2 MULTI-LAYER CIRCUITS In order to save weight and space, and to provide for the interconnection of integrated circuits (which are a feature of a large majority of electronic equipment) the relevant circuits are assembled as a multi-layer moulded package. This consists of three or more single and/or double-sided printed boards and insulating layers of ‘impreg’ material. 1.3 HANDLING PRINTED CIRCUIT BOARDS Since various types of semi-conductor components are mounted on printed circuit boards, care must always be taken in handling techniques. General techniques are as follows: - a) Do not remove or replace units with electrical power applied. b) Do not touch the connectors, leads or edge connectors of circuit

boards unnecessarily. c) Use conductive packaging, shorting plugs, bands or wire when

provided or prescribed by the relevant aircraft Maintenance Manual. d) Pay strict attention to stores procedures to ensure that protective

packaging is not removed during any goods-inwards inspection. Module 5 details procedures for handling “Static Sensitive Devices”.

ELECTRONIC FUNDAMENTALS SERVOMECHANISM PAGE 1 of 23


ENGINEERS


1 SERVOMECHANISMS A servomechanism (servo) is a type of control system whose output is the position of a shaft. They may be controlled remotely when used in conjunction with synchro devices. Synchros themselves transmit position information but cannot amplify torque to move heavy loads. Used with servomechanisms, an output to control such a load can be obtained to give a desired result in relation to an input. 1.1 OPEN LOOP SYSTEM In this system, an input is applied and an output obtained. Figure 1 shows an example; assume an aircraft rudder controlled by an open loop system.

Open Loop System Figure 1

The demand, made by the pilot on the rudder bar, is picked up by the transducer which converts it to an electrical signal; i.e. the demand signal. This signal is amplified and fed to the motor, which responds by moving the load; i.e. the rudder. There is no positional feedback and the pilot does not know if the rudder has adopted the position requested.

INPUTTRANSDUCER

MOTOR LOAD

DEMAND

RESPONSE

DEMANDSIGNAL

AMP



ENGINEERS


1.2 CLOSED LOOP SYSTEM In the closed loop system, the demand is made in the same way. In a basic system, positional feedback would be given to the pilot who would make adjustments accordingly but this is not practical with systems such as aircraft flying controls. Figure 2 shows a closed loop automatic system.

Closed Loop System Figure 2

An output position transducer has been added to the servomotor and this feeds back any difference between input demand and output to an error detector. The error detector outputs an error signal to the amplifier to make any positional corrections necessary at the servo motor and thus the load (or rudder) is positioned as demanded. If for example the pilot wanted to move the rudder 5°, a demand is made at the rudder bar and this is converted to a voltage at the transducer, say +5 volts. The error detector immediately gives an output signal corresponding to +5 volts input and this is amplified to drive the motor, moving the rudder. The output position transducer converts the output position to an electrical signal, which corresponds to the new position of the rudder. As this happens, this signal, (feedback), is fed back to the error detector until the demanded position is achieved and the input is negated. Now, there is no error signal and no output. The feedback has reached -5 volts.

INPUTTRANSDUCER

SERVOMOTOR LOAD

OUTPUTPOSITION

TRANSDUCER

ERRORDETECTOR

POSITIONFEEDBACK

ERRORSIGNAL

AMP



ENGINEERS


1.3 FOLLOW UP If in our example the rudder were to be displaced from its demanded position, or from the optimum speed at which the demanded position may be achieved, an error signal occurs. In the way described, there is a feedback signal and the system returns to its demanded position or speed. This process is called 'follow up'.



ENGINEERS


1.4 FEEDBACK 1.4.1 POSITIONAL FEEDBACK Positional feedback is obtained from transducers positioned at the output. The feedback element, or transducer, converts the output shaft angle into a signal suitable for operating the error detector. In this case a voltage signal. The simplest form of element is a R-pot, or a helical potentiometer similar to that used as a control element. In practice, helical potentiometers are used since they give 360° coverage, which a R-pot cannot provide. Figure 3 shows positional feedback in a dc system.

Positional Feedback Figure 3

ERRORDETECTOR SERVO

MOTOR

LOAD

TACHOGEN

FEEDBACKELEMENT

POSITIONALFEEDBACK

VELOCITYFEEDBACK

CONTROLELEMENT



ENGINEERS


Figure 4 shows a R-Pot & Helical Potentiometer

R-Pot & Helical Potentiometer Figure 4

In ac systems, other components are used to provide positional feedback. Synchros are employed in some servomechanisms. These will be discussed later.

E

θ iPROPORTIONAL

TO θ i

Ei

R-POT

HELICAL POTENTIOMETER

E

θ i

Ei

PROPORTIONAL

TO θ i



ENGINEERS


1.5 ROTARY VARIABLE DIFFERENTIAL TRANSDUCER (RVDT) The RVDT is an inductance transmitter having a primary stator coil, an iron rotor coil and two secondary stator coils. Figure 5 shows the operation of a RVDT.

RVDT Operation Figure 5

The mechanical input changes the position of the iron core. The position of the core changes the magnetic coupling between the primary and the secondary stator coils. When the input rotates, one of the secondary coils receives more magnetic flux and this induces a higher voltage in that coil.

R TS

L3

L1 L2

IRON CORECONNECTED TO

MECHANICALINPUT

PRIMARYCOIL

R TS

R TS

1. ZERO POSITION 2. ROTATED CLOCKWISE

3. ROTATED COUNTER CLOCKWISE



ENGINEERS


The other secondary coil receives less magnetic flux, so a lower voltage is induced. The difference between voltages induced in the secondary stator coils is proportional to the rotated angle. This is an AC Ratio Signal. Figure 5.1: The position of the iron core is zero. The magnetic field induced

by primary coil L3 is equally divided between L1 and L2. Therefore the voltage R-T is zero.

Figure 5.2: The iron core is turned clockwise. Now there is more coupling

between L3 and L2, and less coupling between L3 and L1. The voltage between T and S increases and the voltage between R and S decreases.

Figure 5.3: The iron core turned counter-clockwise. Now there is more

coupling between L3 and L1, and less coupling between L3 and L2. The voltage between T and S decreases, while the voltage between R and S increases.

The difference between figure 5.2 and 5.3 is that the output-voltage between R and T is of opposite phase. The output measured between R and T is an AC RATIO signal. The Linear Variable Differential Transducer (LVDT) is also an inductance transmitter with similar components and similar in operation but of course, the movement detected is linear and not rotary.



ENGINEERS


1.6 CAPACITANCE TRANSMITTER An example of a capacitance transmitter can be seen in a simple fuel gauging system as in Figure 6.

Capacitance Transmitter

Figure 6 This system depends upon the comparison of two capacitance values. One in Loop A, which is the variable capacitance of a tank unit and the other in Loop B, which is fixed. A current is developed in each loop; IS in loop A; IB in loop B. The two loops form a bridge with resistor R across it. If the tank is full, then current IS is the greater. With the tank empty, IS falls so that IB is the greater. Note: The currents act in opposite directions so that a potential is developed across resistor R of a polarity dependent on the direction of current flow and of a magnitude dependent on the size of the current. This signal is transmitted to an amplifier, which powers a 2-phase motor to drive an indicator and a balance potentiometer.

TANK UNIT

FULL

EMPTY

AMPLIFIERSTAGE

REF C

2 - PHASEMOTOR

AMPLIFIER UNITREF

PHASE

INDICATOR

LOOPA

LOOPB

IS

IB

DISCRIMINATIONSTAGE



ENGINEERS


When the balance potentiometer moves as a result of change in fuel level, it adjusts IB, rebalancing the bridge formed by loop A and loop B. Now, no current flows through resistor R, no signal is developed across R and the new fuel level is displayed at the indicator.

1.7 SYNCHROS

1.7.1 INTRODUCTION AC transmission systems are generally known as synchros because of their synchronous action in reproducing the angular movement of a shaft. As mentioned previously, they cannot transmit torque to any appreciable degree but can be used in conjunction with servomechanisms. 1.8 TORQUE SYNCHRO 1.8.1 PRINCIPLE OF OPERATION The principle of a synchro is that of the transformer, where the primary winding is wound onto a rotor and is rotated with respect to a fixed stator winding. The size and phase of the output voltage is dependent on the direction and angular displacement between the primary and secondary windings. The torque synchro comprises two electrically similar units: the transmitter (TX) and the receiver (TR) which are interconnected by transmission lines. The TX and TR have very similar construction. Each has a rotor carrying a single winding concentrically mounted in a stator of three windings, the axes of which are 120° apart. It should be noted that the TX and TR torque synchros are not identical. The difference is that the TR synchro has an oscillation damper added, so that when its rotor rotates to a given position, it does not oscillate as it comes to rest. The rotors of both TX and TR synchros are energized from the ac supply and produce an alternating flux which links with their corresponding stators S1, S2 and S3. This process is the normal transformer action, with the rotors corresponding to the transformer primary winding and the stators to the secondary windings. Consider the case when the two rotors are not aligned. The three voltages induced in each of the two sets of stator windings are different. Currents therefore flow between the two stators and a torque is produced in each synchro which is directed in such a way that the two rotors must align themselves. Normally, the TX rotor position is controlled by the input shaft, while the TR rotor is free to turn, so it is the one which aligns itself with the TX rotor. In this way, any movement of the TX rotor due to movement of the input shaft is repeated synchronously by movement of the receiver rotor.



ENGINEERS


Torque synchros are used for the transmission of angular position information and flight instrument systems is a typical application. Figure 9 shows a Torque Synchro and circuit symbol.

Torque Synchro Figure 9

ROTORFIELD

CURRENTFLOW

STATORFIELD

S1

S2

S3

R1

R2

INPUTSHAFT

OUTPUTSHAFT

S1 S1

S3 S3S2 S2

CIRCUIT SYMBOL



ENGINEERS


Figure 10 shows the construction of a torque synchro.

Torque Synchro Construction Figure 10

STATOR ROTOR COMPLETEASEMBLY

STATORWINDINGS

SHELL

LOWER ENDCAP

SHAFTBEARING

COILS

CORE

LEADS TOSLIP RINGS

SLIPRINGS

STATORLEADS

ROTORLEADS



ENGINEERS


1.9 CONTROL SYNCHRO The basic control synchro system has two units; a synchro control transmitter (CX) and a synchro control transformer (CT) connected as shown in Figure

11.

Control Synchro Figure 11

1.9.1 PRINCIPLE OF OPERATION The CX synchro is similar to that used in the torque synchro system. The control transformer has a stator, which in design and appearance resemble the synchro units already discussed but with high impedance coils to limit the alternating currents through the coils. Further differences in the CT are that the rotor winding has its coils wound so that no torque is produced between it and the stator magnetic fields and the rotor is not energized by the supply voltage applied to the rotor of the control synchro. The CT rotor acts as an inductive winding for determining the phase and magnitude of error signal voltages. The signals, after amplification, are fed to a two-phase motor, which is mechanically coupled to the CT rotor. A control synchro system is at electrical zero when the rotor of the CT is at 90° with respect to the CX rotor. This is the situation as shown in Figure 10 above.

INPUTSHAFT

S1 S1

S3 S3S2 S2

A.C.SUPPLY

M

SERVOMOTOR

A.C.SUPPLY

CX CT



ENGINEERS


If the input shaft is rotated and the CX rotor is disturbed, voltages are induced in the CX stator and currents flow down the transmission lines to the stator windings S1, S2 and S3 of the CT. A magnetic flux is produced, depending on the amount of displacement of the CX rotor and the orientation of its displacement. This flux links with the rotor of CT, inducing a voltage into it, again depending on the amount, or rate of displacement, and its orientation. The voltage, or error voltage, representing the electrical difference between the rotors of CX and CT, is then amplified and passed to the control phase of a two-phase motor. The ac reference phase supply is fixed. The motor now rotates. Its direction depends on the phase of the error signal, as can be seen from Figure 12.

Phase Error Signal

Figure 12 As it rotates, the motor drives the rotor of CT in such a direction as to reduce the error voltage to zero and the new position is reached. By using the error signal amplified by a servo amplifier, a servomotor can be driven to move a control surface as in Figure 11.

APPLIED VOLTAGE

CLOCKWISE ROTATIONVOLTAGE IN-PHASE

ANTI-CLOCKWISE ROTATIONVOLTAGE OUT-OF-PHASE



ENGINEERS


1.10 DIFFERENTIAL SYNCHRO There are two types of differential synchro system:

♦ Torque.

♦ Control. In each, a special type of synchro is inserted between the synchros of the basic torque or control systems. It is called a ‘differential synchro’ and differs from the basic synchros in that it has a three-phase stator and rotor. In a torque differential system it is abbreviated to TDX and in a control differential system, CDX. The inclusion of this synchro between a torque transmitter and receiver or control transmitter and transformer permits an additional input to be algebraically added to, or subtracted from, the system. The layout of a differential synchro and its circuit symbol are shown at Figure 13.

Differential Synchro

Figure 13

S2

S1 S3

R2

R1 R3

ROTOR

STATOR

CIRCUIT SYMBOL

S1

S2

S3

R1

R2

R3



ENGINEERS


Figure 14 shows the construction of a differential synchro

Differential Synchro Construction

Figure 14

STATORASSEMBLY

ROTORASSEMBLY

SKEW CUT TOENABLE SMOOTHER

RUNNING

STATORCONNECTIONS

STATORWINDINGS

ROTORCOILS



ENGINEERS


1.11 TORQUE DIFFERENTIAL SYNCHRO Figure 15 shows a differential synchro system set up for the SUBTRACTION of two inputs.

Torque Differential Synchro

Figure 15 Note that the rotors of the normal transmitter TX and receiver TR are supplied in parallel with the single-phase ac supply. The stator windings of the TX are connected to the stator windings of the TDX and its three rotor windings are connected to the three-stator windings of the TR. The rotor of the TDX is not energized by the ac supply. The circuit is such that one input shaft turns the TX rotor and the second input shaft drives the TDX rotor. The TDX receives an electrical signal corresponding to a particular angular position of the TX rotor, which it modifies by an amount corresponding to the angular position of its own rotor. This modified signal appears at the TDX output and is transmitted to the receiver, where it produces an angular flux, which is the difference of the rotor angles of the two transmitters TX and TDX. If the TDX rotor is locked in one position, the TX/TR chain acts as a normal torque synchro system with a transformer placed between TX and TR.

INPUTSHAFT 60º

INPUTSHAFT 15º

OUTPUTSHAFTθ1 – θ2

TX

TDXTR

60º 45º15º

60º

45º



ENGINEERS


1.12 CONTROL DIFFERENTIAL SYNCHRO Figure 16 illustrates a control differential synchro system.

Control Differential Synchro

Figure 16 As with the straight control synchro system, the ac supply is only applied to the transmitter rotor. The transformer rotor produces an error signal, which after amplification is applied to a motor, causing the CT rotor to move. Apart from these differences the action of the control differential transmitter is the same as for the torque differential synchro system. Torque differential synchros have been used to combine a direction finding loop reading and a compass reading, in navigation systems, to give a true bearing. Control differential synchros, combined with servomotors, are used for moving much heavier loads such as radar scanners where the subtraction or addition of two inputs may be necessary.

INPUTSHAFTθ1

INPUTSHAFTθ2

OUTPUTSHAFTθ1 – θ2

CX CDX CT

ERRORSIGNAL



ENGINEERS


1.13 RESOLVER SYNCHRO This type of synchro is used to convert voltages, which represent the CARTESIAN co-ordinates of a point, into POLAR co-ordinates and vice versa. 1.13.1 POLAR AND CARTESIAN CO-ORDINATES A vector, representing an alternating voltage, can be defined in terms of ‘r’ and the angle it makes with the X-axis: angle (θ). These are the polar co-ordinates of the vector written as r/θ. Figure 17 shows the vector diagram for Polar and Cartesian co-ordinates.

Polar & Cartesian Co-ordinates

Figure 17

θ

r

X

Y

POLAR CO-ORDINATES = r/θ

CARTESIAN CO-ORDINATES X = r COS θ

CARTESIAN CO-ORDINATES Y = r SIN θ



ENGINEERS


1.13.2 RESOLVER SYNCHRO OPERATION The resolver synchro consists of a stator and rotor, each having two windings arranged in phase quadrature as shown in Figure 18.

Resolver Synchro Figure 18

Figure 16b represents the resolver differently for ease of explanation. The resolver has two coils, R1 R2 and R3 R4 at right angles to each other and attached to an input shaft. The stator consists of two coils S1 S2 and S3 S4, also placed at right angles to each other.

INPUT SHAFT S2

S1

S3

S4

R2

R1

R4R3

ROTOR STATORR1

R2

R3 R4

S1

S4S3

S2

a

b



ENGINEERS


1.13.3 CONVERSION FROM POLAR TO CARTESIAN CO-ORDINATES For this purpose, one of the resolver coils is short-circuited, say R3 R4, and the other, R1 R2, has an alternating voltage applied to it. The magnitude of this voltage (r) and the angle (θ) through which both rotor coils are turned, represent the polar co-ordinates r/θ. Figure 19 shows a resolver synchro to carry out this function.

Polar to Cartesian Co-ordinates Figure 19

Consider firstly that the rotor shaft position is such that the R1 R2 coil magnetic field links completely with the stator winding S1 S2, i.e. the coils are aligned. The maximum voltage will therefore be induced in coil S1 S2. Since the stator coil S3 S4 is at right angle to stator coil S1 S2, there will be no voltage developed across it due to R1 R2 coil's magnetic field. When the shaft is rotated at constant speed through 90°, the rotor coil R1 R2 is now in phase quadrature to stator S1 S2, which has zero volts induced in it. However, R1 R2 rotor coil is now aligned with stator coil S3 S4 and this now has maximum voltage induced in it. As the shaft continues to rotate, a cosine voltage wave

θ

ROTOR FLUXR1R1

R2R2

R3R3 R4R4

S1S1

S4S4S3S3

S2S2

R SIN θ

90º 180º 360º270ºθ

R

R COS θ

MAXVOLTS

NOVOLTS

STATOR



ENGINEERS


is developed across S1 S2 stator and a sine voltage wave across S3 S4 stator coil. ‘r cos’ and ‘r sin’ summed together result from the input voltage at R1 R2 and rotor rotation r/. The result represents the cartesian co-ordinates. 1.13.4 CONVERSION FROM CARTESIAN TO POLAR CO-ORDINATES In this arrangement, there are two voltage inputs and these represent the cartesian co-ordinates. They are VX = r cos and VY = r sin θ (Refer Figure 15). VX is input to S1 S2; VY is input to S3 S4. The two together develop an alternating magnetic flux representing the cartesian co-ordinates in the stator. R1 R2 is connected to an amplifier, which drives the output load and the rotor in such a direction as to null the rotor and stop the motor. R3 R4 has a voltage induced in it dependent on the value of the alternating flux. Its value may be calculated using Pythagoras' Theorum √VY² + VX² . Figure 20 shows the layout for performing the above.

Cartesian to Polar Co-ordinates Figure 20

R4R3

S1S1

S2S2

θ θS3 S4

R1

R2 SM

TO LOADVX = r COS θ

VY = r SIN θ VY 2 + VX2

R4 R2

CIRCUIT SYMBOL

R1

R3

S1

S3

S4 S2



ENGINEERS


1.13.5 USE OF RESOLVER SYNCHROS The ability to develop receiver signals at 90° is used, for example, in VOR systems, ADF systems using a non-rotating loop, in autopilots and in flight directors. 1.14 E AND I BAR TRANSMITTER Figure 21a shows an E and I bar transmitter. These devices convert mechanical movements into electrical signals (transducer) and are used in various systems as required. Figure 19a shows an E and I bar as applied to a servo-altimeter.

E & I Bar Transmitter

Figure 21 The ‘E’-bar has a coil wound round the centre limb. This coil is supplied by an ac excitation supply. A magnetic flux is set up within the ‘E’-bar and when the ‘I’-bar is equidistant from the outer limbs of the ‘E’-bar, the waveforms transmitted are equal and opposite (Figure 21b). No output results. If the ‘I’-

A.C. EXCITATION

SUPPLY

RESULTANTWAVEFORM

a b



ENGINEERS


bar is moved (in this case by capsules) one end of the ‘I’-bar is brought in closer proximity to the opposite limb of the ‘E’-bar. The air gap here is reduced, the magnetic field strengthens and the signal from the upper limb coil is increased. (Figure 21b). The opposite end of the ‘I’-Bar moves further away from its associated ‘E’-bar limb, and the resultant signal is weaker. In the case of the servo-altimeter, moving the ‘E’ -bar back to the position nulls the signal so that no signal is produced.

ELECTRONIC INSTRUMENT SYSTEM EIS PAGE 1 of 6



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1 ELECTRONIC INSTRUMENT SYSTEMS All instruments essential to the operation of an aircraft are located on panels, the number of which vary in accordance with the number of instruments required for the appropriate type of aircraft and its flight deck layout. The front instrument panel, positioned in the normal line of sight of the pilots, contains all instruments critical for the safe flight of the aircraft. This panel is normally sloped forward 15° from the vertical to minimize parallax errors. Other panels within the flight deck are typically positioned; Overhead, left and right side and centrally between the pilots. Figure 1 shows the layout of a Boeing 737 Flightdeck.

Boeing 737 Flight-deck

Figure 1




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1.1 FLIGHT INSTRUMENTS There are six flight instruments whose indications are so coordinated as to create a “Picture” of an aircraft’s flight condition and required control movements. These instruments are:

1. Airspeed Indicator. 2. Altimeter. 3. Gyro Horizon Indicator. 4. Direction Indicator 5. Vertical Speed Indicator. 6. Turn & Bank Indicator.

The first real attempt at establishing a standard method of grouping was the “Blind Flying Panel” or “Basic Six”. The “Gyro Horizon Unit (HGU) occupies the top centre position, and since it provides positive and direct indications of the aircraft’s attitude, it is utilized as the “Master Instrument”. As control of airspeed and altitude is directly related to attitude, the “Indicated Air-Speed (IAS), Indicator, Altimeter and Vertical Speed Indicator (VSI) flank the HGU. Changes in direction are initiated by banking the aircraft, and the degree of heading change is obtained from the “Direction Indicator” (DI). The DI supports the interpretation of the roll attitude and is positioned directly below the HGU. The “Turn & Bank Indicator” serves as a secondary reference instrument for heading changes, so it also supports the interpretation of roll attitude. With the development and introduction of new types of aircraft with more comprehensive display presentation, afforded by the indicators of flight director systems, a review of the functions of certain instruments and their relative positions within the group resulted in the adoption of the “Basic T” arrangement as the current standard.




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There are now four key indicators:

1. Attitude Director Indicator. 2. Horizontal Situation Indicator. 3. Combined Speed indicator. 4. Altimeter.

Figure 2 shows the layout of the basic 6 and T instrument groupings.

Basic “Six” and “T” Flight Instrument Grouping Figure 2

GYROHORIZON

DIRECTIONINDICATOR

VERTICALSPEED

INDICATOR

ALTIMETER TURN & BANKINDICATOR

AIRSPEEDINDICATOR

BASIC 6 GROUPING

ATTITUDEDIRECTORINDICATOR

HORIZONTALSITUATIONINDICATOR

ALTIMETER

RADIOMAGNETICINDICATOR

VERTICALSPEED

INDICATOR

COMBINEDAIRSPEEDINDICATOR

BASIC T GROUPING




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1.2 ELECTRONIC INSTRUMENT SYSTEMS Modern technology has enabled some significant changes in the layout of flight instrumentation on most aircraft currently in service. The biggest change has been the introduction of Electronic Instrument systems. These systems have meant that many complex Electro-mechanical instruments have now been replaced by TV type colour displays. These systems also allow the exchange of images between display units in the case of display failures. There are many different Electronic Instrument Systems, including:

1. Electronic Flight Instrument System (EFIS).

2. Engine Indicating & Crew Alerting System (EICAS).

3. Electronic Centralised Aircraft Monitoring (ECAM). Figure 3 shows a typical flight deck layout of an Airbus A320.

Flight Deck Electronic Instrumentation Layout

Figure 3

EFISPFD

EFISPFD

EFISND

EFISND

ECAMENGINE

WARNINGS

ECAMSYSTEMS




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The Electronic Instrument System (EIS) also allows the flight crew to configure the instrument layout by allowing manual transfer of the Primary Flight Display (PFD) with the Navigation Display (ND) and the secondary Electronic Centralised Aircraft Monitoring (ECAM) display with the ND. Figure 4 shows the switching panel from Airbus A320.

A320 EIS Switching Panel Figure 4

ATT HDG

CAPT 3

F/O3

NORM

AIR DATA

CAPT3

F/O3

NORM

E/S DMC

CAPT3

F/O3

NORM

ECAM / ND XFR

CAPT F/ONORM




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As well as a manual transfer, the system will automatically transfer displays when either the PFD or the primary ECAM display fails. The PFD is automatically transferred onto the corresponding ND, with the ECAM secondary display used for the primary ECAM display. The system will also automatically transfer the primary ECAM information onto the ND if a double failure of the ECAM display system occurs. Figure 5 shows a block schematic of the EIS for the Airbus 320.

Electronic Instrument System (EIS) Figure 5

DISPLAYMANAGEMENT

SYSTEMDMS No 1

DISPLAYMANAGEMENT

SYSTEMDMS No 3

DISPLAYMANAGEMENT

SYSTEMDMS No 2

DIGITAL TECHNIQUES NUMBERING SYSTEM PAGE 1 of 21



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1 NUMBERING SYSTEMS The majority of digital computers are wired to understand one particular code. This code usually is not the English language or the decimal numbering system but is instead the binary numbering system. A binary code capable of representing letters of the alphabet, decimal numbers, punctuation marks and special control symbols is used by most digital computers on the market today. Before discussing the binary numbering system and its use in computers, a few rules concerning all numbering systems will be presented. There are three basic characteristics of any number system;

BASE (OR RADIX). POSITION VALUE. DIGIT VALUE.

The base of a numbering system is the total number of unique characters or marks within that system. In the decimal system the base is 10 since there are 10 digits (or characters) which make up the system -0, 1, 2, 3, 4, 5, 6, 7, 8, 9. Each position in a number has a value of BX where B is the base and X is some exponent. For example, the decimal numbers 365 and 653 have two different values even though they are composed of the same digits. The reason that the numbers have different values is that digits of different values occupy positions of different weights:

102 101 100 3 6 5 The first position 100 carries a weight of one. (Any number, except zero, when raised to the zero power is equal to one). The second position 101 carries a weight of 10 and the third position 102 carries a weight of 100 etc. Note that each position is ten times greater than the preceding position. Each digit in a number has a value which exists between zero and the value of the base minus one. For example in the decimal system, the digits range in value from zero to nine. Nine is one less that the base of the system which is ten.




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1.1 GENERAL In describing numbers, one takes into account the value of the various digits and the weight of their respective positions. 102 101 100 3 6 5 is equivalent to: 3 x 102 + 6 x 101 + 5 x 100 = 3 x 100 + 6 x 10 + 5 x 1 =

300 + 60 + 5 = 365

Thus the decimal number 365 is read as three hundred sixty five. Fractional numbers follow the same rules. For example take the decimal number 1402.35 103 102 101 100 10-1 10-2 1 4 0 2 3 5 1 x 103 + 4 x 102 + 0 x 101 + 2 x 100 + 3 x 10-1 + 5 x 10-2 = 1 x 1000 + 4 x 100 + 0 x 10 + 2 x 1 + 3 x 1/10 + 5 x 1/100 = 1000 + 400 + 2 + 3/10 + 5/100 or 1000 + 400 + 2 + 35/100 Note: There is an algebraic rule which states that a number raised to a negative

exponent is equivalent to one over that number raised to a positive exponent.

10-2 = 1/102 or 1/100




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1.2 BINARY NUMBERING SYSTEM The prefix 'BI’ indicates two of something such as bicycle, bifocal, bi-plane etc. The binary numbering system is named after its base, which is two. Since the base is two there are two digits in the system 0 and 1. Position values for a binary number are 2X where x is some exponent and each position will be two times greater in weight than that of the preceding position. Consider the binary number 10110. 24 23 22 21 20 1 0 1 1 0 1 x 24 + 0 x 23 + 1 x 22 + 1 x 21 + 0 x 20 = (1 x 16) + (0 x 8) + (1 x 4) + (1 x 2) + (0 x 1) = 16 + 0 + 4 + 2 + 0 = 22 In describing a binary number in terms of decimal values for the positions, one converts from binary to decimal. Thus a binary 10110 is equivalent to a decimal 22. Often the base of a numbering system is indicated by a subscript in parenthesis. 10110(2) = 22(10) Since the binary system uses only digits 0 and 1 all that one needs to do when converting from binary to decimal is to add the weights of those positions which contain ones. For example consider the number 1101001(2) BIT POSITION 26 25 24 23 22 21 20 POSITION WEIGHT 64 32 16 8 4 2 1 1 1 0 1 0 0 1 64 + 32 + 8 + 1 = 105(10) Therefore: 1101001(2) = 105(10)




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When one desires to convert from decimal to binary there are several methods that may be employed. One method is to use a table. (See table 1). 1024 512 256 128 64 32 16 8 4 2 1 WEIGHT 210 29 28 27 26 25 24 23 22 21 20 BIT POS

Decimal to Binary Conversion

Table 1 Assume the following conversion was desired. 212(10) = ?(2) The method of using the table is to find the largest number in the table, which does not exceed the decimal number that is being converted. The number 128 is the largest possible in this case hence a 'one' bit in the 27 position is required. This immediately defines the size of the binary number as 8 positions (From 27 to 20). Subtracting 128 from 212 leaves a remainder of 84 to be represented by the remaining binary positions. Since 84 is larger than 64 (which is the weight of the 26 position) a 'one' bit is required for the 26 position. Subtracting 64 from 84 leaves a remainder of 20. A 'one' bit in the 25 position would be equivalent to 32, which is too large, thus zero bit must be used for the 25 bit position. So far the binary result is as follows: 27 26 25 24 23 22 21 20 1 1 0 A 'one' bit in the 24 position represents a weight of 16. Sixteen from twenty leaves a remainder of four. Four can be represented in its entirety by a 'one' bit in the 22 position. Therefore the 23, 21 and 20 positions should hold zeros. 27 26 25 24 23 22 21 20 1 1 0 1 0 1 0 0 A re-conversion to decimal would prove the answer's validity. 128 + 64 + 16 + 4 = 212 Therefore: 212(10) = 1 1 0 1 0 1 0 0(2) Another method of converting from decimal to binary is to divide the decimal number by 2 (which is the base of the new number) a successive number of times using the remainders as the digits of the new number. For example consider the following: 28(10) = ?(2)




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0 R = 1 (MSD) 2 1 R = 1 2 3 R = 1 2 7 R = 0 2 14 R = 0 (RIGHT MOST DIGIT OR LSD) 2 28 Division must continue until a zero quotient is obtained. The first remainder is the rightmost digit or least significant digit (LSD) of the new number. Therefore: 28(10) = 1 1 1 0 0(2) A re-conversion to decimal serve as a check. 16 + 8 + 4 = 28




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1.2.1 BINARY FRACTIONS Although many digital computers do not make use of binary fractions, conversion techniques involving them are relatively simple. Some of these techniques will be presented in order to complete the picture of conversion between the binary and decimal systems. The position notation method of converting from binary to decimal can include fractions. Example: 1001.101(2) = ?(10) 23 22 21 20 2-1 2-2 2-3 1 0 0 1. 1 0 1 1 x 23 + 0 x 22 + 0 x 21 + 1 x 20 + 1 x 2-1 + 0 x 2-2 + 1 x 2-3 = 1 x 8 + 0 x 4 + 0 x 2 + 1 x 1 + 1 x 1/2 + 0 x 1/4 + 1 x 1/8 = 8 + 1 + 1/2 + 1/8 or 8 + 1 + .5 + .125 = 9.625 thus: 1001.101(2) = 9.625(10) NOTE: 2-1 = 1/21 = 1/2, 2-2 = 1/22 = 1/4, 2-3 = 1/23 = 1/8 An abbreviated table of decimal equivalents to binary fractions is shown in table 2:

Binary Fraction Conversion 2-1 0.5 2-2 0.25 2-3 0.125 2-4 0.0625 2-5 0.03125 2-6 0.015625

Decimal to Binary Conversion

Table 2




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Just as positions to the left of the binary point were two times greater than that of the preceding position, so the positions to the right of the binary point are two times smaller. Conversion from a decimal fraction to a binary fraction may be done in several ways. One method is to use table 5.2.2. Example: .375(10) = ?(2) Since .5 is greater than .375 a zero bit should be placed in the 2-1 position. A one bit should exist in the 2-2 position, however, since .25 is less than .375. Subtracting .25 from .375 leaves a remainder of .125, which can be fully represented by a one bit in the 2-3 position. Final result is: 2-1 2-2 2-3 0 1 1 THUS: .375(10) = .011(2) A second technique of converting decimal fractions to binary is to multiply the decimal fraction by 2 and look for a carry beyond the decimal point. A carry will indicate a one bit for the 2-1 position; no carry a zero bit. The next step is to again multiply only the fraction portion by 2 and look for a carry. A carry means a one bit for the 2-2 position and no carry indicates a zero bit. The process is continued for as many positions as desired. Example: .375(10) = ?(2) .375 x 2 0.750 2-1 position should hold a zero .750 x 2 1.500 2-2 position should hold a one .500 x 2 1.000 2-3 position should hold a one THUS: .375(10) = .011(2) If a whole number conversion is required in addition to the fraction conversion, the whole number is converted by dividing by two while the fraction is converted by multiplying by 2.




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Example: .205 18.205(10) = ?(2) x 2 .410 2-1 is 0 0 R = 1 (24) x 2 2 1 R = 0 (23) . 820 2-2 is 0 2 2 R = 0 (22) x 2 2 4 R = 1 (21) 1.640 2-3 is 1 2 9 R = 0 (20) x 2 2 18 1.280 2-4 is 1 Accuracy to four places gives the following result: 18.205(10) = 1 0 0 1 0. 0 0 1 1(2) Re-conversion would show that the binary number was not carried out to enough places beyond the binary point to create an exact equivalent. However the number of places of accuracy is up to individual preference. 1.3 ADVANTAGES/DISADVANTAGES OF THE BINARY SYSTEM The binary numbering system is very applicable to computer hardware design. Since there are only two binary digits 0 and 1 these bits (contraction of BINARY DIGITS) can be represented by a switch being open or closed, a light being off or on, a relay being de-energised or energised, a transistor not conducting or conducting, no hole or a hole on paper tape, no magnetized spot or a magnetized spot on magnetic tape or a core being magnetized in one direction or the other. It would require very complicated and expensive circuits in the computer to handle pure decimal numbers and letters of the alphabet whereas very simple circuits handle binary numbers. The speed at which binary arithmetic operations can be performed is also quite desirable in computer operation. Therefore, all incoming data must be converted to a binary code before entering the computer's memory and must be reconverted for outputs that humans recognise. A big disadvantage of the binary numbering system is that it is awkward to use in programming or in computer monitoring operations. Thus it is quite common to use an abbreviated code when dealing with binary numbers. A good short hand system for binary is the octal numbering system.




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1.4 OCTAL NUMBERING SYSTEM The prefix 'OCT' implies eight of something such as octagon, octopus, etc. The base of the octal system is eight since there are eight digits 0, 1, 2, 3, 4, 5, 6, 7. Each position of an octal number carries a value of 8X where x is some exponent. Consider the following octal number: 327(8) Conversion to decimal would be as follows: 82 81 80 3 2 7 3 x 82 + 2 x 81 + 7 x 80 = 3 x 64 + 2 x 8 + 7 x 1 = 192 + 16 + 7 = 215(10) One should note that there are no 8's or 9's in the octal system and that each position of an octal number is 8 times greater in weight than the weight of the preceding position. In converting from decimal to octal one may use a table, such as Table 3, or one may use the 'division by new base' technique.

32768 4096 512 64 8 1 WEIGHT 85 84 83 82 81 80 POS

Decimal to Octal Conversion Table 3




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The later of the two techniques is easier to use. Example: 169(10) = ?(8) 0 R = 2 8 2 R = 5 8 21 R = 1 8 169 Therefore: 169(10) = 251(8) A re-conversion would check the result. 2 x 82 + 5 x 81 + 1 x 80 = 2 x 64 + 5 x 8 + 1 x 1 = 128 + 40 + 1 = 169(10) 1.4.1 OCTAL FRACTIONS Just as in binary fractions many digital computers do not use octal fractions but the rules of conversion will be presented. The following abbreviated table of decimal equivalents for octal positions simplifies conversion. 8-1 = 1/81 = 1/8 = .125 8-2 = 1/82 = 1/64 = .015625 8-3 = 1/83 = 1/152 = .001953125 8-4 = 1/84 = 1/4096 = .000244140625 Example: 37.25(8) = ?(10)

81 80 8-1 8-2 3 7 . 2 5 3 x 81 + 7 x 80 + 2 x 8-1 + 5 x 8-2 = 24 + 7 + .250 + .078125 Therefore: 37.25(8) = 31.328125(10) or 31.33(10) (rounded off) Conversion from a decimal fraction to an octal fraction can also be done by the 'multiply by new base' technique as was done with binary fractions.




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Example: 88.49(10) = ?(8) 0 R = 1 (82) 8 1 R = 3 (81) .49 8 11 R = 0 (80) x 8 8 88 3.92 8-1 is a 3 x 8 7.36 8-2 is a 7 Thus: 88.49(10) o 130.37(8) Note that only the decimal fraction is multiplied by 8 each time. Also note that rounding off was done. 1.5 OCTAL - BINARY CONVERSIONS Since there are only 8 digits in the octal system, each octal digit can be represented by some combination of three binary digits. In fact there are only 8 possible combinations for three binary digits. Octal Binary 0 000 1 001 2 010 3 011 4 100 5 101 6 110 7 111 Conversion between the octal and binary systems then is quite simple since a direct substitution of 3 binary digits for each octal digit is all that is required. Example: 715(8) = ?(2) 7 1 5 111 001 101 Therefore: 715(8) = 111001101(2)




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When converting from binary to octal one marks off groups of three bits from right to left. Example: 11011100(2) = ?(8) 011 011 100 3 3 4 Therefore: 11011100(2) = 334(8) Note that leading zeros are supplied to fill out 3 digits if necessary. When dealing with fractions the only rule other than direct substitution is that groups of three binary digits are marked off from left to right in the binary fraction. Example: 1000111.0101101(2) = ?(8)

001 000 111. 010 110 100 1 0 7. 2 6 4 Therefore: 1000111.0101101(2) = 107.264(8) Note that zeroes are added to the rightmost end of a fraction to fill out the number to three digits. Example: 137.05(8) = ?(2) 1 3 7 . 0 5 001 011 111 . 000 101 or 137.05(8) = 1011111.000101(2) Note that leading zeros may be truncated.




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1.6 ADVANTAGES/DISADVANTAGES OF THE OCTAL SYSTEM Because the conversion between binary and octal is so simple the octal system is often used as shorthand for binary. For example, a particular computer instruction code might be as follows in binary:

0110001101110110

A programmer could write the operation in octal notation thereby reducing some of the cumbersome notation.

061566

The input device or medium would convert the octal digits to binary prior to entering the combination into the computer's memory. Another problem in some computers is reading binary numbers on the console (a monitoring device) or instructing someone to set up a binary code from the console. Octal notation can alleviate the problem to a great extent. In fact, there are a number of computers on the market today which require octal notation in programming and/or console display. Octal techniques in logic design likewise simplify and even save on the number of required circuits as compared to straight binary decoding. The big disadvantage of the octal system is the fact that humans still prefer decimal notation in the end and thus the use of octal might require multiple conversion facilities for data going into or coming out of the computer. Memory dumps (print outs) often are available in a choice of codes, one of which is usually octal.




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1.7 HEXADECIMAL Just as octal is a shorthand for binary because three binary digits can be directly substituted by one octal digit, another numbering system known as hexadecimal, is also a shorthand for binary because of its base. The prefix hexa implies 6 of something and since decimal represents 10, the word hexadecimal means 6 + 10 or 16. Thus the base of the hexadecimal system is 16. By definition of the word 'base' the total number of characters in the system must also be 16. These characters include the ten decimal digits 0-9 and six letters of the alphabet A-F. Table 4 shows decimal-hexadecimal conversions.

HEX 0 1 2 3 4 5 6 7 8 9 A B C D E F DECIMAL 0 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15

Hexadecimal-Decimal Table 4

A hexadecimal number therefore is one whose position values are 16X. The methods of conversion discussed previously still apply. 6 x 162 + A x 161 + F x 160 = 6AF(16) = ?(10) 6 x 256 + 10 x 16 + 15 x 1. = 1536 + 160 + 15 = 1711(10) Decimal-Hexadecimal Example 1: 108(10) = ?(16) 0 R = 13 (equivalent to D) 16 13 R = 0 16 208 Note: Each remainder must be represented by one hexadecimal digit. Therefore: 208(10) = D0(16)




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Decimal-Hexadecimal Example 2: 1834(10) = ?(16) 0 R = 7 16 7 R = 2 16 114 R = 10 16 1834 16 23 16 74 64 1834(10) = 72A(16)




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1.8 BINARY-HEXADECIMAL Four binary digits can form sixteen combinations thereby providing an exact equivalent to the hexadecimal system. This is shown in Table 5 BINARY HEXADECIMAL

0000 0001 0010 0011 0100 0101 0110 0111 1000 1001 1010 1011 1100 1101 1110 1111

0 1 2 3 4 5 6 7 8 9 A B C D E F

Binary – Hexadecimal

Table 5

Therefore, direct substitution can take place between hexadecimal and binary. For every 4 binary digits, one hexadecimal digit can be substituted or vice versa. 1001101(2) = ?(16) 0100 1101 4 D 1001101(2) = 4D(16) CBF(16 = ?(2) C B F 1100 1011 1111 CBF(16) = 110010111111(2)




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Fractions are handled in the same manner: 1101110.01111(2) = ?(16) 0110 1110. 0111 1000 6 E . 7 8 Therefore: 1101110.01111(2) = 6E. 78(16) Note that zeros are added to fill out to multiples of 4 binary digits. The ease with which a binary number can be expressed as a hexadecimal, enables some computer systems to conveniently identify the contents of registers or words in memory. Also it is desirable in business data processing operations to work with decimal numbers. To do this requires a code known as BCD (Binary Coded Decimal). The BCD code is encompassed by the hexadecimal numbering system and thus one may use decimal notation if one desires to do so or hexadecimal and assume that four binary digits represent one decimal or hexadecimal digit. 1.9 BINARY CODED DECIMAL NOTATION If the binary code is to be used in a computer that can handle commercial data processing as well as communications or scientific processing, there has to be a means of representing decimal numbers, letters of the alphabet, punctuation marks and special symbols. It is desirable that this special binary code is also easy to handle in terms of decimal arithmetic. The BCD or binary coded decimal notation solves part of this problem. Below is a chart of the BCD code as applied to decimal numbers. Decimal BCD 0 0000 1 0001 2 0010 3 0011 4 0100 5 0101 6 0110 7 0111 8 1000 9 1001 Direct conversion of any BCD configuration gives the decimal equivalent. BCD notation however does not make use of all 16 possible combinations for four




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binary digits and is therefore susceptible to wasting storage space. The decimal number 15 for example in BCD code would be 0001 0101 while the pure binary equivalent for 15 would be 1111. However, as was stated earlier, letters of the alphabet as well as punctuation marks and special symbols are needed in some form of a binary code. Therefore, a number of computer manufacturers use a modified BCD code. 1.10 BINARY ARITHMETIC One of the tasks a digital computer must be able to perform is to solve complex problems. Some problems require more complex operations than the fundamental operation of addition, subtraction, divide and multiplication. Complex problem solving is achieved by writing it into the computers program (software), however digital circuits (hardware) achieve the fundamental function. 1.11 BINARY ADDITION In the decimal system, the sum of 11 + 3 is 14 and it is not until the sum of the column is greater than 9 that there is a carry from one column of the addition to the next.. Arithmetic operation are very simple in the binary system because as the base of the system is 2, the carry occurs much earlier, so that a sum of two digits resulting in 2 will involve a carry function. As a result there are only four rules to consider when adding binary numbers, which are:

1. 0 + 0 = 0. 2. 0 + 1 = 1. 3. 1 + 1 = 0 carry 1. 4. 1 + 1 + carry 1 = 1 and carry 1.

Example 1 Addition of 10112 (decimal 11) and 00112 (decimal 3). 1011 0011 1110 When adding three or more rows of binary numbers, the addition of all the binary numbers in one column could be carried out as in decimal addition, however, this becomes difficult in remembering how many carries have been made. An easier way is to add two rows at a time, adding the result to the next row and so on. Example 2




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Addition of 1101 + 0111 + 1001 + 0101 a. 1101 0111 10100 b. 10100

01001 11101




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DIGITAL TECHNIQUES DATA CONVERSION PAGE 1 of 10


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1 DATA CONVERSION 1.1 ANALOGUE COMPUTERS Analogue computers operate by using voltages, currents, shaft angles etc to represent physical quantities. The basic concept of the analogue computer is as follows: 1. Physical variables, usually voltages, are used to represent the

magnitudes of all the variables contained within the equation or problem.

2. Computer "building blocks", each performing a single mathematical

function, are interconnected in such a manner that the relationships between the input and output variables correspond to the desired mathematical relationship.

3. The voltage solution exists at a specific point within the system and is

made available to the operator in some form. Generally, there are two types of analogue circuit arrangements in use. The first is a 'general purpose' computing arrangement consisting of a large number of networks, which are capable of providing solutions to a range of problems. The second type is a 'special purpose' arrangement, which is capable of serving as a model for, or simulating, a specific condition. Since the analogue computer operates by a process of measurement, it is best suited to applications where continually varying quantities are to be dealt with. Although computation involving measurement usually introduces errors, it is possible to attain accuracy of better than 0.1%. This is adequate for many applications and, since small analogue computers can deal with relatively simple problems, this type of computer will be met in some equipment carried in aircraft. 1.2 DIGITAL COMPUTERS Digital computers are arithmetic machines: that is, they operate by a process of counting numbers or digits (hence their name). The basic operation that a digital computer can perform is addition.



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The digital computer is, therefore, used when the problem to be solved is of an arithmetical nature and where an exact answer is required. Digital processing errors are very low, with accuracy in the order of 0.001% being possible, although a digital computer operating in a controlling role will have inputs derived from some form of measurement with consequent errors. For specific tasks, the programme of instructions, which supplies the computer with the information on which it operates, can be built in to the machine; digital computers of this type have many aircraft applications. 1.3 ANALOGUE AND DIGITAL SIGNALS Analogue (continuous) information is made available in virtually all aircraft equipment. Figure 1 shows the analogue signal created by a variable resistor. In the circuit +0V is present at the output “A” when the potentiometer is at position 1 and +5V when at position 2. These values would represent either a 1 (+5V) or a 0 (+0V). However, it can be seen from the graph of the analogue signal that it does produce distinct values of +5V and +0V as the potentiometer moves from one end to the other.

Analogue Signal Representation Figure 1

A digital signal is one that contains two distinct values (1 and 0). Figure 2 shows a digital signal being produced by use of a switch. With the switch in the open position, +0V will be present at point (logic 0). When the switch closes, +5V will be present at point (logic 1). Digital signals are often considered to be either “ON” or “OFF” (logic 1 or 0).

+5V

A

POSITION1

POSITION2

+0V

+5V

TIME

O/P A



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Digital Signal Representation Figure 2

Signals in analogue form can be processed using operational amplifiers and other devices in various configurations and ultimately converted to an observable output by a suitable output device. Systems, which are completely analogue, are limited in the accuracy that can be achieved both physically and economically, they also suffer from error and distortion for various reasons such as non-linearity, drift, crosstalk, noise etc. Digital systems, especially since the advent of integrated circuits, offer improvements over analogue systems in most respects, thus modern processing systems employ fixed analogue and digital circuitry (hybrid systems) in which, of course, conversion from one form to the other must take place at certain points within the system. Hybrid systems are more common than all digital systems presumably because of the simplicity of analogue transducers, and the nature of the information to be processed lends itself more readily to analogue representation. For example it would be difficult to digitize an audio signal without converting it from changing air pressure to an electrical analogue by means of a microphone (transducer). For further computing such an electrical analogue signal would be converted into digital form.

+5V A

TIME

O/P A

+0V

+5V



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1.4 ANALOGUE TO DIGITAL CONVERTER In an ADC a range of input values must correspond to a unique digital word. The type of code used depends on the system but here only binary coding will be considered. Consider an analogue signal, which can take on any value between 0 and 7 volts. For any particular voltage there is a corresponding binary code word. For example, using 3-bit words, the voltage analogue value between 4 and 5 volts would be represented in binary code by the word 100, which would change to 101, when the analogue value passed through 5 volts. Figure 3 shows digital representation of an analogue input signals.

Digital Representation of Analogue Signals

Figure 3

8

7

6

5

4

3

2

1

AN

ALO

GU

E SI

GN

AL

000

001

010

011

100

101

110

111

DIGITAL SIGNAL

3 BITWORD



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The levels at which the code changes are known as quantisation levels, and the intervals between them as quantisation intervals. In the example given in Figure 5.3.3, the quantisation levels are 0, 1, 2, 3, 4, 5, 6 and 7 volts, and the quantisation interval is 1 volt. Using a 3-bit word gives 23 = 8 different quantisation levels. With a 4-bit word we would have 24 = 16 quantisation levels with 0.5 volt quantisation intervals giving improved resolution over the same range of input voltage. Thus the more bits available the greater the resolution for a given range of analogue signal input. It can be seen from the above that an ADC using an n-bit word would have a resolution of one part in 2n. 1.5 ANALOGUE TO DIGITAL CONVERSION In order to convert the analogue signal into a digital signal, an Operational Amplifier is used as a comparator. Figure 4 shows an Op amp comparator.

Comparator Circuit Figure 4

The output of the comparator will be logic “0” when the reference voltage is greater than the analogue input, changing to logic “1” when the analogue voltage is greater than the reference voltage.

+ VOUTVIN

+VE

VREF



ENGINEERS


Figure 5 shows the resultant digital waveforms from an analogue input signal using an Op Amp comparator.

Analogue/Digital waveforms Figure 5

In the example in figure 3, the quantisation level was 0 – 7 with a quantisation interval of 1 volt. To convert this range to digital a total of 7 comparator Op Amps would be required. This however would give a word length of 7 bits. We know to represent the range 0 – 7 with an interval of 1 volt will only require a 3-bit word.

VIN 0

VREF

+VMAX

-VMAX

VOUT 0

WHEN VIN < VREF THEN VOUT = -V MAX

WHEN VIN > VREF THEN VOUT = +V MAX



ENGINEERS


To convert the seven bit word to a 3-bit word an encoder circuit is used. The circuit contains a number of logic gates that will convert the 7-bit word down to the required 3-bit notation. Figure 6 shows the layout of an encoder circuit.

Encoder Circuit Figure 6

X

Y

Z

AB

CD

EF

G

LSB

MSB



ENGINEERS


1.6 DIGITAL TO ANALOGUE CONVERSION (DAC) Since many systems used on aircraft will require outputs in analogue form, it will be necessary to be able to convert the digital information back into analogue. The input to the DAC is effectively a number, usually binary coded. This number must be converted to a corresponding number of units of voltage (or current) by the DAC. The output of the DAC will thus be stepped as the digital input changes, taking on a series of discrete values. The spacing between these values (quantisation levels) will depend on the length of the input digital word and the maximum range of the output voltage. For example, a DAC, which can provide an output voltage of between 0 and 16 volts, will, with 4-bit word input, have 1 volt between quantisation levels and is illustrated in Figure 8.

DAC Output Figure 8

16

14

12

10

8

6

4

2

AN

ALO

GU

E O

/P S

IGN

AL

0000

0001

0010

0011

0100

0101

0110

0111

DIGITAL I/P SIGNAL

1000

1001

1010

1011

1100

1101

1110

1111



ENGINEERS


Similarly, an output voltage range of 0 to 10 volts with 10-bit word input will give spacing between quantisation levels of approximately 0.01 volts. The stepped nature of the output can of course be smoothed. To change a digital word into an analogue signal we require a circuit capable of carrying out this function. One method would be to apply the digital word to a corresponding number of resistors (4-bit word – 4 resistors), connected as a potential divider. Figure 9 shows a circuit that would carry out the function of Digital to Analogue conversion.

DAC Weighted Circuit

Figure 9

V OUT

MSB

LSB

4

BIT

WORD

R

2R

4R

8R



ENGINEERS


Figure 10 shows a Digital to analogue converter.

Digital – Analogue Converter Figure 10

MSB

LSB

R

2R

4R

8R

S1

S2

S3

S4

0V

-

+ ANALOGUEOUTPUT

VOLTAGE

V REF

4 BITDIGITALINPUT

DIGITAL TECHNIQUES /ELECTRONIC INSTRUMENT SYSTEM DATA BUSES PAGE 1 of 21


ENGINEERS


1 DATA BUSES The availability of reliable digital semi-conductor technology has enabled the inter-communication task between different equipment to be significantly improved. Previously, large amounts of aircraft wiring were required to connect each signal with all the other equipment. As systems became more complex and more integrated so this problem was aggravated. Digital data transmission techniques use links, which send streams of digital data between equipment. These data links may only comprise two or four wires and therefore the inter-connecting wiring is very much reduced. Recognition of the advantages offered by digital data transmission has led to standardization in both civil and military fields. The most widely used digital data transmission standards are ARINC 429 for civil and MIL-STD-1553B for military systems. 1.1 AERONAUTICAL RADIO INCORPORATED (ARINC) 429 ARINC specification 429 is titled "MARK 33 Digital Information Transfer System" (DITS). We refer to it as ARINC 429 bus, DITS bus, Mark 33 bus or just ‘bus’. 1.1.1 OPERATION An equipment transmits data, via a 429 transmitter, to other equipment. The information flow is uni-directional. One 429 transmitter supplies the data to a pair of wires that we call the bus. One or more ARINC 429 receivers can be connected to the bus. The ARINC 429 bus is a twisted and shielded pair of wires and the shield is connected to ground. The data wires are white and blue. The ground connection is a black wire. If the bus runs through a feed-through plug (for instance on a bulkhead), then the shield is also connected to a black wire that runs through the plug.



ENGINEERS


Figure 1 shows ARINC bus interconnections.

ARINC Bus Interconnection

Figure 1

ARINC 429 BUS

INFORMATION FLOW

TWISTED AND SHIELDED WIRES

ARINC 429TRANSMITTER

ARINC 429RECEIVER

DATAINPUT

DATAINPUT

TX RX

RX



ENGINEERS


1.1.2 DATA BUS CABLE Data bus cable typically consists of a twisted pair of wires surrounded by electrical shielding and insulators. Digital systems operate on different frequencies, voltages and current levels. It is extremely important to ensure that the correct cable is used for the system installed. The cable should not be pinched or bent during installation and data bus cable lengths may also be critical. Refer to current manufacturer’s manuals for cable specifications. Figure 2 shows an example of a data bus cable.

Data Bus (Twisted Pair) Cable Figure 2

TINNED COPPERCONDUCTORS

DATA BUSCABLE “A”

DATA BUSCABLE “B” TINNED COPPER

BRAID SHIELD

ETFE TEFZEL®

INSULATION ETFE TEFZEL®

JACKET



ENGINEERS


1.2 THE ARINC 429 DATA BUS Data words contain the information. An example is Indicated Airspeed (IAS). Another example is Total Air Temperature (TAT). A 429 transmitter transmits IAS, then pauses a moment, and then transmits TAT. 255 different data words can be transmitted on one 429 bus. The information is transmitted at high or low speed:

Low speed is 12 to 14.5 Kbytes/second.

High speed is 100 Kbytes/second. Figure 3 shows the ARINC Dataword format.

ARINC 429 Data Word Formats

Figure 3

DATA WORD 32 BITS DATA WORD 32 BITS DATA WORD 32 BITS

INDICATED AIR SPEED (IAS)TRANSMITTED EITHER:

12 - 14 KBYTES/SEC - LOW SPEED100 KBYTES/SEC - HIGH SPEED

TOTAL AIR TEMPERATURE (TAT)TRANSMITTED EITHER:

12 - 14 KBYTES/SEC - LOW SPEED100 KBYTES/SEC - HIGH SPEED

PAUSE BETWEENDIFFERENT TYPES

OF DATA BEING TRANSMITTED



ENGINEERS


1.2.1 ARINC 429 SPECIFICATIONS ARINC 429 sets specifications for the transfer of digital data between aircraft electronic system components and is a “One-way” communication link between a single transmitter and multiple receivers. ARINC 429 system provides for the transmission of up to 32 bits of data. One of three languages must be used to conform to the ARINC 429 standards:

1. Binary.

2. Binary Coded Decimal (BCD).

3. Discrete. ARINC 429 assigns the first 8 bits as the word label; bits 9 and 10 are the “Source-Destination Indicator” (SDI), bits 11 through to 28 provide data information; bits 29 through to 31 are the “Sign-Status Matrix” (SSM), and bit 32 is a “Parity Bit. There are 256 combinations of word label in the ARINC 429 code. Each word is coded in an octal notation language and is written in reverse order. The source-destination indicator serves as the address of the 32-bit word. That is, the SDI identifies the source or destination of the word. All information sent to a common serial bus is received by any receiver connected to that bus. Each receiver accepts only that information labelled with its particular address; the receiver ignores all other information. The information data of an ARINC 429 coded transmission must be contained within the bus numbered 11 through to 28. This data is the actual message that is to be transmitted. For example, a Digital Air Data Computer (DADC) may transmit the binary message 0110101001 for Indicated Airspeed. Translated into decimal form, this means 425, or an airspeed of 425 knots. The sign-status matrix provides information that might be common to several peripherals (plus or minus, north or south, right or left etc). The parity bit of ARINC 429 code is included to permit error checking by the ARINC receiver. The receiver also performs a “Reasonableness Check”, which deletes any unreasonable information. This ensures that if a momentary defect occurs in the transmission system resulting in unreasonable data, the receiver will ignore that signal and wait for the next transmission. The parity bit will either be set to 1 or 0 depending on the parity used. The parity used in ARINC 429 is “Odd Parity”. If there is an even number of 1 bits in a transmitted word (bits 1 through 31), the parity bit must be 1 to ensure the whole word contains an odd number of 1 bits in the word. Figure 4 shows the layout of a 32-bit word.



ENGINEERS


32 Dataword Format Figure 4

8 ------ 110 / 928 - - - - - - - - - - - - - - - 1131 - 2932

DATAWORD LABEL8 BITS - OCTAL 000 - 377

SOURCE DESTINATIONIDENTIFIER (SDI

0 0 - ALL SYSTEMS0 1 - SYSTEM 11 0 - SYSTEM 21 1 - SYSTEM 3

DATA FIELD 18 BITSBINARY CODED DECIMAL

(BCD)OR

BINARY FORMAT(BNR)

ORDISCRETE FORMAT

SIGN & STATUS MATRIX (SMM)MEANING RELATED TO

FORMAT

PARITY BITEITHER

ODD/EVEN



ENGINEERS


1.3 ARINC 429 WORD REPRESENTING AIRSPEED Figure 5 represents an ARINC 429 code for a DADC word giving information on the aircraft’s indicated airspeed.

ARINC 429 word 206 Indicated Airspeed Figure 5

The word label for airspeed is 206 and it is transmitted using the octal notation code, which is read in reverse to achieve the word label. E.g. word label 602 would be 011 000 01 (bits 1,6 and 7 set to logic 1), 206 in reverse. The SDI label 00 indicates transmission of this data to all receivers connected to the serial bus. The data segment is read left to right, 0110101001 representing the sum of; 1 x 256 (28) + 1 x 128 (27) + 1 x 32 (25) + 1 x 8 (23) + 1 x 1 (20). In decimal form this represents 425. The SMM 011 represents a normal operation of a plus value data; that is, airspeed data is a positive value. The parity bit is set to 1, which denotes an even number of 1s in the transmitted word and no errors are present according to the parity bit.

32 31 30 29 28 27 26 25 24 23 22 21 20 19 18 17 16 15 14 13 12 11 10 9 8 7 6 5 4 3 2 1

1 1 1 0 0 1 1 0 1 0 1 0 0 1 0 0 0 1 1 0 0 0 0 1

PARITY

SIGN STATUSMATRIX

DATA FIELD

SOURCE DESTINATIONIDENTIFIER

WORD LABEL



ENGINEERS


1.4 THE ARINC 429 FORMAT ARINC has a return to zero format. After a bit is transmitted, the voltage returns to zero. If logic 1 is transmitted, line A has a voltage of +5 volts and line B has a voltage of -5 volts with respect to ground. This means that the voltage on line A is 10 volts higher than the voltage on line B. If logic 0 is transmitted, line A has a voltage of -5 volts and line B has a voltage of +5 volts with respect to ground. This means that the voltage on line A is 10 volts lower than the voltage on line B. Spikes caused by interference make the voltage on both wires increase or decrease but have no effect on the voltage of line A with respect to line B. Therefore interference has less effect on the bus. Figure 6 shows the ARINC 429 dataword format.

ARINC 429 Dataword Format

Figure 6

1 1 1 1 1 1 10 0 0 0 0 0 0

HIGH +10v

LOW -10v

NULLLINEA TO B

+5v

-5v

0LINE A

TOGROUND

+5v

-5v

0LINE B

TOGROUND

1 2 3 4 5 6 7 8 27 28 29 30 31 32

RETURN TO ZERO (RZ) FORMAT



ENGINEERS


1.5 DATA TRANSMISSION Most digital communication data is transmitted in a serial form, that is, only one bit at a time. Transmission of data in serial form means each bit is transmitted for only a very short time period. In most systems, the data transmitted requires less than a milli-second. After one bit is sent, the next bit follows; this process is repeated until all the desired bits have been transmitted. This type of system is often referred to as “Time Sharing”, because each transmitted signal shares the wires for a short time interval. Parallel data transmission is a continuous-type of transmission requiring two wires (or one wire and ground) for each bit to be sent. Parallel transmission is so called because each circuit is wired in parallel with respect to the next circuit. With serial data, one pair of transmitting wires can be used to send enormous amounts of serial data. If the data were sent using the parallel method, then hundreds of wires would be required. Most computer systems use the parallel method to transmit data within them, however if the data must be sent to another system, serial data transmission is used. An interpretation circuit is required to convert all parallel data to serial-type data prior to transmission. The device for sending serial data is called a “Multiplexer (MUX), and the device for receiving serial data is called a “Demultiplexer” (DEMUX). Figure 7 shows a data transfer system.



ENGINEERS


Data Transfer System Figure 7

123456789101112

123456789101112

MU

LTIP

LEXE

R

DEM

ULT

IPLE

XER

DATA TRANSFER 00110

SERIAL DATATRANSMISSION

BIT NUMBER

01100

01100

PARALLELDATA

PARALLELDATA

TO CENTRALCONTROL UNIT



ENGINEERS


The MUX circuit operation is shown in Figure 8.

Multiplexer Circuit Operation

Figure 8

The X and Y inputs are the control inputs selecting the data to be multiplexed. Table 1 shows the logic table for X and Y.

X Y 0 0 1 0 0 1 1 1

Multiplexer Control logic table

Table 1

OUTPUT

CONTROLSIGNALS

X Y

D

A

B

C



ENGINEERS


Figure 9 shows the DEMUX logic circuit.

Demultiplexer Logic Circuit Figure 9

0

1

2

3

4

5

6

7

S1

S2

S0

DATAINPUT

BIT 1

BIT 2

BIT 3

BIT 4

BIT 5

BIT 6

BIT 7

BIT 8



ENGINEERS


1.6 ARINC 573 FORMAT The ARINC 573 format has been established for “Digital Flight Data Recorder” (DFDR). It uses the Harvard bi-phase code, containing the bits in bit-cells. Because each bit-cell is a phase transition, the ARINC 573 is self-clocking. If the logic = 1, then the bit-cell will have a phase transition: for a logic 0, there is no phase transition. If the DFDR gives no information, the ARINC 537 output is a symmetric square wave. Figure 10 shows ARINC signal format.

ARINC 537 Signal Format Figure 10

1 1 111 00 0 000 0

+5v

-5v

0v

DATA

ARINC 573HARVARD

BI-PHASE CODE

1 2 3 4 5 61 62 63 64

1 6 1192 43 5 1287 10ONEWORD

ONESUBFRAMESYNCWORD

64 WORDS

SUBFRAME 1 SUBFRAME 2 SUBFRAME 3 SUBFRAME 4

4 SUBFRAMESONE

FRAME

4 SEC 4 SECFRAMES

12 BITS



ENGINEERS


1.7 CONVERTERS In analogue circuits we cannot use digital signals and in digital circuits we cannot use analogue signals. For that reason there are analogue to digital converters and digital to analogue converters. Also there are converters that change analogue signals into other analogue signals, e.g. a pressure to frequency converter, which is used in the air data computer. 1.7.1 EXAMPLES OF CONVERTERS Figure 11 shows three different types of converters.

Converters Figure 11

A

D

AD

PF

ANALOGUE TO DIGITAL CONVERTER

DIGITAL TO ANALOGUE CONVERTER

PRESSURE TO FREQUENCY CONVERTER



ENGINEERS


1.8 THE ARINC 629 DATA BUS The ARINC 629 is a new digital data bus format that offers more flexibility and greater speed than the ARINC 429 system. ARINC 629 permits up to 120 devices to share a “Bi-directional serial data bus”, which can be up to 100M long. The data bus can be either a twisted pair, or a fibre-optic cable. ARINC 629 has two major improvements over the 429 system; firstly there is a substantial weight savings. The ARINC 429 system requires a separate wire pair for each data transmitter. With the increased number of digital systems on modern aircraft, the ARINC 629 system will save hundreds of pounds by using one data bus for all transmitters. Secondly, the ARINC 629 bus operates at speeds up to 2 Mbits/sec; the ARINC 429 is only cables of 100Kbits/sec. Figure 21 shows simplified diagrams of ARINC 429 and 629 bus structures.

ARINC 429/629 Bus Structures Figure 21

ARINC 429 STRUCTURE ARINC 629 STRUCTURE



ENGINEERS


The ARINC 629 system can be thought of as a party line for the various electronic systems on the aircraft. Any particular unit can transmit on the bus or listen for information. At any given time, only one user can transmit, and one or more units can receive data. This “Open Bus” scenario poses some interesting problems for the ARINC 629 system:

1. How to ensure that no single transmitter dominates the use of the bus.

2. How to ensure that the higher-priority systems have a

chance to talk first.

3. How to make the bus compatible with a variety of systems.

The answer is found in a system called “Periodic/Aperiodic Multi-transmitter Bus”. Figure 22 shows ARINC 629 bus structure.

ARINC 629 Bus Structure Figure 22

TERMINALINTERVAL

TERMINALGAPS

1 2 3 4 1

SYNCHRONIZATIONGAP

PERIODIC INTERVAL

1 2 3 4 1

TERMINALINTERVAL

TERMINALGAPS SYNCHRONIZATION

GAP

APERIODIC INTERVAL



ENGINEERS


Each transmitter can use the bus, provided it meets a certain set of conditions.

1. Any transmitter can make only one transmission per terminal interval.

2. Each transmitter is inactive until the terminal gap

time for that transmitter has ended.

3. Each transmitter can make only one transmission; then it must wait until the synchronization gap has occurred before it can make a second transmission.

1.8.1 TERMINAL INTERVAL The Terminal Interval (TI) is a time period common to all transmitters. The TI begins immediately after any user starts a transmission. The TI inhibits another transmission from the same user until after the TI time period. 1.8.2 PERIODIC & APERIODIC INTERVAL A Periodic Interval occurs when all users complete their desired transmission prior to the completion of the TI. If the TI is exceeded, an Aperiodic Interval occurs when one or more users have transmitted a longer than average message. 1.8.3 TERMINAL GAP The Terminal Gap (TG) is a unique time period for each user. The TG time determines the priority for user transmissions. Users with a high priority have a short TG. Users with a lesser need to communicate (lower priority) have a longer TG. No two terminals can ever have the same terminal gap. The TG priority is flexible and can be determined through software changes in the receivers/transmitters. 1.8.4 SYNCHRONIZATION GAP The Synchronization Gap (SG) is a time period common to all users. This gap is a reset signal for the transmitters. Since the Synchronization gap is longer than the terminal gap, the SG will occur on the bus only after each user has had a chance to transmit. If a user chooses not to transmit for a time equal to, or longer than, the SG, the bus is open to all transmitters once again.



ENGINEERS


1.9 MESSAGE FORMATS The data is transmitted in groups called “Messages”. Messages are comprised of “Word Strings” and up to 31 word strings can be in a message. Word strings begin with a label, followed by up to 256 data words. Each label and data word is 20 bits long (3 bits for synchronization, 16 data bits and 1 parity bit). Figure 23 shows the complete structure of the ARINC 629 message.

ARINC 629 Message Structure

Figure 23

START NEXT NEXT NEXT

TERMINAL INTERVAL

LABEL DATA WORD DATA WORD DATA WORDWORD STRINGS

LABEL P DATA P

HI - LOSYNCH

20 BITS

HI - LOSYNCH

20 BITSUPTO

256 DATAWORDS



ENGINEERS


1.10 ARINC 629 DATA BUS COUPLING Another unique feature of the ARINC 629 bus is the “Inductive Coupling” technique used to connect the bus to receivers/transmitters. The bus wires are fed through an inductive pick-up, which uses electromagnetic induction to transfer current from the bus to the user, or from the user to the bus. This system improves reliability, since no break in the bus wiring is required to/from connections. Figure 24 shows an example of Inductive Coupling.

ARINC 629 - Inductive Coupling Technique Figure 24

INDUCTIVEPICK-UP

COUPLINGOUTPUT

DATA

ARINC 629DATA BUS



ENGINEERS


1.11 STUB CABLES The stub cables are for bi-directional data movement between LRU and current mode coupler. The stub cables also supply power from the LRUs to the current couplers. The stub cable has four wires, two to transmit and two to receive. These cables are in the normal aircraft wiring bundles. Figure 25 shows the basic layout for connecting LRUs to the 629 data bus using stub cables. The stub cable length is up to 50ft for TX/RX cable and 75ft for RX only cable.

ARINC 629 Connection Figure 25

LRU TRAY

ARINC 600CONNECTOR

STUB CABLES(TWO SHIELDEDTWISTED PAIRS)1 PAIR RECEIVE

1 PAIR TRANSMIT

ARINC 629 CURRENTMODE COUPLER

STANCHIONDISCONNECT

STUB CABLE(FOUR CONDUCTORS

WITH OVERALL SHIELD)



ENGINEERS


Figure 26 shows ARINC 629 system layout.

ARINC 629 System Layout Figure 26

LRU NO 1 LRU NO 3 LRU NO 5

LRU NO 2 LRU NO 4 LRU NO 6

OVERHEADPANEL

OPAS

DIGITAL TECHNIQUES LOGIC CIRCUITS PAGE 1 of 20


ENGINEERS


1 LOGIC CIRCUITS The term logic in electronics refers to the representation and logical manipulation of numbers usually in a code employing two symbols. i.e., bits. An electronic logic circuit is one whose inputs and outputs can take only one of two states. Where the output of such a circuit depends only on the present state of the input to the circuit, it is called a COMBINATIONAL LOGIC CIRCUIT. Logic circuits may have many inputs and many outputs and be made up of a large number of elements called LOGIC GATES. Most modern electronic logic networks are constructed from two state components in the form of integrated circuits fabricated in a single piece of pure silicon and often referred to as a CHIP. They are available as transistor-transistor logic (TTL) and complementary symmetry metal oxide semiconductor (CMOS or COSMOS) which supersede earlier resistor-transistor logic (RTL) and diode-transistor logic (DTL). Logic circuits are most widely used in computers and calculators, but their use also extends to a wide range of control and test equipment. Figure 1 shows the logic convention.

Logic Conventions Figure 1

POSITIVE LOGIC NEGATIVE LOGIC

0V

5V

0 0

0 01

1

POSITIVE LOGIC : 0 - LOW VOLTAGE: 1 - HIGH VOLTAGE

NEGATIVE LOGIC : 0 - HIGH VOLTAGE: 1 - LOW VOLTAGE



ENGINEERS


As the 'positive logic' representation is favoured by the majority of designers and manufacturers, it is intended to adopt this representation throughout this section. Positive logic refers to the use of a 1 to represent the true or more positive level (e.g. +5v) and 0 to represent the fault, or less positive level (e.g. 0v). 1.1 GATES The word GATE suggests some kind of forceful control, and LOGIC GATES are the basic elements which actively route the flow of digital information through the logic circuits. In a logic circuit, groups of gates working together are able to send particular bits of information to specified locations. A logic gate is a device (usually electronic) that has a single output terminal and a number of inputs, or control terminals. If voltage levels representing the binary states of 1 or 0 are fed to the input terminals, the output terminal will adopt a voltage level equivalent to 1 or 0, depending upon the particular function of the gate. The basic logic gates provide the functions of AND and OR, each being represented by a distinctive symbol. It is sometimes convenient to show the circuit action of the gates by an equivalent contact switching circuit, and these will occasionally be employed to assist in describing the function of a logic gate element.



ENGINEERS


1.2 BASIC 'AND' GATE Figure 2 shows the symbol that represents 2 input AND gate together with its truth table. This gate will only adopt a 1 state at its output terminal when both the inputs A and B, are at the 1 state. This function can be represented by two switches, A and B, connected in series such that the circuit is made only when both switches are CLOSED. (i.e., both in the 1 state).

Basic 'AND' Gate

Figure 2

A B

A

BA.B

SYMBOL

A

0

1

1

0

B

0

0

1

1

A.B

0

0

1

0



ENGINEERS


1.3 BASIC “OR” GATE Figure 3 shows the symbol that represents a 2 input OR gate together with its truth table. This gate will adopt a 1 state at its output terminal when either input A or B or both are at the 1 state. This function can be represented by two switches A and B connected in parallel. Because this gate also performs the AND function (i.e. 1.1 = 1) it is often referred to as an INCLUSIVE OR gate.

Basic 'OR' Gate Figure 3

A

B

A

BA+B

SYMBOL

A

0

1

1

0

B

0

0

1

1

A+B

0

1

1

1



ENGINEERS


1.4 THE 'NAND' GATE When constructing a NAND gate using transistors as the switching devices, the output often represents the 'inversion' of the “AND” gate. Figure 4 shows an example of a 2 input digital gate consisting of two NPN transistors, TR1 and TR2, which are assumed to be perfect switches. In a positive logic system, when input A and input B are both at the 0 state (0v), both transistors are biased OFF and the output will adopt the 1 state (+ 5v). If input A only is now given the 1 state, transistor TR1 is biased ON but no collector current can flow as TR2 is still OFF. Similarly, if input B only is given the 1 state then transistor TR2 is biased ON but again no current can flow as TR1 is OFF. Only when both input A and input B are at the 1 state together, with both transistors ON, will current be allowed to flow taking the output to the 0 state.

The 'NAND' Gate Figure 4

+5V

A.B

A

B

TR1

TR2

A

0

1

1

0

B

0

0

1

1

A.B

1

1

0

1

A

BA.B

SYMBOL



ENGINEERS


1.5 THE 'NOR' GATE Figure 5 shows a further example of a 2 input digital gate, again consisting of two NPN transistors, TR1 and TR2, in a different configuration. When input A and input B are both at the 0 state (0v), both transistors are biased OFF and the output will adopt the 1 state (+ 5v). If input A only is given the 1 state, transistor TR1 will be biased ON and current will flow, making the output take up the 0 state. Similarly, if input B only is given the 1 state, transistor TR2 will be biased ON, taking the output to the 0 state. Finally if both input A and input B are at the 1 state together, the output will again adopt the 0 state.

The 'NOR' Gate Figure 5

+5V

A+BA

B

TR1

TR2

A

0

1

1

0

B

0

0

1

1

A+B

1

0

0

0

A

BA+B

SYMBOL



ENGINEERS


1.6 'EXCLUSIVE OR' GATE The basic OR gate illustrated previously in Figure 3 was seen to include the AND operation in that its output will adopt the 1 state not only when either input A or input B is at the 1 state but also when BOTH inputs are at 1. There are many occasions in logic circuits when it is required to perform the OR operation only when input A or input B are exclusively at the 1 state. In other words, a gate is required whose output adopts the 1 state only when the two input states are not identical, and such a device is known as the EXCLUSIVE OR gate. As an example, suppose the problem is to implement the following logical statement: "A room has two doors and a central light, and switches are to be fitted at each door such that either switch will turn the light on and off". By fitting double-pole changeover switches at each door, a switching circuit could be wired to perform the required operation as shown in Figure 6. If each switch position is designated 'down' for the 1 state and 'up' for the 0 state, then symbols can be allocated to each switch position as shown in the diagram. If the lamp L is designated 1 for ON and 0 for OFF, then the truth table will show the circuit conditions for the switching combinations. As the EXCLUSIVE OR gate can occur frequently in a logic circuit, it has been allocated its own special symbol, as shown in Figure 6, with an equivalent circuit shown at Figure 7. Also, in Boolean algebra expressions, a CIRCLE SUM ⊕ symbol is often employed to signify that a particular expression represents the EXCLUSIVE OR operation. i.e.: A ⊕ B = AB + AB



ENGINEERS


“EXCLUSIVE OR” Represented by Switches

Figure 6

XOR Circuit

Figure 7

UP

DOWN

UP

DOWN

A

A

B

B A

0

1

1

0

B

0

0

1

1

L

0

1

0

1

A

BL

SYMBOL

A

B

Q



ENGINEERS


1.7 THE INVERTER ('NOT' GATE) Figure 8 shows the symbol for an inverter, where the output will produce the complement of the input. This device is often employed when the complement of a particular signal is required at some point in the logic circuit.

The Inverter

Figure 8

A A

A A1 00 1



ENGINEERS


1.8 INVERTING WITH LOGIC GATES Either the NAND or the NOR gate can be connected to operate as a simple inverter as illustrated in Figure 9. In diagram (a) a 2 input NAND gate is shown with one input permanently held at the 1 state (+ 5v), and the resulting output will be the inversion of the single input A. Diagram (b) shows a 2 input NOR gate with one input permanently held at the 0 state (0v) again resulting in an output which will be the inversion of the single input A. These configurations can be particularly useful in logic circuits where the inversion of a variable is required without the need for power amplification.

Logic Gate Inverters Figure 9

(a)NAND INVERTER

(b)NOR INVERTER

A A

+5V OV

A A



ENGINEERS


1.9 MULTIPLE INPUT GATE SYMBOLS Digital integrated circuits are manufactured with multiple inputs to a single gate operation, and the approved symbols to be used to illustrate these types are shown in Figure 10. Diagram (a) shows a multiple input NAND gate symbol, whilst diagram (b) shows the symbol for a multiple input NOR gate.

Multiple Gate Symbols

Figure 10

(a)NAND SYMBOL

(b)NOR SYMBOL

A

B

C

D

E

F

G

A

B

C

D

E

F

G



ENGINEERS


1.10 TIME DELAY ELEMENTS Delay elements are used to 'delay' the travel of a pulse along a line for a short period of time. This is occasionally necessary to ensure that one bit of information does not arrive at some point in the circuit earlier than another. Most delay times are relatively small and only amount to few milli-seconds. Most delay elements have one input terminal and one output terminal, and if a pulse is fed to the input a similar pulse will appear at the output after the specified time period. Figure 12 shows two types of time delay elements.

Time Delay Elements Figure 12

The symbols shown in Figure12 are those used to represent delay elements, and twin vertical lines on the symbol indicate the input side. If the element provides a single delay the duration is included on the symbol as shown in symbol (a). If the delay is tapped to provide multiple outputs, the delay time with respect to the input is included adjacent to the particular tapped output as shown in symbol (b).

5mS

(a) - SINGLE OUTPUT

5mS

2mS

5mS

3mS

(b) - MULTIPLE OUTPUT



ENGINEERS


1.11 ACTIVE STATE INDICATORS Logic diagrams make extensive use of the 'active state indicator' which takes the form of a small circle at the input or output terminals of a logic symbol. It is used to indicate that the normal active state of the particular logic level has been inverted at that point in the symbol. Throughout this section the 'positive logic' convention has been adopted and the 1 state has been used to signify the 'active' state with regard to the symbols and the truth tables. In this instance therefore, the significance of an active state indicator attached to a symbol can be defined as follows: (1) A small circle at the input to any element indicates that a 0 state will

now activate the element at that particular input only. (2) A small circle at the output of any element indicates that the output

terminal of that element will adopt the 0 state when activated.

Figure 13 shows examples of Active State Indicators

Active State Indicators Figure 13

A

1

1

0

0

B

0

1

1

0

AB

0

0

1

0

ABA

B

A

1

1

0

0

B

0

1

1

0

A+B

0

1

1

1

A+BA

B



ENGINEERS


1.12 THE 'INHIBIT' GATE Occasionally it is required to 'hold' one input to an AND gate at a particular logic level in order to disable the entire gate. One method of representing this symbolically is shown in Figure 14, which illustrates a two input gate with an 'INHIBIT' input C carrying an indicator. In this case, with a 1 state at the inhibit input C, the gate is disabled irrespective of the input conditions at A and B. With a 0 state at the inhibit input C however, the gate is now 'enabled' and the output will adopt the 1 state when both input A and input B are at the

1 state.

The 'INHIBIT' Gate Figure 14

ABC

A

B

C

A

0

1

0

1

0

1

0

1

B

0

0

1

1

0

0

1

1

C

1

1

1

1

0

0

0

0

ABC

0

0

0

0

0

0

0

1



ENGINEERS


1.13 AIRCRAFT APPLICATIONS Logic circuits have many uses within aircraft systems, form some simple circuits controlling landing gear selection to complex circuits within systems controlling navigation and system operation. Figure 24 shows a simple logic circuit for an aircraft landing gear system.

Landing Gear Logic Circuit Figure 24

+vTHROTTLE

SWITCH

DOWNRIGHT MAIN GEARDOWN SWITCH

LEFT MAIN GEARDOWN SWITCH

NOSE GEARDOWN SWITCH

+v

+v

+v

WARNINGHORN



ENGINEERS


1.13.1 CIRCUIT OPERATION In order for the “DOWN” light to illuminate, all three landing gear legs must be down and locked, for this function an “AND” gate is used. If all three gears are not down and locked and the throttle is moved back to approach, then the “NOR” gate will activate the horn to warn the crew that they have not selected the gear “DOWN”, with the throttle at approach. 1.13.2 ENGINE STARTING LOGIC CIRCUIT OPERATION The logic circuit at Figure 25 details the various means of starting an engine.

Engine Starting Logic Circuit

Figure 25

OR

OR

ANDAND

AND

AND

2-3 VALVE

PNEUMATIC OVER-PRESSURE (ENG 3)

ENG 3 AIR

No 2 ENGINEGROUND PNEUMATICCONNECTION 1

PNEUMATIC OVER-PRESSURE (ENG 1)

ENG 1 AIR

1-2 VALVE

GROUND PNEUMATICCONNECTION 2

APU LOAD CONTROL VALVEAUXILIARY POWER UNIT (APU)



ENGINEERS


1.14 FOKKER 50 MINI AIDS 1.14.1 TAKE OFF REPORT A take-off report is automatically generated under specific conditions. These are:

GND/FLT switch is in the “Flight” condition (Logic 0). IAS >60kts. Propeller running with at least 675 RPM.

When these conditions are met, a time delay of 5 seconds ensures the aircraft is airborne sufficiently to make a report with relevant “Take-off” information. Figure 26 shows the layout of F50 Mini Aids take-off report.

F50 Mini Aids Take-off Report Figure 26

TAKE-OFFREPORT

PROP 1 > 675 RPMPROP 2 > 675 RPM

IAS > 60 kts

(GND/FLT) FLT = 0

5 SECS

TAKE-OFFREPORT

NONVOLITILEMEMORY

NONVOLATILEMEMORY



ENGINEERS


1.14.2 STABLE CRIUSE REPORTS There are two stable cruise reports, Stable Cruise 1 and Stable Cruise 2. The mini AIDS makes these reports under different conditions. The conditions of stable cruise 2 are more critical than the conditions of stable cruise 1. Both cruise reports require the need for the following conditions:

Altitude of at least 8,000 ft IAS of at least 145 kts. No change in the Air Conditioning system. Both pressure regulating shut-off valves are open (or

bleed air valves closed). In addition Stable cruise 1 requires the following conditions for automatic report generation.

Air temperature may only vary within 2°C. Altitude may only vary within 300 ft. IAS may only vary within 3 kts.

These variations may not exceed these limits for a time period of 64 seconds. The more critical conditions for an automatic stable cruise 2 report generation are:

Altitude may only vary within 100 ft. IAS may only vary within 2 kts. Both high and low-pressure turbines may not exceed a

variation in RPM of more than 0.5%. Both torque forces of the engines may not exceed a variation

of 1%. The mini AIDS also monitors the stable cruise 2 variation for a time period of 64 seconds.



ENGINEERS


Figure 27 shows a block schematic diagram of the mini AIDS cruise reporting.

F50 Mini AIDS Block Schematic Figure 27

NO

NVO

LATI

LEM

EMO

RY

TIM

EDE

LAY

2X32

SEC

15 M

INCO

UNTE

RDE

LAY

STA

BLE

CR

UIS

E1

STA

BLE

CR

UIS

E2

TIM

EDE

LAY

2X32

SEC

TIM

EDE

LAY

X SE

C

STA

BLE

CRUI

SE2

STA

BLE

CRUI

SE1

ENA

BLE

EN

GIN

E TO

RQ

UE

<±

1%

HIG

H P

RE

SS

TUR

B <

± 0

.5%

RP

M

A

IRS

PEE

D

<±

2 kt

s

A

LTIT

UD

E

<±

100

ft

PR

ESS

UR

E R

EGU

LATI

ON

SH

UT

OFF

VA

LVES

OP

EN

NO

CH

AN

GE

AIR

CO

ND

A

IRS

PEE

D

145

kts

A

LTIT

UD

E

> 8

000

ft

A

LTIT

UD

E <

± 30

0 ft

AIR

TEM

P <

± 2

ºC

A

IRS

PEE

D <

± 3

kts

PWR

INTE

RRU

PT

TIM

ED

ELA

Y3

SEC

TIM

ED

ELA

Y30

SEC

TIM

EDE

LAY

X 1

SEC

FLIG

HT/

GRO

UND

GR

OU

ND =

1

LAND

ING

MO

DE

ON

GR

OU

ND M

ODE

RE

SET

AFT

ER

LAN

DIN

G

WR

ITE

INH

IBIT

AFT

ER

RE

POR

TS

TOR

EAG

E

COLL

ECTE

DIN

FOR

MAT

ION

COLL

ECTE

DIN

FOR

MAT

ION



ENGINEERS


1.14.3 OPERATION So that the aircraft first meets the conditions for stable cruise 1, the mini AIDS collects the stable cruise1 report information but does not store it in the non-volatile memory. A 15-minute counter starts to count at the moment the aircraft meets the stable cruise 1 conditions. When the aircraft meets the more critical condition of the stable cruise 2 within the 15 minutes stable cruise 1 is counting, the mini AIDS stores the stable cruise 2 information in the non-volatile memory. When the aircraft does not meet the stable cruise 2 conditions within the 15 minutes, the mini AIDS finally stores stable cruise 1 into the non-volatile memory. If the aircraft does not fly for a total of 15 minutes in a stable cruise 1 condition the mini AIDS stores the stable cruise 1 report in the landing phase 33 seconds after touchdown. After storage of a report 1 or 2, further stable cruise reports are inhibited for that flight. There is however an exception;

After a power interrupt, the mini AIDS stores the collected stable cruise 1 report in the non volatile memory but does not inhibit a new storage of a stable cruise 1 or 2.

To retrieve the data within the non-volatile memory, a data collector unit, or Laptop computer downloads the data.

DIGITAL TECHNIQUES /EIS BASIC COMPUTER STRUCTURE PAGE 1 of 29


ENGINEERS


1 BASIC COMPUTER STRUCTURE A computer is an electronic device, which can accept and process data by carrying out a set of stored instructions in sequence. This sequence of mathematical and logic operations is known as a Program. The computer is constructed from electronic circuits, which operate on an ON/OFF principle. The data and instructions, used in the computer, must therefore be in logical form. The computer uses the digits "1" and "0" of the binary numbering system to represent "OFF" and "ON". All data and program information must, therefore, be converted into binary form, before being fed into the computer circuitry. One of the most important characteristics of a computer is that it is a general-purpose device, capable of being used in a number of different applications. By changing the stored program, the same machine can be used to implement totally different tasks. In general, aircraft computers only have to perform one particular task so that fixed programs can be used. 1.1 ANALOGUE COMPUTERS A computer is basically a problem-solving device. In aircraft radio systems the problem to be solved is concerned with navigation, in that given certain information, such as range and bearing to a fixed known point, steering commands need to be computed to fly the aircraft to the same, or some other fixed point. Since the input and output information is continuously changing during flight, analogue computation provides an obvious means of solving the navigation problems.



ENGINEERS


A block schematic diagram of an analogue computer is shown in Figure 1.

Analogue Computer Block Diagram

Figure 1

The input devices are radio sensors such as VOR, DME, Omega, ADF, Doppler, Loran, Decca, ILS, and non-radio sensors such as the Air Data Unit and Inertial Navigation System. The output of such sensors will be electrical analogues of the quantities being monitored. The electrical signals contain the necessary information needed to solve the navigation problem, the solution being achieved by the computer. The computer consists of a variety of analogue circuits such as summing amplifiers, integrators, comparators, sine cosine resolvers, servo systems, etc. The patching network determines the way, in which the analogue circuits are interconnected, which will be such as to achieve the required outputs for given inputs. There is a disadvantage of analogue computers in that different patching is needed for different applications. Thus aircraft analogue computers are purpose built to solve one particular problem and as such usually form an integral part of a particular equipment.

INPUTDEVICES

OUTPUTDEVICES

ANALOGUECOMPUTINGELEMENTS



ENGINEERS


This lack of flexibility, together with limited accuracy and susceptibility to noise and drift, has led to the introduction of digital computers, made possible by integrated circuits. Even so, the analogue computer, or rather analogue computing circuits, are still extensively used because as stated above, the sensors produce analogue signals. 1.2 ANALOGUE COMPUTER EXAMPLE Consider an aircraft approaching a DME beacon. The distance to go is given as an electrical analogue signal at the output of the aircraft's DME equipment. By using an analogue computer, this signal can be used to provide an indication to the pilot of his ground speed. As the input signal represents distance, a sample of change in distance divided by the lapsed time will provide ground speed. A suitable block diagram to carry out this calculation is shown in Figure 2.

Computing Groundspeed from 'Distance to Go'

Figure 2

DME O/PDISTANCE

TO GO

GROUNDSPEED

INDICATOR

ANALOGUECOMPUTING

DISTANCE

÷TIME

TIMING



ENGINEERS


1.3 DIGITAL COMPUTERS In the digital computer there are basically two types of input, namely Instructions, and Data from the various radio and non-radio sensors, which will be referred to collectively as information. Information must of course, be coded into a form, which the rest of the computer can understand, such as digital form. The essential components of a digital computer are shown in Figure 3.

Digital Computer Block Diagram Figure 3

CONTROL

ARITHMETIC

MEMORYINPUT

CENTRAL PROCESSOR UNIT (CPU)

OUTPUT



ENGINEERS


Coded information is passed to the memory in which it is stored until needed by the other units. The memory is divided into a large number of cells, each of which can store a word representing a piece of information. Each cell has a unique address, through which access to the information contained within that cell can be obtained. There are usually two types of memory, long term and temporary stores. The latter, often termed registers, will be used to hold intermediate results in calculations and data, which is to be processed next in the calculating sequence. The arithmetic unit performs the actual arithmetic operations called for by instructions. It can be compared with a calculator. The results of the calculations must be displayed in a suitable form easily interpreted by the pilot. This is the function of the output unit, which reads from the store. The control unit directs the overall functioning of the computer according to the program of instructions in store. This program is known as software as opposed to the actual circuitry, which is termed hardware. Although control is drawn as a separate unit in the functional block diagram, the control hardware, which comprises timing circuits and electronic switches, is spread throughout the computer. Information is read into the appropriate address of the store under the control of the software. In aircraft navigation applications, incoming data from sensors updates the contents of the store at a rate dependent upon the timing of the computer control. The control acts on instructions held in store in the appropriate sequence. The basic task will be to transfer data from store to the arithmetic unit, to carry out the necessary calculations using registers to store the intermediate results, then writing the final result into the store. The final control function will be to transfer data from store to the output as a result of built in instructions, or on specific instructions from the pilot. This process of input - store - calculate - store - output is carried out sequentially in accordance with software requirements.



ENGINEERS


1.4 BUSES It can be seen from Figure 4 that there are three buses - the data bus, the address bus, and the control bus. Each bus consists of a group of parallel wires. The data bus transfers data between memory, CPU and I/O units, under the control of signals sent through the control bus. For example, if data is to be transferred (sent) from the CPU to a memory location, the control unit within the CPU places an output instruction on the CPU, and write instruction on the memory unit. When the data arrives at the memory, it must be written into the memory at a given address. The address is already present, having been sent by the CPU along the address bus. Hence, data is stored at the memory address given. Note that if the transfer had been from the CPU to an I/O device, the address of the I/O device would have been given. The address bus is one-way only. The control bus usually has one set of wires for input sensing lines, and one set for output controls. Data buses are usually bi-directional; that is, data is either transferred, or fetched along the same set of wires. The control unit usually decides in which direction data will travel. If there are several peripherals, and these all wish to use the CPU at the same time, some method of priority must be established. There are various ways of achieving this. One method uses the control unit to select the lucky peripheral, whilst another method lets the peripherals themselves automatically decide which peripheral takes control.



ENGINEERS


Computer Buses Figure 4

CPU

CONTROL &ARITHMETIC

UNIT

CONTROL &ARITHMETIC

UNIT

I/P

O/PMEMORYMEMORYINPUT/OUTPUT

UNITINPUT/OUTPUT

UNIT

CONTROL BUS

DATA BUS

CLOCKCLOCK

ADDRESS



ENGINEERS


1.5 INPUT/OUTPUT (I/O) UNIT This unit provides the interface between the computer and the computer peripherals. A computer peripheral is any unit, which is attached to, but is not part of, the computer - e.g. visual display units, teleprinters, etc. A simple computing system may have only one input and one output. In such cases, an analogue-to-digital converter (ADC) may suffice for the input, and a digital-to-analogue converter (DAC) for the output. Alternatively, complex-computing systems can literally service thousands of peripherals. Figure 5 illustrates a simple I/O unit. The I/O unit can be described as a fan-out (and fan-in) device. The computer's 8-bit bi-directional data bus can be connected to port 1, 2 or 3. The port chosen is dependent upon the address, on the address bus. The system illustrated allows three peripherals to communicate with the computer. Only one peripheral at a time can send data to the computer, or receive data from the computer. However, this is not a problem, because the computer works very much faster than the peripheral, and hence, it appears that the computer services all three peripherals

simultaneously.

Input/Output Unit Figure 5

INPUT/OUTPUTUNIT

COMPUTER DATA BUS (8 BITS)

ADDRESS BUS

PERIPHERAL 1 PERIPHERAL 2 PERIPHERAL 3

PORT 1 PORT 2 PORT 3

CONTROL BUS



ENGINEERS


Peripherals can have either serial or parallel outputs. Also, as stated previously, peripherals work at a much slower speed than that of the computer. The I/O system must, therefore, be capable of 'conditioning' the data received from the peripherals to a form, which is readily digestible by the computer, and vice versa. 1.6 MEMORY The memory unit is used for the storage of binary coded information. Information consists of instructions and data where:

• Instructions are the coded pieces of information that direct the activities of the CPU.

• Data is the information that is processed by the CPU.

The memory hardware contains a large number of cells or locations. Each location may store a single binary digit or a group of binary digits. The cells are grouped so that a complete binary word is always accessed. Word length varies typically from 4-bits up to 64-bits depending upon machine size. Each location in the memory is identified by a unique address, which then allows access to the word. Consequently, to obtain information from the memory, the correct address must be placed onto the address bus. There are fundamentally two types of memory - primary memory and secondary memory. Primary memory is essential; no computer can operate without this. Secondary memory is necessary to supplement, or back, the primary memory on large computing systems; hence, it is often called backing memory. There are two types of semi-conductor primary memory: ROM (Read Only Memory) and RAM (Random Access memory). Both types employ solid state circuitry, and are packaged in IC form.



ENGINEERS


Figure 6 shows how these primary memories are connected to a simple computer bus.

ROM and RAM Connection to Buses Figure 6

1.7 RANDOM ACCESS MEMORY (RAM) The RAM-type memory will allow data to be written into it, as well as read

from it. With very few exceptions, RAMS lose their contents when the power is

removed and are thus known as “Volatile” memory devices. All computers use RAM to store data and programs written into it either from keyboard, or external

sources such as magnetic tape/disk devices. RAMs are often described in terms of the number of bits, i.e. 1s and 0s, of

data that they hold, or in terms of the number of data words, i.e. groups of bits,

they can hold. Thus a 16384 bit ram can hold 16384 1s and 0s. This data could

be arranged as 16384 1-bit words, 4096 4-bit words or 2084 8-bit words. Semiconductor memories vary in size, e.g. 4K, 64K, 128K, etc. Hence we are

DATA BUS

ROM RAM

MEMORY ADDRESS REGISTER& CHIP SELECT DECODER

ADDRESS BUS

TO INPUT/OUTPUT

DEVICE

TO INPUT/OUTPUT

DEVICE

FROMCPU

TOCPU

NOTE: CONTROL BUSOMITTED FOR SIMPLICITY



ENGINEERS


using K defined as:

K =210 = 1024

Thus a 16K memory has a storage capacity of 16 X 1024 = 16384 words, a 128K

memory 0f 1310672 words and so on. There are two main members of the RAM family:

Static RAM. Dynamic RAM.

The essential difference between them is the way in which bits are stored in the RAM chips. In a static RAM, the bits of data are written in the RAM just once and then left until the data is either read or changed. In a dynamic RAM, the bits of data are repeatedly rewritten in the RAM to ensure that the data is not forgotten. 1.7.1 STATIC RAM Flip-Flops are the basic memory cells in a static RAM. Each flip-flop is based on either two bipolar transistors or two Metal Oxide Semiconductors Field-Effect Transistors (MOSFETS). As many of these memory cells are needed as there are bits to be stored. Thus, in a 16K-bit static memory there are 16384 flip-flops, i.e. 32768 transistors. All these transistors are accommodated on a single silicon chip approximately 4mm2. Figure 7 shows a basic memory cell in a static RAM

Static RAM Cell Figure 7

TR1 TR2

CELL SELECT LINE

+5V

LOGIC 1OUTPUT/INPUT

LOGIC 0OUTPUT/INPUT

16K MEMORY= 16,384 FLIP-FLOPS

= 32,768 TRANSISTORS



ENGINEERS


1.7.2 7489 TTL RAM DEVICE The 7489 TTL Ram package has 64 memory cells, each cell is capable of holding a single bit of data. The cells are organised into locations, and each location is capable of holding a 4-bit word. Thus the 7489 is capable of storing 4-sixteen 4-bit words, i.e. four memory cells are used at each location. Figure 8 shows the memory organisation of the 7489 static RAM.

7489 RAM Device Figure 8

Each location is identified by a unique 4-bit address so that data can only be written or read from that location. The number of words stored in the memory determines the length of the address word. I.E. 16 = 24.

1 3 4 5 6 7 82

16 14 13 12 11 10 915

A

ME WE D1 S1 D2 S2

B C D D4 S4 D3

S3

Vee

Vcc

DA

TA

INPU

T 1

DA

TA

INPU

T 2

DA

TA

INPU

T 3

DA

TA

INPU

T 4

SEN

SEO

UTP

UT

1

SEN

SEO

UTP

UT

2

SEN

SEO

UTP

UT

3

SEN

SEO

UTP

UT

4ADDRESSB,C & D

AD

DR

ESS

A MEM

WR

ITE

ENABLES

4 BITDATA OUT

FOUR MEMORY CELLS

1 0 1 10 1 1 00 0 1 10 1 0 0

16 LOCATIONSEACH HOLDING

FOUR BITS

0

1

2

3

4

5

6

7

8

9

10

11

12

13

14

15

READ/WRITESIGNALS

4 BITDATA IN

1 1 0 1

4 BITADDRESS

1 1 10

1 1 10 1 1 10



ENGINEERS


1.8 READ ONLY MEMORY (ROM) The problem with RAM is that its memory is volatile, i.e. it loses all its data when the power supply is removed. A non-volatile memory is a permanent memory that never forgets its data. One type of non-volatile memory is the Read Only Memory (ROM). A ROM has a pattern of 0s and 1s imprinted in its memory by the manufacturer. It is not possible to write new data into a ROM, which is why it is called a Read-Only Memory. The organisation of data in a ROM is similar to that of a RAM. Thus a 256-bit ROM might be organised as a 256 X 4-bit memory, and so on. The ROM may be regarded as the “Reference Library” of a computer. 1.9 MAGNETIC CORE MEMORY This type of memory is used extensively in airborne digital systems, although integrated circuits are being developed with most modern aircraft systems. This system works by a Ferro-magnetic material will become magnetized if placed in the proximity to an electric current. Each bit in the magnetic core memory is a ferrite ring in which a magnetic field can be induced by a current flowing in a wire. Figure 9 shows typical ferrite ring for storing a single bit.

Ferrite Ring Memory Figure 9



ENGINEERS


Although the wire carrying the current is wound round the ring, the same effect is obtained if the wire passes through the ring. This is a more convenient way to set the magnetic state of each ring when a plane of cores is built. The advantage of this type of memory is that when the power is removed it holds its state, i.e. it is a non-volatile memory. A matrix of cores containing 16 bits of information is shown in Figure 10.

16 Bit Ferrite Memory Figure 10

Y1 Y2 Y3 Y4

X1

X2

X3

X4

X1 & Y1 CURRENTMAGNETIZES THE

CORE

CURRENT ISINSUFFICIENT

TO MAGNETIZECORE WITH ONLY

ONE CURRENT



ENGINEERS


1.10 PROGRAMMABLE ROM (PROM) The user can program a PROM after purchase. Each memory bit element in the PROM contains a nichrome or silicon link that acts as a fuse. The user can selectively 'but out' or 'blow' these fuses by applying pulses of current to the appropriate pins of the IC. A memory element with a non-ruptured fuse stores a 1 and a ruptured fuse stores a 0. The programming is irreversible, so it must be right first time. Figure 11 shows the circuit for a PROM.

PROM Circuit Figure 11

PROMs are capable of high operating speeds, but consume a relatively large amount of power. However, since they are non-volatile, they can be switched off when not being accessed.

+5V +5VSENSE(HIGH)

SENSE(LOW)

“0”FUSELINK

“1”NO FUSE

LINK

LOGIC0

LOGIC1

0V

ADDRESSLINE

TR1

TR2



ENGINEERS


1.11 ERASABLE PROGRAMMABLE READ ONLY MEMORY These memory devices can be programmed, erased and then reprogrammed by the user as often as required. In some devices, the information can be erased by flooding them with ultraviolet light, whilst in others, voltages are applied to the appropriate pins of the device. 1.12 ELECTRICAL ALTERED READ ONLY MEMORY This memory device combines the non-volatility of the ROM with the electrically alterable characteristic of the RAM. It is, therefore, considered as a non-volatile RAM.



ENGINEERS


1.13 THE CENTRAL PROCESSING UNIT (CPU) The CPU is the heart of any computing system. It executes the individual machine instructions, which make up a program. The CPU is formed from the following interconnected units:

1. ALU (Arithmetic Logic Unit).

2. Registers.

3. Control Unit. These units are shown as part of a computer system in Figure 13.

Central Processing Unit Figure 13

MEMORY(REGISTERS)

ARITHMETICUNIT

COMPUTER

HIGHWAY

MEMORY

INPUTOUTPUT

UNIT

CPU

CONTROL

CLOCK



ENGINEERS


ALU. This is where the mathematics and logic functions are implemented. It is not essential for the ALU to subtract, divide, or multiply, as these functions are easily achieved by using addition in conjunction with 2's complement arithmetic. However, more powerful processors include sophisticated arithmetic hardware capable of division, multiplication, fixed and floating point arithmetic etc. Large processors also employ parallel operation for high speed. Registers. These are temporary storage units within the CPU. Some registers have dedicated uses, such as the program counter register and the instruction register. Other registers may be used for storing either data or program information. Figure 14 illustrates the principal registers within the CPU.

The CPUs Internal Registers Figure 14

MEMORY

MEMORYADDRESSREGISTER

INPUTOUTPUT

ADDRESSDECODE

PROGRAMCOUNTERREGISTER

INSTRUCTIONDECODE

REGISTER

CONTROLUNIT

TIMING

STATUS FLAGREGISTER

TEMPORARY REGISTER

ACCUMULATORREGISTER

INTERNAL

HIGHWAY

PORT 1

PORT 2

PORT 3



ENGINEERS


Program counter register. The instructions that comprise a program are stored in the computer's memory. Consequently, the computer must be able to sequentially access each instruction. The address of the first instruction is loaded into the program register, whereupon the instruction is fetched and loaded into another register, appropriately called the instruction decode register. Whilst the CPU is implementing the fetched instruction (e.g. Add, Shift, etc), the program counter register is incremented by 1 to indicate the address of the next instruction to be executed. This system, therefore, provides sequential execution of a program, provided that the program is written and stored sequentially in the memory. The instruction decode register. As stated above, the program counter register locates the address at which the next instruction is to be found. The instruction itself is then transferred from memory into the instruction decode register. As the name implies, this register also incorporates a decoder. the output from the decoder places the necessary logic demands onto the ALU - i.e. shift, add, etc. The accumulator register. This register is really part of the ALU, and it is the main register used for calculations. Consequently, it always stores one of the operands, which is to be operated on by the ALU. The other operand may be stored in any temporary register. The status register. This register is a set of bistables which operate independently of each other. The bistables independently monitor the accumulator to detect such occurrences as a negative result of a calculation, a zero result, an overflow, etc. When such an occurrence arises, the output of the respective bistable is set (logic 1). It is then said to signal or flag the event. It is this register that gives a computer its decision-making capability. For example, if the result of a calculation in a navigational computer is zero, the program could instruct the autopilot to hold its present course. Alternatively, if the zero flag was not set, the computer would then decide to take corrective action. There are many other registers within a CPU, some of which are general-purpose registers. These can be used to store operands or intermediate data within the CPU, thus eliminating the need to pass intermediate results back and forth between memory and accumulator. The control unit. This unit is responsible for the overall action of the computer. It coordinates the units, so that events take place in the correct sequence and at the right time. Because it is responsible for timing operations it includes a clock (normally crystal controlled), so that instructions and data can be transferred between units under strict timing control (synchronous operation). The crystal and the clock generator may either be contained within the CPU, or supplied as separate components.



ENGINEERS


1.14 THE MICROPROCESSOR The three fundamental units, which comprise a CPU, have now been discussed in general terms. So too has a microprocessor, because a microprocessor can be defined as the central processing part of a computer contained within an IC (Integrated Circuit). Figure 15 illustrates how a microprocessor can be used as part of a microcomputer. The microprocessor is small, lightweight, and relatively cheap when compared to any CPU. But it is also relatively slow, capable of processing only hundreds of instructions per second, compared to a large CPU which can process thousands of instructions per second, or a very fast CPU which can process millions of instructions per second (mips). However, many computing applications can tolerate the relative speed disadvantage of the microprocessor hence, its popularity. Microprocessors are typically available in 4, 8 and 16-bit word lengths.

Elementary Microcomputer Figure 15

MICROPROCESSOR(CPU) COMPUTER HIGHWAY

ROM

RAM

OUTPUT

INPUT

INPUT/OUTPUTPORTS



ENGINEERS


The preceding paragraphs defined a microprocessor as a CPU within an IC. This is true of all microprocessors; however, many go beyond this 'minimum' definition. Microprocessors for machine control (lathes, robots, petrol pumps, etc) often incorporate ADC and DAC on the same chip, plus a small amount ROM and RAM. Some microprocessors incorporate all the elements of a total computing system: I/O, ROM, RAM and CPU. Manufacturers designate these as single chip microcomputers. Obviously, their computing power is somewhat limited, because there is a limited amount of space available in just one IC. 1.15 AIRBORNE DIGITAL COMPUTER OPERATION 1.15.1 FLIGHT MANAGEMENT SYSTEM (FMS) A Flight Management System (FMS) is a computer-based flight control system and is capable of four main functions:

1. Automatic Flight Control.

2. Performance Management.

3. Navigation and Guidance.

4. Status and Warning Displays. The FMS utilizes two Flight Management Computers (FMC) for redundancy purposes. During normal operation both computers crosstalk; that is, they share and compare information through the data bus. Each computer is capable of operating completely independently in the event of one failed unit. The FMC receives input data from four sub-system computers:

1. Flight Control Computer (FCC).

2. Thrust Management Computer (TMC).

3. Digital Air Data Computer (DADC).

4. Engine Indicating & Crew Alerting System (EICAS). The communication between these computers is typically ARINC 429 data format. Other parallel and serial data inputs are received from flight deck controls, navigation aids and various airframe and engine sensors. Figure 16 shows a block schematic of the FMS.



ENGINEERS


Flight Management System (FMS)

Figure 16

The FMC contains a large nonvolatile memory that stores performance and navigation data along with the necessary operating programs. Portions of the nonvolatile memory are used to store information concerning: a. Airports. b. Standard Flight Routes.

c. Nav Aid Data. Since this information changes, the FMS incorporates a “Data Loader”. The data loader is either a tape or disk drive that can be plugged into the FMC. This data is updated every 28 days.

AFCAS

NAVIGATIONAL SYSTEMS

FMC 2FMC 1

EFISEFIS

FMSCDU 1

FMSCDU 2

EICAS

TMS



ENGINEERS


Figure 17 shows the layout of FMC memory.

FMC Memory Locations. Figure 17

Variable parameters for a specific flight are entered into the FMS by either data loader, or “Control Display Unit” (CDU). This data will set the required performance for least-cost or least-time en-route configuration.

NAV DATABASE

OPERATIONPROGRAM

PERF DATA

BUFFER

STORAGE

FMC

INITIAL AIRLINEBASE & 28 DAY

UPDATES

RAW DATA FORCOMPUTATIONS

ROLLCHANNEL

PITCHCHANNEL

MODETARGET

REQUESTS

DISPLAYS

AILERONCONTROL

ELEVATORCONTROL

THRUSTLEVER

CONTROL

REQUESTEDROUTE

LATERALVERTICAL

MEMORY STORAGE16 BIT WORDS



ENGINEERS


1.15.2 FMS CONTROL/DISPLAY UNIT (CDU) The CDU provides a means for the crew to communicate with the FMC. It contains pushbutton key controllers and a display screen. The keys are of two types: -

1. Alphanumeric keys, which can be used to enter departure and destination points and also Waypoint if not already stored on tape; they will also be used if the flight plan needs to be changed during the flight.

2. Dedicated keys, which are used for specific functions

usually connected with display. For example, by using the appropriate key the pilot can call up flight plan, Waypoint data, flight progress, present position, etc.

When, for example, a departure point is entered using alphanumeric keys, the information is often held in a temporary register and displayed to the pilot; this is known as a scratchpad display. Once the pilot has checked the information is correct, he can enter the data into the computer store by pressing the appropriate dedicated key typically labelled "Load" or "Enter".



ENGINEERS


Figure 18 shows an FMS Control/Display Unit (CDU).

FMS CDU. Figure 18

During a normal flight, the FMS sends navigation data to the EFIS, which can then display a route map on the EHSI. If the flight plan is altered by the flight crew en-route, then the EHSI map will change automatically.

DISPLAYSCREEN

1 2 3

4 5 6

7 8 9

0

EXEC

MSG

CLEAR

A B C D E F GH I J K L M NO P Q R S T UV W X Y Z /

PPOS NEXTPHASE PERF

FUEL AIRPORTS

DATA FIX

START

DIR

HDGSEL

ENGOUT

SPECF-PLN

ALPHANUMERICKEYPADFUNCTION

SELECTKEYS

LINESELECT

KEYS

DISPLAYBRIGHTNESS

CONTROL



ENGINEERS


1.16 COMPUTER INPUT Figure 19 shows input information for a typical airborne digital computer.

Computer Inputs

Figure 19

The sensors in Figure 19 develop analogue electrical signals representing: -

Bearing and distance to fixed point (VOR/DME). Hyperbolic co-ordinates (Omega). Ground speed and drift angle (Doppler). Aircraft heading (Compass), etc.

These analogue signals must be converted into digital signals before being fed to the computer memory. ADCs, which may be an integral part of the

REGISTERSSEQUENCING

&ADDRESSING

SENSORS:

VOR/DME - OMEGADOPPLER - COMPASS

ETC

MAGNETIC TAPECASSETTE/CARTRIDGE

PUSH BUTTONCONTROLLER

ALPHANUMERICDEDICATED

MAGNETIC CARD READER

TO CONTROL

FROM CONTROLTO

STORE

FROM CONTROL

AD



ENGINEERS


sensor equipment achieve this, or alternatively a converter unit may be installed, which carries out all necessary analogue to digital conversion. 1.16.1 COMPUTER OUTPUT Many different kinds of output device are used, including traditional devices such as relative bearing indicators and steering indicators. With these, suitably designed digital to analogue converters must be used. Similar outputs could be fed to an autopilot. Digital read out can be obtained by use of hybrid (digital and analogue) servo systems, which position an output counter drum or alternatively by use of 7 segment indicators. A ROM, which has the wired in program to convert from binary code to the appropriate drive, drives the segments, which may be light emitting diodes (LED) or liquid crystals (LCD). Cathode ray tubes (CRT) are being increasingly used as output devices both for display of alphanumeric information and, less commonly, electronic maps. CRTs are essentially analogue devices and as such require DACs, which will provide the necessary fairly, complicated drives. Moving map displays may also be used as a means of presenting navigation information to the pilot. The map itself may be an actual chart fitted on rollers, or alternatively projected film. Closed loop servos, which drive the map, are fed from the computer via DACs. 1.17 COMPUTER TERMS 1.17.1 ACCESS TIME The time interval required to communicate with the memory, or storage unit of a digital computer, or the time interval between the instant at which the arithmetic unit calls for information from the memory and the instant at which this information is delivered. 1.17.2 ADDRESS A name or number that designates the location of information in a storage or memory device.



ENGINEERS


1.17.3 COMPUTER LANGUAGE A computer language system is made up of various sub routines that have been evaluated and compiled into one routine that the computer can handle. FORTRAN, COBOL and ALGOL are computer language systems of this type. 1.17.4 CORE MEMORY A programmable, random access memory consisting of many ferromagnetic cores arranged in matrices. 1.17.5 DATA PROCESSING The handling, storage and analysis of information in a sequence of systematic and logical operations by a computer. 1.17.6 DECODER A circuit network in which a combination of inputs produces a single output. 1.17.7 FLOPPY DISC A backing storage facility for microcomputer systems. 1.17.8 INSTRUCTION A machine word or set of characters in machine language that directs a computer to take a certain action. Part of the instruction specifies the operation to be performed, and another part specifies the address. 1.17.9 LANGUAGE A defined group of representative characters of symbols combined with specific rules necessary for their interpretation. The rules enable the translation of the characters into forms (such as digits) which are meaningful to a machine. 1.17.10 MACHINE CODE A program written in machine code consists of a list of instructions in binary form to be loaded into the computer memory for the computer to obey directly.



ENGINEERS


1.17.11 MAGNETIC CORE A form of storage in which information is represented by the direction of magnetization of a core. Advantage of this kind of memory store is it will retain its contents even if electrical power is removed (Non-volatile). 1.17.12 PROGRAMME A plan for the solution of a problem. A precise sequence of coded instructions or a routine for solving a problem with a computer. 1.17.13 REAL TIME The actual time during which a physical process takes place and a computation related to it, resulting in its guidance: or, ‘As it happens’. 1.17.14 ROUTINE A set of coded instructions that direct a computer to perform a certain task. 1.17.15 TIME SHARING Using a device, such as a computer, to work on two or more tasks, alternating the work from one task to the other. Thus the total operating time available is divided amongst several tasks, using the full capacity of the device. 1.17.16 WORD (OR BYTE) An ordered set of characters which has at least one meaning and is stored, transferred, or operated upon by the computer circuits as a unit.

DIGITAL TECHNIQUES /ELECTRONIC INSTRUMENT SYSTEM FIBER OPTICS PAGE 1 of 18


ENGINEERS


1 FIBRE OPTICS Light travels in straight lines, even though lenses and mirrors can deflect it, light still travels in a straight line between optical devices. This is fine for most purposes; cameras, binoculars, etc. wouldn’t form images correctly if light didn’t travel in a straight line. However, there are times when we need to look round corners, or probe inside places that are not in a straight line from our eyes. That is why “FIBRE OPTICS” have been developed. The working of optical fibres depend on the basic principle of optics and the interaction of light with matter. From a physical standpoint, light can be seen either as “Electromagnetic Waves” or as “Photons”. For optics, light should be considered as rays travelling in straight lines between optical elements, which can reflect or refract (bend) them. Light is only a small part of the entire spectrum of electromagnetic radiation. The fundamental nature of all electromagnetic radiation is the same: it can be viewed as photons or waves travelling at the speed of light (300,000 km/s) or 180,000 miles/sec). 1.1 REFRACTIVE INDEX (N) The most important optical measurement for any transparent material is its refractive index (n). The refractive index is the ratio of the speed of light (c) in a vacuum to the speed of light in the medium: The speed of light in a material is always slower than in a vacuum, so the refractive index is always greater than one in the optical part of the spectrum. Although light travels in straight lines through optical materials, something different happens at the surface. Light is bent as it passes through a surface where the refractive index changes. The amount of bending depends on the refractive indexes of the two materials and the angle at which the light strikes the surface between them. The angle of incidence and refraction are measured not from the plane of the surfaces but from a line perpendicular to the surfaces. The relationship is known as “Snells Law”, which is written; ni sin I = nr sin R, where ni and nr are the refractive indexes of the initial medium and the medium into which the light is refracted. I and R are the angles of incidence and refraction.



ENGINEERS


Figure 1 shows an example of light going from air into glass.

Snell’s Law on Refraction (Air into Glass) Figure 1

AIR

GLASS

LIGHT

I

ANGLE OFINCIDENCE

RANGLE OF

REFRACTION

NORMAL LINEPERPENDICULAR

TO GLASS SURFACE



ENGINEERS


Snell’s law indicates that refraction can’t take place when the angle of incidence is too large. If the angle of incidence exceeds a critical angle, where the sine of the angle of refraction would equal one, light cannot get out of the medium. Instead the light undergoes total internal reflection and bounces back into the medium. Figure 2 illustrates the law that the angle of incidence equals the angle of reflection. It is this phenomenon of total internal reflection that keeps light confined within a fibre optic.

Critical Angle Figure 2

41.9º

1.5 SIN 41.9º= 1.00174

θº1 θº2

TOTALINTERNAL

REFLECTION

θº1 = θº2



ENGINEERS


1.2 LIGHT GUIDING The two key elements of an optical fibre are its “Core” and “Cladding”. The core is the inner part of the fibre, through which light is guided. The cladding surrounds it completely. The refractive index of the core is higher than that of the cladding, so light in the core that strikes the boundary with cladding at a glancing angle is confined in the core by total internal reflection. Figure 3 shows the make up of a fibre optic.

Fibre Optic (Core and Cladding) Figure 3

CLADDING

CORE

LIGHT RAY

LIGHT RAY STRIKES THE CLADDINGAT AN ANGLE GREATER THAN THECRITICAL ANGLE, THEREFORE THELIGHT RAY IS REFLECTED RATHER

THAN BEING REFRACTED.



ENGINEERS


1.3 LIGHT COUPLING Another way to look at light guiding in a fibre is to measure the fibre’s acceptance angle. This angle is the angle within which the light should enter the fibre optic to ensure it is guided through it. The acceptance angle is normally measured as a numerical aperture (NA). The numerical aperture and acceptance angle measurements are a critical concern in practical fibre optics. Getting light into a fibre is known as “Coupling”. When fibre optics were first developed in the 1950s, no one believed that much light could be coupled into a single fibre. Instead they grouped fibres into bundles to collect a reasonable amount of light. Only when “LASERS” made highly directional beams possible did researchers seriously begin to consider using single optical fibres. Figure 4 shows light coupling into a fibre optic and the construction of a fibre optic cable.

Light Coupling (Critical Angle) Figure 4

FIBRE OPTIC CABLE

ARAMID YARN

SEPARATORTAPE

OPTICALFIBRES

FILLERSTRANDS

OPTICALFIBRES

FILLERSTRANDS

SEPARATORTAPE

ARAMIDYARN OUTER

JACKET

ENDVIEW

ACCEPTANCE

ANGLE

LIGHT MUST FALLINSIDE THIS ANGLE

TO BE GUIDED THROUGHTHE CORE



ENGINEERS


1.4 ALIGNMENT Coupling light between fibres requires careful alignment and tight tolerances. The highest efficiency comes when the ends of the two fibres are permanently joined. Temporary junctions between two fibre ends, made by connectors, have a slightly higher loss but allow much greater flexibility in reconfiguring a fibre optic network. Figure 5 shows the problems associated with incorrect alignment.

Fibre Optic Alignment Figure 5

LATERAL MISALIGNMENT

ANGULAR MISALIGNMENT

AXIAL MISALIGNMENT

POOR END FINISH



ENGINEERS


1.5 FIBRE OPTIC CONNECTORS Boeing uses three types of connectors: Type A Connector, Type B Connector and Type C Connector. 1.5.1 TYPE “A” CONNECTOR The type “A” connector has these technical qualities:

• A threaded coupling mechanism.

• A butt type connector with ceramic terminuses.

• The transmission of a light beam from the end of one optical fibre into the end of another optical fibre.

Figure 6 shows example of “A” type receptacle and plug connectors.

Type “A” Connector

Figure 6

FIBRE OPTICCABLE

STRAIN RELIEFBOOT

BACKSHELL

RECEPTACLE

ALIGNMENTHOLE

FIBREOPTIC

SLEEVES

THREADEDCOUPLING

JACKSCREW

PLUG

COUPLINGRING

CERAMICTERMINUSALIGNMENT

PINS



ENGINEERS


1.5.2 TYPE “B” CONNECTOR The type “B” connector has these technical qualities:

• A threaded coupling mechanism.

• An extended beam connector that contains a miniature lens behind a protective window.

• The transmission of a light beam by the miniature lens from

an optical fibre through the protective window to the opposite miniature lens into the opposite fibre optic.

Figure 7 shows example of a “B” type receptacle and plug connectors.

Type “B” Connector

Figure 7

FIBRE OPTICCABLE

STRAIN RELIEFBOOT

BACKSHELL

RECEPTACLE

THREADEDCOUPLING

PLUG

COUPLINGRING

ALIGNMENTPINS

PROTECTIVEWINDOW

ALIGNMENTHOLE



ENGINEERS


1.5.3 TYPE “C” CONNECTOR The type “C” connector has these technical qualities:

• A push-pull coupling mechanism.

• An extended beam connector that contains a miniature lens behind a protective window.

• The transmission of a light beam by the miniature lens from

an optical fibre through the protective window to the opposite miniature lens into the opposite fibre optic.

Figure 8 shows example of a “C” type receptacle connector.

Type “C” Connector Figure 8

MOUNTINGFLANGE

PROTECTIVEWINDOW

BACKSHELL

STRAIN RELIEFBOOT

FIBRE OPTICCABLE

RECEPTACLECONNECTOR



ENGINEERS


Figure 9 shows how the light is transferred in the type B and C connectors using miniature lenses and protective window.

Fibre Optic Connection Figure 9

Coupling losses can cause substantial attenuation. Dead space at the emitter/fibre and fibre/receiver junctions and (unless optically corrected) the beam spreads of 7° associated with semi-conducting lasers, are the usual sources of launching problems. To limit this light loss a ball lens is used. These lenses (within the connector) focus the light into another fibre optic cable or an optical receiver. Mono-made fibres are particularly prone to launching losses because it is difficult to produce an accurate square end. Jointing and cabling, in order to produce longer lengths, are currently receiving development attention.



ENGINEERS


Figure 10 shows example of type “A”, “B” and “C” connectors and identification labels.

Fibre Optic Connectors

Figure 10

BOEING TYPE “A” PLUG CONNECTOR

TYPE “A” PLUGTYPE “A” PLUGCONNECTORCONNECTOR

BACKSHELLBACKSHELL

STRAINSTRAINRELIEFRELIEF

MATE WITHMATE WITHIDENTIFICATIONIDENTIFICATION

SLEEVESLEEVE

ASSEMBLYASSEMBLYIDENTIFICATIONIDENTIFICATION

SLEEVESLEEVE

BOEING TYPE “B” PLUG CONNECTOR

TYPE “B” PLUGTYPE “B” PLUGCONNECTORCONNECTOR

BACKSHELLBACKSHELL



SLEEVESLEEVE


SLEEVESLEEVE

BOEING TYPE “C” PLUG CONNECTOR

TYPE “C” PLUGTYPE “C” PLUGCONNECTORCONNECTOR

BACKSHELLBACKSHELL



SLEEVESLEEVE


SLEEVESLEEVE

MOUNTINGMOUNTINGFLANGEFLANGE



ENGINEERS


1.6 ADVANTAGES OF FIBRE OPTICS Fibre-optic communications systems have a large bandwidth, e.g. 1 GHz. The bandwidth is the maximum rate at which information can be transmitted. It has the benefit of:

♦ Immunity to electromagnetic interference in electrically noisy situations.

♦ High security against 'tapping'.

♦ Much greater flexibility than the majority of waveguides.

♦ Low weight when compared with copper - 60 per cent less.

♦ Ability to resist vibration.

♦ Glass fibres have no fire risk.

♦ Inability to form unwanted earth loops.

♦ Inability to short-circuit adjacent filaments when fractured.

♦ High data capacity (>10Gbits/s with a single fibre).

1.7 DISADVANTAGES OF FIBRE OPTICS

♦ Difficult to join.

♦ No transfer of D.C. Power.



ENGINEERS


1.8 SAFETY When working on Fibre Optic connected equipment, care is required when handling cables. If the equipment is energised, invisible light form the fibre optic cable can be sufficient to cause damage to the eyes. Before the face of the connector is examined either one of these conditions must be satisfied:

• The connectors are disconnected from equipment at both ends of the cable.

• The power to the equipment is set to “OFF”.



ENGINEERS


1.9 BASIC OPERATION 1. The input is converted by the encoder to electrical signals, which represent

either the sound waves of the voice, or the scanning of visible media. 2. The emitter sends out probes of infra-red light corresponding to the

electrical values, in strength and duration. 3. The infra-red light is launched into the fibres, which conduct it to the

receiver. 4. The receiver re-converts the light to electrical values. Figure 14 shows fibre optic connection.

Fibre Optic Connection

Figure 5.10.14

EQUIPMENT

STRAINRELIEF

TYPE “B”PLUG

1" MIN

BENDRADIUS

>1.5" FIBREOPTICCABLE



ENGINEERS


1.10 AIRCRAFT APPLICATIONS 1.10.1 OPTICAL DATA BUS Data transmission systems generally utilise a twisted cable pair as a bus. This has its limitations and fibre optics is under active development as the next step for use in aircraft digital systems. 1.10.2 STANAG 3910 DATA BUS SYSTEM This is the European standard data bus with a 20 Mbit/sec data rate and will enter service with the new Eurofighter 2000. This advanced data bus system provides an evolutionary increase in capability by using MIL STD 1553B as the controlling protocol for high speed (20Mbit/sec), message transfer over a fibre optic network. Figure 17 shows the architecture of the STANAG 3910 data bus system.

STANAG 3910 Data Bus System

Figure 17

FIBRE OPTICSTAR

COUPLER

BUSCONTROLLER

SUBSYSTEM

1

SUBSYSTEM

2

SUBSYSTEM

N

CONTROL &LOW SPEEDDATA BUS

HIGH SPEEDDATA BUS

UPTO 31SUB-SYSTEMS



ENGINEERS


The optical star coupler allows light signals from each fibre stub to be coupled into the other fibre stubs and then to the other sub-systems. The data bus also has the normal operation of the MIL STD 1553B data bus. The USA is developing its own version of a fibre optic data bus system. This is a High Speed Data Bus (HSDB), and uses Linear Token Passing as its controlling protocol. It operates at 50 Mbits/sec and operates to connect up to 128 sub-systems. Figure 18 shows the architecture of the Linear Token Passing High Speed Data Bus (LTPHSDB).

Linear Token Passing High Speed Data Bus

Figure 18

FIBRE OPTICSTAR

COUPLER

SUBSYSTEM

1

SUBSYSTEM

2

SUBSYSTEM

3

SUBSYSTEM

N

UPTO 128SUB-SYSTEMS



ENGINEERS


1.10.3 FLY-BY-LIGHT FLIGHT CONTROL SYSTEM Extensive tests have been carried out using the Fly-by-Light technology. It has huge advantages over the current Fly-by-Wire systems. Fibre optic cabling is unaffected by EMI and has a considerably faster data transfer rate (20 Mbit/sec to 100 Mbit/sec). The systems are also lighter than conventional screened cabled systems, since fibre optic cable is lighter than conventional cable and offers great weight saving. Figure 19 shows the configuration of a fly-by-light system

Fly-By-Light System Figure 19

AIRDATA

COMPUTER

FLIGHTCONTROL

COMPUTER

ACTUATORCONTROL

ELECTRONICS

LRGMOTION

SENSORS

ACTUATOR

FIBRE OPTIC CABLE

ELECTRICAL CABLE

CONTROLSURFACE



ENGINEERS


1.10.4 OPERATION Fibre optic cable interconnects the units of the flight control system and eliminates the possibility of propagating electrical faults between units. They are bi-directional and can be used to convey the system status to the flight crews’ control and display panel. A further advantage of fibre optic data transmission is the ability to use “Wavelength Division Multiplexing” (WDM) whereby a single fibre can be used to transmit several channels of information as coded light pulses of different wavelengths (or colours) simultaneously. The individual data channels are then recovered from the optically mixed data by passing the light signal through wavelength selective optical filters, which are tuned to the respective wavelengths. The WDM has a very high integrity, as the multiplexed channels are effectively optically isolated.

ELECTRONIC INSTRUMENT SYSTEM ELECTRONIC DISPLAYS PAGE 1 of 21


ENGINEERS


1 ELECTRONIC DISPLAYS 1.1 GENERAL With the introduction of digital signal-processing technology, it has become possible for drastic changes to both quantitative and qualitative data display methods. This technology has enabled the simplification of many flight deck-instrument layouts, allowing the replacement of complex analogue instruments with state of the art digital instrumentation. This "Glass Cockpit" concept has allowed many instruments to be replaced by one TV type display that can display a large and varied range of information as required. There are three different methods for displaying digital data, these are: 1. Light-Emitting Diodes (LED). 2. Liquid Crystal Display (LCD). 3. Cathode Ray Tube (CRT). 1.2 DISPLAY CONFIGURATIONS Displays of LED and LCD types are usually limited to the application in which a single register of alphanumeric values is required, and are based on the seven segment or the dot matrix configuration. CRT type displays have a wider use and can display navigation, engine performance and system status information. Table 1 shows the different applications for electronic displays.

Display Type Application Light-Emitting Diode Digital counter displays of engine performance.

Liquid Crystal Display Monitoring indicators; Radio frequency selector indicators; Distance Measuring indicators; Control display units of Inertial Navigation Systems, etc.

Cathode Ray Tube Weather radar indicators; display of navigational data; engine performance data; system status;

Electronic Display Applications Table1



ENGINEERS


Figure 1 shows a typical flight-deck instrument panel and the different types of display used.

BAe 146 Electronic Instrument Layout Figure 1

CRTDISPLAYS

LCDDISPLAYS

LEDDISPLAYS



ENGINEERS


1.2.1 SEGMENT DISPLAYS Seven-segment configuration will allow the display of decimal numbers 0-9; it also has the capability to display certain alphabetic characters. To display all alphabetic characters requires an increase in the number of segments from seven to thirteen, and in some cases sixteen segments. Figure 2 shows both seven and thirteen segment display configurations.

Seven and Thirteen Segment Display Formats Figure 2

SEVEN-SEGMENT CONFIGURATION

THIRTEEN-SEGMENT CONFIGURATION



ENGINEERS


In the dot matrix display the patterns generated for each individual character is made up of a specific number of illuminated dots arranged in columns and rows. Figure 3 shows the arrangement for a 4 X 7 configuration (4 columns and 7 rows).

Dot Matrix Configuration Figure 3

7ROWS

4 COLUMNS



ENGINEERS


1.3 LIGHT-EMITTING DIODE (LEDS) One of the most common light sources used in electronics is the “Light Emitting Diode” (LED). A LED is a two terminal semiconductor device comprising a p-n junction, which conducts in one direction only. This semiconductor material emits light when the p-n junction is forward biased and a current is flowing through it. LEDs can be manufactured to emit visible or invisible (infra-red) light. Visible LEDs are often used as indicators in electronic equipment either singly, for indicating ‘power on’ for instance, or in arrays for alpha/numeric displays. LEDs are reliable and have a very long life if treated carefully. Light emission in different colours of the spectrum can, when required, be obtained by varying the proportions of the elements comprising the chip, and also by a technique of "doping" with other elements, i.e. nitrogen. Current consumption (typically about 5 – 20 mA) generally limits the usefulness of a LED to equipment that is not battery powered. 1.3.1 OPERATION The phenomenon which results in the emission of light from a LED is called “Electroluminescence”, or “Injection Luminescence”, and is due to the hole/electron recombinations that take place near a forward biased p-n junction. When electrons are injected into the n region of a p-n diode and are swept through the region near the junction, they recombine with holes in the region. This generates electromagnetic waves of a frequency determined by the difference in the energy levels of the electron and the hole. In order for this recombination to result in luminescence, there must be a net change in the energy levels, and the proton generated must not be recaptured in the material. Figure 4 shows the operation of a LED.



ENGINEERS


LED Operation Figure 4

n

p

JUNCTION

INJECTEDELECTRONSCONTACT

BIASRECOMBINATIONS



ENGINEERS


Figure 5 shows the construction of a LED.

Light-Emitting Diode (LED) Figure 5



ENGINEERS


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 plastic overlay and a diffuser plate. The segments are formed as a sealed integrated circuit pack. The connecting pins of the LEDs are soldered to an associated printed circuit board. Depending on the application and the number of digits comprising the appropriate quantitative display, they will use either independent digit packs, or combined multiple digit packs may be used. Figure 6 shows an LED single digit pack construction.

LED Digit Pack Figure 6



ENGINEERS


LEDs can also be used in a dot-matrix configuration. Each dot making up the decimal numbers is an individual LED and can be arranged either in a 4 X 7 or 5 X 9 configuration. Figure 7 shows an Engine Speed Indicator, the dial portion of the indicator is an analogue type, however it uses an LED dot-matrix configuration for the digital readout of engine speed.

Engine Speed Indicator Figure 7

The digital counter is of unique design in that its signal drive circuit causes an apparent "rolling" effect of the digits which simulates the action of a mechanical drum-type counter as it responds to the changes in engine speed.

0

20 4060

80

100

% RPMN1

Smith's

ANALOGUEENGINESPEED

INDICATORDOT MATRIXLED DISPLAY

ENGINE SPEED



ENGINEERS


Figure 8 shows a Power Plant Instrument Group from a Boeing 737-400, which has both LED and dot matrix, displays.

Boeing 737-400 Power Plant Instrument Group Figure 8

4 X 7 MATRIXDISPLAY

LEDDISPLAY

6 0

5 1

4 23

6 0

5 1

4 23

12 0

10 2

84

6

12 0

10 2

8 46

MAN SET

% RPM

N1

°C

EGT

% RPM

X1000

N2

FF/FU

KGPH/KG

PUSH FUELUSED

RESETFUELUSED

PULLTO

SETN1

PULLTO

SETN1

7265

7265

84 87 84 8

7

100541005

4

2721 272

1



ENGINEERS


1.4 LIQUID CRYSTAL DISPLAY (LCD) Liquid Crystal Displays (LCD) are not actually light sources - they generate no light, merely filtering incident light, in a controlled manner. The LCDs seen in watches, clocks and calculators etc, all work by the same principle. Two transparent but conductive plates sandwich a layer of liquid crystals, which normally all face in the same direction. See Figure 9. Incident light passes through the liquid crystals of polarised particles fairly easily, and is reflected back through the crystals so that an observer sees a light coloured area. However, a voltage applied across the plates causes the liquid crystals to change direction in an attempt to repolarise themselves with the applied voltage. As they turn, they interact with the current flowing between the plates and a state of turbulence is created. The moving particles scatter the incident light, randomly reflecting and refracting it. Little light is reflected back to the observer, so the area between the transparent plates appears dark. Selection of the areas, which are turned dark by using a number of plates and different shaped plates, means that practically any shape of character may be displayed.



ENGINEERS


Figure 9 shows the operation of a LCD.

LED Operation Figure 9

INCIDENT LIGHT REFLECTED LIGHT

TRANSPARENTCONDUCTIVE

PLATES

INCIDENT LIGHT

REFLECTED LIGHT

TRANSPARENT OPAQUE



ENGINEERS


Figure 10 shows the structure of a seven-segment LCD.

Seven-Segment LCD Figure 10

The space between the plates is filled with a liquid crystal compound, and the complete assembly is hermetically sealed with a special thermoplastic material to prevent contamination. When a low-voltage, low-current signal is applied to the segments, the polarisation of the compound is changed together with a change in its optical appearance from transparent to reflective. The 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 a LED, it does not emit light, but merely acts on light passing through it. Depending on the polarisation film orientation, and whether the display is reflective or transmissive, the segment may appear dark on a light background (such as in digital watches and pocket calculators) or light on a dark background.

SEVEN SEGMENTELECTRODE

FRONT PLATE

SEGMENTCONTACTS

COMMONRETURN

CONTACT

BACK PLATE

MIRROR IMAGE(NOT SEGMENTED)

LIQUID CRYSTALLAYER

(TYPICAL SPACING = 10 MICRONS)



ENGINEERS


Figure 11 shows a BCD to seven-segment decoder.

BCD – Seven-Segment decoder Figure 11

BCD TO 7SEGMENTDECODER

LOW VOLTAGE POWERSUPPLY TO EACH

SEGMENT

1

2

4

8

1

1

0

0

0

00

00

1

1



ENGINEERS


1.5 CATHODE RAY TUBE (CRT) Displays of this type, which are based on the electron beam scanning technique, have been used in aircraft for many years. They were first used to display weather radar information and have continued to be an essential part of the “Avionics Fit” in today’s modern aircraft. The CRT is a thermionic device, i.e. one in which electrons are liberated as a result of heat energy. It consists of an evacuated glass envelope inside which are positioned an “Electron Gun”, “Beam-Focusing” and “Beam-Deflection” system. The inside surface of the screen is coated with a crystalline solid material known as a phosphor. Figure 12 shows a cross-section of a CRT.

Cathode Ray Tube CRT Cross-Section

Figure 12

CATHODE ANODE

DEFLECTINGCOILS

GRAPHITE COATING(COLLECTS SECONDARY ELECTRONS

TO PREVENT SCREEN BECOMINGNEGATIVELY CHARGED)

PERMANENTMAGNETS

(BEAM FOCUSING)

HEATER

GRID GLASSENVELOPE ELECTRON

BEAM

SCREEN



ENGINEERS


1.5.1 ELECTRON GUN The electron gun consists of the following:

1. Cathode: An indirectly heated cathode (negatively biased w.r.t. the screen).

2. Grid: A cylindrical grid surrounding the cathode. 3. Anode: Two (sometimes three) anodes. The cathode is a tube of metal closed at one end, with a coating of material that will emit electrons when heated, covering the closed end. To operate the cathode needs to be heated; this is achieved using a coil of insulated wire connected to the cathode. Because the screen of the CRT contains conducting material at a high voltage (5 - 15kV), electrons will be attracted away from the cathode. The free electrons have to pass through a pinhole in a metal plate (Control Grid). Altering the voltage of the grid can control the movement of the electrons through this hole. The voltage of the grid is always negative w.r.t. Cathode. The free electrons are then formed into a beam by the action of the first anode. The anode is of a cylindrical shape and by adjusting the voltage on the anode, the beam can be made to come to a small point at the screen end of the CRT. The screen end of the CRT is coated with a material called a “Phosphor”, which will glow when struck by electrons. The phosphor is usually coated with a thin film of aluminum so that it can be connected to the final accelerating (anode) voltage. The whole tube is formed as a vacuum.



ENGINEERS


Figure 13 shows the typical voltages used in a small CRT.

CRT Voltages Figure 13

GRID FIRST ANODE SECOND ANODE

CATHODE

HEATER

0V -50V +300V +5 kV

CONNECTED TO CONDUCTIVE COATING

ON GLASS



ENGINEERS


This arrangement will produce a point of light at the centre of the screen, but to make the CRT useful for displaying data, this beam of electrons must be able to be moved around the screen. For this, two sets of metal plates are used and if a voltage is passed through them, then the beam will deflect on the screen. These plates are called “Deflection Plates”. These plates are arranged at right angles to each other. The beam can be deflected if a voltage is applied to these plates; this is called “Electrostatic” deflection. Movement of the beam left/right is controlled by the “X” Plates, with the “Y” Plates controlling movement up/down. Figure 14 shows the arrangement for the deflection plates.

X and Y Deflection Plates Figure 14

Y DEFLECTIONPLATES

X DEFLECTIONPLATES

ANODE



ENGINEERS


The other method used for deflection is Electromagnetic. This method is used for TV, computer monitors and most aircraft CRT displays. As an electron moves, it constitutes an electric current, and so a magnetic field will exist around it in the same way as a field around a current-carrying conductor. In the same way that a conductor will experience a deflecting force when placed in a permanent magnetic field, so an electron beam can be forced to move when subjected to electromagnetic fields acting across the space within the tube. Coils are therefore provided around the neck of the tube, and are configured so that fields are produced horizontally (Y-axis field) and vertically (X-axis field). The coils are connected to the signal sources whose variables are to be displayed. The electron beam can be deflected to the left or right, up or down or along a resultant direction depending on the polarities produced by the coils, and on whether one alone is energised, or both are energised simultaneously. Figure 15 shows electromagnetic coil configuration and resultant deflections.

Electromagnetic Deflection

Figure 15

NN

SS

NNSS

MAGNETIC FIELD

ELECTRON BEAMCOMING OUT OF

THE PAPER VERTICALLY DISPOSED MAGNETICCOIL PRODUCES HORIZONTAL

DEFLECTION OF THE BEAM

NECK OFTHE TUBE

HORIZONTALLY DISPOSED MAGNETICCOIL PRODUCES VERTICALDEFLECTION OF THE BEAM

RESULTANT DEFLECTIONOF THE BEAM



ENGINEERS


The most common form of deflection for CRT is a “Linear Sweep”. This means that the beam is taken across the screen at a steady rate from one edge to the other, and is then returned very rapidly (an action called “Fly Back”). To generate such a linear sweep in electrostatic deflection, a Saw-tooth Waveform is used. . Figure 16 shows a Saw-tooth Waveform.

Saw-tooth Waveform

Figure 16 The sawtooth voltage waveform derived for the electrostatic time base is no use for electromagnetic coil deflection because a voltage sawtooth will not produce a linear rise of current through the deflection coils. A practical deflection, or scan coil, will have resistance as well as inductance. The voltage across the resistance of a coil “R” is proportional to the current through it. A linear current ramp in a resistance can only be produced by a steadily rising voltage. Inductor voltage is proportional to the rate of change of current and since the rate of change of current is constant, then the voltage across the inductor must also be constant. A constant applied voltage, therefore, will produce a linear current ramp in an inductor. To provide for both resistance and inductance, the voltage applied to the scan coils to produce a linear current ramp must be a constant value for the inductance and a voltage ramp for the resistance, giving the distinctive

TIME

CURRENT RAMP ORSWEEP FLYBACK



ENGINEERS


“Trapezoidal” shape. Figure 17 shows the scan coil graphs for electromagnetic deflection.

Scan Coil Graphs

Figure 17

MAX

MAX

MAX

MAX

0

0

0

0IDEAL

CURRENT

VOLTAGEACROSS R

VOLTAGEACROSS L

RESULTANTTRAPEZOIDAL

VOLTAGE

ELECTRONIC INSTRUMENT SYSTEM SOFTWARE MANAGEMENT CONTROL PAGE 1 of 4


ENGINEERS


1 SOFTWARE MANAGEMENT CONTROL In the normal maintaining of aircraft, an assessment of system and function criticality is made. With the increasing role of computers in today's aircraft, responsible Design Organisations assign, to each software-based system or equipment, software levels relating to the severity of the effect of possible software errors within user systems or equipments. Table 1 shows the relationship between function criticality category and software level. Effect on Aircraft

and occupants of

failure conditions

or design error

FAR 25.1309 &

JAR 25.1309

definitions

No significant

degradation of

aircraft capability

or crew ability

Reduction of the aircraft capability or

of the crew ability to cope with

adverse operating conditions

Prevention of

continued safe

flight and landing

of the aircraft

ACJ No 1

Jar 25.1309

definitions

Slight reduction

of safety

margins,

Slight increase in

workload, e.g.

routine changes

in flight or plan or

Physical effects

but no injury to

occupants

Significant

reduction in

safety margins

Reduction in the

ability of the flight

crew such that

they cannot be

relied upon to

perform their

tasks accurately,

or injury to

occupants

Large reduction

in safety margins

Physical distress

or workload such

that the flight

crew cannot be

relied upon to

perform their

tasks accurately

or completely, or

serious injury to

or death of a

relatively small

proportion of the

occupants

Loss of aircraft

and/or fatalities

ACJ No 1 to JAR 25.1309

Definition of Criticality Category

Minor Effect Major Effect Hazardous Effect Catastrophic

Effect

FAA Advisory Circular 25.1409-1

definition of Criticality Category

Non-essential Essential Critical

DO-178A/ED-12A

Software level*

Level 3 Level 2 Level 1

Table 1

* Using appropriate design and/or implementation techniques, it may be

possible to use a software level lower than the functional categorisation. Refer to Section 5 of DO-178A/ED-12A, which provides further guidance.



ENGINEERS


1.1 CERTIFICATION OF SOFTWARE For initial certification of a software-based system or equipment, the responsible Design Organisation provides evidence to the CAA that the software has been designed, tested and integrated with the hardware in a manner which ensures compliance with the relevant requirements of BCAR. The primary document for use by certifying authorities is the Software Accomplishment Summary. Its content is listed below to demonstrate the stringency of software control both during certification and continued use when it may be subject to further development and modification. The following is taken from AWN 45A. Related document references have been left in but not clarified. 1.2 CONTENT OF SOFTWARE ACCOMPLISHMENT SUMMARY As a minimum, information relevant to the particular software version should be included in the summary under the following headings: - (a) i) System and Equipment Description This section should

briefly describe the equipment functions and hardware including safety features, which rely on hardware devices or system architecture.

ii) Organisation of Software This section should identify the

particular software version and briefly describe the software functions and architecture with particular emphasis on the safety and partitioning concepts used.

The size of the final software design should be stated, e.g. in terms of

memory bytes, number of modules. The language(s) used should also be stated.

(b) Criticality Categories and Software Levels This section should state

the software levels applicable to the various parts of the software. The rationale for their choice should be stated, either directly, or by reference to other documents.

(c) Design Disciplines This section should briefly describe the design

procedures and associated disciplines, which were applied to ensure the quality of the software. The Organisations which were involved in the production and testing (including flight-testing) of the software should be identified and their responsibilities stated.



ENGINEERS


(d) Development Phases The development phases of the project should be summarised. This information could be included in sub-paragraph (h) below.

(e) Software Verification Plan This section should briefly summarise the

plan (Document No. 11 as defined in DO-178A/ED-12A) and the test results.

(f) Configuration Management The principles adopted for software

identification, modification, storage and release should be briefly summarised.

(g) Quality Assurance The procedures relating to quality assurance of

the software should be summarised including, where applicable, those procedures which applied to liaison between the equipment manufacturer and the aircraft, engine or propeller constructor, as appropriate.

(h) Certification Plan This section should provide a schedule detailing

major milestones achieved and their relationship to the various software releases.

(j) Organisation and Identification of Documents This section should

identify the documents, which satisfy, paragraph 8.1 of DO-178A/ED-12A.

(k) Software Status Any known errors, temporary patches, functional

limitations or similar shortcomings associated with the delivered software should be declared and the proposed timescale for corrective action stated.



ENGINEERS


1.3 MODIFICATION OF SOFTWARE In respect of systems and equipment with Level 1 or Level 2 software, a modification, which affects software, shall not be embodied unless it has been approved by the responsible Design Organisation. Modifications to software will be subject to the same approval procedures as are applied to hardware modifications. Modified software will need to be identified and controlled in accordance with the procedures stated in the software configuration management plan. The CAA will require the design and investigation of modifications, including those proposed by the aircraft operator, to involve the support service provided by the responsible Design Organisation. The re-certification effort will need to be related to the software levels. Aircraft operators will need to ensure that their defect reporting procedures will report software problems to the responsible Design Organisation.

DIGITAL TECHNIQUES /EIS ELECTROMAGNETIC ENVIRONMENT PAGE 1 of 7


ENGINEERS


1 ELECTROMAGNETIC ENVIRONMENT With the development of electronics and digital systems in aviation, aircraft are becoming increasingly susceptible to High Intensity Radio Frequencies (HIRF). Design philosophies in the area of aircraft bonding for protection against HIRF employ methods which may not have been encountered previously by maintenance personnel. Because of this, HIRF protection can be unintentionally compromised during normal maintenance, repair and modification. It is therefore critical that procedures contained in assembly and repair manuals contain reliable procedures to detect any incorrect installation, which could degrade the HIRF protection features. 1.1 PROTECTION AGAINST HIRF There are three primary areas to be considered for aircraft operating in HIRF environments. Aircraft Structure - (aircraft skin and frame).

Electrical Wiring Installation Protection - (Solid or braided

shielding/connectors).

Equipment Protection - (LRU case, electronics input/output protection).



ENGINEERS


Table 1 gives some indication as to the maintenance tasks which may be applied to certain types of electro magnetic protection features: PROTECTION

TYPE CABLE

SHIELDING AIRCRAFT STRUCTURE SHIELDING CIRCUIT

PROTECTION DEVICES

Description Over braid shield, critical individual cable shield

Raceway, conduits

RF gaskets Shield for non conductive surfaces

Structural bonding HIRF protection devices

Examples Metallic conduit, braid

Raceway, conduits

Removable panels

Conductive coating

Contact bonds, rivet

joints

Bonding lead and straps,

pigtails

Resistors, Zener diodes, EMI filters,

filter pins. Degradation or Failure Mode

Corrosion, damage

Corrosion, damage

Corrosion, damage,

deformation

Damage, erosion

Corrosion, damage

Corrosion, damage,

security of attachment

Short circuit, open circuit

Maintenance Operations

Visual inspection, measurement of cable shielding bonding

Visual inspection,

bonding measurement

Visual inspection of

gaskets, bonding leads

and straps

Visual inspection,

measurement of shielding

effectiveness

Visual inspection,

bonding measurement

Visual inspection for

corrosion attachment

and condition, bonding

measurement

Check at test/repair facility in

accordance with maintenance or

surveillance plan.

Applicable Maintenance Tasks for HIRF Protection Measures

Table 1

Note: “Raceway conduits” refers to separate conduits used to route individual cables to the various areas of an aircraft system. “RF gaskets” are gaskets having conductive properties to maintain the bonding integrity of a system. 1.2 TESTING TECHNIQUES Tests of HIRF protection carried out depend upon the criticality of the system under test. Types of test are as follows. 1.3 VISUAL INSPECTION The protection feature should be inspected for damage and corrosion. Degradation may be found in this way but where integrity cannot be assured, other tests may be carried out. 1.4 DC RESISTANCE The milliohm meter is often used to measure the ground path resistance of ground straps or bonding. This technique is limited to the indication of only single path resistance values.



ENGINEERS


1.5 LOW FREQUENCY LOOP IMPEDANCE Low frequency loop impedance testing is a useful method complementary to DC bonding testing. A visual inspection of cable bundle shields, complemented by a low frequency loop impedance test, gives good confidence in the integrity of the shielding provisions. Low frequency loop impedance testing is a method developed to check that adequate bonding exists between over braid (conduit) shields and structure. To achieve the shielding performance required, it is often necessary that both ends of a cable bundle shield be bonded to aircraft structure. In such cases, it is hard to check bonding integrity by the standard DC bonding test method. If the bond between shield and structure at one end is degraded while the other one is still good, there is little chance to find this defect by performing DC bonding measurements. The remaining bond still ensures a low resistance to ground but the current loop through the shield is interrupted, causing degradation of shielding performance. The fault can easily be detected by performing a low frequency loop impedance test. The test set-up requires simple test equipment, refer to Figure 1. A current of about 1 kHz is fed into the conduit under test while measuring the voltage necessary to drive that current. Other versions of the loop impedance test arrangement use different frequencies (200 Hz is typical), and provide the resistive and reactive parts of the loop impedance.



ENGINEERS


Loop Impedance Test Figure 1

The test equipment consists of a generator operating at 1 kHz feeding an injection probe and a current monitoring probe, connected to an AC millivoltmeter. A voltmeter connected to the generator enables the voltage necessary to drive the current to be measured. 1 kHz is a high enough frequency to drive the injection and the monitoring probes and is also enough to avoid specific RF effects, like non-uniform current distribution along the loop under test.

VOLTAGEGENERATOR

CURRENTMONITOR

(AC MILLI-VIOLTS)

STRUCTURE

CONDUIT

LOOP UNDER TEST

FIXING HARDWAREPROVIDING ELECTRICAL

BONDING

CLAMP-ON CURRENTTRANSFORMER

CLAMP-ON CURRENTTRANSFORMER

V1 II

ZCONDUIT + ZSTRUCTURE = V1/II



ENGINEERS


If, in practice, the current is set to 1A, the voltage figure, when expressed in millivolts, gives the loop impedance in milliohms directly. The loop impedance is normally in the range 1-100 milliohms. In this range, accurate results can easily be achieved. If too high loop impedance is found, the joint determining the problem has to be identified. This can be performed by measuring the voltage drop across each joint. The joint with the high voltage drop across it is the defective one, refer to Figure 2.

Identification of A Bad Joint Figure 2

STRUCTURE

CONDUIT

LOOP UNDER TEST

VOLTAGEGENERATOR

CLAMP-ON CURRENTTRANSFORMER V1

V2 = V1 ACROSS BAD JOINT

BRACKET

FIXING NUT

FERRULE

VOLTAGEMONITOR

V2

BAD JOINT



ENGINEERS


As there is no need for a wide band swept RF generator, the test equipment can be quite simple and easy to handle. Hand held battery powered test equipment, especially designed for production monitoring and routine maintenance, is available on the market. 1.6 ELECTRO MAGNETIC INTERFERENCE (EMI) EMI is a subject closely allied to HIRF. Interference can occur in systems from internal sources and external sources. Its prevention and maintenance of measures taken is described under High Intensity Radio Frequencies. 1.7 ELECTRO MAGNETIC COMPATIBILITY (EMC) A further allied subject is EMC. If a new avionics system is introduced into an aircraft, it must be operated at its full range of operating frequencies to ensure no interference to other systems is caused. Similarly, other systems must be operated across their full range to ensure no interference occurs to that system introduced. Full tests to be carried out are normally stipulated by the manufacturer or design organisation. 1.8 LIGHTNING/LIGHTNING PROTECTION Lightning protection is given by the primary and secondary conductors of an aircraft's bonding system. The system is enhanced by the methods discussed under HIRF.



ENGINEERS


1.9 DEGAUSSING If an aircraft is struck by lightning, structural damage can occur and parts of the aircraft may remain magnetised. This magnetic force remaining is called 'Residual Magnetism', and since it could adversely effect some aircraft systems, areas affected must be de-magnetised. The process of de-magnetising is called 'degaussing'. Effected areas are detected using a hand held compass, then an ac electromagnet is passed over these areas to disperse the residual magnetism. A discrepancy between an Aircraft’s main compass and standby compass of (typically) 8° indicates that degaussing is necessary.

EIS TYPICAL ELECTRONIC /DIGITAL A/C SYSTEM PAGE 1 of 54


ENGINEERS


1 ELECTRONIC/DIGITAL AIRCRAFT SYSTEMS Electronic and digital processes are used in many of today's aircraft for a variety of purposes: navigation, dissemination of information, flying and controlling the aircraft. It should be borne in mind that as each manufacturer introduces such a system to the market the chances are that new names for it are added to the dictionary of terms. For instance, an Engine Indication and Crew Alerting System (EICAS) is much the same as a Multi-Function Display System (MFDS), the main difference being the manufacturer. This module will deal with the following Electronic/Digital Systems:

1. Electronic Centralized Monitoring System (ECAM).

2. Electronic Flight Instrument System (EFIS).

3. Engine Indicating & Crew Alerting System (EICAS). 4. Flight Data Recorder System (FDRS).



ENGINEERS


1.1 ELECTRONIC CENTRALIZED AIRCRAFT MONITORING 1.1.1 INTRODUCTION In the ECAM system (originally developed for Airbus aircraft), data relating to the primary system is displayed in checklist, pictorial or abbreviated form on two Cathode Ray Tube (CRT) units. Figure 5 shows the ECAM system functional diagram.

ECAM Functional Diagram Figure 5

DMC 1 DMC 3 DMC 2

FWC 1 FWC 2

CAUT

WARN

CAUT

WARN

ECAMCONTROL PANEL

A/C SYSTEM SENSORSRED WARNINGSSYSTEM PAGESFLIGHT PHASE

A/C SYSTEM SENSORSAMBER WARNINGS

SYSTEM PAGESNAV & AFS SENSORS

SDAC 1 SDAC 1



ENGINEERS


1.2 ECAM SYSTEM COMPONENTS 1.2.1 FLIGHT WARNING COMPUTER (FWC) The two FWCs acquire all data necessary for the generation of alert messages associated with the relevant system failures: Directly form the aircraft sensors or systems for warnings (mainly identified

by red colour).

Through the SDACs for cautions from the aircraft systems (mainly identified by amber colour).

The FWCs generate alphanumeric codes corresponding to all texts/messages to be displayed on the ECAM display units. These can be either be: Procedures associated to failures.

Status functions (giving the operational status of the aircraft and

postponable procedures).

Memo function (giving a reminder of functions/systems, which are temporarily used or items of normal checklist).

1.2.2 SYSTEM DATA ACQUISITION CONCENTRATORS (SDAC) The two SDACs acquire from the aircraft systems malfunctions/failure data corresponding to caution situations and send them to the FWCs for generation of the corresponding alert and procedure messages. The two SDACs acquire then send to the 3 DMCs all aircraft system signals necessary for display of the system information and engine monitoring secondary parameters through animated synoptic diagrams. All signals (discrete, analog, digital) entering the SDACs are concentrated and converted into digital format. 1.2.3 DISPLAY MANAGEMENT COMPUTERS (DMC) The 3 DMCs are identical. Each integrates the EFIS/ECAM functions and is able to drive either ECAM display units (engine/warning or system/status). The DMCs acquire and process all the signals received from various aircraft sensors and computers in order to generate proper codes of graphic instructions corresponding to the images to be displayed.



ENGINEERS


1.2.4 DISPLAY UNITS These can be mounted either side-by-side or top/bottom. The left-hand/top unit is dedicated to information on the status of the system; warnings and corrective action in a sequenced checklist format, while the right-hand/bottom unit is dedicated to associated information in pictorial or synoptic format. Figure 6 shows the layout of ECAM displays.

ECAM Display Layout Figure 6

5

5

MACH8 4

109

IASKNOTS

60

80

120

180200220

240

250

300

350400

140

LDG GEARGRVTY EXTN

RESET

OFF

DOWN



ENGINEERS


1.2.5 ECAM DISPLAY MODES There are four display modes, three of which are automatically selected and referred to as phase-related, advisory (mode and status), and failure-related modes. The fourth mode is manual and permits the selection of diagrams related to any one of 12 of the aircraft’s systems for routine checking, and the selection of status messages, provided no warnings have been triggered for display. Selection of displays is by means of a system control panel. See Figure 14. 1.2.6 FLIGHT PHASE RELATED MODE In normal operation the automatic flight phase-related mode is used, and the displays will be appropriate to the current phase of aircraft operation, i.e. Pre-flight, Take-off, Climb, Cruise, Descent, Approach, and post landing. Figure 7 shows display modes. The upper display shows the display for pre-take off, the lower is that displayed for the cruise.

ECAM Upper and Lower Display (Cruise Mode) Figure 7

8 7 . 05 10

6 5 . 05 10

N1%

6 5 05 10

4 8 05 10

EGTºC

NO SMOKING: ONSEAT BELTS: ONSPLRS: FULLFLAPS: FULL

FOB : 14000KG

LDG INHIBITAPU BLEED

80 80.2N2%

1500 1500FFKG/H

FULL

FLAPS F

ECAM UPPER DISPLAY ECAM LOWER DISPLAY - CRUISE

TAT +19 ºCSAT +17 ºC

G.W. 60300 KGC.G. 28.1 %23 H 56

CKPT 20 FWD 22 AFT 23

24 22 24

AIR

ENGINE

LDG ELEV AUTO 500FT

CAB V/S FT/MIN250

CAB ALT FT4150

VIB (N1)0.8 0.9

VIB (N2)1.2 1.3

F.USEDKG

OILQTY

1530 1530

11.5 11.5



ENGINEERS


1.2.7 ADVISORY MODE This mode provides the flight crew with a summary of the aircraft’s condition following a failure and the possible downgrading of systems. Figure 8 shows an advisory message following a Blue Hydraulic failure.

ECAM Advisory Mode Figure 8

8 7 . 05 10

6 5 . 05 10

N1%

6 5 05 10

4 8 05 10

EGTºC

FOB : 14000KG

80 80.2N2%

1500 1500FFKG/H

FULLFULL

FLAPFLAPSS FF

ADVISORYMESSAGES

FLT CTLSPOILERS SLOW

1 FUEL TANK PUMP LH

HYD B RSVR OVHT B SYS LO PR

FAILUREMESSAGES



ENGINEERS


1.2.8 ECAM FAILURE MODE The failure-related mode takes precedence over the other modes. Failures are classified in 3 levels Level 3: Warning This corresponds to an emergency configuration. This requires the flight crew to carry out corrective action immediately. This warning has an associated aural warning (fire bell type) and a visual warning (Master Warning), on the glare shield panel. Level 2: Caution This corresponds to an abnormal configuration of the aircraft, where the flight crew must be made aware of the caution immediately but does not require immediate corrective action. This gives the flight crew the decision on whether action should be carried out. These cautions are associated to an aural caution (single chime) and a steady (Master Caution), on the glare shield panel. Level 1: Advisory This gives the flight crew information on aircraft configuration that requires the monitoring, mainly failures leading to a loss of redundancy or degradation of a system, e.g. Loss of 1 FUEL TANK PUMP LH or RH but not both. The advisory mode will not trigger any aural warning or ‘attention getters’ but a message appears on the primary ECAM display.



ENGINEERS


Figures 9 – 13 show the 12-system and status pages available.

ECAM System Displays Figure 9

Note; These pages are displayed: Automatically due to an advisory or failure related to the system. Whenever called manually.

AIR CONDITIONING SYSTEM PAGE


G.W. 60300 KGC.G. 28.1 %23 H 56

CKPT 20 FWD 22 AFT 23

FANFANALTN MODE

COND TEMP ºC

C H C H C H

24 22 24

HOTAIR

PRESSURIZATION SYSTEM PAGE


G.W. 60300 KGC.G. 28.1 %23 H 56

CAB PRESS LDG ELEV MAN 500FT

PACK 1 PACK 2

SAFETY

EXTRACTINLETVENT

MAN SYST 2SYST 1

4.10

8

APPSI

11502

02

V/S FT/MIN

DN

UP

4150010

CAB ALTFT



ENGINEERS




ELECTRICAL SYSTEM PAGE


G.W. 60300 KGC.G. 28.1 %23 H 56

ELEC

DC 1 DC 2

DC BATBAT 128V150A

BAT 228V150A

DC ESS

EMERG GEN116V400HZ

ESS TR28V

130A

GEN 126%116V

400HZ

GEN 226%116V

400HZ

APU26%116V

400HZ

EXT PWR116V

400HZ

TR 128V150A

TR 228V150A

AC ESSAC 1 AC 2

FLIGHT CONTROL SYSTEM PAGE


G.W. 60300 KGC.G. 28.1 %23 H 56

F/CTR G B Y

SPD BRKL

AILB G

RAILG B

RELEVY B

LELEVB G

RUDG B Y

PITCH TRIM G Y3.2º UP



ENGINEERS


ECAM System Displays

Figure 11


FUEL SYSTEM PAGE


G.W. 60300 KGC.G. 28.1 %23 H 56

FOB

28750

FUEL KGF.USED 11550

F.USED 21550

CTR

APU

LEFT RIGHT

550 5505600 1075010750

HYDRAULIC SYSTEM PAGE


G.W. 60300 KGC.G. 28.1 %23 H 56

HYDGREEN

3000

YELLOW

3000PSI PSI

BLUE

3000



ENGINEERS



Note; These pages are displayed: Automatically due to an advisory or failure related to the system. Whenever called manually. The Gear/Wheel page is displayed at the related flight phase.

AIR BLEED SYSTEM PAGE


G.W. 60300 KGC.G. 28.1 %23 H 56

BLEED

RAM AIR

20 ºC

230 ºCLO HI

C H

LP HP

1

LPHP

2

24 ºC

50 ºCLO HI

C H

APU

GND

LANDING GEAR/WHEEL/BRAKE SYSTEM PAGE


G.W. 60300 KGC.G. 28.1 %23 H 56

WHEEL

1170

2140

REL

ºC

3140

4140

REL

ºC

AUTO BRK



ENGINEERS



Note; These pages are displayed: Automatically due to an advisory or failure related to the system. Whenever called manually. Related flight phase.

DOOR/OXY SYSTEM PAGE

DOOR OXY 1850 PSI

EMEREXIT

CABIN

CARGOBULK

CARGO

CABINFWD COMPT

AVIONICARM ARM

ARMARM

ARMARM

TAT +19 ºCSAT +17 ºC C.G. 28.1 %23 H 56

APU SYSTEM PAGE

APU

800

10N%

5803

75EGT

ºC

FLAP OPEN

BLEED35 PSI

APU26%116V

400HZ

TAT +19 ºCSAT +17 ºC C.G. 28.1 %23 H 56



ENGINEERS


1.2.9 CONTROL PANEL The layout of the control panel is shown in Figure 14.

ECAM Control Panel Figure 14

CLR

STS

RCL

ENG HYD AC DC

BLEED COND PRESS FUEL

APU F/CTL DOOR WHEEL

OFF OFF

FAULT FAULT

OFF BRT OFF BRT

LEFT DISPLAY RIGHT DISPLAY1 ECAM SGU 2

SYSTEM SYNOPTICDISPLAY SWITCHES

RECALLSWITCH

STATUSMESSAGESWITCH

MESSAGECLEARANCE

SWITCH

SGU SELECTSWITCHES

DISPLAY ON &BRIGHTNESS

CONTROL

DISPLAY ON &BRIGHTNESS

CONTROL



ENGINEERS


1.2.10 ECAM CONTROL PANEL SGU Selector Switches: Controls the respective symbol generator units. Lights are off in normal operation of the system. The “FAULT” caption is illuminated amber if the SGU’s internal self-test circuit detects a failure. Releasing the switch isolates the corresponding SGU and causes the “FAULT” caption to extinguish, and the “OFF” caption to illuminate white. System Synoptic Display Switches: Permit individual selection of synoptic diagrams corresponding to each of the 12 systems, and illuminate white when pressed. A display is automatically cancelled whenever a warning or advisory occurs. CLR Switch: Light illuminates white whenever a warning or status message is displayed on the left-hand display unit. Press to clear messages. STS Switch: Permits manual selection of an aircraft’s status message if no warning is displayed. Illuminates white when pressed also illuminates the CLR switch. Status messages are suppressed if a warning occurs or if the CLR switch is pressed. RCL Switch: Enables previously cleared warning messages to be recalled provided the failure conditions which initiated the warnings still exists. Pressing this switch also illuminates the CLR switch. If a failure no longer exists, the message “NO WARNING PRESENT” is displayed on the left-hand display unit.



ENGINEERS


PAGE INTENTIONALLY

BLANK



ENGINEERS


1.3 ELECTRONIC FLIGHT INSTRUMENT SYSTEM (EFIS) With the introduction of fully integrated, computer-based navigation system, most electro/mechanical instrumentation has been replaced with TV type colour displays. The EFIS system provides the crew with two displays:

1. Electronic Attitude Direction Indicator (EADI). 2. Electronic Horizontal Situation Indicator (EHSI).

The EADI is often referred to as the Primary Flight Display (PFD) and the EHSI as the Navigation Display (ND). The EADI and EHSI are arranged either side by side, with the EADI positioned on the left, or vertically, with the EADI on the top. 1.3.1 SYSTEM LAYOUT As is the case with conventional flight director systems, a complete EFIS installation consists of two systems. The Captain’s EFIS on the left and the First Officer’s on the right. The EFIS comprises the following units:

1. Symbol Generator (SG).

2. Display units X 2 (EADI & EHSI).

3. Control Panel. 4. Remote Light Sensor.

1.3.2 SYMBOL GENERATOR These provide the analog, discrete and digital signal interfaces between the aircraft’s systems, the display units and the control panel. They provide symbol generation, system monitoring, power control and the main control functions of the EFIS overall.



ENGINEERS


Figure 15 shows the interface between the modules within the SG.

Symbol Generator Module Interface Figure 15

MAINPROM

MAINRAM

MAINPROCESSOR

INPUT1

INPUT 2

DISPLAYCONTROL

WXINPUT

DISPLAYSEQUENCER

WX MEMORY2 X 16K RAMS

RASTERGENERATOR

STROKEGENERATOR

DISPLAYDRIVER

FMCRAD ALT

VOREFIS

CONTROL

IRSILS

DMEVOR

WEATHER RADAR DATA

WXRASTER

DISPLAYUNIT

VIDEO

DISPLAYUNIT

RASTER/STROKESELECT

STROKEPOSITION

DATA

DISPLAYUNIT

DEFLECTIONSIGNALS

STR

OK

E/VI

DEO

& P

RIO

RIT

Y D

ATA

CHARACTERDATA

TRA

NSF

ER B

US

DIS

PLA

Y C

OU

NTE

R I/

O B

US

DIS

PLA

Y SE

QU

ENC

ER D

ATA

BU

S



ENGINEERS


Table 1 gives details of the functions of the SG modules.

Module Function Input 1 & 2 Supply of data for use by the main computer.

Main Processor Carries the main control and data processing of the SG. Main RAM Address decoding, read/write memory and input/output

functions for the system. Main PROM Read-only memory for the system.

Display Control Master transfer bus interface. WX Input Time scheduling and interleaving for raster, refresh, input and

standby function of weather radar input data. WX Memory RAM selection for single input data, row and column shifters for

rotate/translate algorithm, and shift registers for video output. Display

Sequencer Loads data into registers on stroke and raster generator cards.

Stroke Generator

Generates all single characters, special symbols, straight and curved lines and arcs on display units.

Raster Generator

Generates master timing signals for raster, stroke, EADI and EHSI functions.

Display Driver Converts and multiplexes X and Y digital stroke and raster inputs into analog for driver operation, and also monitors deflection outputs for correct operation.

Symbol Generator Module Functions

Table 1

1.3.3 DISPLAY UNITS Each display unit consists of the following modules:

1. Cathode Ray Tube. 2. Video Monitor Card. 3. Power Supply Unit. 4. Digital Line Receivers. 5. Analog Line Receivers. 6. Convergence Card.



ENGINEERS


Figure 16 shows a block schematic of the display unit.

EFIS Display Unit Block Schematic Figure 16

CRTVIDEO MONITORCARDDIGITAL LINE

RECEIVERS

ANALOG LINERECEIVERS

DEFLECTIONCARD

CONVERGENCECARD

LOW VOLTAGEPOWER SUPPLY

HIGH VOLTAGEPOWER SUPPLY

115V 4OOHz

LIGHT SENSOR

DISPLAY UNIT BRIGHTNESS

RASTER BRIGHTNESS

X DEFLECTION

Y DEFLECTION

RED

GREEN

BLUE

BEAM TEST

SYNCHRONIZING

INTENSITY

RASTER/STROKE

DAY/NIGHT



ENGINEERS


1.3.4 LOW/HIGH POWER SUPPLIES All a.c. and d.c. power requirements for the overall operation of the DU is provided by a low power supply and a high power supply. They are supplied by 115V 400Hz from the aircraft power supplies. Supplies are automatically regulated and monitored for under/over voltage conditions. 1.3.5 DIGITAL LINE RECEIVERS Receives digital signals from the SG (R,G,B control, test signal, raster and stroke signals and beam intensity). It contains a Digital/Analog converter so that it can provide analog signals to the Video Monitor card. 1.3.6 ANALOG LINE RECEIVERS Receive analog inputs form the SG representing the required X and Y deflections for display writing. 1.3.7 VIDEO MONITOR CARD Contains a video control microprocessor, video amplifiers and monitoring logic for the display unit. It calculates the gain factors for the three-video amplifiers (R, G and B). It also performs input, sensor and display unit monitoring. 1.3.8 DEFLECTION CARD Provides X and Y beam deflection signals for stroke and raster scanning. 1.3.9 CONVERGENCE CARD Takes X and Y deflection signals and develops drive signals for the three radial convergence coils (R, G and B) of the CRT. Voltage compensators monitor the deflection signals in order to establish on which part of the CRT screen the beams are located. Right or left for the X comparator: top or bottom for the Y comparator.



ENGINEERS


Figure 17 shows the EFIS units and signal interface in block schematic form.

EFIS Block Schematic Figure 17

I

CRS

TEST

DIM DH BOT TOP

RASTER DIM

HDG

GSTTG

FULLARC

WX ET MAPSCCP

REV

OFFAUTO

ADF 1ADF 2

BRG BRG

NAV VLF FMS INS 1 INS 2 HDG ATT

ADF 2ADF 1

OFF

VOR 1VOR 2

AIRDATACOMP

INERTIALREF

SYSTEM

NAV AIDILS/VOR

RAD ALT

WEATHERRADAR

DME

FMS

AFCS

GPWS

EFIS SG No 2

EFIS SG No 3

EFIS SG No 1

Honeywell

N

S

W E

3

612

1521

2430

33

GSPD130 KTS

HDG013

NAV 12.1 NMH

CRS345+0

VOR 1

ADF 1

Honeywell

ATT 2AOA

CMDM .99200DH DH 140

RA

G

GS

2020

10 10

20 20

10 10

F

S

CRS

TEST

DIM DH BOT TOP

RASTER DIM

HDG

GSTTG

FULLARC

WX ET MAPSCCP

REV

OFFAUTO

ADF 1ADF 2

BRG BRG

NAV VLF FMS INS 1 INS 2 HDG ATT

ADF 2ADF 1

OFF

VOR 1VOR 2

Honeywell

N

SW E

3

612

1521

2430

33

GSPD130 KTS

HDG013

NAV 12.1 NMH

CRS345+0

VOR 1

ADF 1

Honeywell

ATT 2AOA

CMDM .99200DH DH 140

RA

G

GS

2020

10 10

20 20

10 10

F

S



ENGINEERS


1.3.10 CONTROL PANEL Allows the crew to select the required display configuration and what information is to be displayed. Both Captain and Co-Pilot have their own display controllers. The controllers have two main functions:

Display Controller: Selects the display format for EHSI as FULL, ARC, WX or MAP.

Source Select: Selects the system that will provide information required for display. The source information will be VOR, ADF, INS, FMS, VHF and NAV.

EFIS Display Controller is shown at Figure 18, and the Source Controller is at 19.

EFIS Display Controller Figure 18

CRS

HDG

TEST

WXFULLARC

GSTTG ET MAP SC

CP REV

DIM DH BOT TOP



ENGINEERS


EFIS Source Controller

Figure 19

VHFNAV FMS INS 1 INS 2 HDG ATT

OFF

AUTO

ADF 1

ADF 2

BRGOFF

ADF 1

ADF 2VOR 1 VOR 2

BRG



ENGINEERS


1.3.11 ELECTRONIC ATTITUDE DIRECTOR INDICATOR (EADI) The EADI displays traditional attitude information (Pitch & Roll) against a two-colour sphere representing the horizon (Ground/Sky) with an aircraft symbol as a reference. Attitude information is normally supplied from an Attitude Reference System (ARS). The EADI will also display further flight information. Flight Director commands right/left to capture the flight path to Waypoints: airports and NAVAIDS and up/down to fly to set altitudes: information related to the aircraft’s position w.r.t. Localizer (LOC) and Glideslope (GS) beams transmitted by an ILS. Auto Flight Control System (AFCS) deviations and Autothrottle mode, selected airspeed (Indicated or Mach No) Groundspeed, Radio Altitude and Decision Height information are also shown. Figure 20 shows a typical EADI display

Electronic Attitude Director Indicator (EADI) Display

Figure 20

140 RA

AP ENG200 DH

LOC

ATT 2

F

S

GSHDG

M

Honeywell

M .99

20 20

10 10

20 20

10 10



ENGINEERS


1.3.12 ELECTRONIC HORIZONTAL SITUATION INDICATOR (EHSI)

The EHSI presents a selectable, dynamic colour display of flight progress with plan view orientation. The EHSI has a number of different modes of operation, these are selectable by the flight crew and the number will be dependent on the system fitted.

Figure 21 shows an EHSI display.

Electronic Horizontal Situation Indicator (EHSI) Display

Figure 21

Honeywell

N

S

W

E3

6

1215

2124

3033

WPT

VOR 1

ADF 1

GSPD130 KTS

HDG350

NAV 1

2.1 NMH

CRS315+0

G



ENGINEERS


1.3.13 PARTIAL COMPASS FORMAT

The partial compass mode displays a 90° ARC of compass coordinates. It allows

other features, such as MAP and Weather Radar displays, to be selected. Figure 22 shows a Partial EHSI display (Compass Mode).

EHSI Partial Compass Mode Display

Figure 22

Honeywell

VOR 1

ADF 1

GSPD130 KTS

HDG

350

DTRK317

V

320

30 33

N

2515

50

FMS1 30 NM



ENGINEERS


Figure 23 shows an EHSI partial format with Weather Radar information.

EHSI Weather Radar Display Figure 23

1.3.14 MAP MODE

The MAP mode will allow the display of more navigational information in the partial compass mode. Information on the location of Waypoints, airports, NAVAIDs and the planned route can be overlaid. Weather information can also be displayed in the MAP mode to give a very comprehensive display.

Honeywell

VOR 1

ADF 1

GSPD130 KTS

HDG

350

DTRK317

V

320

30 33

N

25

50

FMS1 30 NM



ENGINEERS


Figure 24 shows an EHSI MAP mode display.

EHSI MAP Mode Display. Figure 24

Honeywell

VOR 1

ADF 1

GSPD130 KTS

HDG

350

DTRK

317

V

320

30 33

N

25

50

FMS1 30 NM

03

0405

05



ENGINEERS


1.3.15 COMPOSITE DISPLAY

In the event of a display unit failure, the remaining unit can display a “Composite Display”. This display is selected via the Display Controller and it consists of elements from an EADI and EHSI display. Figure 25 shows a typical composite display.

EFIS Composite Display Figure 25

140 RADH200 DH

ATT 2

F

S

120 NMHDGILS

M

Honeywell

M .99

010

0033 03

000

20 20

10 10

CRS FR

10 10



ENGINEERS


1.3.16 TESTING

Test is controlled from the DH/TEST knob located on the EFIS control panel. The test, if carried out using the First Officer’s control panel, will have the following effect on the Captain’s EADI:

Runway symbol will fall. Rad Alt digital display indicates 95 to 100 feet.

The First Officer’s EADI warning will be activated:

Amber dashes are displayed on the Rad Alt digital

display. Amber dashes are displayed on the

selected DH digital display. When the TEST button is pressed on the Captain’s EFIS control panel the same test sequence takes place. The test altitude value remains displayed as long as the TEST button is pressed. Releasing the knob causes actual altitude to be displayed and digits of the DH display to show the selected value at the end of the test. The test sequence can be initiated during flight except during APP (Approach).



ENGINEERS


1.3.17 SYMBOL GENERATOR TEST Some EFIS systems have the capability of carrying out a comprehensive Symbol Generator BITE. As an example, the BAe 146 EFIS SG Self-test is described. Initiated by selecting SELF-TEST on the dimming panel and pressing the verifying (DATA), button on the EFIS Control panel. Refer to Figure 26

BAe 146 EFIS Control & Dimming Panels Figure 26

N-AID ARPT GRP DATA

WPT

ADF VOR

OFF

BRG 320

160 80 20

10

RANGE FORMAT

PLAN

MAP ARC

ROSECRS

LNAV V/L

OFF

BACKSPACE FORWARD SPACE VERIFY

WXND

WX OFF

PFD

BRT

COMPACT

DH

TEST

EFISSELF-TEST

BUTTON

DIMMING PANEL

EFIS CONTROL PANEL



ENGINEERS


The Display unit will now display the “Maintenance Master Menu” format as shown in Figure 27. Using the backspace – forward space controls on the EFIS control panel, select “SG SELF TEST”.

Maintenance Master Menu Display Figure 27

FAULT REVIEWFAULT ERASETEST PATTERNSG SELF TESTOPTIONS/CONFIG



ENGINEERS


The Symbol Generator Self-Test sequence is automatic and the process is as shown in Figure 28.

SG Self-Test Process Figure 28

FAULT REVIEWFAULT ERASETEST PATTERNSG SELF TEST

OPTIONS/CONFIG

SELF TEST IN PROGRESS

SYMBOL GENERATOR SELF TESTAIRCRAFT CONFIGURATION YYDP SOFTWARE PART NUMBER:

XXXXXXXXX-XXSMP SOFTWARE PART NUMBER

XXXXXXXX-XXTESTFAIL

SELF TEST FAILURES

FAILURE 1FAILURE 2FAILURE 3FAILURE 4FAILURE 5FAILURE 6

FAIL

PASS

INTERFACE STATUS

STATUS 1STATUS 2STATUS 3STATUS 4STATUS 5STATUS 6

SYMBOL GENERATOR SELF TESTAIRCRAFT CONFIGURATION YYDP SOFTWARE PART NUMBER:

XXXXXXXXX-XXSMP SOFTWARE PART NUMBER

XXXXXXXX-XXTESTPASS



ENGINEERS


The test fail message will appear if any failures internal to EFIS are detected. Depressing the “Forward Space” key after “FAIL”, on completion of the self-test, brings up a self-test failure page that lists the first test that failed. Depressing the “Forward Space” key again brings up the Interface Status page. Depressing the “Forward Space” after “PASS”, on completion of the self-test, brings up the Interface Status page. This page lists any interfaces that are not valid. After confirming the status of the “Self-test Failures” and “Interface Status”, then the operator can reselect the Maintenance Format page to carry out further testing.



ENGINEERS


PAGE INTENTIONALLY

BLANK



ENGINEERS


1.4 ENGINE INDICATION AND CREW ALERTING SYSTEM

1.4.1 INTRODUCTION

EICAS is a further system to indicate parameters associated with engine performance and airframe control by means of CRT display units. This particular variation first appeared on Boeing 757 and 767 aircraft. 1.4.2 SYSTEM LAYOUT EICAS comprises two display units, a control panel and two computers, which receive analogue and digital signals from engine and system sensors. Only one computer is in control, the other being on standby in the event of failure occurring. It may be selected automatically or manually. A functional diagram of an EICAS layout is shown at Figure 29.

EICAS Block Schematic Figure 29

ENGINE&

AIRCRAFTSYSTEMINPUTS

PERF

APU

ELEC

HYD

ECS

MSG

ENGEXCD

CONF

MCDP

DISPLAY SELECT

EICAS MAINT EVENTREAD

AUTO MAN

REC ERASE

TEST

EPCS

MAINTENANCE PANEL

ENGINE STATUS MAX INDRESET

L AUTO R L REVENTRECORD

DISPLAY COMPUTERBRTBAL

BRT THRUST REF SET

BOTH

DISPLAY SELECT PANEL

CAUTION

CANCEL RESET

ENGINE SECONDARYDISPLAY

ORSTATUS DISPLAY

ORMAINTENANCE DISPLAY

ENGINEPRIMARYDISPLAY

&WARNINGSCAUTIONS

ADVISORIES

EICAS COMPUTER No 1

EICAS COMPUTER No 2



ENGINEERS


1.4.3 DESCRIPTION Referring to Figure 29, the upper DU displays warnings and cautions and the engine primary parameters:

N1 Speed.

EGT. If required, program pinning enables EPR to be displayed also. Secondary engine parameters are displayed on the lower DU:

N2 Speed. Fuel Flow.

Oil Quantity Pressure

Engine Temperature

Engine Vibration.

Other system status messages can also be presented on the lower DU for example:

Flight Control Position.

Hydraulic system status.



ENGINEERS


1.4.4 DISPLAYS Figure 30 shows displays presented on the Primary and Secondary DUs.

EICAS Primary & Secondary Displays

Figure 30

TAT 15°c

N1

EGT

V V V V V V V

0 0

10

62

0.010

62

0.0

CAUTION

CANCEL RECALL

3.1 1.9

18 18

120 120

50 50

OIL QTY

VIB

N1 FAN

OIL TEMP

OIL PRESS

88.00 88

86 86

4.4

N2

N3

FF

4.4



ENGINEERS


1.4.5 DISPLAY MODES There are three modes of displaying information:

Operation Mode.

Status Mode.

Maintenance Mode. 1.4.6 OPERATION MODE The Operational Mode is selected by the crew and displays engine operating information and any alerts requiring action by the crew in flight. Normally only the upper unit displays information. The lower unit remains blank and can be selected to display secondary information as required.

1.4.7 STATUS MODE When selected this mode displays data to determine the dispatch readiness of an aircraft, and is closely associated with details contained in an aircraft’s “Minimum Equipment List”. Shown on the lower display unit is the position of the flight control surfaces (Elevator, Ailerons and Rudder), in the form of pointers registered against vertical and horizontal scales. Also displayed are selected sub-system parameters, and equipment status messages. Selection is normally done on the ground, either as part of the Pre-flight checks of dispatch items, or prior to shut-down of electrical power to aid the flight crew in making entries in the aircraft’s technical log.



ENGINEERS


Figure 31 shows a status mode display.

AICAS Status Mode Display Figure 31

1.4.8 MAINTENANCE MODE

Used by maintenance engineers with information in five different display formats to aid troubleshooting and test verification of the major sub-systems. These displays appear on the lower DU and are not available in flight.

L C RHYD QTY 0.99 1.00 0.98

HYD PRESS 2975 3010 3000

APU EGT 440 RPM 103 OIL 0.75

OXY PRESS 1750

AIL ELEV AIL

RUD

0.0 0.0FF

CABIN ALT AUTO 1ELEV FEEL



ENGINEERS


1.4.9 SELECTION PANEL Control of EICAS functions and displays is via the EICAS Control Panel. This can be used both in flight and on the ground. It is normally located on the centre pedestal of an aircraft's flight deck, and its controls are as follows:

• Engine Display Switch: This is of the momentary-push type for removing

or presenting the display of secondary information on the lower display unit.

• Status Display Switch: Also of the momentary-push type, this is used for

displaying the status mode information, referred to earlier, on the lower display unit.

• Event Record Switch: This is of the momentary-push type and is used in

the air or on the ground, to activate the recording of fault data relevant to the environmental control system, electrical power, hydraulic system, performance and APU. Normally, if any malfunction occurs in a system, it is recorded automatically (called an 'auto event') and stored in a non-volatile memory of the EICAS computer. The push switch enables the flight crew to record a suspect malfunction for storage, and this is called a 'manual event'. The relevant data can only be retrieved from memory and displayed when the aircraft is on the ground and by operating switches on the maintenance control panel.

• Computer Select Switch: In the 'AUTO' position it selects the left, or primary, computer and automatically switches to the other computer in the event of failure. The other positions are for the manual selection of left or right computers.

• Display Brightness Control: The inner knob controls the intensity of the displays, and the outer knob controls brightness balance between displays.

• Thrust Reference Set Switch: Pulling and rotating the inner knob

positions the reference cursor on the thrust indicator display (either EPR or NI) for the engine(s) selected by the outer knob.

• Maximum Indicator Reset Switch: If any one of the measured

parameters, e.g. Oil Pressure, EGT, should exceed normal operating limits, it will be automatically alerted on the display units. The purpose of the reset switch is to clear the alerts from the display when the limit exceedance no longer exists.



ENGINEERS


Figure 32 shows an EICAS Control Panel

EICAS Control Panel Figure 32

ENGINE STATUS MAX INDRESETL AUTO R

L BOTH REVENTRECORD

DISPLAY COMPUTERBRT

BAL

BRT THRUST REF SET



ENGINEERS


1.4.10 ALERT MESSAGES Up to eleven alert messages can be displayed on the upper display. They appear in order of priority and in appropriate colour.

Level A - Red - Warnings. Level B - Amber - Cautions.

Level C - Amber - Advisory.

Level A These warnings require “immediate action” by the crew to correct the failure. Master warning lights are also illuminated along with corresponding aural alerts from the central warning system. Level B These cautions require “immediate awareness” of the crew and also may require possible corrective action. Caution lights and aural tones, were applicable, may accompany the caution. Level C These advisories require “awareness” of the crew. No other warnings/cautions are given and no aural tones are associated with this level. The messages appear on the top line at the left of the display screen. In order to differentiate between a caution and an advisory, the advisory is always indented one space to the right.



ENGINEERS


Figure 33. shows EICAS alert messages Level A, B and C.

EICAS Alert Messages Figure 33

The master warning and caution lights are located adjacent to the display units together with a “Cancel” and “Recall” switch (see Figure 29). Pushing the “Cancel” switch removes only the caution and advisory messages, warning messages cannot be cancelled. The “Recall” switch is used to recall the previously cancelled caution and advisory messages for display. On the display, the word RECALL appears on the bottom of the display.

WARNING

CAUTION

CANCEL

RECALL

MASTER WARNING & CAUTION LIGHTS

TAT 15°c

N1

EGT

V V V V V V V

775 999

106 2

110.010

6 2

70.0APU FIRER ENGINE FIRECABIN ALTITUDEC SYS HYD PRESSR ENG OVHTAUTOPILOT C HYD QTY R YAW DAMPER

L UTIL BUS OFF

LEVEL AWARNING

LEVEL BCAUTION

LEVEL CADVISORY

A - WARNING (RED)

B - CAUTION (AMBER)

C - ADVISORY (AMBER)

RED

AMBER



ENGINEERS


Messages are automatically removed from the display when the associated condition no longer exists. If more than one message is being displayed, then as a message is automatically removed, all messages below it will move up one line. If a new fault appears, its associated message is inserted on the appropriate line of the display. This will cause old messages to move down one line. If there are more messages than can be displayed at one time, the whole list forms what is termed a “Page”, and the lower messages are removed and a page number appears on the lower right-hand side of the list. Additional pages are selected by pressing the “Cancel” switch on the Master Warning/Caution panel.

1.4.11 FAILURE OF DU/DISPLAY SELECT PANEL Should a DU fail, all messages, primary and secondary, appear on the remaining DU. Secondary messages may be removed by pressing the 'ENGINE' switch on the display select panel. They may be re-established by pressing the same switch. The format displaying all information is referred to as 'Compact Format'. Should the display select panel fail, status information cannot be displayed.



ENGINEERS


1.4.12 MAINTENANCE FORMAT

Maintenance pages can be called forward on the ground using the Maintenance Panel, refer to Figure 34.

EICAS Maintenance Panel Figure 34

PERF

APU

ELEC

HYD

ECS

MSG

ENGEXCD

CONF

MCDP

DISPLAY SELECT

EICAS MAINT EVENTREAD

AUTO MAN

REC ERASE

TEST

ENVIRONMENTAL CONTROLSYSTEM AND MAINTENANCE

MESSAGE FORMATS

ELECTRICAL AND HYDRAULICSYSTEM FORMAT

PERFORMANCE ANDAUXILLIARY POWER

UNIT FORMATSSELECTS DATA FROM

AUTO OR MANUAL EVENTIN MEMORY

ERASES STORED DATACURRENTLY DISPLAYED

RECORDS REAL-TIMEDATA CURRENTLY DISPLAYED

(IN MANUAL EVENT)

BITE TEST SWITCHFOR SELF-TEST ROUTINEENGINE

EXCEEDANCES

CONFIGURATION ANDMAINTENANCE

CONTROL/DISPLAYPANEL



ENGINEERS


Maintenance pages appear on the lower DU and include system failures, which have occurred in flight or during ground operations. While these pages are selected, the upper DU displays a 'Compact Format' with the message 'PARKING BRAKE' in the top left of the screen. A self-test of the whole system, which can only be activated when an aircraft is on the ground and the parking brake set, is performed by means of the “TEST” switch on the maintenance panel. When the switch is momentarily pressed, a complete test routine of the system, including interface and all signal-processing circuits and power supplies, is automatically performed. For this purpose an initial test pattern is displayed on both display units with a message in white to indicate the system being tested, i.e. 'L or R EICAS' depending on the setting of the selector switch on the display select panel. During the test, the master caution and warning lights and aural devices are activated, and the standby engine indicator is turned on if its display control switch is at 'AUTO'. The message 'TEST IN PROGRESS' appears at the top left of display unit screens and remains in view while testing is in progress. On satisfactory completion of the test, the message 'TEST OK' will appear. If a computer or display unit failure has occurred, the message 'TEST FAIL' will appear followed by messages indicating which of the units has failed.

A test may be terminated by pressing the 'TEST' switch a second time or, if it is safe to do so, by releasing an aircraft's parking brake. The display units revert to their normal primary and secondary information displays.



ENGINEERS


Figure 35 shows the display formats seen during the Maintenance format.

Maintenance Mode Displays Figure 35

10

62

10

62

N1

EGT

85.0 85.0

450 450

PARKING BRAKE

50 OIL PRESS 50105 OIL TEMP 10020 OIL QTY 201.9 N2 VIB 1.9

96.1 96.1

97.0 N2 97.08.4 FF 8.4

ELEC/HYD

LOADAC-VFREQDC-ADC-V

00

1028

0.7812040214028

0.8512539815027

0.00000

28

0.0000

STBYBAT

APUBAT

GNDPWRL R

HYD QTYHYD PRESSHYD TEMP

0.823230

50

O/FULL3210

47

0.722140

115

L C R

AUTO EVENT R HYD QTY

AUTO EVENTSYSTEM FAILURESAUTOMATICALLY

RECORDED DURINGFLIGHT

INDICATED WHENEICAS IN

MAINTENANCE FORMAT



ENGINEERS


PAGE INTENTIONALLY

BLANK



ENGINEERS


1.5 FLIGHT DATA RECORDER SYSTEM (FDRS) The flight data recorder receives and stores selected aircraft parameters from various aircraft systems and sensors in a crash-protected solid state memory. The Digital Flight Data Acquisition Unit (DFDAU) of the Aircraft Information Management System (AIMS) receives all the FDR data. The DFDAU then processes the data and sends it to the FDR, where it is stored. The FDRS operates during any engine start, while the engine is running, during test, or when the aircraft is in the air. The FDR records the most recent 25 hours of flight. In addition to the data recording function, the FDR also has monitor circuits, which send fault information back to the DFDAU. Note: FDRS fitted to a Helicopter start recording only when the rotors turn (i.e. take-off). 1.5.1 OPERATION The AIMS receives power control data from several aircraft systems, power goes to the FDR when the logic is valid. Power control data includes:

Engine Start.

Engine Running.

Air/Ground Logic.

Test. 1.5.2 ANALOGUE DATA The DFDAU receives status and maintenance flag data from the FDR. The DFDAUs receive key events from the VHF and HF LRUs and variable analogue data from the TAT, AOA and engine RPM sensors.



ENGINEERS


1.5.3 DIGITAL DATA The ARINC 429/629 buses provide engine, airframe data and air/ground logic. Engine data includes:

Engine parameters, normal and exceedances.

Commands.

Actual Thrust. Airframe data includes:

Flight deck switch position

Flight control positions

Mode selections on control panels in the flight deck. The DFDAU receives status from the engine and airframe sensors. The DFDAU also receives data and status from the electrical power system. The flight controls ARINC629 buses provide flight data and navigational data. Flight data includes:

Flight control position.

Commands

Status. Navigation data includes:

Pitch, Roll and Yaw attitude.

Acceleration data.

Status.



ENGINEERS


ARINC 429 bus provides navigational (NAV) radio/NAV data and communication (COMM) radio data. Radio data includes:

Radio Frequencies.

Mode.

Parameters.

Status. NAV data is the aircraft’s present position (LAT/LONG) and sensor status. COMM data is radio control panel frequencies and sensor status. The left AIMS cabinet sends left/right DFDAU data on the ARINC 573 data bus to the FDR. The DFDAU sends fault data, status and ground test results to the Central Maintenance Computer. Figure 113 shows a FDR.

Flight Data Recorder Figure 113

UNDERWATERLOCATING

DEVICE



ENGINEERS


Figure 114 shows FDR block schematic diagram.

Flight Data Recorder Block Schematic Figure 114

ARINC 429

ARINC 629

ANALOGUE

ANALOGUEDISCRETES

AIRCRAFTSYSTEMS

AIMS

AR

INC

573

FAULTMONITORING

DFDAU

FDR



ENGINEERS


The following is taken from ANO Section 1, order 53. 1.5.4 USE OF FLIGHT RECORDING SYSTEMS 1. On any flight on which a FDR, a cockpit voice recorder or a combined

cockpit voice recorder/flight data recorder is required to be carried in an airplane, it shall always be in use from the beginning of the take-off run to the end of the landing run.

2. On helicopters, it shall always be in use from the time the rotors first turn

for the purpose of taking off until the rotors are next stopped.

SYSTEMS INSTRUMENTS PAGE 1 of 59


ENGINEERS


1 INSTRUMENT SYSTEMS (ATA 31) Aircraft instruments can, on initial observation, appear a bewildering mass of dials or 'TV ' type screens. The different types of instrumentation required fall into one of the following types:

Pressure instruments Gyroscopic instruments Compasses Mechanical indicators Electronic instruments

1.1 PRESSURE INSTRUMENTS 1.1.1 Air Data Instruments An Air Data system of an aircraft is one which the total pressure created by the forward motion of an aircraft, and the static pressure of the atmosphere surrounding it, are sensed and measured in terms of speed, altitude and rate of change of altitude. The measurement and indication of these three parameters may be achieved by connecting the appropriate sensors, either directly to mechanical-type instruments, or to a remotely-located Air Data Computer (ADC), which then transmits the data in electrical signal format to electro-mechanical or servo-type instruments. The basic Air Data Instruments display airspeed, altitude, Mach number and vertical speed. All are calculated from air pressure received from a Pitot/Static source. 1. Static air pressure, which is simply the outside air pressure at the instant of

measuring.

2. Pitot pressure is the dynamic pressure of the air due to the forward motion of the aircraft and is measured using a tube, which faces the direction of travel.



ENGINEERS


Figure 1 shows a Pressure head as fitted to aircraft to allow Pitot and Static pressures to the relevant indicators.

Aircraft Pressure Head Figure 1

Indicated Airspeed (IAS), Mach No, Barometric Height (Height above sea level), and Vertical speed (Rate of climb/dive) are derived from the Pitot/Static inputs.

IAS = Pitot minus Static - (In knots). Mach No = Pitot - Static divided by Static. Baro Ht = Static - (In feet). Vertical Speed = Change in Static pressure - (X 1000ft/min).

FORWARD

PITOT PROBE STATIC VENTS

PITOT LINE STATIC LINE

HEATER CONNECTION



ENGINEERS


Figure 2 shows typical aircraft static vent:

Aircraft Static Vent Figure 2

1.1.2 Location Of Probes and Static Vents The choice of probe/vent locations is largely dependent on the type of aircraft, speed range and aerodynamic characteristics, and as result there is no common standard for all aircraft. On larger aircraft it is normal to have standby probes and static vents. These are always located one on each side of the fuselage and are interconnected so as to balance out dynamic pressure effects resulting from any Yawing or side-slip motion of the aircraft.

STATICVENT

STATICPIPE

FUSELAGE



ENGINEERS


Figure 3 shows the location of probes and vents on a Boeing 737.

Boeing 737 Air Data Probe and Vent Location Figure 3

Pitot and static pressures are transmitted through seamless and corrosion-resistant metal (light alloy) pipelines. Flexible pipelines are also used when connections to components mounted on anti-vibration mountings is required. In order for an Air Data System to operate effectively under all flight conditions, provision must also be made for the elimination of water that may enter the system as a result of condensation, rain, snow, etc. This will reduce the probability of “Slugs” of water blocking the lines. This provision takes the form of drain holes in the probes, drain taps and valves in the system’s pipelines. The drain holes within the probes are of diameter so as not to introduce errors into the system.



ENGINEERS


Methods of draining the pipelines varies between aircraft types and are designed to have a capacity sufficient to allow for the accumulation of the maximum amount of water that could enter the system between maintenance periods. Figure 4 shows a typical water drain valve.

Water Drain Valve Figure 4

The three primary instruments in the Air Data System are:

Altimeter (Baro Ht). Indicated Air Speed (IAS) Indicator. Vertical Speed Indicator.

The IAS is often combined to display Mach No as well as indicated airspeed and is referred to as the “Combined Speed Indicator”. Figure 5 shows the connection and equations for the primary Air Data instruments.

ORANGEFLOAT

INDICATORTRANSPARENTPLASTIC PIPE

DRAINVALVE

BAYONETFITTING

CAP(SELF SEALING)



ENGINEERS


Air Data Instrumentation Figure 5



ENGINEERS


1.2 ALTIMETER This operates on the aneroid barometric principle, i.e. responds to changes in atmospheric pressure, and is calibrated to indicate these changes in terms of equivalent altitude values. Figure 6 shows a typical altimeter.

Altimeter Figure 6

The pressure sensing element consists of an aneroid capsule, which transmits deflections in response to pressure changes. The capsule is contained within a sealed container that is evacuated to the static pressure. A mechanical linkage connects the capsule to a pointer, which indicates the aircraft’s height above sea level. There is a facility to set the correct pressure of the day in millibars so that the instrument displays the correct height.

5

1 0 1 3SBY

01

46

37

28

X 100 ft

3 5 0 0

5

9

MB

0



ENGINEERS


Figure 7 shows the simplified operation of the altimeter.

Simplified Altimeter operation

Figure 7

1.3 “Q” CODE SETTINGS FOR ALTIMETERS The setting of altimeters to the barometric pressures prevailing at various flight levels and airports is part of the flight operating techniques. It is essential for maintaining adequate separation between aircraft and for terrain clearance during take-off and landing. In order to make the settings, flight crews are dependant on observed meteorological data which is requested and transmitted from ATC and form part of the ICAO “Q” code of communication. There are three code letter groups commonly used in connection with altimeter setting procedures:

1. QNH.

2. QFE.

3. QNE. QNH: Setting the barometric pressure to make the altimeter read airport elevation above-sea level on landing and take-off. When used for landing and take-off, the setting is generally known as “Airport QNH”. Any value set is only valid in the immediate vicinity of the airport concerned.



ENGINEERS


Since an altimeter with a QNH setting reads altitude above sea level, the setting is also useful in determining terrain clearance when an aircraft is en-route. Fir this purpose, the UK and surrounding seas are divided into fourteen “Altimeter Setting Regions”, each transmitting an hourly “Regional QNH” forecast. QFE: Setting the barometric pressure prevailing at an airport to make the altimeter read zero on landing at, or taking off from, that airport. The zero reading is regardless of the airport’s elevation above sea level. QNE: Also known as the “Standard Altimeter Setting (SAS)”. The barometric pressure is set to 1013.25 mb and is used for flights above a prescribed “Transmission Height” and has the advantage that with all aircraft using the same airspace and flying on the same altimeter setting, the requisite separation between aircraft can more readily be maintained. The transition altitude within the UK airspace is usually 3000 - 6000'. Figure 8 shows QNH, QFE and QNE definitions.

QNH, QFE and QNE Definitions Figure 8

SEA LEVEL

QNHHEIGHT ABOVE

SEA LEVEL

QNEFLIGHT LEVEL

STANDARD SETTING1013.25 MILLIBARS

QFEHEIGHT ABOVE

AIRFIELD



ENGINEERS


1.4 COMBINED SPEED INDICATOR This indicator is one, which combines the functions of both a conventional indicator and a Machmeter. Figure 9 shows a typical Combined Speed Indicator (CSI).

Combined Speed Indicator

Figure 9 The internal mechanism consists of two elements (pointer and fixed scale for IAS and a digital readout for Mach No). There is also a second pointer on the IAS scale, this is known as the “Velocity Maximum Operating (Vmo)”. It indicates the aircraft’s maximum safe operating speed over its operating altitude range. To set the desired speed for operation, the flight crew uses the command bug. This speed in turn is the datum speed for the Autothrottle or Fast/Slow speed indicator. The external index bugs are used to set various reference speeds (take-off, flap retract speeds etc.).



ENGINEERS


Figure 10 shows a simplified IAS operation.

IAS Operation Figure 10

STATIC

PITOT



ENGINEERS


1.5 VERTICAL SPEED INDICATOR (VSI) These indicators (also known as Rate-of Speed indicators) are very sensitive differential pressure gauges, designed to indicate the rate of altitude change from variations in static pressure alone. Figure 11 shows a VSI.

Vertical Speed Indicator (VSI) Figure 11

Since the rate at which the static pressure changes is involved in determining vertical speed, a time factor has to be incorporated as a pressure function. This is accomplished by using a special air-metering unit in the sensing system. Its purpose is to create a lag in static pressure across the system and so establish the required pressure difference.

0

1.5

24

6

1.5

24

6

VERTICALSPEED

DOWN

UP

1000FT PER MIN

VSI

RATE OFDIVE SCALE

1,000 ft per sec

RATE OFCLIMB SCALE1,000 ft per sec

MAX INDICATED6,000 ft per sec

RATE OFCLIMB/DIVE

POINTER



ENGINEERS


Figure 12 shows a simplified VSI operation.

VSI Operation Figure 12

1.6 AIR DATA SYSTEMS The complexity of an Air Data System depends primarily upon the type and size of the aircraft, the number of locations at which primary air data is to be displayed, the type of instruments installed, and the number of other systems requiring air data inputs.

0

CLIMB

DIVE

STATICVENT

METERINGUNIT

CAPSULEMECHANICAL

LINKAGE

POINTERAND

SCALE



ENGINEERS


Figure 13 shows a typical air data system for a large aircraft.

Air Data System Figure 13

FLTREC

PC MS 1 A/S 1 ADC 1

DIFFPRESS

MS 2

A/S 2 ADC 2

PRESSURE HEADS

UPPERLOWER

UPPERLOWER

STA

TIC

PIT

OT

ALT

VS

IAS

STA

TIC

PIT

OT

IAS

VS

ALT

F/O

CAPT

PRESSURE HEADS



ENGINEERS


1.7 GYROSCOPIC INSTRUMENTS A number of instruments depend on the use of gyroscopes for their correct operation. It is useful to know the basic principles of how they work, before describing, in some depth, what they do. 1.7.1 Gyroscopic Properties As mechanical device a gyroscope may be defined as a system containing a heavy metal wheel (rotor), universally mounted so that it has three degrees of freedom: Spinning freedom: About an axis perpendicular through its center (axis of

spin XX). Tilting Freedom: About a horizontal axis at right angles to the spin axis

(axis of tilt YY). Veering Freedom: About a vertical axis perpendicular to both the other

two axes (axis of veer ZZ). The three degrees of freedom are obtained by mounting the rotor in two concentrically pivoted rings, called inner and outer rings. The whole assembly is known as the gimbal system of a free or space gyroscope. The gimbal system is mounted in a frame so that in its normal operating position, all the axes are mutually at right angles to one another and intersect at the center of gravity of the rotor. The system will not exhibit gyroscopic properties unless the rotor is spinning. When the rotor is spinning at high speed the device becomes a true gyroscope possessing two important fundamental properties:

Gyroscopic Inertia (Rigidity). Precession.

1.7.2 Rigitity The property, which resists any, force tending to change the plane of rotor rotation. It is dependent on:

1. The mass of the rotor. 2. The speed of rotation.



ENGINEERS


1.7.3 Precession The angular change in direction of the plane of rotation under the influence of an applied force. The change in direction takes place, not in line with the force, but always at a point 90º away in the direction of rotation. The rate of precession also depends on:

1. The strength and direction of the applied force. 2. The angular velocity of the rotor.

Figure 14 shows a gyroscope.

Gyroscope. Figure 14

OUTERRING

FRAME

INNERRING

ROTOR

Y

Y

Z

Z

X

X



ENGINEERS


Figure 15 shows the characteristics of gyro rigidity.

Gyro Rigidity Figure 15

Gyro A has its spin axes parallel with the Earth's spin axes, located at the North Pole. It could hold this position indefinitely. Gyro B has its spin axes parallel to the Earth's spin axes, but located at the Equator. As the Earth rotates, it would appear to continually point North. Gyro C is also situated at the Equator. As the Earth rotates, it appears to rotate about its axes, however it is the Earth that is rotating and not the gyro.

B

A

C



ENGINEERS


This rigidity can be used in a number of gyro instruments including the directional gyro. If an external force is applied to a spinning gyro, its effect will be felt at 900 from the point of application, in the direction of gyro rotation. This is known as precession. It can be seen in Figure 16, that if a force is applied to the bottom of the rotating wheel, it will rotate about its horizontal axis. This property is not wanted in some instruments, such as directional gyros. The use of precession is used in turn indicators, which will be covered later.

Gyro Precession Figure 16

SPIN AXIS

DIRECTIONOF

ROTATION

APPLIEDFORCE

90º

DIRECTIONOF

PRECESSION

PRECESSION RATE= APPLIED FORCE

90º IN THE DIRECTION OF SPIN



ENGINEERS


1.7.4 Vertical Gyro Figure 17 shows the effects on a free gyro in an aircraft circling the earth. As can be seen, it would only be perpendicular to the earth's surface at two points.

Behaviour of a Vertical Gyro Figure 17



ENGINEERS


In order for the gyro to be used to indicate the aircraft's attitude, it has to be corrected to continually be aligned to the vertical. These corrections are very slow and gentle, since the amount of correction needed, for example, in a ten-minute period is small. Figure 18 shows a vertical gyro corrected to the local vertical.

Corrected Vertical Gyro Figure 18



ENGINEERS


Instruments that use either the rigidity or the precession of gyros are:

Gyro Horizon Unit. Attitude Director Indicator. Standby Horizon Unit. Direction Indicator. Turn and Slip Indicator. Turn Co-ordinator.

1.8 GYRO HORIZON UNIT The Gyro Horizon Unit gives a representation of the aircraft’s pitch and roll attitudes relative to its vertical axis. For this it uses a displacement gyroscope whose spin axis is vertical. Figure 19 shows a displacement gyro and the two axis of displacement.

Displacement Gyro Figure 19

PITCHROLL



ENGINEERS


Indications of attitude are presented by the relative positions of two elements, one symbolizing the aircraft itself, the other in the form of a bar stabilized by the gyroscope and symbolizing the natural horizon. Figure 20 shows a typical Gyro Horizon Unit.

Gyro Horizon Unit Figure 20

The gimbal system is so arranged so that the inner ring forms the rotor casing and is pivoted parallel to an aircraft’s lateral axis (YY1); the outer ring is pivoted at the front and rear ends of the instrument case, parallel to the longitudinal axis (ZZ1). The element symbolizing the aircraft may either be rigidly fixed to the case, or it may be externally adjustable for setting a particular pitch trim reference.

SPERRY

3 3

6 6

AIRCRAFTSYMBOL

ROLLSCALE

HORIZONBAR

ROLLPOINTER\



ENGINEERS


Figure 21 shows the construction of the Gyro Horizon unit.

Construction of a Gyro Horizon Unit Figure 21

In operation the gimbal system is stabilized so that in level flight the three axes are mutually at right angles. When there is a change in the aircraft’s attitude, example climbing, the instrument case and outer ring will move about the YY1 of the stabilized inner ring.

X1

X

Y1

Y Z1

ZPIVOTPOINT

ROTOR

SYMBOLICAIRCRAFT

ROLLPOINTER& SCALE

HORIZONBAR

OUTERRING

BALANCEWEIGHT



ENGINEERS


The horizon bar is pivoted at the side and to the rear of the outer ring and engages an actuating pin fixed to the inner ring, thus forming a magnifying lever system. The pin passes through a curved slit in the outer ring. In a climb attitude the pivot carries the rear end of the bar upwards so that it pivots about the stabilized actuating pin. The front end of the bar is therefore moved downwards through a greater angle than that of the outer ring, and since the movement is relative to the symbolic aircraft element, the bar will indicate a climb attitude. Figure 22 shows climb attitude operation.

Climb Attitude operation. Figure 22

ZZ11

ZZ

XX11

XX

HORIZON BARHORIZON BAR



ENGINEERS


Changes in the lateral attitude of an aircraft, i.e. rolling, displaces the instrument case about the axis (ZZ1), and the whole stabilized gimbal system. Hence, lateral attitude changes are indicated by movement of the symbolic aircraft element relative to the horizon bar, and also by relative movement between the roll angle scale and pointer. Figure 23 shows roll attitude operation.

Roll attitude operation Figure 23

Freedom of gimbal system movement is 360º for roll axis and 85º for the and pitch axis. The pitch scale is restricted by means of a resilient stop. This will prevent gimbal lock.

DATUMDATUMXX11

XX

YY YY11

BANK TOBANK TOPORTPORT



ENGINEERS


1.9 ATTITUDE DIRECTOR INDICATOR This unit performs the same functions as a Gyro Horizon unit; i.e. it establishes a stabilized reference about the pitch and roll axes of an aircraft. Instead, however, of providing attitude displays by direct means, it is designed to be operated via a synchro system, which produces and transmits attitude-related signals to the indicator. The synchro system includes a attitude reference source and a computer linked into the aircraft’s navigational system to produce flight director signals for the flight crew to follow to ensure the aircraft follows the required course. Figure 24 shows a typical Attitude Director Indicator (ADI)

Attitude Director Indicator (ADI) Figure 24

TEST

1

2

1

2F

S

A T T YWR

GSLFD



ENGINEERS


1.10 STANDBY HORIZON UNIT Most aircraft currently in service use Flight Director systems, or more sophisticated electronic flight instrument systems, all of which comprise indicators displaying not only attitude data, but navigational data as well. In such aircraft, the role of the conventional gyro horizon is mainly used as a standby instrument located on the center instrument panel. It is used as a reference in the event of a failure that might occur in the attitude display systems. Figure 25 shows a Standby Horizon Unit (SHU).

Standby Horizon Unit

Figure 25

The gyro is powered by 115V; three phase ac supplied from a static inverter, which in turn is supplied by 28V from the battery busbar. In place of the stabilized horizon bar a stabilized attitude sphere is used as the reference. The upper element is coloured blue to display climb attitudes, and black/brown for descending attitudes.

20 20

20 20

C

DPITCHSCALE

AIRCRAFTSYMBOL

ROLLSCALE

POWER“OFF”FLAG

PITCH ERECTION/TRIM KNOB



ENGINEERS


A pitch trim adjustment and fast erection facility is provided, both being controlled by a knob on the lower right-hand corner of the indictor. When the knob is rotated the aircraft symbol can be positioned through !5º, thereby establishing a variable pitch trim reference. Pulling the knob out and holding it actuates the fast-erection circuit. 1.11 DIRECTION INDICATORS This indicator was the first gyroscopic instrument to be introduced as a “Heading Indicator” and although for most aircraft currently in service it has been superseded by remote-indicating compass systems (see later). The instrument uses a horizontal axis gyroscope and, being non-magnetic, is used in conjunction with a magnetic compass. In its basic form, the outer ring of the gyro carries a circular card, graduated in degrees, and referenced against a lubber line fixed to the gyro frame. When the rotor is spinning, the gimbal system and card are stabilized so that, by turning the frame, the number of degrees through which it is turning may be read on the card. Figure 26 shows a Directional Indicator.

Directional Indicator Figure 26

180 170

HEADINGSCALE LUBBER

LINE

CAGING/SETTINGKNOB



ENGINEERS


In the directional gyro, the rotor is enclosed in a case, or shroud, and supported in an inner gimbal which is mounted in an outer gimbal, the bearings of which are located top and bottom on the indicator case. The front of the case contains a cut-out through which the card is visible, and also a lubber line reference. The caging/setting knob is provided at the front of the case to set the indicator onto the correct heading (magnetic). When the setting the heading, the inner gimbal has to be caged to prevent it from precessing as the outer gimbal is rotated. Figure 27 shows the construction of a directional gyro.

Directional Gyro Figure 27

COMPASSCARD

ROTORASSEMBLY

INNERGIMBAL

RING

VERTICAL GIMBALRING

CAGING/SETTING

KNOB

SYNCHRONISERRING



ENGINEERS


1.12 TURN & SLIP INDICATOR This indicator contains two independent mechanisms:

1. A gyroscopically controlled pointer mechanism for the detection and indication of the rate at which an aircraft turns.

2. A mechanism for the detection and indication of slip/slide.

A gimbal ring and magnifying system, which moves the pointer in the correct sense over a scale calibrated in what is termed “Standard Rates”, actuate the rate of turn pointer. Although they are not always marked on a scale, they are classified as follows:

Rate 1 - Turn Rate 180º per minute.



Rate 4 - Turn Rate 720º per minute. Figure 28 shows a typical Turn & Slip indicator.

Turn & Slip Indicator

Figure 28

2 MIN

RATE OF TURNINDICATOR

SLIP/SLIDEINDICATORRATE OF

TURN2 MIN - 360º



ENGINEERS


For the detection of rates of turn, a rate gyroscope is used and is arranged in the manner shown in figure 29.

Rate Gyro Turn Indicator Figure 29

It differs in two respects from the displacement gyro as it only has one gimbal ring and a calibrated spring restraining in the longitudinal axis YY1. When the indicator is in its normal operating position the rotor spin axis, due to the spring restraint, will always be horizontal and the turn pointer at the zero datum. With the rotor spinning, its rigidity will further ensure that the zero position is maintained.

Y

Y1X

X1P

F

FWD

INPUTAXIS



ENGINEERS


When the aircraft turns to the left about the vertical input axis the rigidity of the rotor will resist the turning movement, which it detects as an equivalent force being applied to its rim at point F. The gimbal ring and rotor will therefore be tilted about the longitudinal axis as a result of precession at point P. As the gimbal ring tilts, it stretches the calibrated spring until the force it exerts prevents further deflection of the gimbal ring. Since precession of a rate gyro is equal to its angular momentum and the rate of turn, then the spring force is a measure of the rate of turn. Actual movement of the gimbal ring from its zero position can, therefore, be taken as the required measure of turn rate. 1.12.1 Bank Indication In addition to the primary indication of turn rate, it is also necessary to have an indication that an aircraft is correctly banked for the particular turn. A secondary indicating mechanism is therefore provided, which, depends for its operation on the effect of gravitational and centrifugal forces. A method commonly used for bank indication is one utilizing a ball in a curved liquid-filled glass tube as shown in Figure 26. In the normal level flight the ball is held at the center of the tube by the force of gravity. Let us assume the aircraft turns left at a certain airspeed and bank angle. The indicator case and the tube move with the aircraft and centrifugal force (CF) in addition to that of gravity acts upon the ball and tends to displace it outwards from the center of the tube. However, when the turn is executed at the correct bank angle and matched with airspeed, then there is a balanced condition between the two forces and so the resultant force (R) hold the ball in the center of the tube. If the airspeed were to be increased during the turn, then the bank angle and centrifugal force would also be increased. As long as the bank angle is correct for the appropriate conditions, the new resultant force will still hold the ball central. If the bank angle for a particular rate of turn is not correct (under-banked/over-banked), then the aircraft will tend to either skid or slip. In the skid condition the centrifugal force will be the greatest, whereas in the slip condition the force of gravity is greatest.



ENGINEERS


Figure 30 shows bank indication for various aircraft bank conditions.

Bnk Indications Figure 30



ENGINEERS


1.13 TURN CO-ORDINATOR The final instrument in this group is the turn co-ordinator. Basically, its mechanism is changed slightly from the turn and slip indicator, so that it senses rotation about the longitudinal axis, (bank) as well as the vertical axis, (turn). This gives a more accurate indication to the pilot, of the turning of the aircraft. Figure 31 shows a Turn co-ordinator indicator.

Turn co-ordinator Indicator

Figure 31

2 MINNO PITCH

INFORMATION

TURN COORDINATION

L RRATE OF

TURN

AIRCRAFTSYMBOL

TURNCOORDINATOR



ENGINEERS


1.14 HORIZONTAL SITUATION INDICATOR This indicator derives its name from the fact that its display presents a pictorial plan of the aircraft’s situation in the horizontal plane in the form of its heading, VOR/LOC deviation and other data relating to navigation. Figure 32 shows a typical HSI.

Horizontal Situation Indicator Figure 32



ENGINEERS


The aircraft symbol is fixed at the center of the instrument and displays the heading of the aircraft in relation to a rotating compass card and the VOR/LOC deviation bar (lateral bar). The selector knobs at the bottom corners of the instrument permit the setting of desired magnetic heading and VOR course.



ENGINEERS


1.14.1 Compass Systems The compass has, since the earliest times, given information to travellers with regards to the direction to go. Mounting a compass on a moving object, whether it was a vehicle, a ship or an aircraft poses certain problems. This includes how to mount the compass without the, motion (maybe violent), upsetting the device. Another problem that besets compasses is the fact that they usually point to magnetic north, which slowly moves, and not true north, the difference between the two is something like 1,300-miles/2,000 km. This is of little concern if we are moving slowly, on a boat, in the vicinity of the equator, but vital in an aircraft flying what is known as a 'Trans-polar route' from say, New York to Tokyo. The effect this has on navigational charts is referred to as 'variation'. Figure 33 shows the difference between True North and Magnetic North.

True North & Magnetic North Figure 33

MAGNETICNORTH POLE

GEOGRAPHICALNORTH POLE

17.5º EVARIATION

0º EVARIATION

11º WVARIATION



ENGINEERS


1.15 DIRECT READING COMPASS Direct-reading compasses have the following common principal features:

1. Magnet system housed in a bowl. 2. Liquid damping and liquid expansion compensation.

Figure 34 shows a direct reading compass used as a standby compass in most aircraft.

Standby Compass Figure 34



ENGINEERS


The magnet system comprises an annular cobalt-steel magnet to which is attached a light-alloy card. The card is graduated in increments of 10º, and referenced against a lubber line fixed to the interior of the bowl. The system is pendulously suspended by an iridium-tipped pivot resting in a sapphire cup supported in a holder or stem. The bowl is of a plastic (diakon) and so moulded that it has a magnifying effect on the card and its graduations. It is filled with a silicone fluid to prevent the card oscillating or overshooting after changes of heading. The fluid also provides the system with a certain buoyancy, thereby reducing the weight on the pivot and so diminishing the effects of friction and wear. Changes in the volume of the fluid due to temperature changes, and their resulting effects on damping efficiency, are compensated by a bellows type of expansion device secured to the rear of the bowl. Compensation of the effects of deviation due to longitudinal and lateral components of aircraft magnetism is provided by permanent magnet coefficient “B” and “C” corrector assemblies secured to the compass mounting plate. A small lamp is also provided for illuminating the card. Figure 35 shows a complete standby compass indicator.

Standby Compass Figure 35

S 15 1221

B C

ELECTRICALCONNECTIONFOR LIGHTING

CO-EFFICIENT “A”ADJUSTMENT

CO-EFFICIENT “B”ADJUSTMENT

CO-EFFICIENT “C”ADJUSTMENT

LUBBERLINE



ENGINEERS


1.16 REMOTE READING COMPASS A remote reading compass, is basically one in which an element detects an aircraft’s heading with respect to the horizontal component of the earth’s magnetic field in terms of flux and induced changes in voltage. It then transmits these changes via a synchronous/servo system to a heading indicator. There are two types of remote reading compass systems:

1. The detector element monitors a directional gyro unit linked with a

heading indicator. 2. The detector element operates in conjunction with the platform of an

inertial navigation system (INS). 1.16.1 Detector Unit (Flux Valve) The detector unit detects the effect of the earth’s magnetic field as an electromagnetically induced voltage and controls the heading indicator by means of a variable secondary output voltage signal. The construction of the element takes the form of a three-spoked wheel, slit through the rim between the spokes so that they, and their section of rim, act as three individual flux collectors. Figure 36 shows the construction of a flux valve.

Flux Valve Construction Figure 36

AC POWER

A

B

C

LAMINATEDCOLLECTOR

HORNS

EXCITERCOIL

SECONDARYPICK-OFF

COILS

A

BC



ENGINEERS


The paths taken by the earth’s magnetic field through the spokes for different headings is shown at Figure 37.

Earth’s Flux path Figure 37

The detector unit on its own is not very accurate by virtue of its limited pendulous suspension arrangement. Errors will occur as a result of its tilting under the influence of acceleration forces, e.g. during speed changes on a constant heading and during turns. It is necessary to incorporate within the system a means of monitoring the detector’s output. The horizontal directional gyro is used to give the system short-term accuracy with the detector unit providing long-term accuracy.

PATH OFEARTH’S

FIELD

C

BA

C

BA

C

BA

C

BA



ENGINEERS


Figure 38 shows the arrangement of a remote reading gyro compass system.

Gyro Magnetic Compass System Figure 38

115v 400 Hz

B

+ _

C

+ _

VOR/ADF SYNCDG

SLAVED

N 3

6EW

S

33

30

1512

21

24 ADF

ADF

VOR VOR



ENGINEERS


Figure 39 shows a schematic of a Gyro Magnetic Compass system.

Gyro Magnetic Compass System Schematic Figure 39

N

S

W E 26V AC400 Hz

GYRO

CT

CT

CX

M

M

TG

DETECTORUNIT



ENGINEERS


1.17 ANGLE OF ATTACK (AOA) Apart from the main flight instruments, one item of information that the pilot needs to know at various stages of flight is the angle of attack. Earlier aircraft had a range of devices that gave the pilot indication of an approaching stall, which was an essential indicator but knowing the angle of attack has become an essential part of flying modern, larger aircraft. The simplest forms of angle of attack indicators are the AOA probe and the stall vane. The probe contains slots on the leading edge of the probe itself and, depending on the angle of attack; the air flowing through the different slots move a 'paddle' which indicates the AOA electrically in the cockpit. The stall vane is rather like a small weather vane mounted on the side of the aircraft. The vane follows the airflow, much like the weather vane, but indicating, not pitch angle, but the angle of the airflow relative to the aircraft centerline. i.e. the angle of attack. Figure 40 shows a vane type Angle of Attack transducer.

Angle of Attack Transducer

Figure 40

∝

ANGLEOF

ATTACK

FLIGHT PATH

AIRFLOW

AIRCRAFTLONGITUDINAL

AXIS

VANE ARMANGLE OF ATTACK

TRANSDUCER



ENGINEERS


1.18 STALL WARNING INDICATION To maintain lift at low airspeed, the angle of attack is increased. When this angle is above a critical angle, the aircraft wings will not produce enough lift to support the aircraft, which will begin to stall. Before this situation occurs, the aircraft will shake heavily, this being a natural alert to the pilot. If, however, the aircraft is configured for an approach (Wheels & Flaps down), the airspeed difference between the natural warning and the actual stall is very small, so an alert must be generated before the stall occurs. Modern performance aircraft use the output from an Angle of Attack probe, connected to a Stall Warning system. The stall warning system also has other sensor inputs (Flap, Slat positions). Once the critical angle prior to actual stall is reached, the stall warning system initiates a "Audio warning" and operates a "Stick Shaker", which actually shakes the control column. Figure 41 shows simple stall warning system.

Stall Warning System

Figure 41

M

28V DCSUPPLY

>17.5º

STICKSHAKER

ANGLEOF

ATTACK



ENGINEERS


1.19 OTHER SYSTEM INDICATIONS There are endless different instrument displays, which show the pilot's or flight engineer, the condition of the aircraft's many systems, the range of instruments depending on the size of the aircraft. On earlier airliners there could have been dozens of instruments on the panels to pass on information regarding, for example, oil temperature & pressure, cabin altitude, hydraulic oil quantity, electrical power being used, etc. 1.20 POWERPLANT INSTRUMENTATION Information required by the flight crew to enable them to monitor the engines include:

1. Fuel Contents. 2. Fuel Flow. 3. Engine RPM. 4. Engine Temperature. 5. Engine pressure.

1.21 FUEL CONTENTS GAUGE Most modern aircraft have a number of fuel tanks within the wing structure and each individual tank's contents must be known. There are two main methods of indicating fuel contents:

Resistance Gauges. Capacitance Quantity Indicators.

1.21.1 Resistance Gauges This type of gauge tends to found on smaller aircraft. It has a float in the fuel tank that is connected to a variable resistor. As the fuel level changes, the float will move, thus changing the resistance, which in turn will alter the current flow through a DC circuit, which in turn will operate a meter indicating fuel contents.



ENGINEERS


Figure 53 shows a simplified resistance gauge.

Resistance Gauge Figure 53

1.21.2 Capacitance Quantity Indicators This has the advantage over other quantity systems in that it can give accurate readings in very large or unusually shaped tanks. The probes within the fuel tank are actually capacitors. The two plates of the capacitor will be separated by fuel on the lower end and air on the upper end. Since fuel and air have different dielectric constant values, the amount of capacitance will change as the fuel level rises and falls. The probes will then send signals to the flight deck gauges to indicate fuel contents. This system usually includes a totalizer, which will give a reading of the total fuel on board. Some fuel systems will also include indications of fuel used since take-off.

N

S TANKRESISTOR

+ DCPOWER

INDICATOR

FUEL TANK



ENGINEERS


Figure 54 shows a circuit of a capacitance quantity system.

Capacitance Quantity Indicating System Figure 54

TANK UNIT

FULL

EMPTY

AMPLIFIERSTAGE

REF C

2 - PHASEMOTOR

AMPLIFIER UNITREF

PHASE

INDICATOR

LOOPA

LOOPB

IS

IB

DISCRIMINATIONSTAGE



ENGINEERS


1.22 FUEL FLOW INDICATOR As the name suggests, these indicators show the amount of fuel flowing into the engines. Fuel flow information can be represented as either LBS/HR, Gallons/HR or PSI. Some indicators will show both PSI and either LBS/HR or Gallons/HR. Figure 55 shows a fuel flow indicator.

Fuel Flow Indicator Figure 55

FUELFLOWLBS/HR

195 PSI

2.5 PSI50

100

150 170

R

4555

6575

80

95T.O.

LBS/HRSCALE

PSI SCALE

LEFT ENGINEFUEL FLOW

RIGHT ENGINEFUEL FLOW

L



ENGINEERS


1.23 FUEL PRESSURE INDICATOR Some engines have a fuel pressure gauge that displays the pressure of the fuel supplied to the fuel control unit. Most display the pressure in pounds per square inch (psi) and provide indications to the pilot that the engine is receiving the fuel required for a given power setting. Figure 56 shows a fuel pressure gauge.

Fuel Pressure Gauge Figure 56

There are two types of pressure gauge:

Bourbon Tube type. Pressure Capsule type.

FUELPRESS

125 PSI

10 PSI30

50

80 100

POINTER

PSISCALE



ENGINEERS


1.23.1 Bourbon Tube Fuel Pressure Indicator Is made with a metal tube that is formed in a circular shape with a flattened cross-section. One end is open while the other is sealed. The open end of the bourbon tube is connected to a capillary tube containing pressurized fuel. As the pressurized fuel enters the bourbon tube, the tube tends to straighten. Through a series of gears, this movement is used to move the indicating pointer on the instrument face. Figure 57 shows a Bourbon type fuel pressure gauge and its operation.

Bourbon Tube Fuel Pressure Gauge Figure 57

ANCHORPOINT

POINTERSTAFF

BOURBONTUBE

GEARING



ENGINEERS


1.23.2 Capsule Fuel Pressure Indicator This type of indicator utilizes a “pressure capsule” or “diaphragm”. Like the bourbon tube, a diaphragm type pressure indictor is attached to a capillary tube, which attaches to the fuel system and carries pressurized fuel to the diaphragm. As the diaphragm becomes pressurized it expands, causing an indicator pointer to rotate. Figure 58 shows a pressure capsule type fuel pressure indicator.

Pressure Capsule Fuel Pressure Gauge Figure 58

DIAPHRAGM



ENGINEERS


1.24 ENGINE RPM INDICATORS These instruments indicate the rotational speed of the engine. Low Pressure Compressor (N1), Intermediate Pressure Compressor (N2) and High Pressure Compressor (N3). Figure 59 a RPM gauge for N1 measurement.

N1 RPM Gauge Figure 59



ENGINEERS


The indicator use electromagnetic sensors (which contains a coil of wire that generates a magnetic field) to measure the RPM of the respective compressor blades. The sensor is mounted in the shroud around the fan so, when each fan blade passes the sensor, the magnetic field is interrupted. The frequency at which the fan blades cut across the field is measured by an electronic circuit and then transmitted to a RPM gauge in the cockpit. Figure 60 shows the operation of a N1 & N2 gauges.

N1 & N2 Pressure Gauges Operation Figure 60



ENGINEERS


1.25 ENGINE TEMPERATURE GAUGES Because turbine engines can be severely damaged by high temperature in the turbine sections, a means of measuring the temperature is required. Because of the high temperatures involved, this is carried out using thermocouples. There are a number of different terms and abbreviations used for the gas temperature in turbine engines, these are:

Turbine Inlet Temperature (TIT). Inter Turbine Temperature (ITT). Turbine Outlet Temperature (TOT). Engine Gas Temperature (EGT). Measured Gas Temperature (MGT). Jet Pipe Temperature (JPT).

Figure 61 shows a typical EGT indicator

EGT Indicator Figure 61

7 6 5

EGT°C X 100

1

35

7

9

OVER-TEMPERATUREWARNING LIGHT

TEMERATURESCALE

POINTER

OVER-TEMPLIMIT POINTER

DIGITALREAD-OUT



ENGINEERS


Each type of EGT system consists of several thermocouples spaced at intervals around the circumference of the engine exhaust section casing. The EGT indicator in the cockpit displays the average temperature measured by the individual thermocouples. Figure 62 shows EGT indicator operation.

EGT Indicator Operation Figure 62



ENGINEERS


1.26 ENGINE PRESSURE INDICATORS Engine pressure indicators provide indications of the thrust being produced by a turbojet or turbofan engine. Figure 63 shows an EPR indicator.

EPR Indication System

Figure 63



ENGINEERS


The EPR is the ratio of turbine discharge pressure to compressor inlet pressure. Pressure measurements are recorded by total pressure pickups, or EPR probes, installed in the engine inlet Pt2 section and at the exhaust Pt7 section. Once collected, the data is sent to a differential pressure transducer, which drives a cockpit EPR gauge. Figure 64 shows the operation of an EPR indicator.

EPR Indicator Operation Figure 64



ENGINEERS


Figure 65 shows the engine instrument grouping for a twin engine aircraft.

Power plant instrument grouping Figure 65

EPR

%RPM

EGT

FF

020 40 60

80

1009 2

% RPMN1

1.0

EPR

0.81 5 01.0

1.4

1.6

1.2

7 6 5

EGT°C X 100

1

35

7

9

FF

X 10001

23 4 5

6

86 5 8

1 5 0

EPR

%RPM

EGT

FF

0204060

80

1009 2

% RPMN1

1.0

EPR

0.81 5 0

1.0

1.4

1.6

1.2

7 6 5

EGT°C X 100

1

35

7

9

FF

X 10001

2345

6

8658

1 5 0

A319/A320/A321TECHNICAL TRAINING MANUAL

GENERAL FAMILIARIZATION COURSE

22 AUTO FLIGHT SYSTEM

This document must be used for training purpose only

Under no circumstances should this document be used as a reference.

It will not be updated.

All rights reserved.No part of this manual may be reproduced in any form,

by photostat, microfilm, retrieval system, or any other means,without the prior written permission of Airbus Industrie.

AUTO FLIGHT SYSTEM GENERALSystem Design Philosophy (1) 1.....................** System Presentation (1) 5.......................** System Control and Indicating (1) 9............Basic Operational Principles (1) 23................

_A319/A320/A321 TECHNICAL TRAINING MANUAL22 AUTO FLIGHT SYSTEMGENERAL FAMILIARIZATION COURSE

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22 - AUTO FLIGHT SYSTEM

22-00-00 AUTO FLIGHT SYSTEM DESIGNPHILOSOPHY

CONTENTS:General ConceptNavigationFlight PlanOperationAFS/Fly by WireSystem Design

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AUTO FLIGHT SYSTEM DESIGN PHILOSOPHY

GENERAL CONCEPT

The Auto Flight System (AFS) calculates orders toautomatically control the flight controls and theengines.The Auto Flight System computes orders and sends themto the Electrical Flight Control System (EFCS) and tothe Full Authority Digital Engine Control (FADEC) tocontrol flying surfaces and engines.When the AFS is not active, the above mentionedcomponents are controled by the same systems but ordersare generated by specific devices (i.e. side sticksand thrust levers).

NAVIGATION

A fundamental function of the Auto Flight System isto calculate the position of the aircraft.When computing the aircraft position, the system usesseveral aircraft sensors giving useful information forthis purpose.

FLIGHT PLAN

The system has several flight plans in its memory.These are predetermined by the airline.A flight plan describes a complete flight fromdeparture to arrival, it includes vertical informationand all intermediate waypoints.It can be displayed on the instruments (CRTs).

OPERATION

There are several ways to use the Auto Flight System.The normal and recommended way to use the Auto FlightSystem is to use it to follow the flight planautomatically.Knowing the position of the aircraft and the desiredflight plan (chosen by the pilot), the system is ableto compute the orders sent to the surfaces and enginesso that the aircraft follows the flight plan.The pilot has an important monitoring role.

NOTE: During Auto Flight System operation, side sticksand thrust levers do not move automatically.

AFS/FLY BY WIRE

If the pilot moves the side stick when the Auto FlightSystem is active, it disengages the autopilot.Back to manual flight, when the side stick is released,the Electrical Flight Control System maintains theactual aircraft attitude.

SYSTEM DESIGN

To meet the necessary reliability, the AutoFlightSystem is built around four computers:Two interchangeable Flight Management and GuidanceComputers (FMGCs) and two interchangeable FlightAugmentation Computers (FACs).It is a FAIL OPERATIVE system.Each Flight Management and Guidance Computer and eachFlight Augmentation Computer has a command part and amonitor part to be FAIL PASSIVE.T

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22-00-00 AUTO FLIGHT SYSTEM PRESENTATION

CONTENTS:GeneralControlsFMGCsFACsOther SystemsSelf Examination

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AUTO FLIGHT SYSTEM PRESENTATION

GENERAL

The Auto Flight System (AFS) provides the pilots withfunctions reducing their workload and improving thesafety and the regularity of the flight.The Auto Flight System is designed around:- 2 Flight Management and Guidance Computers (FMGCs),- 2 Flight Augmentation Computers (FACs),- 2 Multipurpose Control and Display Units (MCDUs),- 1 Flight Control Unit (FCU).

CONTROLS

The FCU and the MCDUs enable the pilots to control thefunctions of the FMGCs.The FAC engagement pushbuttons and the rudder trimcontrol panel are connected to the FACs.The MCDUs are used for long-term control of theaircraft and provide the interface between the crewand the FMGC allowing the management of the flight.The FCU is used for short term control of the aircraftand provides the interface required for transmissionof engine data from the FMGC to the Full AuthorityDigital Engine Control (FADEC).

FMGCs

There are two interchangeable FMGCs.Each FMGC is made of two parts: the Flight Managementpart called FM part and the Flight Guidance part calledFG part.The Flight Management part provides functions relatedto flight plan definition, revision and monitoring.The Flight Guidance part provides functions relatedto the aircraft control.

FACs

The basic functions of the FACs are the rudder controland the flight envelope protection.

NOTE: The FAC includes an interface between the AutoFlight System and the Centralized Fault DisplaySystem (CFDS) called Fault Isolation andDetection System (FIDS).This function is activated only in position 1(FAC 1).

OTHER SYSTEMS

The Auto Flight System is connected to the majorityof the aircraft systems.Examples of Auto Flight System data exchanges:- Reception of the aircraft altitude and attitude fromthe Air Data and Inertial Reference System (ADIRS).- Transmission of autopilot orders to the Elevator andAileron Computers (ELACs).

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SELF EXAMINATION

What are the basic functions of the FACs?A - Management functions and flight envelope

protection.B - Rudder control and flight envelope

protection.C - Guidance functions and rudder control.

Where are the FMGC functions controlled from?A - The MCDUs and rudder trim control panel.B - The FCU and rudder trim control panel.C - The FCU and MCDUs.

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22-00-00 AUTO FLIGHT SYSTEM CONTROL ANDINDICATING

CONTENTS:FCUMCDUsNDsPFDsThrust LeversSide SticksRudder PedalsResetsRMPsEWD/SDAttention Getters

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AUTO FLIGHT SYSTEM CONTROL AND INDICATING

FCU

The Flight Control Unit (FCU) is installed on theglareshield. The FCU front face includes an Auto FlightSystem (AFS) control panel between two ElectronicFlight Instrument System (EFIS) control panels.The AFS control panel allows and displays theengagement of autopilots (APs) and autothrust (A/THR),and the selection of guidance modes and flightparameters.

NOTE: The EXPEDite pushbutton can be optionallyremoved from the AFS control panel.

The two EFIS control panels control and display, foreach EFIS side (Capt and F/O), the Primary FlightDisplay and Navigation Display functions (respectivelybaro and Flight Director (FD) conditions, andNavigation Display modes).

MCDUs

Two Multipurpose Control and Display Units (MCDUs) arelocated on the center pedestal.The MCDU is the primary entry/display interface betweenthe pilot and the FM part of the FMGC.MCDU allows system control parameters and flight plansto be inserted, and is used for subsequentmodifications and revisions.The MCDU displays information regarding flightprogress and aircraft performances for monitoring andreview by the flight crew.

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NDs

The two Navigation Displays (NDs) are located on themain instrument panel.The Navigation Display is built from:

- flight plan data,- data selected via the FCU,- aircraft present position,- wind speed/direction,- ground speed/track.

PFDs

The two Primary Flight Displays (PFDs) are located onthe main instrument panel.The Flight Mode Annunciator (FMA) is the top part ofthe Primary Flight Display (PFD).Each PFD displays:

- AP/FD/A/THR engagement status on the FMA,- AP/FD and A/THR armed/engaged modes on the FMA,- FD orders,- FAC characteristic speeds on the speedscale.




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THRUST LEVERS

The thrust levers are located on the center pedestal.The thrust levers allow the Take-Off/Go-Around (TO/GA)modes and the autothrust to be engaged.Two autothrust instinctive disconnect pushbuttonslocated on the thrust levers allow the autothrustfunction to be disengaged.

SIDE STICKS

The Capt and F/O side sticks are respectively locatedon the Capt lateral panel and F/O lateral panel.The autopilot is disengaged when the take over prioritypushbutton on the side stick is pressed or when a forceabove a certain threshold is applied on the side stick.

RUDDER PEDALS

The rudder pedals are fitted in the Capt and F/Opositions.Rudder pedals override disconnects the autopilot.




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RESETS

The FMGC, FAC, FCU and MCDU resets are possible in thecockpit.Depending on the computer (1 or 2), the circuitbreakers are located either on the overhead circuitbreakers panel 49VU or on the rear circuit breakerspanel 121VU.




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RMPs

The Radio Management Panels (RMPs) are located on thecenter pedestal near Multipurpose Control and DisplayUnits 1 and 2.The RMPS are used for navaid standby selection.

EWD/SD

The Engine/Warning Display (EWD) and the System Display(SD) are located on the main instrument panel.The EWD displays AFS warning messages. The SD displays AFS information such as inoperativesystems on the STATUS page or landing capabilitiesavailability.




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ATTENTION GETTERS

The attention getters are located on the glareshieldpanel on the Capt and F/O sides.The MASTER CAUTION and/or the MASTER WARNING areactivated when an AFS disconnection occurs.The AUTOLAND warning is activated when a problem occursduring final approach in automatic landing.




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22-00-00 BASIC OPERATIONAL PRINCIPLES

CONTENTS:GeneralData Base LoadingPower-up Test FD EngagementMCDU InitializationA/THR EngagementAP EngagementSelf Examination

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BASIC OPERATIONAL PRINCIPLES

GENERAL

This sequence describes the operational use of theFlight Management and Guidance Computers (FMGCs) in anormal operation with a total availability of theconcerned functions.The short-term pilot orders are entered through theFlight Control Unit (FCU). The long-term pilot ordersare entered through the Multipurpose Control andDisplay Unit (MCDU).Four key-words for the control principle and both typesof guidance are to be kept in mind in order to avoidhandling errors.Aircraft control is AUTOMATIC (Autopilot orautothrust), or MANUAL (Pilot action on side sticksor on thrust levers). Aircraft guidance is MANAGED(Targets are provided by the FMGC), or SELECTED(Guidance targets are selected by the pilot throughthe FCU).

DATA BASE LOADING

The data base must be loaded and updated to keep thesystem operational.

NOTE: Only the navigation data base is periodicallyupdated.

POWER-UP TEST FD ENGAGEMENT

As soon as electrical power is available, the FlightDirector (FD) is automatically engaged provided thatthe power-up test is successful.No guidance symbols are displayed as long as no AP/FDmode is active.

MCDU INITIALIZATION

First, MCDU STATUS page is displayed. Then, the pilotuses the MCDU for flight preparation, which includes:

- choice of the data base,- flight plan initialization,- radio nav entries and checks,- performance data entry (V1, VR, V2 and FLEXTEMP).

V2, at least, must be inserted in the MCDU beforetake-off.Entry of the flight plan (lateral and vertical) andV2 into the MCDU is taken into account by the FlightManagement (FM) part and confirmed by the lighting ofthe associated lights on the FCU.

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A/THR ENGAGEMENT

Autothrust (A/THR) engagement occurs when the pilotmoves the thrust levers to the TO/GA or FLX/MCT gate.Then:. the FMGC automatically engages:

- the take-off modes for yaw and longitudinalguidance (RunWaY (RWY) and Speed ReferenceSystem (SRS)),

- the autothrust function (but it is not active).. the FD symbols appear on the PFD (Green FD yaw barand pitch bar).For take-off, the thrust levers are set to the TO/GAgate or the FLEX/MCT gate if a flexible temperaturehas been entered on the MCDU.At the thrust reduction altitude, the FM part warnsthe pilot to set the thrust levers to CLB gate.

NOTE: The thrust levers normally will not leave thisposition until an audio message "RETARD"requests to the pilot to set the thrust leversto IDLE gate before touchdown.

AP ENGAGEMENT

Either autopilot (AP) can only be engaged 5 secondsafter lift off. Only one autopilot can be engaged ata time, the last in, being the last engaged.After the normal climb, cruise and descent phases,selection of LAND mode (Autoland) allows both APs tobe engaged together.After touchdown, during ROLL OUT mode, APs remainengaged to control the aircraft on the runwaycenterline.Then the pilot disengages the APs at low speed, taxiesand stops the aircraft.

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SELF EXAMINATION

When is FD engaged?A - As soon as at least one AP is engaged.B - As soon as A/THR is engaged.C - At the end of a successful power-up test.

Concerning AP engagement, which of the following istrue?

A - Both APs can be engaged whatever theflight phase.

B - During the approach phase, it isrecommended to engage the second AP.

C - Both APs can never be engaged at thesame time (Last in, last engaged).

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23 COMMUNICATIONS

GENERAL** General System Presentation (1) 1...............** System Control and Indicating (1) 5............

SPEECH COMMUNICATIONSpeech Communication Presentation (1) 13..........Radio Management Panel Presentation (1) 17........Audio System Presentation (1) 23...................** Audio Control Panel Presentation (1) 27........VHF System Presentation (1) 33.....................SELCAL System Presentation (1) 39..................** Ground Crew Call SYS Pres./Operation (1) 43....Static Discharging (1) 47..........................

SATCOMMCS SATCOM Presentation (1) 55.....................

COCKPIT VOICE RECORDER SYSTEM** CVR System Presentation (1) 59..................

CABIN INTERCOMMUNICATION DATA SYSTEM ( CIDS)** CIDS Design Philosophy (1) 65...................** Forward Attendant Panel Presentation (1) 69....** AFT Attendant Panel Presentation (1) 73........** PTP Presentation (1) 77.........................

PAX ENTERTAINMENT SYSTEM** Passenger Entertainment SYS Pres. (1) 81.......

_A319/A320/A321 TECHNICAL TRAINING MANUAL23 COMMUNICATIONSGENERAL FAMILIARIZATION COURSE

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23-00-00 GENERAL SYSTEM PRESENTATION

CONTENTS:VHFHF (Option)SELCAL (SELective CALling)CIDSPassenger AddressInterphoneCockpit Voice Recorder

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GENERAL SYSTEM PRESENTATION

VHF

The Very High Frequency (VHF) system serves for shortrange voice communications.

HF (Option)

The High Frequency (HF) system serves for alllong-distance voice communications between differentaircraft (in flight or on the ground), or between theaircraft and one or several ground stations.

SELCAL (SELective CALling)

The purpose of the SELCAL system is to give visual andaural indications to the crew, concerning callsreceived from ground stations through VHF and HFsystems.

CIDS

The Cabin Intercommunication Data System (CIDS) isdesigned to interface flight crew, cabin attendants,passengers, ground service and various cabin systemsdedicated to cabin attendant or passenger use.The CIDS is used to control, test and monitor variouscabin systems dedicated to cabin attendant or passengeruse.

PASSENGER ADDRESS

The Passenger Address (PA) allows voice announcementsto be broadcast to all passengers, from the cockpitand cabin attendant stations through the CIDS.

INTERPHONE

There are 3 interphone systems on the aircraft:the flight interphone, the cabin interphone and theservice interphone.

- The flight interphone system allowscommunication between the flight crew members,and between the flight crew and the groundmechanic at the external power receptacle orin the avionics bay.

- The cabin interphone system allowscommunication between the cockpit and the cabinattendant stations, and between the cabinattendant stations.

- The service interphone system enablescommunication between the different serviceinterphone jacks, the cockpit and the cabinattendant stations.

COCKPIT VOICE RECORDER

The Cockpit Voice Recorder (CVR) records in-flight andon-ground crew conversations and radio communications.

TMUCOG101-T01 LEVEL 1



UFD0100


GENERAL SYSTEM PRESENTATIONTMUCOG101-P01 LEVEL 1



UFD0100



TMUCOG101 LEVEL 1



UFD0100


23 - COMMUNICATIONS

23-00-00 SYSTEM CONTROL AND INDICATING

CONTENTS:CockpitCabinAvionics BayNose Landing Gear

TMUCOG202 LEVEL 1



UFD0100


SYSTEM CONTROL AND INDICATING

COCKPIT

In the cockpit, we find:- 3 Audio Control Panels (ACPs) for the selectionof communication systems (in transmission andreception) and for the control of the receivedaudio signal levels,

- 3 Radio Management Panels (RMPs) for theselection of radio communication and navigationfrequencies,

- 1 AUDIO SWITCHING selector for thereconfiguration of channels, in case of ACPfailure,

- 1 CALLS panel for flight crew-to-groundmechanic or flight crew-to-cabin attendantcalls,

- and various items of acoustic equipment.The acoustic equipment comprises:2 loudspeakers with volume control (1),2 radio PTT switches (on the side sticks),2 hand microphones (2),headsets (3),boomsets (3),oxygen mask microphones (3).Facilities are provided in the cockpit forheadsets and boomsets.




UFD0100


SYSTEM CONTROL AND INDICATING - COCKPITTMUCOG202-P01 LEVEL 1



UFD0100



CABIN

Panels are installed in the cabin, for the control andmonitoring of the various cabin systems:

- The Forward Attendant Panel (FAP) is locatedin the forward entrance area of the aircraft.The cabin attendants can control the differentcabin systems from here.

- 3 Additional Attendant Panels (AAPs) can beinstalled near the doors and are dedicated tocabin zones.1 AAP is basically installed near the aftpassenger crew door.

Note: The F.A.P. and the A.A.P. are customized perairline request.

- The Programming and Test Panel (PTP) islocated at the forward attendant station,behind a hinged access door next to the FAP.The PTP enables to test and re-program theCabin Intercommunication Data System.




UFD0100


TMUCOG202-P02 LEVEL 1

SYSTEM CONTROL AND INDICATING - CABIN



UFD0100



AVIONICS BAY

In the avionics bay, the SELCAL code panel is installedfor coding the SELCAL code assigned to the aircraft.

NOSE LANDING GEAR

On the external power control panel, some features arededicated to ground mechanic-to-flight crew or flightcrew-to-ground mechanic calls.




UFD0100


SYSTEM CONTROL AND INDICATING - AVIONICS BAY & NOSE LANDING GEARTMUCOG202-P03 LEVEL 1



UFD0100



TMUCOG202 LEVEL 1



UFD0100


23 - COMMUNICATIONS

23-51-00 SPEECH COMMUNICATION PRESENTATION

CONTENTS:COM/NAV SystemsRMPACPAMUSELCALStatic DischargingSelf Examination

TMUCOMA02 LEVEL 1

EFFECTIVITY 23-51-00 Page 1May 31/00


UFD0100


SPEECH COMMUNICATION PRESENTATION

SYSTEMS

Communication and navigation systems are connected tothe AMU for analog inputs and to the RMP for frequencyselection.

RADIO MANAGEMENT PANEL

The radio management panels (RMP) centralize radiocommunication frequency control.RMP 1 and RMP 2 can also serve as backups for theflight management and guidance computers (FMGC) forradio navigation frequency control (VOR, DME, ILS,ADF).The aircraft is equipped with three RMPs which areidentical and interchangeable.The 3rd RMP is optional.

AUDIO CONTROL PANEL

The ACPs supplies the means:- to use the various radio communication andradio navigation facilities installed on theaircraft for transmission and reception of theaudio signals.

- to display the various calls (SELCAL, groundcrew call and calls from the Cabin Attendants).

The ACPs serve only for control and indication.

AUDIO MANAGEMENT UNIT

The audio management unit (AMU) ensures the interfacebetween the user (jack panel and ACP) and the variousradio communication and radio navigation systems.The Audio Management Unit is equipped with a TESTcircuit (BITE) which enables connection to the CFDIU.The AMU ensures the following functions:

- Transmission- Reception- SELCAL and display of ground crew and CabinAttendant calls

- Flight interphone- Emergency function for the Captain and FirstOfficer stations.

SELCAL

The selective calling system provides visual and auralindication of calls received from ground stations.

STATIC DISCHARGING

The purpose of the static discharges is to dischargestatic electricity and to prevent interference ofcommunication systems.

TMUCOMA02-T01 LEVEL 1



UFD0100


SPEECH COMMUNICATION PRESENTATIONTMUCOMA02-P01 LEVEL 1



UFD0100


SELF EXAMINATION

What is the purpose of the RMPs?A - To enable the received audio signals to

be selected.B - To enable the received audio signals and

the frequencies to be selected.C - To enable the frequencies of all the

radio communication systems to beselected.

What is the purpose of the AMU?A - To centralize all the audio signals and

the frequencies of the communicationsystems.

B - To act as an interface between the usersand the various radio communication andradio navigation systems.

C - To receive audio signals only.

TMUCOMA02 LEVEL 1



UFD0100


23 - COMMUNICATIONS

23-13-00 RADIO MANAGEMENT PANELPRESENTATION

CONTENTS:Radio Management Panel (RMP) DescriptionWindowsTransfer P/BCommunication KeysSEL IndicatorDual Selector KnobNavigation KeysON/OFF SwitchSelf Examination

TMUCOMB03 LEVEL 1

EFFECTIVITY 23-13-00 Page 1Jun 30/00


UFD0100


RADIO MANAGEMENT PANEL PRESENTATION

RADIO MANAGEMENT PANEL (RMP) DESCRIPTION

The RMPs are used for the selection of radiocommunication frequencies. They are also used for thestandby selection of radio navigation frequencies inback-up mode.3 RMPs are used for frequency selection, each one cancontrol any VHF (HF) frequency.Note: the third RMP is optional.The 3 RMPs permanently dialog so that each RMP isinformed of the last selection made on any of the otherRMPs. If two RMPs fail, the remaining RMP controls allthe VHF transceivers.The transmission of data to the communication andnavigation systems and the dialog between the RMPs areperformed through ARINC 429 buses.

WINDOWS

There are 2 display windows:- The active window displays the operationalfrequency

- The standby/course window displays the standbyfrequency or the course in back-up navigationmode.

The windows are liquid crystal displays with a highcontrast.

TRANSFER P/B

When the transfer key is pressed, the standby frequencybecomes the operational frequency, and the operationalfrequency becomes the standby frequency.

COMMUNICATION KEYS

There are 6 pushbutton keys for the radio communicationsystems. 3 of them are used for VHF, the 3 others forHF. The AM key controls the selection of the AM mode forHF transceivers provided that HF1 or HF2 is selected.When a key is pressed, the relevant active and thestandby frequencies are automatically displayed in thededicated windows.

SEL INDICATOR

Although one RMP can control frequencies of anytransceiver, each RMP has dedicated systems.The normal configuration is:

- RMP1 allocated with VHF1, VHF3 and HF1, ifinstalled,

- RMP2 allocated with VHF2 and HF2, if installed.If the optional RMP3 is installed, it will be allocatedwith VHF3 and HF systems which are no longer dedicatedto RMP1 or RMP2.The SEL indicator light will come on white on the RMPsinvolved, when an RMP takes control of a non dedicatedsystem frequency selection.For example, if VHF2 is selected on RMP1, the SELindicator lights come on on RMP1 and RMP2.

DUAL SELECTOR KNOB

The dual selector knob is used for the selection ofthe frequency/course displayed in the standby/coursewindow.

TMUCOMB03-T01 LEVEL 1



UFD0100


RADIO MANAGEMENT PANEL PRESENTATIONTMUCOMB03-P01 LEVEL 1



UFD0100


RADIO MANAGEMENT PANEL PRESENTATION

NAVIGATION KEYS

The NAVigation guarded pushbutton key allows the radionavigation systems to be selected, in back-up modeonly, when the Flight Management and Guidance Computers(FMGCs) have failed. In radio navigation back-up mode,only RMP1 and RMP2 can perform navigationfrequency/course selection using the dual selectorknob.

ON/OFF SWITCH

The latching ON/OFF switch allows the crew to set theRMP on or off.

TMUCOMB03-T01 LEVEL 1



UFD0100


SELF EXAMINATION

What happens if RMP2 fails?A - The communication systems are inoperativeB - VHF2 frequencies cannot be controlled.C - All communication frequencies can be

controlled.

TMUCOMB03 LEVEL 1



UFD0100



TMUCOMB03 LEVEL 1



UFD0100


23 - COMMUNICATIONS

23-51-00 AUDIO SYSTEM PRESENTATION

CONTENTS:GeneralTransmissionReceptionFlight InterphoneSelective Calling (SELCAL)CallsSelf Examination

TMUCOMH02 LEVEL 1



UFD0100


AUDIO SYSTEM PRESENTATION

GENERAL

The AMU centralizes the Audio Signals used by the crew.The crew controls and operates these functionsindependently with the Audio Control Panels.The audio management system provides:

- radio communication and navigation for crewutilization

- flight interphone system- selective calling system (SELCAL)- visual indication of ground crew and cabinattendant calls.

Each cockpit occupant Audio Equipment includes:- oxygen mask,- headset,- boomset,- handmicrophone,

except for the 4th occupant which is only equippedwith a jack box.

TRANSMISSION

In transmission mode, the AMU collects microphoneinputs of the various crew stations and directs themto the communication transceivers.

RECEPTION

In reception mode, the AMU collects the audio outputsof the communication transceivers and navigationreceivers and directs them to the various crewstations.

INTERPHONE

The flight interphone function allows interpone linksbetween the various crew stations in the cockpit andwith the groud crew through the jack at the externalpower receptacle panel (108 VU) and the avionicscompartment jack panel (63 VU).

SELCAL

The Selective Calling system enables reception withaural and visual indication of calls from groundstations equipped with a coding device

NOTE: The SELCAL decoding unit is located inside theAMU.

CALLS

Cabin attendant and mechanic calls are indicated onthe Audio Control Panels.

TMUCOMH02-T01 LEVEL 1



UFD0100


AUDIO SYSTEM PRESENTATIONTMUCOMH02-P01 LEVEL 1



UFD0100


SELF EXAMINATION

What is the function of the AMU?A - It monitors the radio frequency

selection.B - It integrates all the crew communication

functions.C - It monitors the NAV frequency selection.

TMUCOMH02 LEVEL 1



UFD0100


23 - COMMUNICATIONS

23-51-00 AUDIO CONTROL PANEL PRESENTATION

CONTENTS:GeneralTransmission KeysReception KnobInterphone/Radio Selector SwitchVoice FilterResetPassenger AddressSelf Examination

TMUCOMI01 LEVEL 1

EFFECTIVITY 23-51-00 Page 1Mar 31/00


UFD0100


AUDIO CONTROL PANEL PRESENTATION

GENERAL

3 Audio Control Panels (ACPs) are provided in thecockpit for the Captain, the First Officer and thethird occupant.Each ACP allows:

- the use of various radio communication andradio navigation facilities installed in theaircraft for transmission and reception of theaudio signals,

- the display of various calls received throughthe SELCAL system, from ground mechanics andfrom cabin attendants,

- the use of flight, cabin and service interphonesystems.

TRANSMISSION KEYS

Eight rectangular electronic keys are used for theselection of the transmission channel and for thedisplay of various calls received through SELCALsystem, from ground mechanics and from cabinattendants.MECH light on the INTerphone key flashes amber toindicate a ground mechanic call.ATT light on the CABin key flashes amber to indicatea cabin attendant call.

NOTE: Only one transmission channel can be selectedat a time.

RECEPTION KNOB

Fifteen pushbutton knobs are used to select receptionand to adjust the volume of received signals.When the reception channel is selected, the pushbuttonknob pops out and comes on white.

INTERPHONE/RADIO SELECTOR SWITCH

The INTerphone/RADio selector switch permits theutilization of the interphone or the radio, when theboomsets or oxygen masks are used by the crew. The INT position allows direct flight interphonetransmission:

- whatever the transmission key selected andprovided no Push-To-Talk switch is activated,

- when no transmission key is selected.The neutral position allows reception only.The RAD position is used as a Push-To-Talk switch whena transmission key is selected.

VOICE FILTER

A voice filter can be used on the ADF and VOR channels.When used, the identification signals transmitted bythe navaids are greatly attenuated (32 dB) so as tohear only voice signals.ON comes on green when the voice filter is in service(ON VOICE key pressed in).

TMUCOMI01-T01 LEVEL 1



UFD0100


AUDIO CONTROL PANEL PRESENTATIONTMUCOMI01-P01 LEVEL 1



UFD0100


AUDIO CONTROL PANEL PRESENTATION

RESET

The RESET key is used to cancel all the lighted calls.

NOTE: MECH and ATT lights go off automatically after60 seconds if the call is not cancelled by theRESET key.

PASSENGER ADDRESS

A key enables the selection of the Passenger Addresstransmission.This key must be pressed in during the wholetransmission.An AMU pin program can inhibit the unstable operationof the PA key.

TMUCOMI01-T01 LEVEL 1



UFD0100


SELF EXAMINATION

What happens in case of a SELCAL call on VHF2?A - CALL light flashes amber on the VHF2 key.B - The three green bars on the VHF2 key

come on.C - CALL light comes on white on the VHF2

key.

On the ACP, is it possible to transmitsimultaneously on Passenger Address and VHFchannels?

A - Yes.B - No.

What is the function of the RESET key?A - The RESET key is used to restart the

system.B - The RESET key is used to cancel the

previous selections.C - The RESET key is used to cancel all the

lighted calls.

TMUCOMI01 LEVEL 1



UFD0100



TMUCOMI01 LEVEL 1



UFD0100


23 - COMMUNICATIONS

23-12-00 VHF SYSTEM PRESENTATION

CONTENTS:PurposePrincipleComponents

TMUCOMF02 LEVEL 1



UFD0100


VHF SYSTEM PRESENTATION

PURPOSE

The VHF system allows short distance voicecommunications between different aircraft (in flightor on ground) or between the aircraft and a groundstation.The VHF is used for short range voice communications.

PRINCIPLE

For voice communications, the crew uses acousticequipment.

- side-stick radio selectors,- loudspeakers,- oxygen-masks,- boomsets, - headsets,- hand-microphones.

The Audio Management Unit (AMU) acts as an interfacebetween the crew and the VHF system.The Audio Control Panels (ACPs) allow selection of theVHF1,VHF2, or VHF3 transceiver in transmission orreception mode and for the control of the receivedaudio signal.The Radio Management Panels (RMPs) serve to select theVHF frequencies.The VHF transceiver, tuned on the frequency selectedby one of the 3 Radio Management Panels (RMPs),transforms the audio signals into VHF signals (intransmission mode) or VHF signals into audio signals(in reception mode).

TMUCOMF02-T01 LEVEL 1



UFD0100


VHF SYSTEM PRESENTATIONTMUCOMF02-P01 LEVEL 1



UFD0100


VHF SYSTEM PRESENTATION

COMPONENTS

Let’s see the main components of the VHF system.The VHF system comprises:

- 3 VHF transceivers (1),- 3 blade antennae,

associated with control systems:- 3 RMPs (2),- 3 ACPs (2),- 1 AMU (1).

NOTE : RMP 3 is optional.

TMUCOMF02-T02 LEVEL 1



UFD0100


VHF SYSTEM PRESENTATION - COMPONENTSTMUCOMF02-P02 LEVEL 1



UFD0100



TMUCOMF02 LEVEL 1



UFD0100


23 - COMMUNICATIONS

23-51-00 SELCAL SYSTEM PRESENTATION

CONTENTS:SELCAL PhilosophySELCAL OperationSelf Examination

TMUCOML01 LEVEL 1

EFFECTIVITY 23-51-00 Page 1Apr 30/96


UFD0100


SELCAL SYSTEM PRESENTATION

SELCAL PHILOSOPHY

The selective calling system provides visual and auralindication of calls received from ground stationsequipped with a coding device.The ground station tone generator provides the assignedaircraft code which modulates a VHF (or an HF)transmitter.In order to receive the SELCAL CALL, the same frequencyas on the ground must be activated in the aircraft.SELCAL: SELective CALling systemThis function is integrated in the AMU. The A/C codecan be set on the SELCAL code panel fitted in theavionics bay.

SELCAL OPERATION

When a selcal call is received, the CALL light flashesamber on the corresponding transmission key and abuzzer sound is heard.The buzzer signal is generated by the Flight WarningComputer (FWC).CALL flashes amber on all the ACPs when a selcal callis received.The CALL indication can be manually cleared by pressingthe RESET key on any ACP or it can be automaticallycleared upon transmission on the called channel.

TMUCOML01-T01 LEVEL 1



UFD0100


SELCAL SYSTEM PRESENTATIONTMUCOML01-P01 LEVEL 1



UFD0100


SELF EXAMINATION

How is the SELCAL CALL light reset?A - By pressing the transmission key on the

ACP.B - By pressing the CLR pushbutton.C - By pressing the RESET key on any ACP.

TMUCOML01 LEVEL 1



UFD0100


23 - COMMUNICATIONS

23-42-00 GROUND CREW CALL SYSTEMPRESENTATION AND OPERATION

CONTENTS:Ground Mechanic to Flight Crew CallFlight Crew to Ground Mechanic CallSelf Examination

TMUCOMO02 LEVEL 1



UFD0100


GROUND CREW CALL SYSTEM PRESENTATION AND OPERATION

The ground crew call system enables flight crew toground mechanic or ground mechanic to flight crewcalls.

GROUND MECHANIC TO FLIGHT CREW CALL

When the COCKPIT CALL pushbutton is pressed in onpanel 108VU, the MECH light flashes amber on allACPs and a buzzer is heard.An action on the RESET key of any ACP will make allthe MECH lights go off.

NOTE: MECH lights go off automatically after 60 secondsif the call is not cancelled by the RESET key.

FLIGHT CREW TO GROUND MECHANIC CALL

The horn sounds in the nosewheel well as long as theMECH pushbutton is pressed in on the cockpit CALLSpanel, and the COCKPIT CALL blue light on panel 108VUstays on.The RESET pushbutton on panel 108VU makes the COCKPITCALL blue light go off.

TMUCOMO02-T01 LEVEL 1



UFD0100


GROUND CREW CALL SYSTEM PRESENTATION AND OPERATIONTMUCOMO02-P01 LEVEL 1



UFD0100


SELF EXAMINATION

How is a ground mechanic to flight crew callindicated in the cockpit?

A - An ECAM message is displayed and abuzzer sounds.

B - The MECH light flashes on the captainACP and a buzzer sounds.

C - The MECH light flashes on all ACPs and abuzzer sounds.

TMUCOMO02 LEVEL 1



UFD0100


23 - COMMUNICATIONS

23-71-00 COCKPIT VOICE RECORDER SYSTEMPRESENTATION

CONTENTS:GeneralComponentsRecorder PanelSelf Examination

TMUCOMY01 LEVEL 1



UFD0100


COCKPIT VOICE RECORDER SYSTEM PRESENTATION

GENERAL

The Cockpit Voice Recorder (CVR) records the last 30minutes of crew conversations and communications.It records automatically in flight and on ground whenat least one engine is running and for 5 minutes afterthe last engine is shut down.The CVR can also operate in manual mode on the ground.

COMPONENTS

The components of the Cockpit Voice Recorder systemare:

- The Cockpit Voice Recorder, located in the aftsection of the aircraft.

- The CVR microphone, used for recording thedirect conversations between crew members inthe cockpit and all aural warnings.It is located at the bottom of the overheadpanel.

- The recorder (RCDR) panel, providing CVRcontrols for manual operation, test and erasureof the recording.It is located on panel 21VU on the overheadpanel.

- The CVR HEADSET jack mounted on the cockpitmaintenance panel 50VU.

TMUCOMY01-T01 LEVEL 1



UFD0100


COCKPIT VOICE RECORDER SYSTEM PRESENTATIONTMUCOMY01-P01 LEVEL 1



UFD0100


COCKPIT VOICE RECORDER SYSTEM PRESENTATION

RECORDER PANEL

GROUND CONTROLThe CVR is automatically energized in flight and onground when at least one engine is running and for 5minutes after the last engine is shut down.For manual control, on the ground, the CVR has to beenergized by pressing the ground control (GND CTL)pushbutton on the recorder (RCDR) panel.

CVR TESTWhen the CVR TEST pushbutton is pressed, either onground or in flight, a test tone is generated 4 timesfor approximately 0.8 seconds.A headset connected to the CVR HEADSET jack mountedon the cockpit maintenance panel enables monitoring.

CVR ERASEThe CVR ERASE pushbutton is used for manual erasureof the recording, only on ground with parking brakeapplied.It must be pressed for at least 2 seconds.For complete manual erasure of the recording, the CVRhas to be energized.

TMUCOMY01-T02 LEVEL 1



UFD0100


COCKPIT VOICE RECORDER SYSTEM PRESENTATIONTMUCOMY01-P02 LEVEL 1



UFD0100


SELF EXAMINATION

What is the purpose of the CVR ?A - To record radio communications during

take off and landing.B - To record crew conversations as soon as

an incident occurs.C - To record crew conversations and

communications.

TMUCOMY01 LEVEL 1



UFD0100


23 - COMMUNICATIONS

23-51-00 FLIGHT INTERPHONE SYSTEMOPERATION

CONTENTSINT SelectionRAD SelectionINT Key and KnobSelf Examination

TMUCOMN01 LEVEL 3

EFFECTIVITY 23-51-00 Page 1Aug 31/94

_A319/A320/A321 TECHNICAL TRAINING MANUAL23 COMMUNICATIONSMECHANICS / ELECTRICS & AVIONICS COURSE

UFD4200


FLIGHT INTERPHONE SYSTEM OPERATION

INT SELECTION

The INT position of the INT/RAD selector switch enablespermanent use of the flight interphone without anyfurther action and whatever the radio key selected(Here VHF 1).This is a stable position.

NOTE : The radio function has priority over the flightinterphone function.So, even with the INT/RAD switch in INTposition, the flight interphone is momentarilycut during a radio emission ( Radio key selectedand hand microphone or side-stick Push To Talkactuated).

RAD SELECTION

The RAD position of the INT/RAD selector switch putsthe preselected channel inemission (Here VHF 1).This is an unstable position.This position acts like the selection of the handmicrophone pushbutton or like the Push To Talkpushbutton of the side-stick.

INT KEY and KNOB

The flight interphone can also be used like a VHFtransceiver. Selection of the INT transmission keylights the green bars, indicating that the flightinterphone is ready to operate.Pressing and releasing the INT reception knob enablesadjustment of the interphone level. If done, the knobcomes on white.Placing and holding the INT/RAD switch in RAD positionenables the operator to talk through the flightinterphone system.

TMUCOMN01-T01 LEVEL 3



UFD4200


FLIGHT INTERPHONE SYSTEM OPERATIONTMUCOMN01-P01 LEVEL 3



UFD4200


FLIGHT INTERPHONE SYSTEM OPERATION

SELF EXAMINATION

Which action must be performed to talk through theflight interphone system ?

A - Pressing the INT transmission key andwith the INT/RAD selector switch toneutral position.

B - Either setting the INT/RAD selector toINT, or pressing the INT transmissionkey and setting the INT/RAD selectorswitch to RAD.

C - Pressing any radio transmission key andINT transmission key together.

TMUCOMN01 LEVEL 3



UFD4200




34 NAVIGATION

NAVIGATION GENERALSystem Presentation (1) 1..........................** Radio Navigation Control Presentation (1) 5....** System Controls Presentation (1) 9.............Standby Instrument Presentation (1) 19............Radio Management Panel (RMP) Presentation (1) 31..DDRMI Presentation (1) 35..........................

ADIRSADIRS Principle (1) 41..............................ADIRS Presentation (1) 55..........................** Air Data Probes Presentation (1) 61............

MULTI MODE RECEIVER (MMR) SYSTEMMMR System Description (1) 65......................

RADIO ALTIMETER (RA) SYSTEMRadio Altimeter System Presentation (1) 81........

TRAFFIC COLLISION AVOIDANCE SYSTEM (TCAS)TCAS Presentation (1) 89............................

ENHANCED GROUND PROXIMITY WARNING SYSTEM(EGPWS)** EGPWS Presentation (1) 97.......................

DISTANCE MEASURING EQUIPMENT (DME) SYSTEMDME System Presentation (1) 105....................

AIR TRAFFIC CONTROL (ATC) SYSTEMATC System Presentation (1) 113....................

AUTOMATIC DIRECTION FINDER (ADF) SYSTEMADF System Presentation (1) 119....................

VOR/MARKERS SYSTEMVOR/MKR System Presentation (1) 129................

_A319/A320/A321 TECHNICAL TRAINING MANUAL34 NAVIGATIONGENERAL FAMILIARIZATION COURSE

UFD0100

EFFECTIVITYALL

34 NAVIGATION


34 CONTENTSPage i

Dec 31/00


UFD0100

EFFECTIVITYALL

34 NAVIGATION


34 CONTENTSPage ii

Dec 31/00


34 - NAVIGATION

34-00-00 SYSTEM PRESENTATION

CONTENTS:GeneralADIRSLanding and Taxiing Aids SystemsIndependent Position Determining SystemsDependent Position Determining Systems

TMUNA2001 LEVEL 1



UFD0100


SYSTEM PRESENTATION

GENERAL

The aircraft navigation systems provide the crew withthe data required for flight within the mostappropriate safety requirements.This data is divided into four groups:

- AIR DATA/INERTIAL REFERENCE SYSTEM (ADIRS),- LANDING AND TAXIING AIDS,- INDEPENDENT POSITION DETERMINING,- DEPENDENT POSITION DETERMINING.

ADIRS

The ADIRS is an integrated Air Data System and anInertial Reference System. One part called Air DataReference mainly computes speed and altitudeinformation from air parameters. The other part calledInertial Reference mainly computes heading, attitudeand position from gyros and accelerometers.The ADIRS is composed of three Air Data/InertialReference Units (ADIRUs).Besides the ADIRUs, there are still standbyinstruments:

- Altimeter and Airspeed indicators directlysupplied by pressure lines,

- Standby Compass,- Standby Horizon.

LANDING AND TAXIING AIDS

The Head-Up Display (HUD) is used as a piloting aidssystem for roll out, take-off and landing (optional).The Instrument Landing System (ILS), is use to obtainthe optimum aircraft position during an approach andlanding phase.The Marker (MKR) system is used to indicate thedistance to the runway threshold during an ILS descent.The aircraft is equipped with:

- 1 HUD (optional),- 2 ILS,- 1 MARKER (Included in the VOR receiver).

Frequency Control is achieved either automatically ormanually (through the MCDU) by the Flight Managementand Guidance Computers (FMGCs) or manually through theRadio Management Panels (RMPs).

TMUNA2001-T01 LEVEL 1



UFD0100


TMUNA2001-P01 LEVEL 1

SYSTEM PRESENTATION



UFD0100


SYSTEM PRESENTATION

INDEPENDENT POSITION DETERMINING SYSTEMS

This part of the navigation systems, called independentsystem, provides information regarding the safety ofthe aircraft without taking reference from any groundstation.The Weather Radar / Predictive Windshear (WR/PWS)system detects the position and intensity ofprecipitations which are shown on the NavigationDisplays (ND’s).The windshear capability serves to detect any suddenchange of wind speed and/or direction (Optional).The Radio Altimeter (RA) system gives the aircraftheight above the ground, independently of theatmospheric pressure.The Traffic Collision Avoidance System (TCAS) detectsthe aircraft in the immediate vicinity.The Enhanced Ground Proximity Warning System (EGPWS)warns the flight crew about the aircraft behaviour indangerous configuration when approaching the ground.This part of the Navigation system includes:

- 1 Weather Radar / Predictive Windshear (WR/PWS)(the second is optional),

- 2 Radio Altimeter (RA),- 1 Traffic Collision Avoidance System (TCAS),- 1 Enhanced Ground Proximity Warning System(EGPWS).

DEPENDENT POSITION DETERMINING SYSTEMS

This part of the navigation system, called dependentsystem, provides various means of navigation throughdata exchange with ground installations or satellites.

The Distance Measuring Equipment (DME) system givesthe aircraft slant distance to a ground station.The Air Traffic Control system (ATC) enables a groundoperator to identify and track the aircraft withouthaving to communicate with the flight crew.The Automatic Direction Finder (ADF) system is a radiocompass system providing the azimuth of a NonDirectional Beacon (NDB) with respect to the aircraftcenter line.The VHF Omni Range (VOR) system gives the bearing ofa ground VOR Station with respect to the magnetic Northand the aircraft angular deviation related to apreselected course.The Global Positioning System (GPS) is based on themeasurement of the transmission time of signalsbroadcast by satellites.This part of the Navigation includes:

- 2 DME,- 2 ATC,- 1 ADF (the second is optional),- 2 VOR,- 2 GPS.

NOTE 1: The VOR or DME frequency control is achievedeither automatically or manually (through theMCDU) by the FMGCs or manually though the RMPs.

NOTE 2: Although the Marker Beacon belongs to theLanding Aids System, it is physicallyintegrated into the VOR receiver.




UFD0100


34 - NAVIGATION

34-00-00 RADIO NAVIGATION CONTROLPRESENTATION

CONTENTS:Automatic TuningManual TuningBack-Up TuningSelf Examination

TMUNAV701 LEVEL 1



UFD0100


RADIO NAVIGATION CONTROL PRESENTATION

AUTOMATIC TUNING

The Automatic Tuning permits the control of VOR / DME,ILS and ADF by the Flight Management and GuidanceSystem.In this case the RMP is transparent to its associatedFMGC.In normal operation FMGC1 tunes receivers 1, FMGC2tunes receivers 2.In case of failure of FMGC 1 or 2, the remaining FMGCcontrols all receivers.

MANUAL TUNING

The manual tuning permits the pilot to select, throughthe Multipurpose Control Display Unit, a specificfrequency for display on the EFIS.

NOTE: To return to the autotuning mode, the manualtuning has to be cleared.

BACK-UP TUNING

Radio Management Panels 1 and 2 located on the pedestalprovide back-up for Radio Navigation tuning.We are in the case of both FMGCs inoperative oremergency electrical supply.The ILS course and frequency are the only RadioNavigation data exchanged.The selected values on RMP 1 and RMP 2 are identicalfor ILS 1 and ILS 2.

TMUNAV701-T01 LEVEL 1



UFD0100


RADIO NAVIGATION CONTROL PRESENTATIONTMUNAV701-P01 LEVEL 1



UFD0100


SELF EXAMINATION

Can VOR 2 frequency be changed through RMP 1?A - Yes.B - No.

In back-up mode, an ILS can be tuned through:A - RMP 1 or 2.B - RMP 1 only.C - The onside RMP only.

TMUNAV701 LEVEL 1



UFD0100


34 - NAVIGATION

34-00-00 SYSTEM CONTROLS PRESENTATION

CONTENTS:Multipurpose Control Display Unit (MCDU)ADIRS Control Display Unit (ADIRS CDU)Radio Management Panel (RMP)Audio Control Panel (ACP)

TMUNA2101 LEVEL 1



UFD0100


SYSTEM CONTROLS PRESENTATION

MULTIPURPOSE CONTROL DISPLAY UNIT (MCDU)

The Multipurpose Control and Display Unit (MCDU) allowsthe crew:

- To display the Radio Navigation frequencies(automatically or manually tuned) on a specificpage called RAD/NAV.

- To align the Inertial Reference systems froma specific page called INIT via the FMGC.

- To initiate tests for all navigationsystems and for troubleshooting via theCFDIU.




UFD0100


SYSTEM CONTROLS PRESENTATION - MCDUTMUNA2101-P01 LEVEL 1



UFD0100



ADIRS CONTROL DISPLAY UNIT (ADIRS CDU)

The ADIRS Control and Display Unit allows the followingfunctions:

- To switch on the ADR and IR by setting a singlecontrol to NAV.

- To disconnect the ADR output bus by a specificpushbutton.

- To check the ADIRU operation- To align the IR instead of using the MCDU.

NOTE: when set to ATT, the systems are still energizedbut the IR is in downgraded operation mode.




UFD0100


SYSTEM CONTROLS PRESENTATION - ADIRS CDUTMUNA2101-P02 LEVEL 1



UFD0100



RADIO MANAGEMENT PANEL (RMP)

The main function of the Radio Management Panels (RMP)is to control all communication frequencies. Howeverthey are also used for standby selection ofRadio/NAV frequencies.The standby operation is used in case of dual FMGCfailure, provided the NAV pushbutton switch has beenpressed.




UFD0100


SYSTEM CONTROLS PRESENTATION - RMPTMUNA2101-P03 LEVEL 1



UFD0100



AUDIO CONTROL PANEL (ACP)

The Audio Control Panels (ACP) enable to control thereception of all audio signals identifying the variousbeacons and stations.

NOTE: DME identification signals can be selected byusing the knob of the colocated VOR or ILS.




UFD0100


SYSTEM CONTROLS PRESENTATION - ACPTMUNA2101-P04 LEVEL 1



UFD0100



TMUNA2101 LEVEL 1



UFD0100


34 - NAVIGATION

34-20-00 STANDBY INSTRUMENT PRESENTATION

CONTENTS:Standby CompassStandby HorizonStandby AltimeterStandby Airspeed Indicator (ASI)Metric Altimeter (Option)

TMUNAVE03 LEVEL 1



UFD0100


STANDBY INSTRUMENT PRESENTATION

STANDBY COMPASS

The standby compass is located on the top windshieldcenter post.It is stowed in normal configuration.A correction card is glued to the side of the compassassembly.Non-magnetic lamp:- A non-magnetic lamp assembly lights the compass card.It is controlled by a STBY COMPASS switch located onthe INT LT panel on the overhead panel.Graduated compass:- The graduated compass card is attached to a magneticelement. It is free to rotate inside the compass bowland is immersed in a damping liquid.Lubber line:- A lubber line indicates the magnetic heading.Compensation holes:- Two holes marked NS and EW, allow compensation bypositioning two small magnetic bars calledcompensators.

TMUNAVE03-T01 LEVEL 1



UFD0100


STANDBY INSTRUMENT PRESENTATION - COMPASSTMUNAVE03-P01 LEVEL 1



UFD0100



STANDBY HORIZON

The Standby horizon is located on the center instrumentpanel and comprises the following elements:Roll pointer:- The roll information is given by a pointer whichmoves in front of a graduated roll scale.Roll scale:- The roll scale is graduated in 10 degree incrementsbetween -30 and +30 degrees and 15 degree incrementsup to 60 degrees.Flag:- The flag comes into view if a failure is detectedin the electrical power supply or if the gyro rotorspeed drops below 18000 RPM.Pitch drum:- The pitch drum is divided into two zones separatedby the reference horizon.The pitch indications are displayed between -80 and +80 degrees.Aircraft symbol:- The aircraft symbol is fixed.Resetting knob:- Fast resetting can be performed by pulling the cagingknob (Also used for shipping to protect the gyro).




UFD0100


STANDBY INSTRUMENT PRESENTATION - HORIZONTMUNAVE03-P02 LEVEL 1



UFD0100



STANDBY ALTIMETER

The Standby Altimeter is located on the centerinstrument panel and comprises the following elements:Adjustable bugs:- Four manually adjustable bugs are provided forreference altitude setting.Altitude counter:- A display counter made up of two drums displays thetens of thousands, and the thousands of feet.When the altitude is below 10000 feet, the left drumdisplays black and white stripes.In case of negative altitude the left drum displaysorange and white stripes.Altitude pointer:- The pointer indicates the hundreds of feet with 20feet increments.To prevent the pointer from sticking, an internalvibrator is installed.It is only supplied in flight.Altitude dial:- The altitude dial is calibrated from O to 1000 feetwith 20 feet graduations.Baro correction counter:- The baro correction is displayed on a countergraduated in hecto Pascals.Adjustment baro setting knob:- The knob enables adjustment of the baro setting inthe range of 750 to 1050 hecto Pascals.




UFD0100


STANDBY INSTRUMENT PRESENTATION - ALTIMETERTMUNAVE03-P03 LEVEL 1



UFD0100



STANDBY AIRSPEED INDICATOR (ASI)

The Standby Airspeed Indicator (ASI) is located on thecenter instrument panel and comprises the followingelements:Ajustable bugs:- Four manually adjustable bugs are provided forreference speed setting.Speed pointer:- The pointer moves on a graduated dial.Speed dial:- The dial is made of two linear scales: one from 60ktto 250kt with 5kt increments, the other from 250 to450kt with 1Okt increments.




UFD0100


STANDBY INSTRUMENT PRESENTATION - AIRSPEED INDICATORTMUNAVE03-P04 LEVEL 1



UFD0100



METRIC ALTIMETER (OPTION)

The Metric Altimeter is located on the centerinstrument panel and comprises the following elements:A display counter, made up of two drums displays thetens of thousands and the thousands of meters.Altitude counter:- When the altitude is below 10,000 meters, the leftdrum displays black and white stripes.In case of negative altitude the left drum displaysorange and white stripes.Altitude pointer:- The pointer indicates the hundreds of meters with50 meter increments.Altitude dial:- The altitude dial is calibrated from O to 1000 meterswith 50 meter graduations.Baro correction counter:- The baro correction is displayed on a the lowercounter and is graduated in hecto Pascals.Adjustment baro setting knob:- The knob enables adjustment of the baro setting inthe range of 870 to 1050 hecto Pascals.




UFD0100


TMUNAVE03-P05 LEVEL 1

STANDBY INSTRUMENT PRESENTATION - METRIC ALTIMETER (OPTION)



UFD0100



TMUNAVE03 LEVEL 1



UFD0100


34 - NAVIGATION

34-00-00 RADIO MANAGEMENT PANEL (RMP)PRESENTATION

CONTENTS:GeneralStandby Navigation KeysRotating KnobStandby/Course (STBY/CRS) WindowActive Window

TMUNA2201 LEVEL 1



UFD0100


RADIO MANAGEMENT PANEL (RMP) PRESENTATION

GENERAL

The radio Nav frequency selection can only be performedwhen the guarded NAV key LED is on, after a dual FMGSlossThe ON/OFF switch controls the power supply of theRMP.

STANDBY NAVIGATION KEYS

As long as the NAV key is the only STBY NAV key selected,the windows still display communication frequencies.Then, pressing the VOR, ILS or ADF key changes thedisplays to the last RMP memorized values (frequencyand course).At any time, communication frequencies are stillselectable, simply by pressing the corresponding key.Beat Frequency Oscillator (BFO) is set on or off bypressing the key.

NOTE: The Microwave Landing System (MLS) key is aprovision.

ROTATING KNOB

Two concentric knobs allow preselection of frequencyfor radio communication and standby navigation systemsand selection of the required course for VORs and ILSs:

- the outer knob controls the most significantdigits,

- the inner knob controls the least significantdigits.

The desired frequency or course is set in the STBY/CRSwindow. Frequency becomes active by pressing thetransfer key.

STANDBY/COURSE (STBY/CRS) WINDOW

The STANDBY/COURSE window displays a standby frequencyor a course.Both can be changed by rotating the knob, but only thestandby frequency can be made active by pressing thetransfer key.If a course is displayed, the associated frequency isdisplayed in the ACTIVE window.

NOTE: If a course is displayed on the STBY/CRS window,pressing the transfer key will display the ACTIVEfrequency in both windows.

ACTIVE WINDOW

The active window shows the frequency in use of thesystem identified by the green LED on the selectedkey.




UFD0100


RADIO MANAGEMENT PANEL (RMP) PRESENTATIONTMUNA2201-P01 LEVEL 1



UFD0100



TMUNA2201 LEVEL 1



UFD0100


34 - NAVIGATION

34-00-00 DDRMI PRESENTATION

CONTENTS:GeneralNormal OperationFailure and Non Computed Data (NCD)Self Examination

TMUNAVF02 LEVEL 1



UFD0100


DDRMI PRESENTATION

GENERAL

The Digital Distance and Radio Magnetic Indicator(DDRMI) is located on the center instrument panel.It’s a combined VOR/ADF/DME RMI.

Note: Some DDRMIs are not equipped with the ADFcapability.

NORMAL OPERATION

The DME 1 Distance is displayed in the left hand window.The DME 2 Distance is displayed in the right handwindow.A single pointer indicates the VOR 1 or ADF 1 bearing.A double pointer indicates the VOR 2 or ADF 2 bearing.The selection of VOR or ADF is provided for each pointerby a selector switch.

TMUNAVF02-T01 LEVEL 1



UFD0100


DDRMI PRESENTATION - NORMAL OPERATIONTMUNAVF02-P01 LEVEL 1



UFD0100


DDRMI PRESENTATION

FAILURE AND NON COMPUTED DATA (NCD)

When a failure is detected by the DME or RMI monitoringcircuits, the corresponding DME display window isblanked.In case of Non Computed Data (NCD), for example:

- out-of-range station, the window shows whitehorizontal dashed lines.

Heading information normally comes from ADIRU 1.If it fails, the heading is provided by ADIRU 3 afterpilot switching.In case of VOR or ADF 1 or 2 receiver failure, a redflag comes into view and the corresponding pointer isset to the 3 o’clock position.In case of Non Computed Data (NCD), no failure flagappears, but the corresponding pointer is set to the3 o’clock position.

TMUNAVF02-T02 LEVEL 1



UFD0100


TMUNAVF02-P02 LEVEL 1

DDRMI PRESENTATION - FAILURE AND NON COMPUTED DATA (NCD)



UFD0100


SELF EXAMINATION

From which ADIRU can the DDRMI receive information ?A - ADIRU 1 or 2.B - ADIRU 2 or 3.C - ADIRU 1 or 3.

TMUNAVF02 LEVEL 1



UFD0100


34 - NAVIGATION

34-10-00 ADIRS PRINCIPLE

CONTENTS:GeneralADM Functional DescriptionADM InputsADM OutputADR ComputationIR StrapdownRing Laser GyroAccelerometerIR Computation

TMUADI001 LEVEL 1



UFD0100


ADIRS PRINCIPLE

GENERAL

The Air Data/Inertial Reference Unit (ADIRU) comprisesan Air Data Reference Unit and an Inertial ReferenceUnit, both included in a single unit.Data from external sensors (Angle of Attack, Total AirTemperature, Air Data Module) are used by the ADIRU.The ADIRUs are interfaced with the ADIRS DisplayControl Unit (CDU) for control and status annunciation.

TMUADI001-T01 LEVEL 1



UFD0100


ADIRS PRINCIPLE - GENERALTMUADI001-P01 LEVEL 1



UFD0100


ADIRS PRINCIPLE

ADM FUNCTIONAL DESCRIPTION

A microcomputer processes an ARINC signal accordingto the discrete inputs and to the digitized pressure.

ADM INPUTS

The ADM Inputs are one pressure input and severaldiscrete inputs.The ADMs are identical.The discrete inputs determine the ADM location and thetype of pressure data (Pitot or Static) provided tothe ADR.

ADM OUTPUT

The ADM output is an ARINC bus which provides digitalpressure information, type of pressure, ADMidentification and BITE status to the ADIRU.




UFD0100


ADIRS PRINCIPLE - ADMTMUADI001-P02 LEVEL 1



UFD0100


ADIRS PRINCIPLE

ADR COMPUTATION

The ADR processes sensor and ADM inputs in order toprovide air data to users.




UFD0100


ADIRS PRINCIPLE - ADRTMUADI001-P03 LEVEL 1



UFD0100


ADIRS PRINCIPLE

IR STRAPDOWN

In a strapdown Inertial Reference System the gyros andthe accelerometers are solidly attached to the aircraftstructure.The strapdown laser gyro provides directlyaccelerations and angular speeds.




UFD0100


ADIRS PRINCIPLE - IRTMUADI001-P04 LEVEL 1



UFD0100


ADIRS PRINCIPLE

RING LASER GYRO

The three ring laser gyros, one for each rotation axis,provide inertial rotation data and are composed of twoopposite laser beams in a ring.At rest, the two beams arrive at the sensor with thesame frequency.An aircraft rotation creates a difference offrequencies between the two beams.The frequency difference is measured by optical meansproviding a digital output which, after computation,will provide rotation information.

NOTE: Light Amplification Stimulated Emission ofRadiation (LASER)




UFD0100


ADIRS PRINCIPLE - GYROTMUADI001-P05 LEVEL 1



UFD0100


ADIRS PRINCIPLE

ACCELEROMETER

Three accelerometers, one for each axis, provide linearaccelerations.The acceleration signal is sent to a processor whichuses this signal to compute the velocity and theposition.

IR COMPUTATION

Each ADIRU computes the laser gyro and theaccelerometer outputs to provide Inertial Referencedata to users.




UFD0100


ADIRS PRINCIPLE - ACCELEROMETERTMUADI001-P06 LEVEL 1



UFD0100



TMUADI001 LEVEL 1



UFD0100


34 - NAVIGATION

34-10-00 ADIRS PRESENTATION

CONTENTS:GeneralMCDUADIRS CDUProbesFCUGPSDMCDMC/PFD & NDADIRS SwitchingUsersSelf Examination

TMUADIA01 LEVEL 1



UFD0100


ADIRS PRESENTATION

GENERAL

The Air Data Inertial Reference System (ADIRS) iscomposed of three Air Data Inertial Reference Units(ADIRU), each having their own set of probes andsensors and a common Control Display Unit (CDU).

MCDU

The Multipurpose Control and Display Units (MCDUs) arenormally used to align the Inertial References, toinitiate the ADIRU tests and to display ADIRUinformation.

ADIRS CDU

The ADIRS Control Display Unit is used as a back-upfor Inertial Reference alignment. It is also used formode selection, information display and statusindication.

PROBES

The Air Data input parameters, such as total and staticpressures, Angle Of Attack (AOA) and Total AirTemperature (TAT) are sent, from the related probesand sensors, to the three ADIRUs.

NOTE : static and total pressure are sent to the ADIRUsvia the ADMs.

FCU

The ADIRUs receive, from the Flight Control Unit (FCU),the Baro correction set by the crew.

GPS

The Global Positioning System (GPS) provides data tothe ADIRS, mainly A/C position and speed. The ADIRSprocesses the GPS data and provides pure GPS data,pure IR data and hybrid GPS/ADIRS data to users.

DMC

The Display Management Computers (DMCs) 1 and 2 receivetheir data from their related ADIRU and from ADIRU 3.The Display Management Computer 3 (DMC3) receivesinformation from all three ADIRUs, to operate as aback-up in case of DMC1 or 2 failure.

DMC/PFD & ND

ADIRU 1 and 2 display information via DMC 1 and 2, onthe corresponding Primary Flight Display (PFD) andNavigation Display (ND).ADIRU 3 operates as a back-up in case of ADIRU 1 or 2failure.

TMUADIA01-T01 LEVEL 1



UFD0100


ADIRS PRESENTATIONTMUADIA01-P01 LEVEL 1



UFD0100


ADIRS PRESENTATION

ADIRS SWITCHING

Basically, ADIRU 1 is associated to the captaininstruments, ADIRU 2 to the first officer instrumentsand ADIRU 3 is in standby.In case of failure of the Air Data Reference (ADR) orInertial Reference (IR) function of ADIRU 1 or 2, theaffected instruments and displays may be manuallyswitched independently to ADIRU 3 by means of selectorswitches.

USERS

The ADIRUs are directly connected to other user system.

TMUADIA01-T01 LEVEL 1



UFD0100


SELF EXAMINATION

Which ADIRU does not supply DMC 1?A - ADIRU 1.B - ADIRU 2.C - ADIRU 3.

TMUADIA01 LEVEL 1



UFD0100



TMUADIA01 LEVEL 1



UFD0100


34 - NAVIGATION

34-13-00 AIR DATA PROBES PRESENTATION

CONTENTS:Pitot ProbesStatic PortsAOA SensorsTAT SensorsWater DrainSelf Examination

TMUADIB01 LEVEL 1



UFD0100


AIR DATA PROBES PRESENTATION

PITOT PROBES

The total pressure is sent from the Pitot Probes tothe Air Data Modules which convert it into ARINC wordsused by the Air Data Inertial Reference Units.Three pitot probes provide total pressure to three AirData Modules (ADM) which convert this pressure intodigital format (ARINC 429).ARINC words are then sent to the corresponding AirData Inertial Reference Unit (ADIRU). The standby pitotprobe supplies the standby Airspeed Indicator (ASI)and ADR3 through its related ADM.

STATIC PORTS

Each Air Data Module transforms the static pressurecoming from the static ports into ARINC words.Six static ports provide static pressure to five ADMswhich convert this pressure into digital format (ARINC429).The standby static ports provide an average pressuredirectly to the standby instruments, and to ADR3through a single ADM.

AOA SENSORS

Each ADIRU receives angle of attack information fromits corresponding Angle Of Attack (AOA) sensors.The Angle Of Attack sensors are also called Alphaprobes.

TAT SENSORS

The three ADIRUs receive Total Air Temperatureinformation from two Total Air Temperature sensors.

NOTE: that ADIRU3 receives the Total Air Temperature(TAT) from the TAT 1 sensor which is composedof two elements.

WATER DRAIN

The probes are installed in such a way that theirpressure lines do not require a water drain, exceptfor that of the standby static ports.

TMUADIB01-T01 LEVEL 1



UFD0100


AIR DATA PROBES PRESENTATIONTMUADIB01-P01 LEVEL 1



UFD0100


SELF EXAMINATION

How do pitot and static probes supply the ADIRUs?A - Using ADMs which convert pressure into

digital format.B - Directly with total and static pressures.C - Directly with digital format.

Where does ADIRU3 receive TAT information from?A - Captain TAT sensor.B - First-Officer TAT sensor.C - Standby TAT sensor.

Which pressure line(s) need(s) to be drained?A - All pitot lines.B - All static lines.C - Only the standby static line.

TMUADIB01 LEVEL 1



UFD0100


34 - NAVIGATION

34-36-00 MMR SYSTEM PRESENTATION

CONTENTS:GeneralILS PrincipleGPS PrincipleComponentsILS IndicatingGPS IndicatingSelf Examination

TMUMMRA01 LEVEL 1

EFFECTIVITY 34-36-00 Page 1Nov 30/99


UFD0100


MMR SYSTEM PRESENTATION

GENERAL

The Multi Mode Receiver (MMR) system is a NavigationSensor with 2 internal receivers.MMR = ILS + GPS

ILS PRINCIPLE

The function of the Instrument Landing System (ILS)is to provide the crew and airborne system users withsignals transmitted by a ground station.A descent axis is determined by the intersection of aLocalizer beam (LOC) and a Glide Slope beam (G/S)created by this ground station at known frequencies.The ILS allows measurement and display of angulardeviations and receives the Morse audio signal whichidentifies the ILS ground station.ILS operational frequency range:

- LOC: between 108.1 and 111.95 MHz,- G/S: between 329.15 and 335 MHz.

TMUMMRA01-T01 LEVEL 1



UFD0100


MMR SYSTEM PRESENTATION - ILS PRINCIPLETMUMMRA01-P01 LEVEL 1



UFD0100



GPS PRINCIPLE

The NAVigation System Time And Ranging GlobalPositioning System ( NAV.S.T.A.R. GPS ) is a worldwidenavigation radio aid which uses satellite signals toprovide accurate navigation information.The architecture of the system is composed of 3 partscalled segments:

- Spatial Segment- Control Segment- User Segment

SPATIAL SEGMENTThe spatial segment is composed of a constellation of24 satellites.These satellites are arranged in six separate orbitalplanes of four satellites each on a circular orbit andhave the following characteristics:

- 55° inclination to the Equator,- an altitude of approx 20200 km with an orbitalperiod of 12 sideral hours.

These satellites give:- the satellite position (Ephemeris of theconstellation),

- the constellation data (Almanach).- the atmospheric corrections.




UFD0100


MMR SYSTEM PRESENTATION - GPS PRINCIPLE - SPATIAL SEGMENTTMUMMRA01-P02 LEVEL 1



UFD0100



GPS PRINCIPLE (Continued)

CONTROL SEGMENTThe control segment is composed of four monitorstations and one master control station which trackthe satellites, compute the ephemeris, clockcorrections and control the navigation parameters andtransmit them to the GPS users.The four monitor stations are located at:

- KWAJALEIN- HAWAII- ASCENCION ISLAND- DIEGO GARCIA

The master control station is located at:- COLORADO SPRINGS.




UFD0100


MMR SYSTEM PRESENTATION - GPS PRINCIPLE - CONTROL SEGMENTTMUMMRA01-P03 LEVEL 1



UFD0100



GPS PRINCIPLE (Continued)

USER SEGMENTThe principle of GPS position computation is based onthe measurement of transmission time of the GPS signalsbroadcast by at least 4 satellites.This segment is constitued by the GPS receiver anddefined as follows:

- signal acquisition,- distance calculation,- navigation computation (Satellite choice,positioning, propagation corrections),

- detection and isolation of failed satellites(GPS PRIMARY).

NOTE: When GPS mode is active, no VOR/DME/ADF data isused for navigation.




UFD0100


TMUMMRA01-P04 LEVEL 1

MMR SYSTEM PRESENTATION - GPS PRINCIPLE - USER SEGMENT



UFD0100



COMPONENTS

The components are two ILS antennas, 2 GPS antennasand two MMR units.The MMR system is also connected to:

- Primary Flight Display (PFD) and NavigationDisplay (ND) for display.

- Electronic Flight Instrument System (EFIS)control unit for display control.

- Flight Management and Guidance Computer (FMGC)for ILS auto-tuning and GPS position.

- Multipurpose Control Display Units (MCDU) forILS manual tuning.

- Captain and First Officer Radio ManagementPanels (RMP) for ILS back-up tuning.

- Audio Control Panels (ACP) for ILS audiosignal.

- Air Data and Inertial Reference Unit forGPIR data.




UFD0100


MMR SYSTEM PRESENTATION - COMPONENTSTMUMMRA01-P05 LEVEL 1



UFD0100



ILS INDICATING

The ILS data appears on the PFD as soon as the ILSpushbutton switch on the EFIS control panel has beenpressed in and on the ND when ROSE/ILS mode has beenselected.ILS information is displayed in magenta.The ILS1 information is displayed on PFD1 and ND2.The ILS2 information is displayed on PFD2 and ND1.




UFD0100


MMR SYSTEM PRESENTATION - ILS INDICATINGTMUMMRA01-P06 LEVEL 1



UFD0100



GPS INDICATING

The GPS data is displayed on the MCDUs and on the NDs.- GPS data on MCDU (GPS MONITOR page): * GPS position (Lat, Long) * True Track * GPS altitude * Figure of Merit * Ground Speed * Number of satellites tracked * Mode.- GPS message on ND: * Availability of the GPS Primary navigation function.




UFD0100


TMUMMRA01-P07 LEVEL 1

MMR SYSTEM PRESENTATION - GPS INDICATING



UFD0100


SELF EXAMINATION

The satellite orbital planes have an inclination of:A - 60°.B - 55° to the Equator.C - 45° to the Equator.

The control segment is composed of:A - 3 monitor stations.B - 5 monitor stations.C - 4 monitor stations and one master

control station.

TMUMMRA01 LEVEL 1



UFD0100


34 - NAVIGATION

34-42-00 RADIO ALTIMETER SYSTEMPRESENTATION

CONTENTS:PrincipleComponentsIndicating

TMURADG01 LEVEL 1



UFD0100


RADIO ALTIMETER SYSTEM PRESENTATION

PRINCIPLE

The Radio Altimeter (RA) System determines the heightof the aircraft above the terrain during initial climb,approach and landing phases.The RA can therefore operate over non-flat groundsurface.The principle of the radio altimeter is to transmit afrequency modulated signal, from the aircraft to theground, and to receive the ground reflected signalafter a certain delay.The time between the transmission and the receptionof the RA signal is proportional to the A/C height.

TMURADG01-T01 LEVEL 1



UFD0100


RADIO ALTIMETER SYSTEM PRESENTATION - PRINCIPLETMURADG01-P01 LEVEL 1



UFD0100



COMPONENTS

The components are two transceivers, two fans, twotransmission antennae and two reception antennae.The RA system is also connected to the DMCs for displayon the Primary Flight Displays (PFDs).




UFD0100


RADIO ALTIMETER SYSTEM PRESENTATION - COMPONENTSTMURADG01-P02 LEVEL 1



UFD0100



INDICATING

The aircraft height data is displayed on the PrimaryFlight Displays for heights less than or equal to 2500ft.The altitude is also shown by means of:

- A ground line rising on to the pitch down area(Below 300 ft).

- A red ribbon next to the altitude scale(Below 500 ft).




UFD0100


RADIO ALTIMETER SYSTEM PRESENTATION - INDICATINGTMURADG01-P03 LEVEL 1



UFD0100



TMURADG01 LEVEL 1



UFD0100


34 - NAVIGATION

34-43-00 TCAS PRESENTATION

CONTENTS:PrincipleComponentsIndicatingSelf Examination

TMUTCAB05 LEVEL 3



UFD0100


TCAS PRESENTATION

PRINCIPLE

The Traffic Collision Avoidance System (TCAS) is asystem whose function is to detect and display aircraftin the immediate vicinity and to provide the flightcrew with indications to avoid these intruders.The TCAS indications for flight plan modifications arein the vertical plane only.The TCAS detects the air traffic control system orTCAS equipped aircraft and maintains surveillancewithin a range determined by its sensivity.To evaluate threat potential of other aircraft thesystem divides the space around aircraft into 4volumes.- OTHER TRAFFIC VOLUME.The OTHER TRAFFIC VOLUME is the first volume providingthe presence and the progress of on intruder.(No collision threat).- PROXIMATE TRAFFIC VOLUME.The proximate traffic volume is defined by a givenvolume around the TCAS equipped aircraft.(No collision threat, but in vicinity).

- TRAFFIC ADVISORY VOLUME (TA).When the intruder is relatively near but does notrepresent an immediate threat, the TCAS provides anaural and visual information known as Traffic Advisory(TA).The TCAS aural messages can be inhibited depending onhigher priority aural messages.- RESOLUTION ADVISORY VOLUME (RA).When the intruder represents a collision threat, theTCAS triggers an aural and visual alarm known asResolution Advisory (RA), which informs the crew aboutavoidance maneuvers.

TMUTCAB05-T01 LEVEL 3



UFD0100


TCAS PRESENTATION - PRINCIPLETMUTCAB05-P01 LEVEL 3



UFD0100


TCAS PRESENTATION

COMPONENTS

The TCAS components are two antennae, one TCAS computerand one TCAS/ATC control panel.Note: The TCAS/ATC control panels shown here after are

given as examples. They may differ according tothe aircraft configuration.




UFD0100


TCAS PRESENTATION - COMPONENTSTMUTCAB05-P02 LEVEL 3



UFD0100


TCAS PRESENTATION

INDICATING

The TCAS indications appear on the PFD and the ND.The visual resolution and traffic advisory indicationsare associated with aural indications such as "TRAFFIC,TRAFFIC", "CLIMB, CLIMB"...The TCAS displays only the most threatening intruders.




UFD0100


TCAS PRESENTATION - INDICATINGTMUTCAB05-P03 LEVEL 3



UFD0100


SELF EXAMINATION

Which aircraft are detected by the TCAS ?A - All.B - Only ATC Mode S equipped A/C.C - ATC equipped A/C.

TMUTCAB05 LEVEL 3



UFD0100


34 - NAVIGATION

34-48-00 ENHANCED GROUND PROXIMITYWARNING SYSTEM PRESENTATION

CONTENTS:GeneralPrincipleComponentsIndicating

TMUEGPA01 LEVEL 1

EFFECTIVITY 34-48-00 Page 1Feb 28/98


UFD0100


ENHANCED GROUND PROXIMITY WARNING SYSTEM PRESENTATION

GENERAL

The Enhanced Ground Proximity Warning System (EGPWS)is built over the current GPWS.

- EGPWS = GPWS + "ENHANCED" functions.

PRINCIPLE

The purpose of the Enhanced Ground Proximity WarningSystem (EGPWS) is to help prevent accidents caused byControlled Flight Into Terrain (CFIT).When boundaries of any alerting envelope are exceeded;aural alert messages, visual annunciations anddisplays are generated.The basic GPWS modes generate aural and visual warningscorresponding to an aircraft behaviour when the alertenvelope is penetrated.The "ENHANCED" features complete the basic GPWS modes:

- Terrain Clearance Floor (TCF): Increase theterrain clearance envelope around the airportrunway.

- Terrain Awareness alerting and Display (TAD):Incorporation of a terrain database to predictconflict between flight path and terrain andto display the conflicting terrain.

TMUEGPA01-T01 LEVEL 1



UFD0100


ENHANCED GROUND PROXIMITY WARNING SYSTEM PRESENTATION - PRINCIPLETMUEGPA01-P01 LEVEL 1



UFD0100



COMPONENTS

The system comprises an EGPWC, a control panel, twowarning lights and two TERRAIN ON ND mode pushbuttonswitches.The EGPWS is connected to various navigation systems(WR, RA, ADIRS, ILS...).It processes the navigation data and generates alarms.




UFD0100


ENHANCED GROUND PROXIMITY WARNING SYSTEM PRESENTATION - COMPONENTS

TMUEGPA01-P02 LEVEL 1



UFD0100



INDICATING

The basic GPWS modes generate visual warnings throughassociated lights and synthetic warnings through theloudspeakers.The "ENHANCED" GPWS functions allow the terrain hazardsto be displayed on the Navigation Display (ND).




UFD0100


ENHANCED GROUND PROXIMITY WARNING SYSTEM PRESENTATION - INDICATINGTMUEGPA01-P03 LEVEL 1



UFD0100



TMUEGPA01 LEVEL 1



UFD0100


34 - NAVIGATION

34-51-00 DME SYSTEM PRESENTATION


TMUDMEH01 LEVEL 1



UFD0100


DME SYSTEM PRESENTATION

PRINCIPLE

The Distance Measuring Equipment (DME) providesdigital readout of the aircraft slant range distancefrom a selected ground station.The system generates interrogation pulses from anonboard interrogator and sends them to a selectedground station.After a 50 micro seconds delay, the ground stationreplies.The interrogator determines the distance in NauticalMiles (NM) between the station and the aircraft.The interrogator detects the Morse audio signal whichidentifies the ground station.

TMUDMEH01-T01 LEVEL 1



UFD0100


DME SYSTEM PRESENTATION - PRINCIPLETMUDMEH01-P01 LEVEL 1



UFD0100



COMPONENTS

The components are two antennae and two interrogators.The DME system is also connected to:

- Primary Flight Displays (PFD), NavigationDisplays (ND) and Digital Distance RadioMagnetic Indicator (DDRMI) for display.

- Electronic Flight Instrument System (EFIS)control unit for display control.

- Flight Management and Guidance Computers (FMGC)for tuning (manual and auto).

- Captain and F/O Radio Management Panels (RMP)for back-up tuning.

- Audio Control Panels (ACPs) for DME audiosignal.




UFD0100


DME SYSTEM PRESENTATION - COMPONENTSTMUDMEH01-P02 LEVEL 1



UFD0100



INDICATING

The DME distance is shown on the Primary Flight Display(PFD) (if ILS/DME) and on the Navigation Display (ND)(if VOR/DME).The DME distance is also shown on the two windows ofthe Digital Distance Radio Magnetic Indicator (DDRMI).




UFD0100


DME SYSTEM PRESENTATION - INDICATINGTMUDMEH01-P03 LEVEL 1



UFD0100



TMUDMEH01 LEVEL 1



UFD0100


34 - NAVIGATION

34-52-00 ATC SYSTEM PRESENTATION

CONTENTS:PrincipleComponents

TMUATCF08 LEVEL 1



UFD0100


ATC SYSTEM PRESENTATION

PRINCIPLE

The Air Traffic Control (ATC) transponder is anintegral part of the Air Traffic Control Radar BeaconSystem (ATCRBS).The transponder is interrogated by radar pulsesreceived from the ground station.It automatically replies by a series of pulses.These reply pulses are coded to supply identification(Mode A) and automatic altitude reporting (Mode C) ofthe aircraft on the ground controller’s radar scope.These replies enable the controller to distinguish theaircraft and to maintain effective ground surveillanceof the air traffic.The ATC transponder also responds to interrogationsfrom aircraft equipped with a Traffic CollisionAvoidance System (TCAS) (Mode S).

TMUATCF08-T01 LEVEL 1



UFD0100


ATC SYSTEM PRESENTATION - PRINCIPLETMUATCF08-P01 LEVEL 1



UFD0100


ATC SYSTEM PRESENTATION

COMPONENTS

The components are two transponders, four antennae,and one ATC/TCAS control panel.Note: The TCAS/ATC control panels shown here after are

given as examples. They may differ according tothe aircraft configuration.

TMUATCF08-T02 LEVEL 1



UFD0100


ATC SYSTEM PRESENTATION - COMPONENTSTMUATCF08-P02 LEVEL 1



UFD0100



TMUATCF08 LEVEL 1



UFD0100


34 - NAVIGATION

34-53-00 ADF SYSTEM PRESENTATION


TMUADFF01 LEVEL 1



UFD0100


ADF SYSTEM PRESENTATION

PRINCIPLE

The Automatic Direction Finder (ADF) is a radionavigation aid.The ADF system provides:

- An identification of the relative bearing toa selected ground station called NonDirectional Beacon (NDB).

- Aural identification of the ground station.The relative bearing is the angle between the aircraftheading and the aircraft/ground station axis.

TMUADFF01-T01 LEVEL 1



UFD0100


ADF SYSTEM PRESENTATION - PRINCIPLETMUADFF01-P01 LEVEL 1



UFD0100



The combination of signals, received from two loopantennae and from one omni-directional sense antenna,provides bearing information.The ground stations operate in a frequency range of190 to 1750 Khz.An additional Morse signal is provided to identify theselected ground station.




UFD0100


TMUADFF01-P02 LEVEL 1

ADF SYSTEM PRESENTATION - PRINCIPLE



UFD0100



COMPONENTS

The Automatic Direction Finder system is composed oftwo receivers and two antennae.

The ADF system is also connected to:- Navigation Displays (ND) and Digital DistanceRadio Magnetic Indicator (DDRMI) for display.

- Electronic Flight Instrument System (EFIS)panels for control display.

- Flight Management and Guidance Computer (FMGC)for auto-tuning.

- Multipurpose Control Display Units (MCDU) formanual tuning.

- Captain and First Officer Radio ManagementPanels (RMP) for back-up tuning.

- Audio Control Panels (ACP) for ADF audiosignal.

Note: ADF 2 system is optional.




UFD0100


ADF SYSTEM PRESENTATION - COMPONENTSTMUADFF01-P03 LEVEL 1



UFD0100



INDICATING

The Automatic Direction Finder (ADF) systeminformation can be displayed on the Navigation Displays(ND) system and on the Digital Distance Radio MagneticIndicator (DDRMI).On the NDs, depending on the position of the ADFselector switch on the EFIS control panel:

- ADF 1 is represented by a single pointer- ADF 2 is represented by a double pointer.

On the DDRMI, depending on the position of the ADFselector switch:

- ADF 1 is represented by a single pointer- ADF 2 is represented by a double pointer.

Note: Some DDRMIs are not equipped with the ADFcapability.




UFD0100


ADF SYSTEM PRESENTATION - INDICATINGTMUADFF01-P04 LEVEL 1



UFD0100



TMUADFF01 LEVEL 1



UFD0100


34 - NAVIGATION

34-55-00 VOR/MARKER SYSTEMS PRESENTATION

CONTENTS:VOR PrincipleMKR PrincipleComponentsVOR IndicatingMKR Indicating

TMUVORG02 LEVEL 1



UFD0100


VOR/MARKER SYSTEMS PRESENTATION

VOR PRINCIPLE

The Very high frequency Omni-directional Range(VOR)system is a medium-range radio navigation aid.The VOR system receives, decodes and processes bearinginformation from the omni-directional ground station(working frequency range: 108 to 117.95 Mhz)The ground VOR station generates a reference phasesignal and a variable phase signal.The phase difference between these signals, calledbearing, is function of the aircraft position withrespect to the ground station.The bearing is the angle between the Magnetic Northand the ground station/aircraft axis.Furthermore, the VOR station provides a Morseidentification which identifies the station.

TMUVORG02-T01 LEVEL 1



UFD0100


VOR/MARKER SYSTEMS PRESENTATION - VOR PRINCIPLETMUVORG02-P01 LEVEL 1



UFD0100



MRK PRINCIPLE

The MARKER (MKR) system is a radio navigation aid whichindicates the distance between the aircraft and therunway threshold.The MARKER (MKR) system is normally used together withthe ILS system during an ILS approach.




UFD0100


VOR/MARKER SYSTEMS PRESENTATION - MKR PRINCIPLETMUVORG02-P02 LEVEL 1



UFD0100



COMPONENTS

The VOR and MKR systems are composed of two receivers,one marker antenna and one dual VOR antenna.The VOR/MKR system is also connected to:

- Navigation Displays (ND), Primary FlightDisplays (PFD) and VOR/ADF/DME Radio MagneticIndicator (VOR/ADF/DME RMI) for display.

- Electronic Flight Instrument System (EFIS)panels for control display.

- Flight Management and Guidance Computers (FMGC)for auto-tuning.

- Multipurpose Control Display Units (MCDU) formanual tuning.

- Captain and First Officer Radio ManagementPanels (RMP) for back-up tuning.

- Audio Control Panels (ACP) for VOR/MKRaudio signal.




UFD0100


VOR/MARKER SYSTEMS PRESENTATION - COMPONENTSTMUVORG02-P03 LEVEL 1



UFD0100



VOR INDICATING

The indicators show that the aircraft is flying fromthe ground station and is on the right,crossing andthen on the left hand side of the course selected bythe pilot.




UFD0100


VOR/MARKER SYSTEMS PRESENTATION - VOR INDICATING (EXAMPLE)TMUVORG02-P04 LEVEL 1



UFD0100



MKR INDICATING

When the aircraft overflies the Marker, the type ofMarker is display on the PFDs in different colors, andis indicated by an aural identification.




UFD0100


VOR/MARKER SYSTEMS PRESENTATION - MKR INDICATINGTMUVORG02-P05 LEVEL 1



UFD0100



TMUVORG02 LEVEL 1



UFD0100


SYSTEMS ON BOARD MAINTENANCEPAGE 1 of 26


ENGINEERS


1 ON BOARD MAINTENANCE SYSTEMS On board maintenance systems enable the engineer to confirm faults and in some cases go straight to the defective item, thus saving time and money in the maintenance of aircraft. There are many different on board maintenance systems in use on modern aircraft, ranging from a simple magnetic indicator on an LRU, to complex systems that allow engineers to connect laptop computers to down load system parameters and fault data.

1.1 MULTI FUNCTION COMPUTER SYSTEM (MFC) In flight monitoring and ground test capabilities are provided by the MFC system (as fitted to the ATR 72). It consists of two independent computers MFC1 and MFC2. The use of these two computers has meant the removal of a total of 9 redundant LRUs. Each computer includes two independent modules, Module A & B. Each Module receives signals from all the various systems and system controls. They also include a self-test capability so that each module can be tested to ensure it is operating correctly.

1.1.1 FUNCTION After processing the input information, the unit will output to the various systems to: 1. Monitor, control and authorize operation of the aircraft systems.

2. Manage system failures and flight envelope anomalies and command

triggering of associated warning in the "Crew Alerting System" (CAS).

3. Provide readout of BITE memory via a maintenance panel on the flight deck, giving information of any system failures.



ENGINEERS


Figure 1 shows a simplified block diagram of the MFC system.

MFC Block Schematic Diagram Figure 1

MFC 1

INPUTS

OUTPUTS OUTPUTS

INPUTS

PRIMARYSECONDARY

PRIMARYSECONDARY

MFC 1ASTATUS

MFC 1BSTATUS

FAULTACTIVATE

FAULTACTIVATE

MFC 1A MFC 1B

ELECTICALPOWER

ELECTICALPOWER



ENGINEERS


1.1.2 MAINTENANCE PANEL The ATR 72 maintenance panel (located right-hand console), enables the operator to identify faults on the system using a rotary switch and a failure display. The control panel (located on the overhead panel) allows the switching on and fault monitoring of the MFC system. Figure 2 shows the MFC Maintenance and control panels.

MFC Maintenance & Control Panels Figure 2

MFC1A 2A1B 2B

FAULT

OFF

FAULT

OFF

FAULT

OFF

FAULT

OFF

F F F F8 4 2 1

NORMFLT

WOW & L/G

DOORS

BOOTS

NAV

BRKFLTCTL

1

2

3

MISC

MFC

ERS

BITE ADV DISPLAY MFCDATABUS

BITELOADED

PTA/ERS

MAGIND

TEST

MFC CONTROL PANEL (OVERHEAD)

MFC MAINTENANCE PANEL (OVERHEAD)



ENGINEERS


The Maintenance panel has the following functions: Bite Loaded Indicator - Indicates when a fault has been recorded by the maintenance system. System Selector Switch - Normally placed in the NORM FLT position. During Bite Advisory Display - Indicates, through illuminated lights, the code of the failure recorded. Combination of illumination of these lights enables up to 14 failures per system to be coded. PTA/ERS push-button - PTA function (push to advance) enables recorded failures on selected system to be run. At the end of the selected system test FFFF is displayed. It also acts as an "Erase" function; this will clear current faults from the syste Test push-button - Used to check operation of the "BITE LOADED" magnetic indicator. Data Bus connector - Enables the connection of the Maintenance Test Set system to be connected. This enables the down load of all faults onto a Notebook type computer.



ENGINEERS


The failure codes are all listed in the aircraft maintenance manual. Table 1 shows an example of the code/failure relationship.

SYSTEM: WOW/L/G

CODE 8 4 2 1 DEFINITION

1 F Right Main Gear Prime DnLk Prox Switch Fail 2 F Nose Gear Prime DnLk Prox Switch Fail 3 F F Left Main Gear Prime DnLk Prox Switch Fail 4 F Right Main Gear Sec DnLk Prox Switch Fail 5 F F Nose Gear Sec DnLk Prox Switch Fail 6 F F Left Main Gear Sec DnLk Prox Switch Fail 7 F F F Left Main Gear WOW 1 Prox Switch Fail 8 F Nose Gear WOW 1 Prox Switch Fail 9 F F Right Main Gear WOW 1 Prox Switch Fail A F F Left Main Gear WOW 2 Prox Switch Fail B F F F Nose Gear WOW 2 Prox Switch Fail C F F Right Main Gear WOW 2 Prox Switch Fail D F F F E F F F

F F F F End of list for selected system

Failure Codes - De-icing Boots System Table 1



ENGINEERS


1.1.3 BUILT-IN TEST EQUIPMENT (BITE) Large aircraft often incorporate "built-in Test Equipment" (BITE) systems to monitor and detect faults in a variety of aircraft systems. Before BITE systems, faults finding often required the connection of special “Test Equipment” then lengthy tests to establish where the fault lay. Then the rectification by replacing the required Line Replacement Unit (LRU) followed by a functional test to confirm the system serviceability, and finally, the removal of the test equipment. The use of BITE systems reduces the time-spent fault finding and thus eliminates the need for specialist test equipment. The BITE continuously tests the various systems and stores all fault information to be recalled later, either by the flight crew or a maintenance team. Once the appropriate repair has been made, the BITE system can then be used to reset the system for operation. Most BITE systems are capable of isolating system faults with at least 95% probability of success on the first attempt. The introduction of digital systems on the aircraft has made BITE systems possible. Discrete digital signals are used as the code language for BITE systems. The BITE system interprets the various combinations of digital signals to determine a system's status. If an incorrect input value is detected, the BITE system records the fault and displays the information upon request. This information may be by illuminating a number of Light Emitting Diodes (LED's), or, as with modern systems, a display on a CRT or TV display. A complex BITE system is capable of testing thousands of input parameters from several different systems. Most BITE systems perform two types of test programs:

Operational Test Maintenance test

Normal operational checks start with initialization upon switch on of system power supplies.



ENGINEERS


Figure 3 shows the BITE flow sequence.

BITE Flow Diagram Figure 3

POWERUP

RESET

POWERUP

RESET

INITIALIZEINITIALIZE

INPUTINPUT

PROTECTIONPROTECTION

CONTROLCONTROL

OUTPUTOUTPUT

OPERATIONALBITE

OPERATIONALBITE



ENGINEERS


The operational BITE program is designed to check:

Input signals.

Protection circuitry.

Control circuitry.

Output signals.

Operational BITE circuitry. During normal system operation, the BITE monitors a "Watchdog" signal initiated by the BITE program. This watchdog routine detects any hardware failure or excessive signal distortion, which may create an operational fault. If the BITE program detects either of these conditions, it automatically provides isolation of the necessary component, initiates warnings and records the fault in a Non-volatile memory. The maintenance program of the BITE is entered into only when the aircraft is on the ground and the "Maintenance Test" routine is requested. On aircraft fitted with Flight Management System FMS, a more complex BITE system is provided. In the Boeing 737, the FMS BITE provides fast and accurate diagnosis of the main FMS components.



ENGINEERS


Figure 4 shows the Boeing 737 FMS Bite System.

Boeing FMS BITE System Figure 4



ENGINEERS


1.1.4 OPERATION Self-contained In-flight monitoring and ground test capabilities are provided for the main FMS components. Each major FMS component contains comprehensive tests for itself, its sensor inputs, and other interfaces. In-flight data is automatically stored for analysis on the ground through the BITE system. BITE is controlled via the FMS Control Display Unit, CDU. The FMS display will display (in plain English), system status for all systems under test. The operator simply selects from a menu of test options and inputs interactive responses via the CDU. BITE runs the test and provides corrective action diagnostics. The system is designed for line maintenance fault isolation to a single line replacement unit (LRU), within minutes. The BITE system will also carry out system verification; to check interfaces after corrective maintenance action.



ENGINEERS


1.2 DATA LOADING Navigation information required by the aircraft systems is loaded using "Data Loaders". These loaders are capable of downloading thousands of bytes of information into the required system in a matter of seconds. The validity of the current data loaded into an aircraft can be checked using the FMS CDU, which will show the current version, loaded into it. Figure 5 shows a Data Loader as fitted to the Boeing 737

Boeing 737 Data Loader Figure 5

POWER

429 BUSINTERFACE

ON/OFF

SPAREFUSE

LINEFUSE

PROG CHNG COMP RDY XFER R/W FAIL

PROG DATA TRANSFER IN PROGRESSCHNG DATA CHANGE IS REQUIREDCOMP DATA TRANSFER IS COMPLETERDY UNIT READY FOR OPERATIONXFER DATA TRANSFER FAILURER/W UNABLE TO ACCESS DISK DATAFAIL SYSTEM TEST FAILURE

DISK STORAGE DISK STORAGE



ENGINEERS


1.2.1 NAVIGATION DATA BASE The Navigation database (NDB) contains data that describes the environment in which the aircraft operates. The type of information loaded includes:

Approaches.

Country Name.

Waypoints.

Airports.

Runways.

Marker Beacons.

Holding Patterns. This information is used by the Flight Management Computer (FMC), to create flight plans that define the aircraft route from origin to destination. The source data and the NDB are updated on a 28-day cycle that it corresponds to the normal revision cycle for navigation charts. Each update disk contains the data for the current cycle and the next one. This arrangement provides the user with greater flexibility since it is not necessary to load a new disk on a specific day. Each PCMCIA card contains 8 megabytes of storage.



ENGINEERS


1.3 CENTRAL MAINTENANCE COMPUTING SYSTEM (CMCS) The CMCS supports both line and extended maintenance functions through menu selections on the Maintenance Access Terminal (MAT) or Portable Maintenance Access Terminal (PMAT). Other menu selections support special maintenance functions, on-line help and report production. Figure 9 shows the location of the MAT.

Maintenance Access Terminal (MAT) Figure 9

MAT KEYBOARD

MAINTENANCE ACCESSTERMINAL (MAT)

MAT KEYBOARDSLOT

FLIGHT COMPARTMENTREAR RIGHT SIDEWALL



ENGINEERS


The CMCS is used for:

Monitoring the aeroplane’s systems for faults.

Processing fault information.

Supplying maintenance messages. Monitoring flight deck effects (FDE).

Maintenance messages give the engineers detailed fault information to help in troubleshooting. The Aeroplane Condition Monitoring System (ACMS) monitors for any system faults, if a fault is detected, a maintenance message is sent to the CMCS. The CMCS processes the data and shows a maintenance message so the maintenance crew can examine it and find a corrective action.

1.3.1 FLIGHT DECK EFFECT (FDE) FDE inform the flight and ground crews of the conditions relating to the safe operation of the aircraft. The ground crew must find the cause of an FDE to find the corrective action. The FDE data is used along with the aircraft’s maintenance manuals to isolate the fault. The ACMS monitor conditions related to the loss of a system or function. If a condition exists that requires repair or deferral, the ACMS sends FDE data to the AIMS Primary Display System (PDS). The PDS will show the FDE on the MAT and PMAT.

1.3.2 MAINTENANCE ACCESS TERMINAL (MAT) The MAT has a display screen and controls for selecting and viewing fault data. A keyboard is also provided (stored when not in use) which allows certain entries and controls displayed data. The MAT also has a cursor control device, which has a power supply module that receives 115V ac via the “MAINT ACCESS TERMINAL” circuit breaker located on the overhead panel. This PSM then distributes power for the remainder of the MAT. The cursor control device contains the following controls:

Track Ball.

Selection Keys.

Brightness Control.



ENGINEERS


Figure 10 shows the MAT and cursor control device.

MAT & Cursor Control Device

Figure 10

MAT DISPLAY

MAT DUALDISK DRIVE

MAT CURSORCONTROL DEVICE

BRIGHTNESSCONTROL

POWER SUPPLYMODULE

SELECTION KEYS (3)TRACK BALL

CURSOR CONTROLDEVICE



ENGINEERS


Figure 11 shows the MAT display showing FDE data.

MAT Displayed Data Figure 11

LINEMAINTENACE

EXTENDEDMAINTENANCE

OTHERFUNCTIONS HELP REPORT

INBOUND FLIGHT DECK EFFECTS N77701TBC1234 KBFI/KMWH

LEG STATRT WAS 1753Z 07 JUL 00THIS DATA IS FROM LEFT CMCF

FDE: F/D ZONE TEMP CTRL STATUS NOT ACTIVE

Fault Code : 216 011 00 1948z 07JUL00

FDE: CAPT RA FLAG PFD FLAG ACTIVE

Fault Code : 343 311 31 1948z 07JUL00

Maintenance Message: 34-42011 ACTIVE

Approach 1941z 07JUL00

Radio Altimeter Transceiver (left) has an internal fault.

Select text of Maintenance Message, then select theMAINTENANCE MESSAGE DATA button to get moredata.

Flight Deck Effects recorded during the present leg MAINTENANCEMESSAGE DATA

ERASEFAULTGO BACK



ENGINEERS


1.4 PORTABLE MAINTENANCE ACCESS DEVICE (PMAT) The PMAT is stored within the electronics bay and has the same functions as the MAT. There is a PMAT terminal receptacle located on the MAT position. There are also four other PMAT receptacles located throughout the aircraft. These are located:

Electronics Bay. Nose Gear. Right Main Gear Bay. Stabilizer Bay.

Figure 12 shows a PMAT and receptacle.

Portable Maintenance Access terminal (PMAT)

Figure 12

LCDDISPLAY

KEYBOARDDISK

DRIVE

CURSORCONTROL

POWERSWITCH

SELECTIONSWITCHESPMAT

PMAT RECEPTACLE



ENGINEERS


1.5 AEROPLANE CONDITION MONITORING SYSTEM (ACMS) The ACMS (Boeing 777) collects monitors and records data from the aircraft’s system. The data collected by the system is used to produce reports. These reports are used to:

Analyze aeroplane performance.

Analyze trends.

Report significant events.

Troubleshoot faults. Figure 13 shows the layout of the Boeing 777 ACMS.

Boeing 777 ACMS

Figure 13

DATA

AIMS

ACMF PDF

QAR CMCF

TMCF FDCF

FMCF DCMF

DFDAF

ACMS REPORTS

XXXX XX X XX XXXXXXXXXXXXXXXXX XXXX XX XXXXXXXXXX X XXXXXXXXXXXXXXXXXXXX XXXXXXXXXXXXXXXXXXXXXX XXXXXXXXXXXXXX XX XXXXXXXXXXXXXXXX

XXXXXXX X X X XXXXXXXXXXXXXXXXXXXXXXXXXXXX

XXXXXXXXXXXX XXXXXXX

ACMS REPORTS




ACMS REPORTS




AIRPLANE CONDITIONMONITORING SYSTEM

(ACMS)



ENGINEERS


The ACMS receives data from the Aeroplane Conditioning Monitoring Function (ACMF) which is located in the left-hand AIMS cabinet. The ACMS is accessed through formats on the Maintenance Access Terminal (MAT), Portable Maintenance Access Terminal (PMAT) or the side displays on the flight deck. The ACMS can if required be programmed by the user airline to carry out custom features. Figure 14 shows the general arrangement of ACMS.

ACMS (Boeing 777) Figure 14

FLIGHT CONTROLARINC 629 BUS (3)

SYSTEMSARINC 629 BUS (4)

ARINC 429ANALOGDISCRETES

AIRCRAFT

SDU

VHFTX/RX

FLIGHT COMPARTMENTPRINTER

QAR

MAT

PMAT

LHDISPLAY

RHDISPLAY

LEFT HAND AIMS CABINET



ENGINEERS


1.5.1 AIRPLANE CONDITION MONITORING FUNCTION (ACMF) The ACMF is a combination of standard and custom software. The custom software is set to the following functions:

Report Format. Report Content. Triggers.

Triggers are logic equations that detect conditions and cause data to be recorded, e.g. engine exceedances. The ACMF sends data to the following units:

Quick Access Recorder (QAR). Maintenance Access Terminal (MAT). Portable Maintenance Access Terminal (PMAT). MAT or PMAT disk drives (to record data onto diskette). Flight deck Side Displays (SD). Data Communication Management Function (DCMF).

Note: The DCMF is used to send data to the airline base while the aircraft is airborne via either the VHF communication or Satellite communication system. The ACMS collects data to record and sends reports to many output devices. The MAT and PMATs allows the user to see the ACMS data and control the function of the ACMS. Aircraft systems send data into the AIMS cabinet input/output modules on:

Flight Control ARINC 629 Buses. System ARINC 629 Buses. ARINC 429 Buses. Analog Inputs. Discrete Inputs.



ENGINEERS


1.5.2 QUICK ACCESS RECORDER (QAR) The QAR records data sent from the ACMF onto a 3.5 inch 128 MB optical disk and holds 41 hours of data. A spare disk is located within the unit should the active disk become full. Figure 15 shows a QAR and optical disk.

Quick Access Recorder (QAR) Figure 15

The optical disk has a magnetic surface with an infrared laser optically tracking the disk. Data from the ACMF (Core Processing Module, CPM) is received by the QARs CPU. The CPU does a self-test to check the validity of the data and then sends control information to the memory device.

PENNY&

GILES

MADE INU.K.

OPTICAL QAR

PRESSSPARE DISK

EJECT

FAIL

LOW CAPACITY

MAINTENANCE

POWER ON

QUICK ACCESS RECORDER

OPTICAL DISK CARTRIDGE

DISPLAY

DISPLAY



ENGINEERS


The QRA memory device contains two memories:

1. Flash memory (non-volatile).

2. Formatter memory. The flash memory holds configuration data, system data and identification files and sends this data to the formatter. The formatter arranges the received data, then sends it to the cartridge drive circuits. The cartridge drive circuits control the position of the laser tracking recording head. They also write data on and read data from the optical disk. The front keyboard is used to read information from the optical disk and to run functional tests. The CPU also sends data to the 16 bit LCD displays. These displays show:

Stored data. QAR menus.

Test results.

Messages.

The QAR sends data and status to the CPM/COMM in the left AIMS cabinet. The ACMF monitors the data and status.

1.6 AEROPLANE INFORMATION MANAGEMENT SYSTEM (AIMS) The AIMS collects and calculates large quantities of data and manages this data for several integrated aircraft systems. The AIMS has software functions that do all the calculations for each aircraft system. The AIMS has two cabinets, which do the calculations for these systems. Each cabinet contains:

Cabinet Chassis. Four input/output Modules (IOM).

Four Core Processor Modules (CPM).

The IOM and CPM are in the cabinet chassis, which has a backplane data bus and a backplane power bus to distribute data and power to the IOMs and CPMs.



ENGINEERS


The IOMs transfer data between the software functions in the AIMS CPMs and external sources. The CPMs supply the software/hardware to do the calculations. There are four types of CPMs:

1. CPM/COMM – Core Processor Module/Communication.

2. CPM/ACMF - Core Processor Module/Aircraft Condition Monitoring Function.

3. CPM/B - Core Processor Module/Basic.

4. CPM/GG - Core Processor Module/Graphics Generator.

Figure 16 shows the AIMS system (Boeing777).

AIMS System

Figure 16

AIRCRAFT CONDITIONMONITORING SYSTEM

(ACMS)

PRIMARY DISPLAYSYSTEM

(PDS)

CENTRAL MAINTENCECOMPUTING SYSTEM

(CMCS)

FLIGHT MANAGEMENTCOMPUTING SYSTEM

(FMCS)

THRUST MANAGEMENTCOMPUTING SYSTEM

(TMCS)

DATA COMMUNICATIONMANAGEMENT SYSTEM

(DCMS)

AIMS LEFT-HAND CABINET

AIMS RIGHT-HAND CABINET

FLIGHT DATARECORDER SYSTEM

(FDRS)



ENGINEERS


1.6.1 FLIGHT COMPARTMENT PRINTING SYSTEM The flight compartment printer supplies high-speed hard copies of text for the following systems:

Primary Display System (PDS).

Aeroplane Condition Monitoring System (ACMS).

Central Maintenance Computing System (CMCS). The flight compartment printer receives data from the print driver partition of the Data Communication Management Function (DCMF). The DCMF is located within the AIMS. The DCMF prioritises data sent to the printer in the following order:

Flight Deck Communication Function (FDCF) of the DCMS. Central Maintenance Computing Function (CMCF) of the

CMCF. Aeroplane Condition Monitoring Function (ACMF) of the

ACMS. Multi Function Display (MFD).

The printer can print at 300 dots per inch (DPI). It uses a roll of paper, which is 125 feet long and is A4 European Air standard paper. The printer contains all mechanical components and electronics necessary for printer operation. The mechanical components include:

Printer head. Rollers to move paper. Motor and drive system.

The electronic components include:

Power supply module. Processor board. Controller board.



ENGINEERS


Interconnection board



ENGINEERS


Figure 17 shows the flight compartment printer.

Flight Compartment Printer

Figure 17

Controller Board – Receives brightness controls from dimmer controls that drive the lights on the front panel. Processing Board – Processes all inputs for the left AIMS cabinet and changes the data signals to control the thermal printer. Interconnection Board – Controls the flow of data between the processor board and the controller board and the mechanical devices that print three paper.

FAIL PAPER CUT SLEW RESET TEST

TOP VIEW

SIDE VIEW

23 - COMMUNICATIONS

23-73-00 CABIN INTERCOMMUNICATION DATASYSTEM DESIGN PHILOSOPHY

CONTENTS:GeneralPrinciplePassengers FunctionsCrew FunctionsCabin Systems FunctionsMonitoring And Test FunctionsAircraft Systems FunctionsCockpit Controls And Indicating

TMUCI2A02 LEVEL 1



UFD4200


CABIN INTERCOMMUNICATION DATA SYSTEM DESIGN PHILOSOPHY

GENERAL

Changing market demands require flexibility incustomized cabin layouts and optional cabin systems.With the Cabin Intercommunication Data System (CIDS),the operator is able to change the cabin layout withouthardware changes (e.g. cabin loudspeakers,PAX-equipment ...).This can be simply done by entering, on board, newcabin parameters in the software.The CIDS is a microprocessor based system. It monitors,tests, operates and provides control and monitoringof the cabin functions.

PRINCIPLE

To manage various functions, the CIDS has a centralunit, the CIDS DIRECTOR. It is linked to the ForwardAttendant Panel (FAP) for control and monitoring ofthe cabin functions.The Director then communicates, through a bus system,with Decoder Encoder Units (DEUs). The DEUs send (andreceive) information to (and from) the cabin, passengerand crew systems.The Director has interfaces to other aircraft systems.Through a Programming and Test Panel (PTP) the CIDScan be programmed to customer demand.The PTP is also used to test the entire CIDS.

PASSENGER FUNCTIONS

- general cabin illumination control,- passenger address,- passenger call,- passenger lighted signs,- passenger reading light switching.

CREW FUNCTIONS

- cabin and flight crew interphone,- service interphone,- emergency evacuation signalling.

CABIN SYSTEMS FUNCTIONS

- boarding music,- pre-recorded announcement,- lavatory smoke warning,- temperature regulated drain mast system,- emergency lighting.

MONITORING AND TEST FUNCTIONS

- system programming and test,- work light test,- escape slide bottle pressure monitoring,- reading lights test,- extended emergency lighting test.

TMUCI2A02-T01 LEVEL 1



UFD4200


CABIN INTERCOMMUNICATION DATA SYSTEM DESIGN PHILOSOPHYTMUCI2A02-P01 LEVEL 1



UFD4200


CABIN INTERCOMMUNICATION DATA SYSTEM DESIGN PHILOSOPHY

AIRCRAFT SYSTEMS FUNCTIONS

- interface with aircraft systems:e.g. FWC, LGCIU, PRAM, SFCC, etc...

COCKPIT CONTROLS AND INDICATING

- call panel,- evac panel,- NS/FSB panel,- PA handset,- service interphone.

TMUCI2A02-T01 LEVEL 1



UFD4200


23 - COMMUNICATION

23-73-00 CABIN INTERCOMMUNICATION DATASYSTEM PRESENTATION

CONTENTS:GeneralDirectorsType A Decoder Encoder UnitsType B Decoder Encoder UnitsForward Attendant PanelProgramming and Test Panel

TMUCI2B02 LEVEL 1



UFD4200


CABIN INTERCOMMUNICATION DATA SYSTEM PRESENTATION

GENERAL

The CIDS consists of the following components:- the Directors including ON Board ReplaceableMemories (OBRM),

- the Programming and Test Panel (PTP) includingCabin Assignment Module (CAM),

- the Forward Attendant Panel (FAP),- the Additional Attendant Panels,- the Type A Decoder Encoder Units (DEUs A),- the Type B Decoder Encoder Units (DEUs B),- Cockpit equipment,- Cabin equipment.

DIRECTORS

For redundancy, two directors are installed.In normal operation of the CIDS, director 2 is in hotstand-by.Both directors receive the same inputs and perform thesame computations. The outputs of the director in hotstand-by are disabled.The directors are connected through two CIDS bussesto the type A and type B DEUs to carry the variousdata to the cabin equipment.The FAP, PTP and other systems are connected directlyto the directors basically for control, indication andtest of the CIDS functions.

TYPE A DECODER ENCODER UNITS

The type A DEUs provide the interface between thedirectors and the passenger related systems.

TYPE B DECODER ENCODER UNITS

The type B DEUs provide the interface between thedirectors and the attendant and cabin related systems.

FORWARD ATTENDANT PANEL

The Forward Attendant Panel (FAP) is installed at theforward attendant station.From the FAP, the various cabin systems can becontrolled and monitored.

PROGRAMMING AND TEST PANEL

The Programming and Test Panel (PTP) is installed atthe forward attendant station next to the FAP.The PTP contains the Cabin Assignment Module (CAM)which is used to store all information for the actualcabin layout.

TMUCI2B02-T01 LEVEL 1



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CABIN INTERCOMMUNICATION DATA SYSTEM PRESENTATIONTMUCI2B02-P01 LEVEL 1



UFD4200



TMUCI2B02 LEVEL 1



UFD4200


23 - COMMUNICATIONS

23-73-00 CIDS - DIRECTOR/DEU ARCHITECTURE

CONTENTS:DEU ADEU BSelf Examination

TMUCI2C03 LEVEL 3



UFD4200


CIDS - DIRECTOR/DEU ARCHITECTURE

DEU A

Twenty six type A Decoder Encoder Units (DEUs) areinstalled above the windows in the cabin ceiling andclose to the center ceiling for the DEUs in the entrancearea.The type A DEUs are connected to the directors via atop-line data bus (i.e. : two wire twisted and shieldedcable).A broken wire in one top-line bus will only affect thetype A DEUs behind the crack on this bus.The type A DEUs of the other top-line bus will workwithout disturbance.

PASSENGER SIGNSThe passenger signs include NO SMOKING or the optionalNO ELECTRONIC DEVICE lights, FASTEN SEAT BELT lights,NON SMOKER ZONE lights and RETURN TO SEAT lights inthe lavatories.Furthermore, for the PAX call system, the seat rowlights are connected to the type A DEUs.

CABIN LIGHTSThe cabin lights include:

- Entrance area lights,- Lavatory lights,- Attendant lights,- Reading lights,- Cabin fluorescent strip lights.

LOUDSPEAKERSThe loudspeakers are installed in the Passenger ServiceUnit (PSU), in each lavatory and close to the attendantstation.They are all identical and are used for:

- Passenger address announcements,- Call chimes (optional).

PASSENGER CALLPushbuttons are fitted in the PSU above each seat rowand in the lavatories.

READING/LIGHT POWER UNITOne R/L power unit for three reading lights isinstalled in each Passenger Service Unit (PSU).

TMUCI2C03-T01 LEVEL 3



UFD4200


CIDS - DIRECTOR/DEU ARCHITECTURETMUCI2C03-P01 LEVEL 3



UFD4200


CIDS - DIRECTOR/DEU ARCHITECTURE

DEU B

Basically 4 type B DEUs (max. 6 - Optional) areinstalled near the exit doors in the center ceiling.They are connected to the directors via a middle linedata bus.There are two supplementary DEU B mounts installed asa provision.The fig. on the next page shows a typical Type B DEUinterface. It may vary with different locations andwith specific airline requirements.

SLIDE PRESSURE SYSTEM (Optional)The directors receive signals from the bottle pressuresensors via type B DEUs.If the pressure is low, the CIDS CAUTION light on theFAP comes on.

DOOR PRESSURE SYSTEM (Optional)The directors receive signals from the bottle pressuresensors via type B DEUs.If the pressure is low, the CIDS CAUTION light on theFAP comes ON.

CREW INTERPHONE SYSTEMThe crew interphone system enables communicationbetween cockpit crew and cabin attendants and betweeneach attendant station.

NOTE: From each attendant station it is possible tocommunicate with personnel at the serviceinterphone connections.

EPSUsThe Emergency Power Supply Units (EPSUs) are connectedto type B DEUs for the emergency lighting system test.

DRAIN MASTThe directors receive signals from the drain mastcontrol unit via type B DEUs.If the drain mast heater or the control unit fails theCIDS CAUTION light on the FAP comes on.

ATTND AND PANELOne Attendant Indication Panel is installed near eachattendant seat for message purposes.

AREA CALL PANELOne basic and one optional ACP can be connected toeach DEU B.

TMUCI2C03-T02 LEVEL 3



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CIDS - DIRECTOR/SYSTEM ARCHITECTURETMUCI2C03-P02 LEVEL 3



UFD4200


SELF EXAMINATION

A break in one top line data bus:A - Disables all DEUs.B - Affects only type B DEUs.C - Only affects the type A DEUs behind the

crack on this bus.

The Area Call Panel (ACP) is connected to:A - The directors directly.B - The type A DEUs.C - The type B DEUs.

TMUCI2C03 LEVEL 3



UFD4200


23 - COMMUNICATIONS

23-30-00 PASSENGER ENTERTAINMENT SYSTEMPRESENTATION

CONTENTS:GeneralPESPES VideoPRAM

TMU23EA01 LEVEL 3



UFD4200


PASSENGER ENTERTAINMENT SYSTEM PRESENTATION

GENERAL

The passenger address and entertainment systemcomprises the following basic functions:

- Passenger Entertainment System (PES),- Passenger Entertainment System Video (PESvideo),

- Pre-Recorded Announcements and boarding Musicsystem (PRAM).

The PES comprises the PES music, the passenger addressand the passenger service.

PES

The PES transmits pre-recorded music programs,passenger address information, video and video soundsto the passengers.The audio signals can be heard through headphonesconnected to the Passenger Control Units (PCU).

The PCU allows several music channels and video audiochannels to be selected and the volume to be adjusted.

The PCU also allows the reading lights and passengercalls to be remotely controlled through the PassengerService System (PSS).

All pre-recorded announcements (video and sound) andthe passenger address messages, heard in the headphonesthrough the PCU, have priority over the music and videosound entertainment channels.

The anouncements and passenger address messages arealso broadcast through the passenger addressloudspeakers, via the CIDS.

The PES audio reproducers supply music channels to theMain Multiplexer and boarding music channels to theCIDS director.The CIDS broadcasts the boarding music through thepassenger address loudspeakers.

Boarding Music (BGM) channel and volume control isperformed on the Forward Attendant Panel (FAP).

The Main Multiplexer is connected to the CFDIU toensure the passenger entertainment BITE function.

PES VIDEO

The PES video shows pre-recorded video movies and videoannouncements through different display units in thepassenger compartment.The video sound is transmitted to the Main Multiplexerand to the CIDS.

Therefore video sounds can be heard from the headsetthrough the PCU or from the cabin passenger addressloudspeakers.

The in-seat video display units are supplied throughthe Main Multiplexer.

TMU23EA01-T01 LEVEL 3



UFD4200


PASSENGER ENTERTAINMENT SYSTEM PRESENTATION - SCHEMATIC

TMU23EA01-P01 LEVEL 3



UFD4200


PASSENGER ENTERTAINMENT SYSTEM PRESENTATION

PRAM

The PRAM is an audio tape reproducer which containspre-recorded announcements and boarding music suppliedto the CIDS director.The announcements are also sent to the MainMultiplexer.

The PRAM is controlled from the FAP.

TMU23EA01-T02 LEVEL 3



UFD4200


SYSTEMS INFORMATION SYSTEM PAGE 1 of 9


ENGINEERS


1.1 ARINC COMMUNICATION, ADDRESSING & REPORTING SYSTEM The ACARS is a digital data link for either ground-air or air-ground connections. The system reduces the flight crew’s workload because it transmits routine reports automatically and simplifies other reporting. The ACARS network is made up of three sections:

Airborne System.

Ground Network.

Airline Operations Centre.

The airborne system has an ACARS Management Computer (MU) which manages the incoming and outgoing messages, and a Multi-Purpose Interactive Display Unit (MPIDU) which is used by the flight crew to interface with the ACARS system. A printer can also be installed to allow incoming messages to be printed for future reference. ACARS operates using the VHF 3 communications system on a frequency of 131.55 MHz. Since ACARS only operates on one frequency, all transmitted messages must be as short as possible. To achieve a short message, a special code block using a maximum of 220 characters is transmitted in a digital format. If longer messages are required, more than one block will be transmitted. Each ACARS message takes approximately 1 second of airtime to be sent. Sending and receiving data over the ACARS network reduces the number of voice contacts required on any one flight, thereby reducing communication workload. ACARS operates in two modes:

Demand Mode.

Polled Mode.



ENGINEERS


1.1.1 DEMAND MODE The demand mode allows the flight crew of airborne equipment to initiate communications. To transmit a message, the MU determines if the ACARS channel is free from other communications from other ACARS, if it is clear, the message is sent. If the ACARS VHF channel is busy, then the MU waits until the frequency is available. The ground station sends a reply to the message transmitted from the aircraft. If an error reply or no reply is received, the MU continues to transmit the message at the next opportunity. After six attempts (and failures), the airborne equipment notifies the flight crew. 1.1.2 POLLED MODE In the polled mode, the ACARS only operates when interrogated by the ground facility. The ground facility routinely uplinks “questions” to the aircraft equipment and when a channel is free the MU responds with a transmitted message. The MU organises and formats flight data prior to transmission and upon request, the flight information is transmitted to the ground facility. The ground station receives and relays messages or reports to the ARINC ACARS Control Centre. The control centre sorts the messages and sends them to the operator's control centre (several airlines participate in the ACARS network). The ACARS also reduces the congestion of the VHF communication channels because transmissions of ACARS take fractions of a second while the same report/message in aural form may have taken in excess of ten seconds.



ENGINEERS


ACARS may be connected to other airplane systems such as the “Digital Flight Data Acquisition Unit” (DFDAU). The DFDAU collects data from many of the aircraft’s systems such as Air Data Computer, Navigation and Engine monitoring systems, and in turn makes this data available to ACARS. More recent ACARS installations have been connected to the “Flight Management Computer” (FMC), permitting flight plan updates, predicated wind data, take-off data and position reports to be sent over the ACARS network. The ACARS in use vary greatly from one airline to another and are tailored to meet each airline’s operational needs. When satellite communication systems are adopted, ACARS will take on a truly global aspect. Figure 1 shows an ACARS network.

ACARS Network

Figure 1

MAINTENANCEOPERATIONS

FLIGHTOPERATIONS

PASSENGERSERVICES

AIRLINECOMPUTER

SYSTEM

A/C SYSTEMS ACARS VHF 3

TRANSMISSIONNETWORK

VHF TRANSMITTER/RECEIVER



ENGINEERS


1.1.3 DESCRIPTION The ACARS is operational as soon as the electrical power is supplied and does not have an ON/OFF switch. The ACARS has the following components:

1. ACARS Management Unit (MU).

2. Mu

lti-Purpose Interactive Display Unit (MPIDU).

3. Ide

nt plug.

4. Program pins.

5. Th

ermal Printer. 1.1.4 MANAGEMENT UNIT (MU) The Management Unit (MU) converts the data from and to the VHF-COMM. Requests from ground-stations for communication or reports go from the MU to the MIDU or Flight Data Acquisition Unit (FDAU). Most of the reports are generated in the FDAU. The MU itself makes the report. The unit uses information from the FWS for this message (parking brake and ground/flight for example). The interface wiring between MU and FDAU/MIDU is ARINC 429. The MU codes the messages for VHF-COMM. The messages contain the aircraft's registration and the airline code. This information comes from the ident plug. The MU also decodes the messages from the VHF-COMM. When there is a message for the crew, the MIDU shows a message annunciation, while the MU also makes a discrete for the Flight Warning System (FWS) to make an alert. The VHF-COMM can be used for data transmissions for the ACARS or normal communication. You can select the voice or data mode on the MIDU. 1.1.5 MULTI-PURPOSE INTERACTIVE DISPLAY UNIT (MPIDU)



ENGINEERS


Displays messages, reports and communication requests to the crew. It incorporates touch-screen control in lieu of external pushbuttons and knobs. The touch-screen control is made possible by the use of infrared sensors along the sides of the display. Control inputs are made from menus displayed on the MIDU.



ENGINEERS


Figure 2 show the display layout of the MIDU.

Multipurpose Interactive Display Unit (MIDU) Figure 2

Collins

DATA

LINK

1 2 3

4 5 6

7 8 9

0

CLR RET DEL

FLT : 0123 0008

IN SENDDFDAU FAILNUMERIC ENTRY 13 : 02 : 58



ENGINEERS


1.1.6 ACARS PRINTER A thermal printer is provided for the printing of ACARS messages. Operation of the printer is optional as all printed information can be viewed on the MIDU. Weather report information is sent directly to the printer from the ACARS ground-station. The printer uses rolls of 4.25” thermal paper. A red stripe appears along the edge of the paper when the supply is low. Figure 3 shows the ACARS Printer.

ACARS Thermal Printer Figure 3

SELFTEST

PPRADV

PWRON

ALERTRESET

PTRBUSY

PUSHBUTTONCONTROLS

DOOR LOCKINGSCREW

PAPER LOADINGDOOR

PAPER CUTTINGEDGE



ENGINEERS


1.1.7 PRINTER OPERATION The printer is normally located aft of the centre pedestal and has a “Self Test” feature for pre-flight operational testing. • SELF TEST PUSH BUTTON: Pushing the “Self Test” pushbutton activates a

printer self test which prints the following: THE QUICK BROWN FOX JUMPED OVER THE 1 2 3 4 5 6 7 8 9 0 LAZY DOGS

• PPR ADV PUSHBUTTON: Used to advance the paper. • DOOR LOCKING SCREW: Secures the paper loading door shut. • PWR ON LIGHT: Illuminates when power is applied to the printer. • ALERT RESET: Resets the printer if an alert is detected. • PTR BUSY LIGHT: Illuminates amber when the printer is printing. Remains

ON until paper advance is complete. • PAPER LOADING DOOR: Printer paper roll is replaced via opening this door. • PAPER CUTTING EDGE: Allows for smooth paper cutting when a printed

message is removed from the printer. ACARS communications are accomplished via the ARINC network and the VHF 3 transceiver. VHF 3 is dedicated to this purpose and is automatically controlled by the ACARS frequency of 131.55 MHz and is tuned remotely by the ground stations if frequency change is necessary.



ENGINEERS


Figure 4 shows a block schematic of the ACARS.

ACARS Schematic Diagram Figure 4

MANAGEMENTUNIT

VHF 3TX/RX

FLIGHT DATAACQUISTION UNIT

AIRCRAFTSYSTEMS

Collins

DATA

LINK

1 2 3

4 5 6

7 8 9

0

CLR RET DEL

FLT : 0123 0008

IN SENDDFDAU FAIL

NUMERIC ENTRY 13 : 02 : 58

MULTIPURPOSE INTERACTIVEDISPLAY UNIT

THERMAL PRINTER

VHF 3ANTENNA

General

The Electronic Flight Bag (EFB) lets the flight crew access to theelectronic flight operation data, general purpose computing andcommunications.

Abbreviations and Acronyms

* AC - advisory circular (FAA)* ACARS - aircraft communication addressing and reportingsystem

* ADC - application dispatch controller* AIMS - airplane information management system* API - application program interface* APU - auxiliary power unit* ARINC - aeronautical radio, incorporated* BCA - Boeing commercial airplanes* BEGGS - Boeing e-plane ground support system* BIT - Built-in test* BITE - built-in test equipment* CAM - CAT application module (e-Plane)* CAT - common administrative tool (e-Plane)* CCA - circuit card assembly* CCD - cursor control device* CDROM - compact disk read only memory* CIU - camera interface unit* CMS - cabin management system* CPU - central processing unit* CRC - cyclic redundancy check* CSS - cabin surveillance system* DDM - distributed data management* DFDAU - digital flight data acquisition unit* DFIM - DDM flight-bag interface module (application)

* DHCP - dynamic host configuration protocol* DNS - domain name server* DSPL - display* DU - display unit* ECMF - eplane communications management function* EFB - electronic flight bag* EFIS - electronic flight instrument system* EICAS - engine indicating and crew alert system* EPT - electronic-enabled portable terminal* EU - electronic unit* FAA - federal aviation administration* FAR - federal aviation regulation* FDEVSS - flight deck entry video surveillance system* FIND - find identification of network devices* FTP - file transfer protocol (application)* FTS - file transfer service (application)* GPS - global positioning system* ICAO - international civil aviation organization* IO - input output* HST - high speed transciever* JAA - joint airworthiness authorities* LAN - local area network* LRU - line replaceable unit* LSAP - loadable software airplane parts* LSK - line select keyEGP 101-999*MAU - microwave antenna unitEGP 101-999; EGP 001-005 POST SB 777-46-0026*MMR - multi-mode receiver* NIC - network interface card* NOTAM - notice to airmen (FAA)* NTP - network time protocol

ELECTRONIC FLIGHT BAG - INTRODUCTION

EFFECTIVITYEGP 101-999; EGP 001-005 POST SB 777-46-0026 46-11-00

Page 5Jan 05/2010D633W101-EGP

BOEING PROPRIETARY - Copyright# - Unpublished Work - See title page for details.

777 AIRCRAFT MAINTENANCE MANUAL

46-11-00-002

* OAS - operationally approved software (FAA)* OS - operating system* PDL - portable data loader* PMAT - portable maintenance access terminal* PPPoE - point to point protocol over Ethernet* PWR - powerEGP 101-999* SAR - staging area reporting (application)EGP 101-999; EGP 001-005 POST SB 777-46-0026* SATCOM - satellite communication* SMF - security management function* TPA - taxi position awareness* TSO - technical service orderEGP 101-999* TWLU - terminal wireless LAN unitEGP 101-999; EGP 001-005 POST SB 777-46-0026* VDC - volts direct current* VPN - virtual private network*WPM - windows print manager* XFR - transfer






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General

The electronic flight bag (EFB) has two display units (DU) andtwo supporting electronics units (EU). The captain’s EFB systemis independent from the first officer’s EFB system. Each EFBsystem consists of a DU and an EU.

Description

The EFB provides the flight crew with a paperless flight deckenvironment and enhance the quality of information available tothe crew.

The flight crew interacts with the EFB via the display unit (DU)either by pushing the buttons on the DU bezel, or by using atouch-screen that is a feature of certain applications (example:electronic logbook).

In addition, the flight crew can also make use of the cursorcontrol device (CCD) and the portable keyboard (optional).

The electronic Unit (EU) has these functions:

* Process aircraft interface signals* Program memory (hard-disk drive)* Ethernet communications network* Video input processing* Convert the digital video output signal to the DU* Supply 28V DC power to the onside DU

EFB - GENERAL DESCRIPTION





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