Avionics Systems
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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
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 semiconductors 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.
ELECTRON
4
4
4
4
4
4
4
4
4
4
4
4
HOLE
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. ELECTRONIC FUNDAMENTALS
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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. ELECTRONS
HOLES
SEMICONDUCTOR MATERIAL
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.
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4
4
4
4
4
5
4
DONOR IMPURITY ATOM 4
5
4
(a)
EXTRA ELECTRON
5
4
ACCEPTOR IMPURITY ATOM
HOLE 4
4
4
3
4
3
4
4
4
4
3
4
(b)
Extrinsic Semiconductor Figure 3
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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 +2 0V 0V
0V P.D.
SEMICONDUCTOR MAT ERIAL
CURRENT FLOW
+10V
20V
+11V
+9V
+11V
POSITIVE CHARGE CARRIERS (HOLES)
+9V
+9V
NEGATIVE CHARGE CARRIERS (ELECTRONS)
+11V
ELECTRONIC FUNDAMENTALS
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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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HALL EFFECT SENSOR
GENERATOR CONTROL UNIT
STARTER GENERATOR
HALL EFFECT SENSOR
CURRENT MEASURING
TO DISTRIBUTION
Hall Effect Sensors Figure 5
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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.
DEPLETION LAYER POSITIVE IONS
NEGATIVE IONS
N-TYPE
P-TYPE
ELECTRONS
HOLES
Junction Diode Figure 6 ELECTRONIC FUNDAMENTALS
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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 ELECTRONIC FUNDAMENTALS
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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.
_
+ ANODE
CATHODE
FORWARD BIASED
REVERSED BIASED
NO CURRENT
CURRENT FLOW
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).
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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.
mA 200 FORWARD BIAS
150 100 50
VOLTS -200V
-150
-100
-50V
0.25V
0.5V
0.75V
1V
-0.02 -0.04 REVERSED BIAS
-0.06 -0.08 µA
Silicon Diode Characteristics Figure 9
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mA 200 FORWARD BIAS
150 100 50
VOLTS -200V
-150
-100
-50V
0.25V
0.5V
0.75V
1V
50 100 REVERSED BIAS
150 200 µA
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.
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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.
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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.
A.C. INPUT
D.C. OUTPUT
+ 0 -
+ INPUT
0
OUTPUT
-
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.
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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.
250mA @ 200V
1A @ 1000V
1000A @ 2500V
10A @ 400V 1A @ 1500V
Silicon Rectifier Diodes Figure 12
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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.
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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
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Figure 15 shows the circuit for a “Full-Wave bridge” rectifier.
Full-Wave Bridge rectifier Figure 15
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1.10 THE LIGHT-EMITTING DIODE (LED) LEDs are made from a semi- conductor 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.
CIRCUIT SYMBOL EMITS LIGHT
EARTH
+5V
DIODE IS FORWARD BIASED
ON
EARTH
+5V
DIODE IS REVERSED BIASED
OFF
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.
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1.11 THE PHOTO CONDUCTIVE DIODE This device is a normal PN junction with a transparent case or window. All semiconductor 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 ELECTRONIC FUNDAMENTALS
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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.
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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.
LOW RESISTANCE
FLUKE 23
SERIES
MUL TIMET ER
000.23 0
OFF
10
20
O HM S
30
V
P
V 300 m V
N
Ω
P RE S S RANGE
A
A UT ORANGE
A
VΩ
10A
CATHODE ! 300 mA
10 00V 75 0V
COM
FUSED
SYMBOL
STRUCTURE
Testing Diodes Figure 27
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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.
COLLECTOR
C
N BASE
P N EMITTER
CIRCUIT SYMBOL
B
N OT P OINTING IN
E
THE NPN TRANSISTOR
COLLECTOR
C
P BASE
N
CIRCUIT SYMBOL
B
P E
EMITTER
THE PNP TRANSISTOR
PNP & NPN Transistors Figure 1 ELECTRONIC FUNDAMENTALS
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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 Ptype 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).
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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.
C
N - TYPE
B P - TYPE
N - TYPE
E DIODE MODEL
NPN Transistor & Diode Circuit Figure 2
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Figure 3 shows a simple transistor circuit using electron flow to explain the operation.
IC HIGH (99%)
C IB LOW (1%)
B
E IE HIGH (100%)
NPN Transistor Operation Figure 3
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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
LAMP
SOLENOID ANALOGY
E
C B
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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).
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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.
C E
C
B OUTPUT INPUT
INPUT
E
B
COMMON BASE
COMMON EMITTER
E B OUTPUT C
INPUT
COMMON COLLECTOR
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.
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OUTPUT
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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
Input B
Output X3
Input A = Input B
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A combined logic circuit that would carry out the function is shown at Figure
A
B
X1
X2
X3
TRUTH TABLE
A X1 (AB)
11. Combination Logic Circuit Figure 11
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1.5 LINEAR (OR ANALOGUE) IC Figure 12 shows the type of analogue signal handled by the Linear Integrated Circuit.
0
TIME
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.
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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.
POWER SUPPLY (+)
INVERTING INPUT
2
7 8 VOLTAGE OUTPUT
6 NON-INVERTING INPUT
1 3 5
4 POWER SUPPLY (–)
V–
NON-INVERTING INPUT
INVERTING INPUT
4
3
2
1
5
6
7
8
VOLTAGE
V+
GROUND
OUTPUT
Op Amp and Circuit Figure 2
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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 noninverting 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.
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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.
+12V
GAIN = 400 VOUT 5.88V 5.87V ZERO LINE
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:……………………….
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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.
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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.
BASE BOARD
FRONT
Printed Circuit Board Figure 1
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CIRCUIT MODULE DESIGNATION (E.G. SIGNAL SELECTOR)
IC1
IC2
IC3
IC4
IC5
CIRCUIT REFERENCE C2
INTEGRATED CIRCUIT CHIPS
IC6
REAR
FINGER OR EDGE CONNECTOR
Printed Circuit Board Figure 2
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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”.
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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.
DEMAND INPUT TRANSDUCER
DEMAND SIGNAL AMP
RESPONSE
LOAD
MOTOR
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.
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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.
ERROR DETECTOR INPUT TRANSDUCER
AMP
SERVO MOTOR
LOAD
ERROR SIGNAL
POSITION FEEDBACK
OUTPUT POSITION TRANSDUCER
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.
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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'.
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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.
CONTROL ELEMENT
ERROR DETECTOR
SERVO MOTOR
LOAD
VELOCITY FEEDBACK
POSITIONAL FEEDBACK
TACHO GEN
FEEDBACK ELEMENT
Positional Feedback Figure 3
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Figure 4 shows a R-Pot & Helical Potentiometer
Ei
θi
PROPORTIONAL TO
θi
E
E
R-POT
Ei PROPORTIONAL TO
θi
θi
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.
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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.
PRIMARY COIL
L3
IRON CORE CONNECTED TO MECHANICAL INPUT
L1
R
L2
R
T
S
S
T
2. ROTATED CLOCKWISE
1. ZERO POSITION
R
S
T
3. ROTATED COUNTER CLOCKWISE
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. ELECTRONIC FUNDAMENTALS
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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.
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1.6 CAPACITANCE TRANSMITTER An example of a capacitance transmitter can be seen in a simple fuel gauging system as in Figure 6. TANK UNIT
EMPTY
IS
LOOP A
IB
LOOP B
REF C FULL
2 - PHASE MOTOR
DISCRIMINATION STAGE AMPLIFIER STAGE
INDICATOR
REF PHASE
AMPLIFIER UNIT
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.
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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.
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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.
S1
S3
INPUT SHAFT
S1
S2
S3
OUTPUT SHAFT
S2
CIRCUIT SYMBOL STATOR FIELD S1
ROTOR FIELD
R1
S2
R2 S3
CURRENT FLOW
Torque Synchro Figure 9
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Figure 10 shows the construction of a torque synchro.
SHAFT
BEARING
STATOR WINDINGS COILS
SHELL CORE
SLIP RINGS
LEADS TO SLIP RINGS STATOR LEADS
LOWER END CAP
ROTOR
STATOR
COMPLETE ASEMBLY
Torque Synchro Construction Figure 10
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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
S1
S1
CT
CX A.C. SUPPLY
S2
S3
S2
S3
A.C. SUPPLY
INPUT SHAFT
M
SERVO MOTOR
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.
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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.
APPLIED VOLTAGE
ANTI-CLOCKWISE ROTATION VOLTAGE OUT-OF-PHASE
CLOCKWISE ROTATION VOLTAGE IN-PHASE
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.
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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.
STATOR S1
S3
R1
R3
R2
ROTOR
CIRCUIT SYMBOL
R1
S2
R2 R3
S1 S2 S3
Differential Synchro Figure 13
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Figure 14 shows the construction of a differential synchro
STATOR CONNECTIONS
STATOR WINDINGS ROTOR ASSEMBLY STATOR ASSEMBLY ROTOR COILS SKEW CUT TO ENABLE SMOOTHER RUNNING
Differential Synchro Construction Figure 14
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1.11 TORQUE DIFFERENTIAL SYNCHRO Figure 15 shows a differential synchro system set up for the SUBTRACTION of two inputs.
60º 15º 60º
45º
TX
45º
TR TDX
INPUT SHAFT 60º
INPUT SHAFT 15º
OUTPUT SHAFT θ1 – θ2
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.
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1.12 CONTROL DIFFERENTIAL SYNCHRO Figure 16 illustrates a control differential synchro system.
CX
CDX
CT
ERROR SIGNAL
INPUT SHAFTθ1
INPUT SHAFTθ2
OUTPUT SHAFT θ1 – θ2
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.
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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 CO-ORDINATES = r/θ CARTESIAN CO-ORDINATES X = r COS θ CARTESIAN CO-ORDINATES Y = r SIN θ
X
r
Y
θ Polar & Cartesian Co-ordinates Figure 17
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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.
S1
R1 R3 S3
S4
R4 R2 ROTOR
INPUT SHAFT
S2
STATOR
R1
S1
R2
S2
a R3
R4
S3
S4
b 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.
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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.
MAX VOLTS
R ROTOR
FLUX
STATOR
R1
S1
R2
S2
R COS θ
R3
R4
S3
θ
S4
90º
NO VOLTS
θ
180º
270º
360º
R SIN θ
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 ELECTRONIC FUNDAMENTALS
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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.
R1 S1
S2 R2
SM
VX = r COS θ S3
S4
VY = r SIN θ
R3
R4
θ
θ
TO LOAD
VY 2 + VX2 S4
S2
S3
R3
S1
R1
CIRCUIT SYMBOL R4
R2
Cartesian to Polar Co-ordinates Figure 20
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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.
A.C. EXCITATION SUPPLY
RESULTANT WAVEFORM
a
b
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’ELECTRONIC FUNDAMENTALS
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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.
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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.
AIRSPEED INDICATOR
GYRO HORIZON
ALTIMETER
DIRECTION INDICATOR
VERTICAL SPEED INDICATOR
TURN & BANK INDICATOR
BASIC 6 GROUPING
COMBINED AIRSPEED INDICATOR
RADIO MAGNETIC INDICATOR
ATTITUDE DIRECTOR INDICATOR
ALTIMETER
VERTICAL SPEED INDICATOR
HORIZONTAL SITUATION INDICATOR
BASIC T GROUPING
Basic “Six” and “T” Flight Instrument Grouping Figure 2
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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.
EFIS PFD
EFIS ND
ECAM ENGINE WARNINGS
EFIS ND
EFIS PFD
ECAM SYSTEMS
Flight Deck Electronic Instrumentation Layout Figure 3
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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.
AIR DATA
ATT HDG NORM CAPT 3
E/S DMC
NORM F/O 3
CAPT 3
ECAM / ND XFR NORM
NORM F/O 3
CAPT 3
F/O 3
CAPT
F/O
A320 EIS Switching Panel Figure 4
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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.
DISPLAY MANAGEMENT SYSTEM DMS No 1
DISPLAY MANAGEMENT SYSTEM DMS No 3
DISPLAY MANAGEMENT SYSTEM DMS No 2
Electronic Instrument System (EIS) Figure 5
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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 POSITION WEIGHT
26 25 24 23 22 21 20 64 32 16 8 4 2 1 1 1 0 1 0 0 1 64 + 32 + 8 + 1 = 105(10)
Therefore:
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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 210
512 256 29 28
128 27
64 26
32 25
16 24
8 23
4 22
2 21
1 20
WEIGHT 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) DIGITAL TECHNIQUES
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0 2 1 2 3 2 7 2 14 2 28
R = 1 (MSD) R=1 R=1 R=0 R = 0 (RIGHT MOST DIGIT OR LSD)
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 x2 0.750 .750 x2 1.500 .500 x2 1.000
THUS:
2-1 position should hold a zero 2-2 position should hold a one 2-3 position should hold a one .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:
18.205(10) = ?(2) 0 2 1 2 2 2 4 2 9 2 18
R=1 R=0 R=0 R=1 R=0
4
(2 ) (23) (22) (21) (20)
.
.205 x2 .410 x2 820 x2 1.640 x2 1.280
2-1 is 0 2-2 is 0 2-3 is 1 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 85
4096 84
512 83
64 82
8 81
1 80
WEIGHT 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 8 2 8 21 8 169
R = 2 R = 5 R = 1
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.
Example:
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
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. DIGITAL TECHNIQUES
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Example:
88.49(10) = ?(8) R = 1 (82) R = 3 (81) R = 0 (80)
0 8 1 8 11 8 88
.49 x8 3.92 8-1 is a 3 x8 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 0 1 2 3 4 5 6 7
Binary 000 001 010 011 100 101 110 111
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:
Therefore:
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715(8) = ?(2) 7 1 5 111 001 101 715(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 3
Therefore:
011 3
100 4
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 1 0
Therefore:
111. 010 110 7. 2 6
100 4
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) 1 001
or
3 011
7 111
. .
0 000
5 101
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 0
DECIMAL
1 1
2 2
3 3
4 4
5 5
6 6
7 7
8 8
9 9
A B C D E F 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.
6AF(16) = ?(10)
6 x 162 + A x 161 + F x 160 6 x 256 + 10 x 16 + 15 x 1. 1536 + 160 + 15
= = = 1711(10)
Decimal-Hexadecimal Example 1: 108(10) = ?(16) 0 16 13 16 208
R = 13 R = 0
(equivalent to D)
Note: Each remainder must be represented by one hexadecimal digit. Therefore:
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208(10) = D0(16)
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Decimal-Hexadecimal Example 2: 1834(10) = ?(16) 0 16 7 16 114 16 1834 16 23 16 74 64
R = 7 R = 2 R = 10
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 0000 0001 0010 0011 0100 0101 0110 0111 1000 1001 1010 1011 1100 1101 1110 1111
HEXADECIMAL 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 4
1101 D
1001101(2) = 4D(16) CBF(16 = ?(2)
C 1100
B 1011
F 1111
CBF(16) = 110010111111(2)
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Fractions are handled in the same manner: 1101110.01111(2) = ?(16) 0110 6
1110. 0111 1000 E .
7
8
1101110.01111(2) = 6E. 78(16)
Therefore:
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 1 2 3 4 5 6 7 8 9
0000 0001 0010 0011 0100 0101 0110 0111 1000 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 DIGITAL TECHNIQUES
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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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ENGINEERS
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.
A
+5V POSITION 2
POSITION 1
+5V O/P A +0V TIME
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). DIGITAL TECHNIQUES
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+5V
A
O/P A +5V +0V TIME
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.
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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.
ANALOGUE SIGNAL
8 7 6 5 4 3 2 1 0 0 0
0 0 1
0 1 0
0 1 1
1 0 0
1 0 1
1 1 0
1 1 1
3 BIT WORD
DIGITAL SIGNAL
Digital Representation of Analogue Signals Figure 3
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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 nbit 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.
+VE VREF
+
VOUT
VIN
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.
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Figure 5 shows the resultant digital waveforms from an analogue input signal using an Op Amp comparator.
VREF VIN
0
+VMAX VOUT 0 -VMAX WHEN VIN < VREF THEN VOUT = -V MAX WHEN VIN > VREF THEN VOUT = +V MAX
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.
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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.
A B
LSB
C
X
D
E F
Y
G
Z MSB
Encoder Circuit Figure 6
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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.
ANALOGUE O/P SIGNAL
16 14 12 10 8 6 4 2
0 0 0 0
0 0 0 1
0 0 1 0
0 0 1 1
0 1 0 0
0 1 0 1
0 1 1 0
0 1 1 1
1 0 0 0
1 0 0 1
1 0 1 0
1 0 1 1
1 1 0 0
1 1 0 1
1 1 1 0
1 1 1 1
DIGITAL I/P SIGNAL
DAC Output Figure 8
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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.
MSB
R
4 B I T
2R
V OUT
W O R D
4R
LSB
8R
DAC Weighted Circuit Figure 9
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Figure 10 shows a Digital to analogue converter.
V REF S1
MSB
S2
R
2R -
4 BIT DIGITAL INPUT
S3
S4
4R
+
ANALOGUE OUTPUT VOLTAGE
8R
LSB 0V
Digital – Analogue Converter Figure 10
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ENGINEERS
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.
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Figure 1 shows ARINC bus interconnections. DATA INPUT
DATA INPUT
ARINC 429 BUS TWISTED AND SHIELDED WIRES
TX
ARINC 429 TRANSMITTER
RX
INFORMATION FLOW
ARINC 429 RECEIVER
RX
ARINC Bus Interconnection Figure 1
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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.
TINNED COPPER CONDUCTORS
DATA BUS CABLE “B” DATA BUS CABLE “A”
ETFE TEFZEL® INSULATION
ETFE TEFZEL® JACKET
TINNED COPPER BRAID SHIELD
Data Bus (Twisted Pair) Cable Figure 2
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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.
PAUSE BETWEEN DIFFERENT TYPES OF DATA BEING TRANSMITTED
DATA WORD 32 BITS
DATA WORD 32 BITS
DATA WORD 32 BITS
INDICATED AIR SPEED (IAS) TRANSMITTED EITHER: 12 - 14 KBYTES/SEC - LOW SPEED 100 KBYTES/SEC - HIGH SPEED TOTAL AIR TEMPERATURE (TAT) TRANSMITTED EITHER: 12 - 14 KBYTES/SEC - LOW SPEED 100 KBYTES/SEC - HIGH SPEED
ARINC 429 Data Word Formats Figure 3
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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.
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32 31 - 29
28 - - - - - - - - - - - - - - - 11
10 / 9
8 ------ 1
DATAWORD LABEL 8 BITS - OCTAL 000 - 377
PARITY BIT EITHER ODD/EVEN
DATA FIELD 18 BITS BINARY CODED DECIMAL (BCD) OR BINARY FORMAT (BNR) OR DISCRETE FORMAT
SOURCE DESTINATION IDENTIFIER (SDI 0 0 - ALL SYSTEMS 0 1 - SYSTEM 1 1 0 - SYSTEM 2 1 1 - SYSTEM 3
SIGN & STATUS MATRIX (SMM) MEANING RELATED TO FORMAT
32 Dataword Format Figure 4
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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.
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
0 0 01100001
1 1 0 0 1 1 0 1 0 1 0 0 1
DATA FIELD PARITY WORD LABEL SIGN STATUS MATRIX
SOURCE DESTINATION IDENTIFIER
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.
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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. RETURN TO ZERO (RZ) FORMAT
HIGH +10v
LINE A TO B
NULL
1
1
1 0
0
3
4
1
1
1
0
0
0
0
27
28
29
1 0
LOW -10v
LINE A TO GROUND
LINE B TO GROUND
+5v 0 -5v
+5v 0 -5v
1
2
5
6
7
8
30
31
ARINC 429 Dataword Format Figure 6
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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.
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PARALLEL DATA
SERIAL DATA TRANSMISSION DATA TRANSFER 00110
DEMULTIPLEXER
1 2 3 4 5 6 7 8 9 10 11 12
MULTIPLEXER
0 1 1 0 0
PARALLEL DATA
BIT NUMBER
1 2 3 4 5 6 7 8 9 10 11 12
TO CENTRAL CONTROL UNIT
Data Transfer System Figure 7
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The MUX circuit operation is shown in Figure 8.
A B OUTPUT
C D
CONTROL SIGNALS
X
Y 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 0 1 0 1
Y 0 0 1 1
Multiplexer Control logic table Table 1
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Figure 9 shows the DEMUX logic circuit.
0
BIT 1
S2
1
BIT 2
S1
2
BIT 3
3
BIT 4
4
BIT 5
5
BIT 6
6
BIT 7
7
BIT 8
S0
DATA INPUT
Demultiplexer Logic Circuit Figure 9
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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.
4 SEC
4 SEC
FRAMES 4 SUBFRAMES ONE FRAME
SUBFRAME 1
SUBFRAME 2
SUBFRAME 3
SUBFRAME 4
64 WORDS ONE SUBFRAME
1
2
3
4
5
61
62
63
64
SYNC WORD 12 BITS ONE WORD
1
2
3
4
5
6
7
8
9
10
11
12
+5v ARINC 573 HARVARD BI-PHASE CODE DATA
0v -5v
1
1
0
0
0
1
0
0
1
0
1
0
ARINC 537 Signal Format Figure 10
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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.
A ANALOGUE TO DIGITAL CONVERTER
D
A DIGITAL TO ANALOGUE CONVERTER
D
PRESSURE TO FREQUENCY CONVERTER
P F
Converters Figure 11
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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 STRUCTURE
ARINC 629 STRUCTURE
ARINC 429/629 Bus Structures Figure 21
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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.
TERMINAL GAPS
1
SYNCHRONIZATION GAP
3
2
4
1
TERMINAL INTERVAL
PERIODIC INTERVAL TERMINAL GAPS
1
2
SYNCHRONIZATION GAP
3
4
1
TERMINAL INTERVAL
APERIODIC INTERVAL
ARINC 629 Bus Structure Figure 22
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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.
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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.
START
NEXT
NEXT
NEXT
TERMINAL INTERVAL
LABEL DATA WORD
DATA WORD DATA WORD
WORD STRINGS
HI - LO SYNCH
20 BITS
20 BITS
LABEL
DATA
P
P
UPTO 256 DATA WORDS
HI - LO SYNCH
ARINC 629 Message Structure Figure 23
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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.
INDUCTIVE PICK-UP ARINC 629 DATA BUS
COUPLING OUTPUT DATA
ARINC 629 - Inductive Coupling Technique Figure 24
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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 600 CONNECTOR
STANCHION DISCONNECT
STUB CABLES (TWO SHIELDED TWISTED PAIRS) 1 PAIR RECEIVE 1 PAIR TRANSMIT
LRU TRAY
STUB CABLE (FOUR CONDUCTORS WITH OVERALL SHIELD)
ARINC 629 CURRENT MODE COUPLER
ARINC 629 Connection Figure 25
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Figure 26 shows ARINC 629 system layout.
OVERHEAD PANEL LRU NO 5
LRU NO 3
LRU NO 1
OPAS
LRU NO 2
LRU NO 4
LRU NO 6
ARINC 629 System Layout Figure 26
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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 transistortransistor logic (TTL) and complementary symmetry metal oxide semiconductor (CMOS or COSMOS) which supersede earlier resistortransistor 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.
POSITIVE LOGIC
: 0 - LOW VOLTAGE : 1 - HIGH VOLTAGE
NEGATIVE LOGIC
: 0 - HIGH VOLTAGE : 1 - LOW VOLTAGE
0
1
5V
0
0
0
1
0V POSITIVE LOGIC
NEGATIVE LOGIC
Logic Conventions Figure 1
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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.
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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).
A
B
A A.B B SYMBOL
A
B
A.B
0
0
0
1
0
0
0
1
0
1
1
1
Basic 'AND' Gate Figure 2
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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.
A
B
A
A
B
A+B
0
0
0
1
0
1
0
1
1
1
1
1
A+B B SYMBOL
Basic 'OR' Gate Figure 3
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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.
+5V
A A.B B
A.B A
B
SYMBOL
TR1
TR2
A
B
A.B
0
0
1
1
0
1
0
1
1
1
1
0
The 'NAND' Gate Figure 4
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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.
A
+5V
A+B B SYMBOL A+B
A
B
TR1
TR2
A
B
A+B
0
0
1
1
0
0
0
1
0
1
1
0
The 'NOR' Gate Figure 5
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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
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A L B SYMBOL
UP
DOWN
UP
A
B
A
B DOWN
A
B
L
0
0
0
1
0
1
0
1
1
1
1
0
“EXCLUSIVE OR” Represented by Switches Figure 6
A
Q
B
XOR Circuit Figure 7
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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.
A
A
A 1 0
A 0 1
The Inverter Figure 8
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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.
A
A
A
A
+5V
OV
(a) NAND INVERTER
(b) NOR INVERTER
Logic Gate Inverters Figure 9
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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.
A A
B
B C
C
D
D
E
E
F
F
G
G
(a) NAND SYMBOL
(b) NOR SYMBOL
Multiple Gate Symbols Figure 10
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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.
5mS (a) - SINGLE OUTPUT
2mS
5mS
5mS
3mS (b) - MULTIPLE OUTPUT
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).
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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.
A
A A+B
AB B
B
A
B
AB
A
B
A+B
1
0
0
1
0
0
1
1
0
1
1
1
0
0
0
0
0
1
0
1
1
0
1
1
Figure 13 shows examples of Active State Indicators Active State Indicators Figure 13
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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
A B
ABC
C
A
B
C
ABC
0
0
1
0
1
0
1
0
0
1
1
0
1
1
1
0
0
0
0
0
1
0
0
0
0
1
0
0
1
1
0
1
1 state.
The 'INHIBIT' Gate Figure 14
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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.
+v DOWN
RIGHT MAIN GEAR DOWN SWITCH
+v NOSE GEAR DOWN SWITCH
+v LEFT MAIN GEAR DOWN SWITCH
WARNING HORN
+v THROTTLE SWITCH
Landing Gear Logic Circuit Figure 24
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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.
AUXILIARY POWER UNIT (APU)
AND
APU LOAD CONTROL VALVE
GROUND PNEUMATIC CONNECTION 2 1-2 VALVE ENG 1 AIR PNEUMATIC OVERPRESSURE (ENG 1)
AND AND
OR
OR
No 2 ENGINE
GROUND PNEUMATIC CONNECTION 1 ENG 3 AIR PNEUMATIC OVERPRESSURE (ENG 3)
AND
2-3 VALVE
Engine Starting Logic Circuit Figure 25
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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.
IAS > 60 kts (GND/FLT) FLT = 0
TAKE-OFF REPORT
NON VOLATILE VOLITILE MEMORY
5 SECS
PROP 1 > 675 RPM PROP 2 > 675 RPM
F50 Mini Aids Take-off Report Figure 26
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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.
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< ± 2 kts
ENGINE TORQUE < ± 1%
HIGH PRESS TURB < ± 0.5% RPM
AIRSPEED
< ± 100 ft
TIME DELAY 3 SEC
LANDING MODE
> 8 000 ft
TIME DELAY 2X32SEC
PRESSURE REGULATION SHUT OFF VALVES OPEN
NO CHANGE AIR COND
AIRSPEED 145 kts
ALTITUDE
TIME DELAY 2X32SEC
TIME DELAY 30 SEC
COLLECTED INFORMATION
PWR INTERRUPT
15 MIN COUNTER DELAY
COLLECTED INFORMATION
ENABLE
STABLE CRUISE 2
STABLE CRUISE 1
ON GROUND MODE
STABLE CRUISE 1
NON VOLATILE MEMORY
WRITE INHIBIT AFTER REPORT STOREAGE
STABLE CRUISE 2
TIME DELAY X SEC
RESET AFTER LANDING
TIME DELAY X 1 SEC
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Figure 27 shows a block schematic diagram of the mini AIDS cruise reporting.
F50 Mini AIDS Block Schematic Figure 27
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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 nonvolatile 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.
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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 generalpurpose 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.
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A block schematic diagram of an analogue computer is shown in Figure 1.
ANALOGUE COMPUTING ELEMENTS
INPUT DEVICES
OUTPUT DEVICES
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. DIGITAL TECHNIQUES /EIS
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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.
ANALOGUE COMPUTING DME O/P DISTANCE TO GO
DISTANCE
÷
GROUND SPEED INDICATOR
TIME
TIMING
Computing Groundspeed from 'Distance to Go' Figure 2
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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.
CONTROL
ARITHMETIC
INPUT
OUTPUT
MEMORY
CENTRAL PROCESSOR UNIT (CPU)
Digital Computer Block Diagram Figure 3
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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.
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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.
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ADDRESS
I/P
INPUT/OUTPUT INPUT/OUTPUT UNIT UNIT
CLOCK CLOCK
MEMORY
O/P
CONTROL CONTROL & & ARITHMETIC ARITHMETIC UNIT UNIT
CPU CONTROL BUS
DATA BUS
Computer Buses Figure 4
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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 digitalto-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 fanout (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 PERIPHERAL 1
PERIPHERAL 2
PERIPHERAL 3
PORT 1
PORT 2
PORT 3
CONTROL BUS
INPUT/OUTPUT UNIT
COMPUTER DATA BUS (8 BITS)
ADDRESS BUS
simultaneously. Input/Output Unit Figure 5
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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.
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Figure 6 shows how these primary memories are connected to a simple computer bus.
DATA BUS TO INPUT/OUTPUT DEVICE
TO CPU
ROM
RAM
MEMORY ADDRESS REGISTER & CHIP SELECT DECODER
TO INPUT/OUTPUT DEVICE
NOTE: CONTROL BUS OMITTED FOR SIMPLICITY
FROM CPU
ADDRESS 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 DIGITAL TECHNIQUES /EIS
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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 FieldEffect 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 +5V
TR1
TR2
CELL SELECT LINE
LOGIC 1 OUTPUT/INPUT
16K MEMORY = 16,384 FLIP-FLOPS = 32,768 TRANSISTORS
LOGIC 0 OUTPUT/INPUT
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1.7.2 7489 TTL RAM DEVICE
16
15
14
13
B
C
D
12
D4
11
10
S4
S3 S2
2
3
4
5
6
7 SENSE OUTPUT 2
D2
DATA INPUT 2
S1
SENSE OUTPUT 1
D1
DATA INPUT 1
WE
WRITE
ME
MEM
ADDRESS A
9
D3
A
1
SENSE OUTPUT 3
DATA INPUT 3
Vcc
SENSE OUTPUT 4
DATA
ADDRESS B,C & D
INPUT 4
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.
8
FOUR MEMORY CELLS
Vee 1 0 0 0
ENABLES
0 1 0 1
1 1 1 0
1 0 1 0
0 1 2 3 4
4 BIT ADDRESS
5 6
1101
16 LOCATIONS EACH HOLDING FOUR BITS
7 8 9 10 11
READ/WRITE SIGNALS
12
1
1
0
1
13 14 15
4 BIT DATA IN
1
1
0
1
1
1
0
1
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.
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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
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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.
Y1
Y2
Y3
Y4
CURRENT IS INSUFFICIENT TO MAGNETIZE CORE WITH ONLY ONE CURRENT
X1
X2
X3
X4 X1 & Y1 CURRENT MAGNETIZES THE CORE
16 Bit Ferrite Memory Figure 10
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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.
+5V
+5V
SENSE (HIGH)
“0” TR1
SENSE (LOW)
“1”
FUSE LINK
NO FUSE LINK
ADDRESS LINE
LOGIC 0
LOGIC 1
TR2 0V
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.
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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.
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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.
CPU ARITHMETIC UNIT
C O M P U T E R
CONTROL
INPUT OUTPUT UNIT
CLOCK
H I G H W A Y
MEMORY (REGISTERS)
MEMORY
Central Processing Unit Figure 13
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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. PROGRAM COUNTER REGISTER
PORT 1 INPUT OUTPUT ADDRESS DECODE
INSTRUCTION DECODE REGISTER
CONTROL UNIT
ACCUMULATOR REGISTER
PORT 3
I N T E R N A L
H I G H W A Y
TIMING
PORT 2
MEMORY ADDRESS REGISTER
TEMPORARY REGISTER
MEMORY
STATUS FLAG REGISTER
The CPUs Internal Registers Figure 14
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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 generalpurpose 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. DIGITAL TECHNIQUES /EIS
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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.
INPUT/ OUTPUT PORTS
OUTPUT
INPUT
ROM
MICROPROCESSOR (CPU)
COMPUTER HIGHWAY
RAM
Elementary Microcomputer Figure 15
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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.
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FMS CDU 1
FMS CDU 2 AFCAS EICAS
FMC 1
FMC 2
TMS
EFIS
NAVIGATIONAL SYSTEMS
EFIS
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. c.
b.
Standard Flight Routes.
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.
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Figure 17 shows the layout of FMC memory.
INITIAL AIRLINE BASE & 28 DAY UPDATES
REQUESTED ROUTE LATERAL VERTICAL
NAV DATA BASE BUFFER
F PER
MEMORY STORAGE 16 BIT WORDS
RAW DATA FOR COMPUTATIONS
ROLL CHANNEL
AILERON CONTROL
PITCH CHANNEL
ELEVATOR CONTROL
MODE TARGET REQUESTS
THRUST LEVER CONTROL
A DAT
OPERATION PROGRAM
DISPLAYS
STORAGE
FMC
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.
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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".
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Figure 18 shows an FMS Control/Display Unit (CDU).
LINE SELECT KEYS DISPLAY SCREEN
ALPHANUMERIC KEYPAD
FUNCTION SELECT KEYS PPOS
NEXT PHASE
1
2
3
DIR
FUEL
AIR PORTS
4
5
6
HDG SEL
DATA
FIX
7
8
9
PERF
0
START ENG OUT SPEC F-PLN
EXEC MSG CLEAR
A
B
C
D
E
F
G
H
I
J
K
L
M
N
O
P
Q
R
S
T
U
V
W
X
Y
Z
/
DISPLAY BRIGHTNESS CONTROL
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.
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1.16 COMPUTER INPUT Figure 19 shows input information for a typical airborne digital computer. FROM CONTROL
SENSORS:
A
VOR/DME - OMEGA DOPPLER - COMPASS ETC
D
MAGNETIC CARD READER
FROM CONTROL
MAGNETIC TAPE CASSETTE/CARTRIDGE
PUSH BUTTON CONTROLLER ALPHANUMERIC DEDICATED
REGISTERS SEQUENCING & ADDRESSING
TO STORE
TO CONTROL
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
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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.
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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.
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1.17.11
ENGINEERS
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.
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1
ENGINEERS
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.
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Figure 1 shows an example of light going from air into glass.
ANGLE OF INCIDENCE
LIGHT
AIR NORMAL LINE PERPENDICULAR TO GLASS SURFACE
I
GLASS
R ANGLE OF REFRACTION
Snell’s Law on Refraction (Air into Glass) Figure 1
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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.
TOTAL INTERNAL REFLECTION
41.9º θº1
θº2
θº1 = θº2
1.5 SIN 41.9º = 1.00174
Critical Angle Figure 2
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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.
CORE LIGHT RAY
CLADDING LIGHT RAY STRIKES THE CLADDING AT AN ANGLE GREATER THAN THE CRITICAL ANGLE, THEREFORE THE LIGHT RAY IS REFLECTED RATHER THAN BEING REFRACTED.
Fibre Optic (Core and Cladding) Figure 3
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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. ACCEPTANCE ANGLE
FILLER LIGHT MUST FALL INSIDE THIS ANGLE
STRANDS ARAMID YARN
TO BE GUIDED THROUGH THE CORE
OPTICAL FIBRES
OPTICAL SEPARATOR
FIBRES
TAPE
ARAMID YARN
OUTER JACKET
SEPARATOR
FILLER STRANDS
END
TAPE
VIEW
FIBRE OPTIC CABLE
Light Coupling (Critical Angle) Figure 4
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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.
LATERAL MISALIGNMENT
ANGULAR MISALIGNMENT
AXIAL MISALIGNMENT
POOR END FINISH
Fibre Optic Alignment Figure 5
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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.
FIBRE OPTIC CABLE STRAIN RELIEF BOOT BACKSHELL
THREADED COUPLING JACK SCREW
COUPLING RING CERAMIC TERMINUS
FIBRE OPTIC ALIGNMENT SLEEVES PINS
ALIGNMENT HOLE
RECEPTACLE
PLUG
Type “A” Connector Figure 6
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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.
FIBRE OPTIC CABLE STRAIN RELIEF BOOT BACKSHELL
THREADED COUPLING COUPLING RING
ALIGNMENT PINS
ALIGNMENT HOLE
PROTECTIVE WINDOW
RECEPTACLE
PLUG
Type “B” Connector Figure 7
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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.
STRAIN RELIEF BOOT MOUNTING FLANGE
FIBRE OPTIC CABLE
BACKSHELL
PROTECTIVE WINDOW
RECEPTACLE CONNECTOR
Type “C” Connector Figure 8
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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.
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Figure 10 shows example of type “A”, “B” and “C” connectors and identification labels.
TYPE “A” PLUG CONNECTOR STRAIN RELIEF
BACKSHELL
ASSEMBLY IDENTIFICATION SLEEVE
MATE WITH IDENTIFICATION SLEEVE
BOEING TYPE “A” PLUG CONNECTOR
TYPE “B” PLUG CONNECTOR STRAIN RELIEF
BACKSHELL
ASSEMBLY IDENTIFICATION SLEEVE
MATE WITH IDENTIFICATION SLEEVE
BOEING TYPE “B” PLUG CONNECTOR
TYPE “C” PLUG CONNECTOR
BACKSHELL MOUNTING FLANGE
STRAIN RELIEF
ASSEMBLY IDENTIFICATION SLEEVE
MATE WITH IDENTIFICATION SLEEVE
BOEING TYPE “C” PLUG CONNECTOR
Fibre Optic Connectors Figure 10
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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.
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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”.
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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.
BEND RADIUS >1.5"
FIBRE OPTIC CABLE
STRAIN RELIEF 1" MIN EQUIPMENT TYPE “B” PLUG
Fibre Optic Connection Figure 5.10.14
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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.
UPTO 31 SUB-SYSTEMS
BUS CONTROLLER
SUB SYSTEM 1
SUB SYSTEM 2
SUB SYSTEM N
CONTROL & LOW SPEED DATA BUS HIGH SPEED DATA BUS
FIBRE OPTIC STAR COUPLER
STANAG 3910 Data Bus System Figure 17
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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).
UPTO 128 SUB-SYSTEMS
SUB SYSTEM 1
SUB SYSTEM 2
SUB SYSTEM 3
SUB SYSTEM N
FIBRE OPTIC STAR COUPLER
Linear Token Passing High Speed Data Bus Figure 18
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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
LRG
FIBRE OPTIC CABLE
MOTION SENSORS
ELECTRICAL CABLE
FLIGHT CONTROL COMPUTER
ACTUATOR CONTROL ELECTRONICS
ACTUATOR AIR DATA COMPUTER CONTROL SURFACE
Fly-By-Light System Figure 19
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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.
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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 deckinstrument 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 Light-Emitting Diode Liquid Crystal Display
Cathode Ray Tube
Application Digital counter displays of engine performance. Monitoring indicators; Radio frequency selector indicators; Distance Measuring indicators; Control display units of Inertial Navigation Systems, etc. Weather radar indicators; display of navigational data; engine performance data; system status;
Electronic Display Applications Table1
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Figure 1 shows a typical flight-deck instrument panel and the different types of display used.
LED DISPLAYS
CRT DISPLAYS
LCD DISPLAYS
BAe 146 Electronic Instrument Layout Figure 1
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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-SEGMENT CONFIGURATION
THIRTEEN-SEGMENT CONFIGURATION
Seven and Thirteen Segment Display Formats Figure 2
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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).
7 ROWS
4 COLUMNS
Dot Matrix Configuration Figure 3
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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.
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BIAS RECOMBINATIONS
p JUNCTION
n INJECTED ELECTRONS
CONTACT
LED Operation Figure 4
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Figure 5 shows the construction of a LED.
Light-Emitting Diode (LED) Figure 5
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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 ELECTRONIC INSTRUMENT SYSTEM
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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 dotmatrix configuration for the digital readout of engine speed.
DOT MATRIX LED DISPLAY ENGINE SPEED
20 0
40
ANALOGUE ENGINE SPEED INDICATOR
60
N1 % RPM 80 100
Smith's
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.
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Figure 8 shows a Power Plant Instrument Group from a Boeing 737-400, which has both LED and dot matrix, displays.
MAN SET
% RPM 12
0
72 65 8
12 2
10
10
N
4
6
8
4 X 7 MATRIX DISPLAY
0 2 72 65 4 6
1
°C
8 7
84
84 87
EGT % RPM
LED DISPLAY
5
5
100 4
100 4
N 2 X1000
6 5
0
2
27 1
4
5
2
3
PULL TO SET N1
6 1
4
FF/FU
0
27 21 3
1
2
KGPH/KG
PUSH
FUEL USED
RESET FUEL USED
PULL TO SET N1
Boeing 737-400 Power Plant Instrument Group Figure 8
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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.
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Figure 9 shows the operation of a LCD.
INCIDENT LIGHT
REFLECTED LIGHT
TRANSPARENT CONDUCTIVE PLATES
INCIDENT LIGHT REFLECTED LIGHT
TRANSPARENT
OPAQUE
LED Operation Figure 9
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Figure 10 shows the structure of a seven-segment LCD.
LIQUID CRYSTAL LAYER (TYPICAL SPACING = 10 MICRONS)
SEVEN SEGMENT ELECTRODE
MIRROR IMAGE (NOT SEGMENTED)
FRONT PLATE
BACK PLATE
SEGMENT CONTACTS COMMON RETURN CONTACT
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. ELECTRONIC INSTRUMENT SYSTEM
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Figure 11 shows a BCD to seven-segment decoder.
LOW VOLTAGE POWER SUPPLY TO EACH SEGMENT
1
0
1
1
2 4 8
0 BCD TO 7 0 SEGMENT 0 1 DECODER 1 0 0
0
BCD – Seven-Segment decoder Figure 11
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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.
GRAPHITE COATING (COLLECTS SECONDARY ELECTRONS TO PREVENT SCREEN BECOMING NEGATIVELY CHARGED)
DEFLECTING COILS CATHODE
ANODE
HEATER
GRID
GLASS ENVELOPE PERMANENT MAGNETS (BEAM FOCUSING)
ELECTRON BEAM
SCREEN
Cathode Ray Tube CRT Cross-Section Figure 12
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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.
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Figure 13 shows the typical voltages used in a small CRT.
CONNECTED TO CONDUCTIVE COATING ON GLASS
CATHODE GRID
FIRST ANODE
SECOND ANODE
HEATER
0V
-50V
+300V
+5 kV
CRT Voltages Figure 13
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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.
Y DEFLECTION PLATES
ANODE
X DEFLECTION PLATES
X and Y Deflection Plates Figure 14
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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. MAGNETIC FIELD
N NECK OF THE TUBE
S
ELECTRON BEAM COMING OUT OF THE PAPER
VERTICALLY DISPOSED MAGNETIC COIL PRODUCES HORIZONTAL DEFLECTION OF THE BEAM
N
S
HORIZONTALLY DISPOSED MAGNETIC COIL PRODUCES VERTICAL DEFLECTION OF THE BEAM
RESULTANT DEFLECTION OF THE BEAM
Electromagnetic Deflection Figure 15 ELECTRONIC INSTRUMENT SYSTEM
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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.
RAMP OR SWEEP
CURRENT
FLYBACK
TIME
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 ELECTRONIC INSTRUMENT SYSTEM
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“Trapezoidal” shape. Figure 17 shows the scan coil graphs for electromagnetic deflection.
MAX IDEAL CURRENT
0 MAX
VOLTAGE ACROSS R
0 MAX
VOLTAGE ACROSS L
0
MAX RESULTANT TRAPEZOIDAL VOLTAGE
0
Scan Coil Graphs Figure 17
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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
FAR 25.1309 &
No significant
Reduction of the aircraft capability or
and occupants of failure conditions
JAR 25.1309
degradation of
of the crew ability to cope with
continued safe
definitions
aircraft capability
adverse operating conditions
flight and landing
or design error
or crew ability
Prevention of
of the aircraft Large reduction
Slight reduction
in safety margins
Significant
of safety
reduction in
Physical distress
margins,
safety margins
or workload such
Slight increase in
Reduction in the
that the flight
ACJ No 1
workload, e.g.
ability of the flight
crew cannot be
Jar 25.1309
routine changes
crew such that
relied upon to
Loss of aircraft
definitions
in flight or plan or
they cannot be
perform their
and/or fatalities
Physical effects
relied upon to
tasks accurately
but no injury to
perform their
or completely, or
occupants
tasks accurately,
serious injury to
or injury to
or death of a relatively small
occupants
proportion of the occupants ACJ No 1 to JAR 25.1309
Minor Effect
Major Effect
Hazardous Effect
Definition of Criticality Category FAA Advisory Circular 25.1409-1
Catastrophic Effect
Non-essential
Essential
Critical
Level 3
Level 2
Level 1
definition of Criticality Category DO-178A/ED-12A Software level*
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.
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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.
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(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/ED12A.
(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.
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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.
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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).
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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
Description
Over braid shield, critical individual cable shield Metallic conduit, braid
Raceway, conduits
RF gaskets
Raceway, conduits
Removable panels
Corrosion, damage
Corrosion, damage
Corrosion, damage, deformation
Damage, erosion
Visual inspection, bonding measurement
Visual inspection of gaskets, bonding leads and straps
Visual inspection, measurement of shielding effectiveness
Examples
Degradation or Failure Mode
Maintenance Operations
Visual inspection, measurement of cable shielding bonding
AIRCRAFT STRUCTURE SHIELDING
Shield for non conductive surfaces Conductive coating
CIRCUIT PROTECTION DEVICES
Structural bonding
Contact bonds, rivet joints Corrosion, damage
Visual inspection, bonding measurement
Bonding lead and straps, pigtails Corrosion, damage, security of attachment Visual inspection for corrosion attachment and condition, bonding measurement
HIRF protection devices Resistors, Zener diodes, EMI filters, filter pins. Short circuit, open circuit
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.
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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.
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CURRENT MONITOR (AC MILLI-VIOLTS)
VOLTAGE GENERATOR
CLAMP-ON CURRENT TRANSFORMER
V1
II
CLAMP-ON CURRENT TRANSFORMER
FIXING HARDWARE PROVIDING ELECTRICAL BONDING
CONDUIT LOOP UNDER TEST
STRUCTURE
ZCONDUIT + ZSTRUCTURE = V1/II
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.
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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.
VOLTAGE GENERATOR VOLTAGE MONITOR
CLAMP-ON CURRENT TRANSFORMER
V1
V2 FIXING NUT BAD JOINT
FERRULE BRACKET CONDUIT
LOOP UNDER TEST
STRUCTURE
V2 = V1 ACROSS BAD JOINT
Identification of A Bad Joint Figure 2
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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.
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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.
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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).
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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.
WARN
WARN
CAUT
CAUT
ECAM CONTROL PANEL
DMC 1
FWC 1
DMC 3
SDAC 1
A/C SYSTEM SENSORS RED WARNINGS SYSTEM PAGES FLIGHT PHASE
DMC 2
SDAC 1
A/C SYSTEM SENSORS AMBER WARNINGS SYSTEM PAGES
FWC 2
NAV & AFS SENSORS
ECAM Functional Diagram Figure 5
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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.
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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.
350 300
400
8 4 MACH
60 1 0 9
80
250
120 IAS KNOTS
240 220
200
140 180
5
LDG GEAR GRVTY EXTN
5
RESET OFF DOWN
ECAM Display Layout Figure 6
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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. Preflight, 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.
ENGINE 10
5
8 7. 0
5
10
F.USED
6 5. 0
N1 %
1530
FOB : 14000KG
KG
1530
OIL 10
5
6 50
80 1500
5
EG T ºC
10
4 80
FLAP
F
11.5
(N1) 0.9
VIB 1.2
(N2) 1.3
11.5
AIR LDG ELEV AUTO
N2 %
80.2
FF KG/H
1500
NO SMOKING: SE AT BE LTS: SP LRS: FLAPS :
S
QTY
VIB 0.8
ON ON FULL FULL
FULL
500FT
CAB V/S FT/MIN CKPT 20
FWD 22
AFT 23
24
22
24
250 CAB ALT FT 4150
LDG INHIBIT APU BLEED
ECAM UPPER DISPLAY
TAT +19 ºC SAT +17 ºC
23 H 56
G.W. 60300 KG C.G. 28.1 %
ECAM LOWER DISPLAY - CRUISE
ECAM Upper and Lower Display (Cruise Mode) Figure 7
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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.
10
5
87.0
650
ADVISORY MESSAGES
80 1500
65.0
N1 %
10
5
10
5
FOB : 14000KG 10
5
EGT ºC
480
N2 %
80.2
FF KG/H
1500
HYD B RSVR OVHT B SYS LO PR
FAILURE MESSAGES
S
FLAP
F
FULL
FLT CTL SPOILERS SLOW
1 FUEL TANK PUMP LH
ECAM Advisory Mode Figure 8
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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.
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Figures 9 – 13 show the 12-system and status pages available.
COND
TEMP ºC
CAB PRESS AP PSI
ALTN MODE FAN
FAN
CKPT 20
FWD 22
24
22
C
H
C
LDG ELEV MAN 500FT V/S FT/MIN
2
8
AFT 23
0
0 4.1
24 H
C
INLET
SAFETY
EXTRACT
PACK 1
23 H 56
SYST 2
VENT
HOT AIR
TAT +19 ºC SAT +17 ºC
10 0 4150
DN
MAN
SYST 1
H
1150 2
CAB ALT FT
UP
G.W. 60300 KG C.G. 28.1 %
AIR CONDITIONING SYSTEM PAGE
TAT +19 ºC SAT +17 ºC
PACK 2
23 H 56
G.W. 60300 KG C.G. 28.1 %
PRESSURIZATION SYSTEM PAGE
ECAM System Displays Figure 9 Note; These pages are displayed: Automatically due to an advisory or failure related to the system. Whenever called manually.
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ELEC
BAT 1 28V 150A
F/CTR
BAT 2 28V 150A
DC BAT
DC 1
GBY
DC 2 DC ESS
TR 1 28V 150A
AC 1 GEN 1 26% 116V 400HZ
TAT +19 ºC SAT +17 ºC
ESS TR 28V 130A
EMERG GEN 116V 400HZ
23 H 56
SPD BRK
L AIL BG
R AIL GB
PITCH TRIM G Y 3.2º UP
AC 2
AC ESS
APU 26% 116V 400HZ
TR 2 28V 150A
EXT PWR 116V 400HZ
L ELEV BG
GEN 2 26% 116V 400HZ
G.W. 60300 KG C.G. 28.1 %
ELECTRICAL SYSTEM PAGE
TAT +19 ºC SAT +17 ºC
RUD GBY
23 H 56
R ELEV YB
G.W. 60300 KG C.G. 28.1 %
FLIGHT CONTROL SYSTEM PAGE
ECAM System Displays Figure 10 Note; These pages are displayed: Automatically due to an advisory or failure related to the system. Whenever called manually.
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FUEL KG
F.USED 1
1550
F.USED 2
APU
HYD
1550
FOB
3000
LEFT
10750
TAT +19 ºC SAT +17 ºC
YE LLOW
5600
23 H 56
PSI
3000
PSI
3000
RIGHT
CTR
550
BLUE
GREE N
28750
10750
550
TAT +19 ºC SAT +17 ºC
G.W. 60300 KG C.G. 28.1 %
FUEL SYSTEM PAGE
23 H 56
G.W. 60300 KG C.G. 28.1 %
HYDRAULIC SYSTEM PAGE
ECAM System Displays Figure 11 Note; These pages are displayed: Automatically due to an advisory or failure related to the system. Whenever called manually.
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BLEED
WHEEL
20 ºC
24 ºC C
C
H RAM AIR
50 ºC
170 1
ºC REL
140
140
2
3
ºC REL
LO
HI
4
AUTO BRK
23 H 56
LO
HI
140 1
TAT +19 ºC SAT +17 ºC
H 230 ºC
LP TAT +19 ºC SAT +17 ºC
G.W. 60300 KG C.G. 28.1 %
LANDING GEAR/WHEEL/BRAKE SYSTEM PAGE
2
GND APU HP HP
23 H 56
LP G.W. 60300 KG C.G. 28.1 %
AIR BLEED SYSTEM PAGE
ECAM System Displays Figure 12 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.
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APU
OXY 1850 PSI
DOOR ARM
ARM
APU 26% 116 V 400 HZ
AVIONIC
CABIN FWD COMPT
BLE ED 35 PSI
CARG O
ARM
EMER EX IT
10
ARM
0
80
FLAP OPEN
CARG O BULK CABIN
TAT +19 ºC SAT +17 ºC
ARM
ARM
23 H 56
N %
5
7
3
580
TAT +19 ºC SAT +17 ºC
C.G. 28.1 %
DOOR/OXY SYSTEM PAGE
EG T ºC
23 H 56
C.G. 28.1 %
APU SYSTEM PAGE
ECAM System Displays Figure 13 Note; These pages are displayed: Automatically due to an advisory or failure related to the system. Whenever called manually. Related flight phase.
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1.2.9 CONTROL PANEL
The layout of the control panel is shown in Figure 14.
DISPLAY ON & BRIGHTNESS CONTROL
DISPLAY ON & BRIGHTNESS CONTROL
SGU SELECT SWITCHES
1
LEFT DISPLAY
OFF
ECAM
SGU
2
FAULT
FAULT
OFF
OFF
RIGHT DISPLAY
BRT
OFF
BRT
MESSAGE CLEARANCE SWITCH CLR
STS
RCL
STATUS MESSAGE SWITCH
RECALL SWITCH
ENG
HYD
AC
DC
BLEED
COND
PRESS
FUEL
APU
F/CTL
DOOR
WHEEL
SYSTEM SYNOPTIC DISPLAY SWITCHES
ECAM Control Panel Figure 14
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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.
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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.
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Figure 15 shows the interface between the modules within the SG.
WEATHER RADAR DATA
MAIN PROM
WX INPUT
MAIN RAM
INPUT 1
INPUT 2
DISPLAY DRIVER
DISPLAY UNIT DEFLECTION SIGNALS
STROKE POSITION DATA
STROKE GENERATOR
DISPLAY SEQUENCER IRS ILS DME VOR
STROKE/VIDEO & PRIORITY DATA
DISPLAY COUNTER I/O BUS
DISPLAY CONTROL
RASTER GENERATOR
WX MEMORY 2 X 16K RAMS DISPLAY SEQUENCER DATA BUS
FMC RAD ALT VOR EFIS CONTROL
TRANSFER BUS
MAIN
PROCESSOR
DISPLAY UNIT VIDEO
WX RASTER
CHARACTER DATA
DISPLAY UNIT RASTER/STROKE SELECT
Symbol Generator Module Interface Figure 15
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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 Loads data into registers on stroke and raster generator cards. Sequencer Stroke Generates all single characters, special symbols, straight and Generator curved lines and arcs on display units. Raster Generates master timing signals for raster, stroke, EADI and Generator 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.
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Figure 16 shows a block schematic of the display unit.
115V 4OOHz
LOW VOLTAGE POWER SUPPLY
HIGH VOLTAGE POWER SUPPLY
LIGHT SENSOR DISPLAY UNIT BRIGHTNESS RASTER BRIGHTNESS
RED GREEN BLUE BEAM TEST SYNCHRONIZING
DIGITAL LINE RECEIVERS
VIDEO MONITOR CARD
CRT
INTENSITY RASTER/STROKE DAY/NIGHT
X DEFLECTION Y DEFLECTION
ANALOG LINE RECEIVERS
DEFLECTION CARD
CONVERGENCE CARD
EFIS Display Unit Block Schematic Figure 16
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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.
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Figure 17 shows the EFIS units and signal interface in block schematic form.
Honeywell GS
ATT 2 AOA F
20
20
10
10
10
10
G GS TTG
WX
DIM
CRS
ET
DH
SC CP
MAP
BOT
REV
TOP
S CMD M .99 200DH
HDG
TEST RASTER DIM
AIR DATA COMP NAV
FMS
INS 1
INS 2
ATT
HDG
I
140RA
Honeywell
VOR 2
CRS +0
OFF
N 33
H 2.1 NM 3
30
BRG
BRG
NAV 1
345
ADF 1
OFF
DH
EFIS SG No 1
VOR 1 ADF 2
AUTO
20
6
VOR 1
ADF 1
E 1 2
INERTIAL REF SYSTEM
VLF
ADF 2 ADF 1
20
W 24
ARC
21
S
HDG
NAV AID ILS/VOR
15
FULL
GSPD
013
130 KTS
EFIS SG No 3 RAD ALT Honeywell GS
ATT 2
WEATHER RADAR
AOA F
20
20
10
10
10
10
G S CMD M .99 200DH
DME FULL ARC
DIM
CRS
FMS
GS TTG
WX
ET
DH
MAP
BOT
SC CP
REV
TOP
20
20 DH
140RA
HDG
TEST RASTER DIM
EFIS SG No 2
AFCS
Honeywell VLF
FMS
INS 1
INS 2
CRS
ATT
HDG
NAV 1
345 +0
AUTO
BRG
30 VOR 1
BRG ADF 1
HDG
S
013
EFIS Block Schematic Figure 17
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H 2.1 NM 3
E 1 2
ADF 1 OFF
OFF
N 33
VOR 2
W 24
VOR 1 ADF 2
6
ADF 2 ADF 1
21
NAV
15
GPWS
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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.
FULL ARC
GS TTG
WX
DIM
MAP
ET
DH
BOT
SC CP
REV
TOP
HDG
CRS TEST
EFIS Display Controller Figure 18
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NAV
VHF
FMS
INS 1
INS 2
HDG
ADF 2
VOR 1
ADF 1
ATT
VOR 2
ADF 2 ADF 1
AUTO OFF
OFF
BRG
BRG
EFIS Source Controller Figure 19
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1.3.11 ELECTRONIC ATTITUDE DIRECTOR INDICATOR (EADI)
The EADI displays traditional attitude information (Pitch & Roll) against a twocolour 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
Honeywell
LOC
HDG
GS
ATT 2 20
20
F
S M .99 200 DH
20
10
10
10
10 20
M AP ENG
140 RA
Electronic Attitude Director Indicator (EADI) Display Figure 20
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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.
Honeywell NAV 1
CRS 315 +0
H
33
6
24
3
WPT
N
W
30
2.1 NM
G
E
21
VOR 1
ADF 1
12
15
S HDG
350 GSPD 130 KTS
Electronic Horizontal Situation Indicator (EHSI) Display Figure 21
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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).
Honeywell
DTRK
317
FMS1 30 NM
320 30
33
N V VOR 1
50 ADF 1
HDG
350
25 15
GSPD 130 KTS
EHSI Partial Compass Mode Display Figure 22
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Figure 23 shows an EHSI partial format with Weather Radar information.
Honeywell
DTRK
317
FMS1 30 NM
320 30
33
N V VOR 1
50 ADF 1
HDG
350
GSPD
25
130 KTS
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.
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Figure 24 shows an EHSI MAP mode display.
Honeywell
DTRK
317
FMS1 30 NM
320 33
30 05
04
N
05
V VOR 1
50
03
ADF 1
HDG
350
GSPD
25
130 KTS
EHSI MAP Mode Display. Figure 24
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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.
Honeywell
120 NM HDG ILS
CRS FR ATT 2
20
20
F
010
10
10
10
000
S M .99 200 DH
10
M 33
00
03 DH
140 RA
EFIS Composite Display Figure 25
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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).
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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
RANGE
WPT
PLAN OFF
ADF
10
320
BRG
FORMAT
160
80
20
ROSE
MAP
ARC
OFF LNAV
VOR
N-AID
BACKSPACE
ARPT
GRP
CRS
V/L
DATA
FORWARD SPACE
VERIFY
EFIS CONTROL PANEL
BRT
EFIS SELF-TEST BUTTON
ND
WX
PFD DH
TEST COMPACT
WX OFF
DIMMING PANEL
BAe 146 EFIS Control & Dimming Panels Figure 26
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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”.
FAULT REVIEW FAULT ERASE TEST PATTERN SG SELF TEST OPTIONS/CONFIG
Maintenance Master Menu Display Figure 27
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The Symbol Generator Self-Test sequence is automatic and the process is as shown in Figure 28.
FAULT REVIEW FAULT ERASE TEST PATTERN SG SELF TEST OPTIONS/CONFIG
SELF TEST IN PROGRESS
PASS FAIL
SYMBOL GENERATOR SELF TEST AIRCRAFT CONFIGURATION YY DP SOFTWARE PART NUMBER: XXXXXXXXX-XX SMP SOFTWARE PART NUMBER XXXXXXXX-XX TEST PASS
SYMBOL GENERATOR SELF TEST AIRCRAFT CONFIGURATION YY DP SOFTWARE PART NUMBER: XXXXXXXXX-XX SMP SOFTWARE PART NUMBER XXXXXXXX-XX TEST FAIL
SELF TEST FAILURES
INTERFACE STATUS
FAILURE 1 FAILURE 2 FAILURE 3 FAILURE 4 FAILURE 5 FAILURE 6
STATUS 1 STATUS 2 STATUS 3 STATUS 4 STATUS 5 STATUS 6
SG Self-Test Process Figure 28
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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 selftest, 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.
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PAGE INTENTIONALLY BLANK
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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.
ENGINE PRIMARY DISPLAY & WARNINGS CAUTIONS ADVISORIES
EICAS COMPUTER No 2
ENGINE & AIRCRAFT SYSTEM INPUTS
CAUTION
COMPUTER
DISPLAY
ENGINE STATUS
CANCEL
ENGINE SECONDARY DISPLAY OR STATUS DISPLAY OR MAINTENANCE DISPLAY
EICAS COMPUTER No 1
EVENT RECORD
L AUTO R
BRT
EICAS MAINT
THRUST REF SET DISPLAY SELECT
BRT BAL
BOTH L
R
MAX IND RESET
RESET
ELEC
PERF
MSG
HYD
APU
CONF
ENG EXCD
EPCS
ECS
MCDP
DISPLAY SELECT PANEL
EVENT READ AUTO
MAN
REC
ERASE
TEST
MAINTENANCE PANEL
EICAS Block Schematic Figure 29
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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:
EIS
Flight Control Position.
Hydraulic system status.
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1.4.4 DISPLAYS
Figure 30 shows displays presented on the Primary and Secondary DUs.
CAUTION
TAT 15°c 0.0
0.0 10
10
CANCEL RECALL
6
2
6
2
N1 0
0
EGT
V VV VV V V
50
50
OIL
PRESS
120
120
OIL
88.00 N2 86
TEMP
18
18
OIL
88
86
N3 4.4
4.4
QTY
N1
FAN
3.1
1.9
FF
VIB
EICAS Primary & Secondary Displays Figure 30
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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.
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Figure 31 shows a status mode display.
HYD QTY
L 0.99
C R 1.00 0.98
HYD PRESS
2975
3010 3000
APU
EGT 440
OXY PRESS
RPM 103
OIL 0.75
0.0
FF
0.0
CABIN ALT AUTO 1 ELEV FEEL
1750
RUD
AIL ELEV AIL
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.
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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 nonvolatile 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.
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Figure 32 shows an EICAS Control Panel
COMPUTER
DISPLAY
BRT
THRUST REF SET
BRT
ENGINE
STATUS
EVENT RECORD
BAL
L AUTO R
L BOTH R
MAX IND RESET
EICAS Control Panel Figure 32
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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.
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Figure 33. shows EICAS alert messages Level A, B and C.
RED WARNING
LEVEL A WARNING
CAUTION
AMBER
CANCEL RECALL
LEVEL B CAUTION
LEVEL C ADVISORY
TAT 15°c APU FIRE R ENGINE FIRE CABIN ALTITUDE C SYS HYD PRESS R ENG OVHT AUTOPILOT C HYD QTY R YAW DAMPER L UTIL BUS OFF
MASTER WARNING & CAUTION LIGHTS
110.0
70.0 10
10 6
2
6
2
N1 999
775
EGT
VVVVVVV
A - WARNING (RED) B - CAUTION (AMBER) C - ADVISORY (AMBER)
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.
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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.
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1.4.12 MAINTENANCE FORMAT
Maintenance pages can be called forward on the ground using the Maintenance Panel, refer to Figure 34.
PERFORMANCE AND AUXILLIARY POWER UNIT FORMATS ENVIRONMENTAL CONTROL SYSTEM AND MAINTENANCE MESSAGE FORMATS
ELECTRICAL AND HYDRAULIC SYSTEM FORMAT
EICAS MAINT DISPLAY SELECT
ECS
ELEC
PERF
MSG
HYD
APU
CONF MCDP
CONFIGURATION AND MAINTENANCE CONTROL/DISPLAY PANEL
SELECTS DATA FROM AUTO OR MANUAL EVENT IN MEMORY
EVENT READ AUTO
MAN
REC
ERASE
ENG EXCD
ENGINE EXCEEDANCES
TEST
BITE TEST SWITCH FOR SELF-TEST ROUTINE
ERASES STORED DATA CURRENTLY DISPLAYED RECORDS REAL-TIME DATA CURRENTLY DISPLAYED (IN MANUAL EVENT)
EICAS Maintenance Panel Figure 34
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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.
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Figure 35 shows the display formats seen during the Maintenance format.
96.1
96.1
PARKING BRAKE
85.0
85.0 10
10 2
2
6
6
N1 450
450
INDICATED WHEN EICAS IN MAINTENANCE FORMAT
EGT
50 OIL PRESS 105 OIL TEMP 20 OIL QTY 1.9 N2 VIB
97.0 8.4
50 100 20 1.9
N2
97.0 8.4
FF
ELEC/HYD
LOAD AC-V FREQ DC-A DC-V
HYD QTY HYD PRESS HYD TEMP
STBY BAT
L
R
APU BAT
GND PWR
0 0 10 28
0.78 120 402 140 28
0.85 125 398 150 27
0.00 0 0 0 28
0.00 0 0
L
C
R
0.82 3230 50
O/FULL 3210 47
0.72 2140 115
AUTO EVENT
AUTO EVENT SYSTEM FAILURES AUTOMATICALLY RECORDED DURING FLIGHT
R HYD QTY
Maintenance Mode Displays Figure 35
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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.
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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.
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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.
UNDERWATER LOCATING DEVICE
Flight Data Recorder Figure 113
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Figure 114 shows FDR block schematic diagram.
FDR ARINC 429
ANALOGUE
AIRCRAFT SYSTEMS
ANALOGUE DISCRETES
ARINC 573
ARINC 629
FAULT MONITORING
DFDAU AIMS
Flight Data Recorder Block Schematic Figure 114
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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.
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1
ENGINEERS
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.
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Figure 1 shows a Pressure head as fitted to aircraft to allow Pitot and Static pressures to the relevant indicators.
PITOT LINE
STATIC LINE
HEATER CONNECTION
FORWARD
PITOT PROBE
STATIC VENTS
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).
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Figure 2 shows typical aircraft static vent:
FUSELAGE
STATIC VENT
STATIC PIPE
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.
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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 corrosionresistant 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.
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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.
ORANGE FLOAT INDICATOR
TRANSPARENT PLASTIC PIPE
DRAIN VALVE
BAYONET FITTING CAP
(SELF SEALING)
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.
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Air Data Instrumentation Figure 5
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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.
0 9 8
1
SBY
1013
2
X 100 ft 7
3 5 0 00
3
5
6
5
4
MB
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.
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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. SYSTEMS
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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.
QNE FLIGHT LEVEL
QNH HEIGHT ABOVE SEA LEVEL
QFE HEIGHT ABOVE AIRFIELD
STANDARD SETTING 1013.25 MILLIBARS SEA LEVEL
QNH, QFE and QNE Definitions Figure 8 SYSTEMS
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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.).
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Figure 10 shows a simplified IAS operation.
PITOT
STATIC
IAS Operation Figure 10
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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.
RATE OF CLIMB SCALE 1,000 ft per sec
RATE OF CLIMB/DIVE POINTER
1
2
1000FT PER MIN
4
VSI
.5
6
UP
0DOWN
6
MAX INDICATED 6,000 ft per sec
VERTICAL SPEED
.5
1
2
4
RATE OF DIVE SCALE 1,000 ft per sec
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.
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Figure 12 shows a simplified VSI operation.
METERING UNIT
STATIC VENT
CLIMB
0
POINTER AND SCALE
DIVE
CAPSULE
MECHANICAL LINKAGE
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.
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Figure 13 shows a typical air data system for a large aircraft.
PRESSURE HEADS
UPPER
LOWER
VS
PC
MS 1
A/S 1
ADC 1
PITOT
IAS
STATIC
F/O
FLT REC
DIFF PRESS
ALT
PITOT
VS
STATIC
MS 2
ALT
IAS
A/S 2
ADC 2
CAPT UPPER
LOWER
PRESSURE HEADS
Air Data System Figure 13
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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:
1.7.2
Gyroscopic Inertia (Rigidity).
Precession.
Rigitity
The property, which resists any, force tending to change the plane of rotor rotation. It is dependent on:
SYSTEMS
1.
The mass of the rotor.
2.
The speed of rotation.
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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.
Z FRAME
Y
X
ROTOR OUTER RING
X Y
INNER RING
Z
Gyroscope. Figure 14
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Figure 15 shows the characteristics of gyro rigidity.
A
B C
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.
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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.
DIRECTION OF ROTATION
PRECESSION RATE = APPLIED FORCE 90º IN THE DIRECTION OF SPIN
SPIN AXIS 90º APPLIED FORCE
DIRECTION OF PRECESSION
Gyro Precession Figure 16
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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
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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 tenminute period is small. Figure 18 shows a vertical gyro corrected to the local vertical.
Corrected Vertical Gyro Figure 18
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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.
ROLL
PITCH
Displacement Gyro Figure 19
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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.
AIRCRAFT SYMBOL
3 ROLL SCALE
6
6
SPERRY
HORIZON BAR
3
ROLL POINTER\
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.
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Figure 21 shows the construction of the Gyro Horizon unit.
X OUTER RING
ROTOR
Y
Z1
SYMBOLIC AIRCRAFT
BALANCE WEIGHT
PIVOT POINT
Z Y1 ROLL POINTER & SCALE
X1
HORIZON BAR
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.
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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.
X
Z
1
Z
HORIZON BAR
X
1
Climb Attitude operation. Figure 22
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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.
X
Y
Y
1
BANK TO PORT DATUM X
1
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.
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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)
FD 2
F
GSL
1
1
S
2
A
TT
RW
Y
TEST
Attitude Director Indicator (ADI) Figure 24
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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).
ROLL SCALE AIRCRAFT SYMBOL
20
20
20
20
POWER “OFF” FLAG
D
PITCH SCALE C
PITCH ERECTION/ TRIM KNOB
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.
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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 fasterection 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.
HEADING SCALE
LUBBER LINE
180
170
CAGING/SETTING KNOB
Directional Indicator Figure 26
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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.
VERTICAL GIMBAL RING
ROTOR ASSEMBLY
INNER GIMBAL RING
COMPASS CARD
SYNCHRONISER RING CAGING/ SETTING KNOB
Directional Gyro Figure 27
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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 2 - Turn Rate 360º per minute. Rate 3 - Turn Rate 540º per minute. Rate 4 - Turn Rate 720º per minute. Figure 28 shows a typical Turn & Slip indicator.
RATE OF TURN INDICATOR
2 MIN SLIP/SLIDE INDICATOR
RATE OF TURN 2 MIN - 360º
Turn & Slip Indicator Figure 28
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For the detection of rates of turn, a rate gyroscope is used and is arranged in the manner shown in figure 29.
INPUT AXIS
FWD Y1
X
F Y
P
X1
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.
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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/overbanked), 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.
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Figure 30 shows bank indication for various aircraft bank conditions.
Bnk Indications Figure 30
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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.
AIRCRAFT SYMBOL
TURN COORDINATION
RATE OF TURN
L
R 2 MIN NO PITCH INFORMATION
TURN COORDINATOR
Turn co-ordinator Indicator Figure 31
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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
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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.
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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.
GEOGRAPHICAL NORTH POLE MAGNETIC NORTH POLE
11º W VARIATION 17.5º E VARIATION
0º E VARIATION
True North & Magnetic North Figure 33
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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
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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.
B
C
CO-EFFICIENT “B” ADJUSTMENT
CO-EFFICIENT “C” ADJUSTMENT
21
S
15 12
CO-EFFICIENT “A” ADJUSTMENT
LUBBER LINE
ELECTRICAL CONNECTION FOR LIGHTING
Standby Compass Figure 35
SYSTEMS
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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.
LAMINATED COLLECTOR HORNS
A
A
AC POWER
B
EXCITER COIL
C
B C
SECONDARY PICK-OFF COILS
Flux Valve Construction Figure 36
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The paths taken by the earth’s magnetic field through the spokes for different headings is shown at Figure 37.
PATH OF EARTH’S FIELD
A
A
C
B
B
C
C
B
C A
B A
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.
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Figure 38 shows the arrangement of a remote reading gyro compass system.
115v 400 Hz
B
+
_
6
E 12
21 VOR
S 15
A D F
_+
3
W 3 0 24
N 33
C
A D F
VOR
SLAVED
DG
VOR/ADF
SYNC
Gyro Magnetic Compass System Figure 38
SYSTEMS
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Figure 39 shows a schematic of a Gyro Magnetic Compass system.
CT N
M
W
E
26V AC 400 Hz
S
TG
CX CT M GYRO
DETECTOR UNIT
Gyro Magnetic Compass System Schematic Figure 39
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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 AIRCRAFT LONGITUDINAL AXIS
VANE ARM ANGLE OF ATTACK TRANSDUCER
∝ FLIGHT PATH
AIRFLOW
Angle of Attack Transducer Figure 40
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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.
28V DC SUPPLY
ANGLE OF ATTACK
>17.5º
M STICK SHAKER
Stall Warning System Figure 41
SYSTEMS
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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.
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Figure 53 shows a simplified resistance gauge.
INDICATOR
N S
TANK RESISTOR
+ DC POWER
FUEL TANK
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.
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Figure 54 shows a circuit of a capacitance quantity system.
TANK UNIT
EMPTY
IS
LOOP A
IB
LOOP B
REF C FULL
2 - PHASE MOTOR
DISCRIMINATION STAGE
AMPLIFIER STAGE INDICATOR
REF PHASE
AMPLIFIER UNIT
Capacitance Quantity Indicating System Figure 54
SYSTEMS
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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.
PSI SCALE 2.5 PSI
LBS/HR SCALE
FUEL FLOW LBS/HR
195 PSI T.O. 170
50 45 55 65
L
75
100 80
R
95
LEFT ENGINE FUEL FLOW
150
RIGHT ENGINE FUEL FLOW
Fuel Flow Indicator Figure 55
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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.
10 PSI
30
POINTER FUEL PRESS
50
125 PSI 100
80
PSI SCALE
Fuel Pressure Gauge Figure 56 There are two types of pressure gauge: Bourbon Tube type. Pressure Capsule type.
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1.23.1
Bourbon Tube Fuel Pressure Indicator
Is made with a metal tube that is formed in a circular shape with a flattened crosssection. 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.
POINTER STAFF
BOURBON TUBE
ANCHOR POINT GEARING
Bourbon Tube Fuel Pressure Gauge Figure 57
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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.
DIAPHRAGM
Pressure Capsule Fuel Pressure Gauge Figure 58
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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
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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
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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
POINTER
TEMERATURE SCALE
5 3
EGT °C X 100
7
OVER-TEMP LIMIT POINTER
1 9
7 6 5 OVER-TEMPERATURE WARNING LIGHT
DIGITAL READ-OUT
EGT Indicator Figure 61
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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
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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
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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
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Figure 65 shows the engine instrument grouping for a twin engine aircraft.
0.8
EPR
EPR
1.0 1.2 1.0 0.8
1.0
1.0 1.2
1 5 0 1.4
1.4
1.6
1.6
40 20 60 N1 80 % RPM 0
20 40 60 N1 0 % RPM 80
5 EGT 3 7 °C X 100
5 EGT °C X 100 7
1
1
7 6 5
9
9
6
6
FF
1
FF X 1000
6 5 8
EGT
7 6 5
5 4 3
3 4 5 2
%RPM
9 2
9 2
EGT
EPR
100
100
3
EPR 1 5 0
1 5 0
%RPM
150
8
8
FF X 1000
8 5 6
2 1
Power plant instrument grouping Figure 65
SYSTEMS
INSTRUMENTS PAGE 59 of 59
FF
A319/A320/A321 TECHNICAL 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.
A319/A320/A321 TECHNICAL TRAINING MANUAL _ GENERAL FAMILIARIZATION COURSE
22 AUTO FLIGHT SYSTEM
22 AUTO FLIGHT SYSTEM UFD0100
TABLE OF CONTENTS
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AUTO FLIGHT SYSTEM GENERAL System Design Philosophy (1) ..................... 1 ** System Presentation (1) ....................... 5 ** System Control and Indicating (1) ............ 9 Basic Operational Principles (1) ................ 23
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22 - AUTO FLIGHT SYSTEM 22-00-00 AUTO FLIGHT SYSTEM DESIGN PHILOSOPHY
TMUFMGS01 LEVEL 1
UFD0100
CONTENTS: General Concept Navigation Flight Plan Operation AFS/Fly by Wire System Design
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22 AUTO FLIGHT SYSTEM
AUTO FLIGHT SYSTEM DESIGN PHILOSOPHY GENERAL CONCEPT
OPERATION
The Auto Flight System (AFS) calculates orders to automatically control the flight controls and the engines. The Auto Flight System computes orders and sends them to the Electrical Flight Control System (EFCS) and to the Full Authority Digital Engine Control (FADEC) to control flying surfaces and engines. When the AFS is not active, the above mentioned components are controled by the same systems but orders are generated by specific devices (i.e. side sticks and thrust levers).
There are several ways to use the Auto Flight System. The normal and recommended way to use the Auto Flight System is to use it to follow the flight plan automatically. Knowing the position of the aircraft and the desired flight plan (chosen by the pilot), the system is able to compute the orders sent to the surfaces and engines so that the aircraft follows the flight plan. The pilot has an important monitoring role. NOTE: During Auto Flight System operation, side sticks and thrust levers do not move automatically.
NAVIGATION AFS/FLY BY WIRE A fundamental function of the Auto Flight System is to calculate the position of the aircraft. When computing the aircraft position, the system uses several aircraft sensors giving useful information for this purpose.
If the pilot moves the side stick when the Auto Flight System is active, it disengages the autopilot. Back to manual flight, when the side stick is released, the Electrical Flight Control System maintains the actual aircraft attitude.
FLIGHT PLAN
TMUFMGS01-T01 LEVEL 1
UFD0100
SYSTEM DESIGN The system has several flight plans in its memory. These are predetermined by the airline. A flight plan describes a complete flight from departure to arrival, it includes vertical information and all intermediate waypoints. It can be displayed on the instruments (CRTs).
EFFECTIVITY EFFECTIVITY ALL
To meet the necessary reliability, the AutoFlight System is built around four computers: Two interchangeable Flight Management and Guidance Computers (FMGCs) and two interchangeable Flight Augmentation Computers (FACs). It is a FAIL OPERATIVE system. Each Flight Management and Guidance Computer and each Flight Augmentation Computer has a command part and a monitor part to be FAIL PASSIVE.
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22 AUTO FLIGHT SYSTEM
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22 - AUTO FLIGHT SYSTEM 22-00-00 AUTO FLIGHT SYSTEM PRESENTATION
TMUFGCA01 LEVEL 1
UFD0100
CONTENTS: General Controls FMGCs FACs Other Systems Self Examination
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22 AUTO FLIGHT SYSTEM
GENERAL
FMGCs
The Auto Flight System (AFS) provides the pilots with functions reducing their workload and improving the safety 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).
There are two interchangeable FMGCs. Each FMGC is made of two parts: the Flight Management part called FM part and the Flight Guidance part called FG part. The Flight Management part provides functions related to flight plan definition, revision and monitoring. The Flight Guidance part provides functions related to the aircraft control.
CONTROLS
FACs
The FCU and the MCDUs enable the pilots to control the functions of the FMGCs. The FAC engagement pushbuttons and the rudder trim control panel are connected to the FACs. The MCDUs are used for long-term control of the aircraft and provide the interface between the crew and the FMGC allowing the management of the flight. The FCU is used for short term control of the aircraft and provides the interface required for transmission of engine data from the FMGC to the Full Authority Digital Engine Control (FADEC).
The basic functions of the FACs are the rudder control and the flight envelope protection. NOTE: The FAC includes an interface between the Auto Flight System and the Centralized Fault Display System (CFDS) called Fault Isolation and Detection System (FIDS). This function is activated only in position 1 (FAC 1). OTHER SYSTEMS The Auto Flight System is connected to the majority of the aircraft systems. Examples of Auto Flight System data exchanges: - Reception of the aircraft altitude and attitude from the Air Data and Inertial Reference System (ADIRS). - Transmission of autopilot orders to the Elevator and Aileron 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. the The The The
FMGC functions controlled from? MCDUs and rudder trim control panel. FCU and rudder trim control panel. FCU and MCDUs.
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Where are A B C -
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22 - AUTO FLIGHT SYSTEM 22-00-00 AUTO FLIGHT SYSTEM CONTROL AND INDICATING
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CONTENTS: FCU MCDUs NDs PFDs Thrust Levers Side Sticks Rudder Pedals Resets RMPs EWD/SD Attention Getters
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AUTO FLIGHT SYSTEM CONTROL AND INDICATING FCU
MCDUs
The Flight Control Unit (FCU) is installed on the glareshield. The FCU front face includes an Auto Flight System (AFS) control panel between two Electronic Flight Instrument System (EFIS) control panels. The AFS control panel allows and displays the engagement of autopilots (APs) and autothrust (A/THR), and the selection of guidance modes and flight parameters.
Two Multipurpose Control and Display Units (MCDUs) are located on the center pedestal. The MCDU is the primary entry/display interface between the pilot and the FM part of the FMGC. MCDU allows system control parameters and flight plans to be inserted, and is used for subsequent modifications and revisions. The MCDU displays information regarding flight progress and aircraft performances for monitoring and review by the flight crew.
NOTE: The EXPEDite pushbutton can be optionally removed from the AFS control panel.
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The two EFIS control panels control and display, for each EFIS side (Capt and F/O), the Primary Flight Display and Navigation Display functions (respectively baro and Flight Director (FD) conditions, and Navigation Display modes).
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NDs
PFDs
The two Navigation Displays (NDs) are located on the main 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.
The two Primary Flight Displays (PFDs) are located on the main instrument panel. The Flight Mode Annunciator (FMA) is the top part of the 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 speed scale.
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AUTO FLIGHT SYSTEM CONTROL AND INDICATING 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 pushbuttons located on the thrust levers allow the autothrust function to be disengaged. SIDE STICKS The Capt and F/O side sticks are respectively located on the Capt lateral panel and F/O lateral panel. The autopilot is disengaged when the take over priority pushbutton on the side stick is pressed or when a force above a certain threshold is applied on the side stick. RUDDER PEDALS
TMUFGCH01-T03 LEVEL 1
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The rudder pedals are fitted in the Capt and F/O positions. Rudder pedals override disconnects the autopilot.
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AUTO FLIGHT SYSTEM CONTROL AND INDICATING RESETS
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The FMGC, FAC, FCU and MCDU resets are possible in the cockpit. Depending on the computer (1 or 2), the circuit breakers are located either on the overhead circuit breakers panel 49VU or on the rear circuit breakers panel 121VU.
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AUTO FLIGHT SYSTEM CONTROL AND INDICATING RMPs The Radio Management Panels (RMPs) are located on the center pedestal near Multipurpose Control and Display Units 1 and 2. The RMPS are used for navaid standby selection. EWD/SD
TMUFGCH01-T05 LEVEL 1
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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 inoperative systems on the STATUS page or landing capabilities availability.
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AUTO FLIGHT SYSTEM CONTROL AND INDICATING ATTENTION GETTERS
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The attention getters are located on the glareshield panel on the Capt and F/O sides. The MASTER CAUTION and/or the MASTER WARNING are activated when an AFS disconnection occurs. The AUTOLAND warning is activated when a problem occurs during final approach in automatic landing.
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22 - AUTO FLIGHT SYSTEM 22-00-00 BASIC OPERATIONAL PRINCIPLES
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CONTENTS: General Data Base Loading Power-up Test FD Engagement MCDU Initialization A/THR Engagement AP Engagement Self Examination
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BASIC OPERATIONAL PRINCIPLES GENERAL
POWER-UP TEST FD ENGAGEMENT
This sequence describes the operational use of the Flight Management and Guidance Computers (FMGCs) in a normal operation with a total availability of the concerned functions. The short-term pilot orders are entered through the Flight Control Unit (FCU). The long-term pilot orders are entered through the Multipurpose Control and Display Unit (MCDU). Four key-words for the control principle and both types of guidance are to be kept in mind in order to avoid handling errors. Aircraft control is AUTOMATIC (Autopilot or autothrust), or MANUAL (Pilot action on side sticks or on thrust levers). Aircraft guidance is MANAGED (Targets are provided by the FMGC), or SELECTED (Guidance targets are selected by the pilot through the FCU).
As soon as electrical power is available, the Flight Director (FD) is automatically engaged provided that the power-up test is successful. No guidance symbols are displayed as long as no AP/FD mode is active.
DATA BASE LOADING
TMUFGCQ01-T01 LEVEL 1
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The data base must be loaded and updated to keep the system operational.
MCDU INITIALIZATION First, MCDU STATUS page is displayed. Then, the pilot uses 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 FLEX TEMP). V2, at least, must be inserted in the MCDU before take-off. Entry of the flight plan (lateral and vertical) and V2 into the MCDU is taken into account by the Flight Management (FM) part and confirmed by the lighting of the associated lights on the FCU.
NOTE: Only the navigation data base is periodically updated.
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BASIC OPERATIONAL PRINCIPLES A/THR ENGAGEMENT
AP ENGAGEMENT
Autothrust (A/THR) engagement occurs when the pilot moves the thrust levers to the TO/GA or FLX/MCT gate. Then: . the FMGC automatically engages: - the take-off modes for yaw and longitudinal guidance (RunWaY (RWY) and Speed Reference System (SRS)), - the autothrust function (but it is not active). . the FD symbols appear on the PFD (Green FD yaw bar and pitch bar). For take-off, the thrust levers are set to the TO/GA gate or the FLEX/MCT gate if a flexible temperature has been entered on the MCDU. At the thrust reduction altitude, the FM part warns the pilot to set the thrust levers to CLB gate.
Either autopilot (AP) can only be engaged 5 seconds after lift off. Only one autopilot can be engaged at a time, the last in, being the last engaged. After the normal climb, cruise and descent phases, selection of LAND mode (Autoland) allows both APs to be engaged together. After touchdown, during ROLL OUT mode, APs remain engaged to control the aircraft on the runway centerline. Then the pilot disengages the APs at low speed, taxies and stops the aircraft.
TMUFGCQ01-T01 LEVEL 1
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NOTE: The thrust levers normally will not leave this position until an audio message "RETARD" requests to the pilot to set the thrust levers to IDLE gate before touchdown.
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SELF EXAMINATION When is A B C
FD engaged? - As soon as at least one AP is engaged. - As soon as A/THR is engaged. - At the end of a successful power-up test.
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Concerning AP engagement, which of the following is true? A - Both APs can be engaged whatever the flight phase. B - During the approach phase, it is recommended to engage the second AP. C - Both APs can never be engaged at the same time (Last in, last engaged).
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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.
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TABLE OF CONTENTS
Page
GENERAL ** General System Presentation (1)............... 1 ** System Control and Indicating (1) ............ 5 SPEECH COMMUNICATION Speech 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 SATCOM MCS 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
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PAX ENTERTAINMENT SYSTEM ** Passenger Entertainment SYS Pres. (1) ....... 81
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23 - COMMUNICATIONS 23-00-00 GENERAL SYSTEM PRESENTATION
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CONTENTS: VHF HF (Option) SELCAL (SELective CALling) CIDS Passenger Address Interphone Cockpit Voice Recorder
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GENERAL SYSTEM PRESENTATION VHF
INTERPHONE
The Very High Frequency (VHF) system serves for short range voice communications.
The purpose of the SELCAL system is to give visual and aural indications to the crew, concerning calls received from ground stations through VHF and HF systems.
There are 3 interphone systems on the aircraft: the flight interphone, the cabin interphone and the service interphone. - The flight interphone system allows communication between the flight crew members, and between the flight crew and the ground mechanic at the external power receptacle or in the avionics bay. - The cabin interphone system allows communication between the cockpit and the cabin attendant stations, and between the cabin attendant stations. - The service interphone system enables communication between the different service interphone jacks, the cockpit and the cabin attendant stations.
CIDS
COCKPIT VOICE RECORDER
The Cabin Intercommunication Data System (CIDS) is designed to interface flight crew, cabin attendants, passengers, ground service and various cabin systems dedicated to cabin attendant or passenger use. The CIDS is used to control, test and monitor various cabin systems dedicated to cabin attendant or passenger use.
The Cockpit Voice Recorder (CVR) records in-flight and on-ground crew conversations and radio communications.
HF (Option) The High Frequency (HF) system serves for all long-distance voice communications between different aircraft (in flight or on the ground), or between the aircraft and one or several ground stations.
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SELCAL (SELective CALling)
PASSENGER ADDRESS The Passenger Address (PA) allows voice announcements to be broadcast to all passengers, from the cockpit and cabin attendant stations through the CIDS.
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23 - COMMUNICATIONS 23-00-00 SYSTEM CONTROL AND INDICATING
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CONTENTS: Cockpit Cabin Avionics Bay Nose Landing Gear
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SYSTEM CONTROL AND INDICATING COCKPIT
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In the cockpit, we find: - 3 Audio Control Panels (ACPs) for the selection of communication systems (in transmission and reception) and for the control of the received audio signal levels, - 3 Radio Management Panels (RMPs) for the selection of radio communication and navigation frequencies, - 1 AUDIO SWITCHING selector for the reconfiguration of channels, in case of ACP failure, - 1 CALLS panel for flight crew-to-ground mechanic or flight crew-to-cabin attendant calls, - 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 for headsets and boomsets.
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SYSTEM CONTROL AND INDICATING CABIN Panels are installed in the cabin, for the control and monitoring of the various cabin systems: - The Forward Attendant Panel (FAP) is located in the forward entrance area of the aircraft. The cabin attendants can control the different cabin systems from here. - 3 Additional Attendant Panels (AAPs) can be installed near the doors and are dedicated to cabin zones. 1 AAP is basically installed near the aft passenger crew door. Note: The F.A.P. and the A.A.P. are customized per airline request.
- The Programming and Test Panel (PTP) is
TMUCOG202-T02 LEVEL 1
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located at the forward attendant station, behind a hinged access door next to the FAP. The PTP enables to test and re-program the Cabin Intercommunication Data System.
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SYSTEM CONTROL AND INDICATING AVIONICS BAY In the avionics bay, the SELCAL code panel is installed for coding the SELCAL code assigned to the aircraft. NOSE LANDING GEAR
TMUCOG202-T03 LEVEL 1
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On the external power control panel, some features are dedicated to ground mechanic-to-flight crew or flight crew-to-ground mechanic calls.
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23 - COMMUNICATIONS 23-51-00 SPEECH COMMUNICATION PRESENTATION
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CONTENTS: COM/NAV Systems RMP ACP AMU SELCAL Static Discharging Self Examination
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SPEECH COMMUNICATION PRESENTATION SYSTEMS
AUDIO MANAGEMENT UNIT
Communication and navigation systems are connected to the AMU for analog inputs and to the RMP for frequency selection.
The audio management unit (AMU) ensures the interface between the user (jack panel and ACP) and the various radio communication and radio navigation systems. The Audio Management Unit is equipped with a TEST circuit (BITE) which enables connection to the CFDIU. The AMU ensures the following functions: - Transmission - Reception - SELCAL and display of ground crew and Cabin Attendant calls - Flight interphone - Emergency function for the Captain and First Officer stations.
RADIO MANAGEMENT PANEL The radio management panels (RMP) centralize radio communication frequency control. RMP 1 and RMP 2 can also serve as backups for the flight management and guidance computers (FMGC) for radio navigation frequency control (VOR, DME, ILS, ADF). The aircraft is equipped with three RMPs which are identical and interchangeable. The 3rd RMP is optional.
TMUCOMA02-T01 LEVEL 1
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AUDIO CONTROL PANEL The ACPs supplies the means: - to use the various radio communication and radio navigation facilities installed on the aircraft for transmission and reception of the audio signals. - to display the various calls (SELCAL, ground crew call and calls from the Cabin Attendants). The ACPs serve only for control and indication.
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SELCAL The selective calling system provides visual and aural indication of calls received from ground stations. STATIC DISCHARGING The purpose of the static discharges is to discharge static electricity and to prevent interference of communication systems.
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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 be selected.
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What is the purpose of the AMU? A - To centralize all the audio signals and the frequencies of the communication systems. B - To act as an interface between the users and the various radio communication and radio navigation systems. C - To receive audio signals only.
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23 - COMMUNICATIONS 23-13-00 RADIO MANAGEMENT PANEL PRESENTATION
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CONTENTS: Radio Management Panel (RMP) Description Windows Transfer P/B Communication Keys SEL Indicator Dual Selector Knob Navigation Keys ON/OFF Switch Self Examination
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RADIO MANAGEMENT PANEL PRESENTATION RADIO MANAGEMENT PANEL (RMP) DESCRIPTION
COMMUNICATION KEYS
The RMPs are used for the selection of radio communication frequencies. They are also used for the standby selection of radio navigation frequencies in back-up mode. 3 RMPs are used for frequency selection, each one can control any VHF (HF) frequency. Note: the third RMP is optional. The 3 RMPs permanently dialog so that each RMP is informed of the last selection made on any of the other RMPs. If two RMPs fail, the remaining RMP controls all the VHF transceivers. The transmission of data to the communication and navigation systems and the dialog between the RMPs are performed through ARINC 429 buses.
There are 6 pushbutton keys for the radio communication systems. 3 of them are used for VHF, the 3 others for HF. The AM key controls the selection of the AM mode for HF transceivers provided that HF1 or HF2 is selected. When a key is pressed, the relevant active and the standby frequencies are automatically displayed in the dedicated windows.
There are 2 display windows: - The active window displays the operational frequency - The standby/course window displays the standby frequency or the course in back-up navigation mode. The windows are liquid crystal displays with a high contrast.
Although one RMP can control frequencies of any transceiver, each RMP has dedicated systems. The normal configuration is: - RMP1 allocated with VHF1, VHF3 and HF1, if installed, - RMP2 allocated with VHF2 and HF2, if installed. If the optional RMP3 is installed, it will be allocated with VHF3 and HF systems which are no longer dedicated to RMP1 or RMP2. The SEL indicator light will come on white on the RMPs involved, when an RMP takes control of a non dedicated system frequency selection. For example, if VHF2 is selected on RMP1, the SEL indicator lights come on on RMP1 and RMP2.
TRANSFER P/B
DUAL SELECTOR KNOB
When the transfer key is pressed, the standby frequency becomes the operational frequency, and the operational frequency becomes the standby frequency.
The dual selector knob is used for the selection of the frequency/course displayed in the standby/course window.
TMUCOMB03-T01 LEVEL 1
WINDOWS
UFD0100
SEL INDICATOR
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RADIO MANAGEMENT PANEL PRESENTATION NAVIGATION KEYS The NAVigation guarded pushbutton key allows the radio navigation systems to be selected, in back-up mode only, when the Flight Management and Guidance Computers (FMGCs) have failed. In radio navigation back-up mode, only RMP1 and RMP2 can perform navigation frequency/course selection using the dual selector knob. ON/OFF SWITCH
TMUCOMB03-T01 LEVEL 1
UFD0100
The latching ON/OFF switch allows the crew to set the RMP on or off.
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SELF EXAMINATION
TMUCOMB03 LEVEL 1
UFD0100
What happens if RMP2 fails? A - The communication systems are inoperative B - VHF2 frequencies cannot be controlled. C - All communication frequencies can be controlled.
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23 - COMMUNICATIONS 23-51-00 AUDIO SYSTEM PRESENTATION
TMUCOMH02 LEVEL 1
UFD0100
CONTENTS: General Transmission Reception Flight Interphone Selective Calling (SELCAL) Calls Self Examination
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AUDIO SYSTEM PRESENTATION GENERAL
INTERPHONE
The AMU centralizes the Audio Signals used by the crew. The crew controls and operates these functions independently with the Audio Control Panels. The audio management system provides: - radio communication and navigation for crew utilization - flight interphone system - selective calling system (SELCAL) - visual indication of ground crew and cabin attendant calls. Each cockpit occupant Audio Equipment includes: - oxygen mask, - headset, - boomset, - handmicrophone, except for the 4th occupant which is only equipped with a jack box.
The flight interphone function allows interpone links between the various crew stations in the cockpit and with the groud crew through the jack at the external power receptacle panel (108 VU) and the avionics compartment jack panel (63 VU). SELCAL The Selective Calling system enables reception with aural and visual indication of calls from ground stations equipped with a coding device NOTE:
The SELCAL decoding unit is located inside the AMU.
CALLS Cabin attendant and mechanic calls are indicated on the Audio Control Panels.
TRANSMISSION
TMUCOMH02-T01 LEVEL 1
UFD0100
In transmission mode, the AMU collects microphone inputs of the various crew stations and directs them to the communication transceivers. RECEPTION In reception mode, the AMU collects the audio outputs of the communication transceivers and navigation receivers and directs them to the various crew stations.
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SELF EXAMINATION
TMUCOMH02 LEVEL 1
UFD0100
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.
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23 - COMMUNICATIONS 23-51-00 AUDIO CONTROL PANEL PRESENTATION
TMUCOMI01 LEVEL 1
UFD0100
CONTENTS: General Transmission Keys Reception Knob Interphone/Radio Selector Switch Voice Filter Reset Passenger Address Self Examination
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AUDIO CONTROL PANEL PRESENTATION GENERAL
RECEPTION KNOB
3 Audio Control Panels (ACPs) are provided in the cockpit for the Captain, the First Officer and the third occupant. Each ACP allows: - the use of various radio communication and radio navigation facilities installed in the aircraft for transmission and reception of the audio signals, - the display of various calls received through the SELCAL system, from ground mechanics and from cabin attendants, - the use of flight, cabin and service interphone systems.
Fifteen pushbutton knobs are used to select reception and to adjust the volume of received signals. When the reception channel is selected, the pushbutton knob pops out and comes on white.
TMUCOMI01-T01 LEVEL 1
UFD0100
TRANSMISSION KEYS Eight rectangular electronic keys are used for the selection of the transmission channel and for the display of various calls received through SELCAL system, from ground mechanics and from cabin attendants. MECH light on the INTerphone key flashes amber to indicate a ground mechanic call. ATT light on the CABin key flashes amber to indicate a cabin attendant call. NOTE: Only one transmission channel can be selected at a time.
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INTERPHONE/RADIO SELECTOR SWITCH The INTerphone/RADio selector switch permits the utilization of the interphone or the radio, when the boomsets or oxygen masks are used by the crew. The INT position allows direct flight interphone transmission: - whatever the transmission key selected and provided 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 when a transmission key is selected. VOICE FILTER A voice filter can be used on the ADF and VOR channels. When used, the identification signals transmitted by the navaids are greatly attenuated (32 dB) so as to hear only voice signals. ON comes on green when the voice filter is in service (ON VOICE key pressed in).
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AUDIO CONTROL PANEL PRESENTATION RESET The RESET key is used to cancel all the lighted calls. NOTE: MECH and ATT lights go off automatically after 60 seconds if the call is not cancelled by the RESET key. PASSENGER ADDRESS
TMUCOMI01-T01 LEVEL 1
UFD0100
A key enables the selection of the Passenger Address transmission. This key must be pressed in during the whole transmission. An AMU pin program can inhibit the unstable operation of the PA key.
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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 transmit simultaneously on Passenger Address and VHF channels? A - Yes. B - No.
TMUCOMI01 LEVEL 1
UFD0100
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.
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23 - COMMUNICATIONS 23-12-00 VHF SYSTEM PRESENTATION
TMUCOMF02 LEVEL 1
UFD0100
CONTENTS: Purpose Principle Components
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VHF SYSTEM PRESENTATION PURPOSE The VHF system allows short distance voice communications between different aircraft (in flight or on ground) or between the aircraft and a ground station. The VHF is used for short range voice communications.
TMUCOMF02-T01 LEVEL 1
UFD0100
PRINCIPLE For voice communications, the crew uses acoustic equipment. - side-stick radio selectors, - loudspeakers, - oxygen-masks, - boomsets, - headsets, - hand-microphones. The Audio Management Unit (AMU) acts as an interface between the crew and the VHF system. The Audio Control Panels (ACPs) allow selection of the VHF1,VHF2, or VHF3 transceiver in transmission or reception mode and for the control of the received audio signal. The Radio Management Panels (RMPs) serve to select the VHF frequencies. The VHF transceiver, tuned on the frequency selected by one of the 3 Radio Management Panels (RMPs), transforms the audio signals into VHF signals (in transmission mode) or VHF signals into audio signals (in reception mode).
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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).
TMUCOMF02-T02 LEVEL 1
UFD0100
NOTE : RMP 3 is optional.
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23 - COMMUNICATIONS 23-51-00 SELCAL SYSTEM PRESENTATION
TMUCOML01 LEVEL 1
UFD0100
CONTENTS: SELCAL Philosophy SELCAL Operation Self Examination
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SELCAL SYSTEM PRESENTATION SELCAL PHILOSOPHY The selective calling system provides visual and aural indication of calls received from ground stations equipped with a coding device. The ground station tone generator provides the assigned aircraft code which modulates a VHF (or an HF) transmitter. In order to receive the SELCAL CALL, the same frequency as on the ground must be activated in the aircraft. SELCAL: SELective CALling system This function is integrated in the AMU. The A/C code can be set on the SELCAL code panel fitted in the avionics bay.
TMUCOML01-T01 LEVEL 1
UFD0100
SELCAL OPERATION When a selcal call is received, the CALL light flashes amber on the corresponding transmission key and a buzzer sound is heard. The buzzer signal is generated by the Flight Warning Computer (FWC). CALL flashes amber on all the ACPs when a selcal call is received. The CALL indication can be manually cleared by pressing the RESET key on any ACP or it can be automatically cleared upon transmission on the called channel.
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SELF EXAMINATION
TMUCOML01 LEVEL 1
UFD0100
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.
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23 - COMMUNICATIONS 23-42-00 GROUND CREW CALL SYSTEM PRESENTATION AND OPERATION
TMUCOMO02 LEVEL 1
UFD0100
CONTENTS: Ground Mechanic to Flight Crew Call Flight Crew to Ground Mechanic Call Self Examination
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GROUND CREW CALL SYSTEM PRESENTATION AND OPERATION The ground crew call system enables flight crew to ground mechanic or ground mechanic to flight crew calls. GROUND MECHANIC TO FLIGHT CREW CALL When the COCKPIT CALL pushbutton is pressed in on panel 108VU, the MECH light flashes amber on all ACPs and a buzzer is heard. An action on the RESET key of any ACP will make all the MECH lights go off. NOTE: MECH lights go off automatically after 60 seconds if the call is not cancelled by the RESET key. FLIGHT CREW TO GROUND MECHANIC CALL
TMUCOMO02-T01 LEVEL 1
UFD0100
The horn sounds in the nosewheel well as long as the MECH pushbutton is pressed in on the cockpit CALLS panel, and the COCKPIT CALL blue light on panel 108VU stays on. The RESET pushbutton on panel 108VU makes the COCKPIT CALL blue light go off.
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SELF EXAMINATION
TMUCOMO02 LEVEL 1
UFD0100
How is a ground mechanic to flight crew call indicated in the cockpit? A - An ECAM message is displayed and a buzzer sounds. B - The MECH light flashes on the captain ACP and a buzzer sounds. C - The MECH light flashes on all ACPs and a buzzer sounds.
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23 - COMMUNICATIONS 23-71-00 COCKPIT VOICE RECORDER SYSTEM PRESENTATION
TMUCOMY01 LEVEL 1
UFD0100
CONTENTS: General Components Recorder Panel Self Examination
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COCKPIT VOICE RECORDER SYSTEM PRESENTATION GENERAL The Cockpit Voice Recorder (CVR) records the last 30 minutes of crew conversations and communications. It records automatically in flight and on ground when at least one engine is running and for 5 minutes after the last engine is shut down. The CVR can also operate in manual mode on the ground.
TMUCOMY01-T01 LEVEL 1
UFD0100
COMPONENTS The components of the Cockpit Voice Recorder system are: - The Cockpit Voice Recorder, located in the aft section of the aircraft. - The CVR microphone, used for recording the direct conversations between crew members in the cockpit and all aural warnings. It is located at the bottom of the overhead panel. - The recorder (RCDR) panel, providing CVR controls for manual operation, test and erasure of the recording. It is located on panel 21VU on the overhead panel. - The CVR HEADSET jack mounted on the cockpit maintenance panel 50VU.
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COCKPIT VOICE RECORDER SYSTEM PRESENTATION RECORDER PANEL GROUND CONTROL The CVR is automatically energized in flight and on ground when at least one engine is running and for 5 minutes after the last engine is shut down. For manual control, on the ground, the CVR has to be energized by pressing the ground control (GND CTL) pushbutton on the recorder (RCDR) panel. CVR TEST When the CVR TEST pushbutton is pressed, either on ground or in flight, a test tone is generated 4 times for approximately 0.8 seconds. A headset connected to the CVR HEADSET jack mounted on the cockpit maintenance panel enables monitoring.
TMUCOMY01-T02 LEVEL 1
UFD0100
CVR ERASE The CVR ERASE pushbutton is used for manual erasure of the recording, only on ground with parking brake applied. It must be pressed for at least 2 seconds. For complete manual erasure of the recording, the CVR has to be energized.
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SELF EXAMINATION
TMUCOMY01 LEVEL 1
UFD0100
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.
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23 - COMMUNICATIONS 23-51-00 FLIGHT INTERPHONE SYSTEM OPERATION
TMUCOMN01 LEVEL 3
UFD4200
CONTENTS INT Selection RAD Selection INT Key and Knob Self Examination
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FLIGHT INTERPHONE SYSTEM OPERATION INT SELECTION
INT KEY and KNOB
The INT position of the INT/RAD selector switch enables permanent use of the flight interphone without any further action and whatever the radio key selected (Here VHF 1). This is a stable position.
The flight interphone can also be used like a VHF transceiver. Selection of the INT transmission key lights the green bars, indicating that the flight interphone is ready to operate. Pressing and releasing the INT reception knob enables adjustment of the interphone level. If done, the knob comes on white. Placing and holding the INT/RAD switch in RAD position enables the operator to talk through the flight interphone system.
NOTE : The radio function has priority over the flight interphone function. So, even with the INT/RAD switch in INT position, the flight interphone is momentarily cut during a radio emission ( Radio key selected and hand microphone or side-stick Push To Talk actuated). RAD SELECTION
TMUCOMN01-T01 LEVEL 3
UFD4200
The RAD position of the INT/RAD selector switch puts the preselected channel in emission (Here VHF 1). This is an unstable position. This position acts like the selection of the hand microphone pushbutton or like the Push To Talk pushbutton of the side-stick.
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UFD4200
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FLIGHT INTERPHONE SYSTEM OPERATION SELF EXAMINATION
TMUCOMN01 LEVEL 3
UFD4200
Which action must be performed to talk through the flight interphone system ? A - Pressing the INT transmission key and with the INT/RAD selector switch to neutral position. B - Either setting the INT/RAD selector to INT, or pressing the INT transmission key and setting the INT/RAD selector switch to RAD. C - Pressing any radio transmission key and INT transmission key together.
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A319/A320/A321 TECHNICAL TRAINING MANUAL GENERAL FAMILIARIZATION COURSE 34 NAVIGATION
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.
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34 NAVIGATION UFD0100
TABLE OF CONTENTS
Page
NAVIGATION GENERAL System 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 ADIRS ADIRS Principle (1) .............................. 41 ADIRS Presentation (1) .......................... 55 ** Air Data Probes Presentation (1) ............ 61 MULTI MODE RECEIVER (MMR) SYSTEM MMR System Description (1) ...................... 65 RADIO ALTIMETER (RA) SYSTEM Radio 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) SYSTEM DME System Presentation (1) .................... 105 AIR TRAFFIC CONTROL (ATC) SYSTEM ATC System Presentation (1) .................... 113
UFD0100
AUTOMATIC DIRECTION FINDER (ADF) SYSTEM ADF System Presentation (1) .................... 119 VOR/MARKERS SYSTEM VOR/MKR System Presentation (1) ................ 129
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34 - NAVIGATION 34-00-00 SYSTEM PRESENTATION
TMUNA2001 LEVEL 1
UFD0100
CONTENTS: General ADIRS Landing and Taxiing Aids Systems Independent Position Determining Systems Dependent Position Determining Systems
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SYSTEM PRESENTATION GENERAL
LANDING AND TAXIING AIDS
The aircraft navigation systems provide the crew with the data required for flight within the most appropriate 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.
The Head-Up Display (HUD) is used as a piloting aids system for roll out, take-off and landing (optional). The Instrument Landing System (ILS), is use to obtain the optimum aircraft position during an approach and landing phase. The Marker (MKR) system is used to indicate the distance 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 or manually (through the MCDU) by the Flight Management and Guidance Computers (FMGCs) or manually through the Radio Management Panels (RMPs).
TMUNA2001-T01 LEVEL 1
UFD0100
ADIRS The ADIRS is an integrated Air Data System and an Inertial Reference System. One part called Air Data Reference mainly computes speed and altitude information from air parameters. The other part called Inertial Reference mainly computes heading, attitude and position from gyros and accelerometers. The ADIRS is composed of three Air Data/Inertial Reference Units (ADIRUs). Besides the ADIRUs, there are still standby instruments: - Altimeter and Airspeed indicators directly supplied by pressure lines, - Standby Compass, - Standby Horizon.
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SYSTEM PRESENTATION
TMUNA2001-T01 LEVEL 1
UFD0100
INDEPENDENT POSITION DETERMINING SYSTEMS This part of the navigation systems, called independent system, provides information regarding the safety of the aircraft without taking reference from any ground station. The Weather Radar / Predictive Windshear (WR/PWS) system detects the position and intensity of precipitations which are shown on the Navigation Displays (ND’s). The windshear capability serves to detect any sudden change of wind speed and/or direction (Optional). The Radio Altimeter (RA) system gives the aircraft height above the ground, independently of the atmospheric pressure. The Traffic Collision Avoidance System (TCAS) detects the aircraft in the immediate vicinity. The Enhanced Ground Proximity Warning System (EGPWS) warns the flight crew about the aircraft behaviour in dangerous 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
The Distance Measuring Equipment (DME) system gives the aircraft slant distance to a ground station. The Air Traffic Control system (ATC) enables a ground operator to identify and track the aircraft without having to communicate with the flight crew. The Automatic Direction Finder (ADF) system is a radio compass system providing the azimuth of a Non Directional Beacon (NDB) with respect to the aircraft center line. The VHF Omni Range (VOR) system gives the bearing of a ground VOR Station with respect to the magnetic North and the aircraft angular deviation related to a preselected course. The Global Positioning System (GPS) is based on the measurement of the transmission time of signals broadcast 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 achieved either automatically or manually (through the MCDU) by the FMGCs or manually though the RMPs. NOTE 2: Although the Marker Beacon belongs to the Landing Aids System, it is physically integrated into the VOR receiver.
This part of the navigation system, called dependent system, provides various means of navigation through data exchange with ground installations or satellites.
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34 - NAVIGATION 34-00-00 RADIO NAVIGATION CONTROL PRESENTATION
TMUNAV701 LEVEL 1
UFD0100
CONTENTS: Automatic Tuning Manual Tuning Back-Up Tuning Self Examination
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RADIO NAVIGATION CONTROL PRESENTATION AUTOMATIC TUNING The Automatic Tuning permits the control of VOR / DME, ILS and ADF by the Flight Management and Guidance System. In this case the RMP is transparent to its associated FMGC. In normal operation FMGC1 tunes receivers 1, FMGC2 tunes receivers 2. In case of failure of FMGC 1 or 2, the remaining FMGC controls all receivers. MANUAL TUNING The manual tuning permits the pilot to select, through the Multipurpose Control Display Unit, a specific frequency for display on the EFIS. NOTE: To return to the autotuning mode, the manual tuning has to be cleared.
TMUNAV701-T01 LEVEL 1
UFD0100
BACK-UP TUNING Radio Management Panels 1 and 2 located on the pedestal provide back-up for Radio Navigation tuning. We are in the case of both FMGCs inoperative or emergency electrical supply. The ILS course and frequency are the only Radio Navigation data exchanged. The selected values on RMP 1 and RMP 2 are identical for ILS 1 and ILS 2.
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SELF EXAMINATION Can VOR 2 frequency be changed through RMP 1? A - Yes. B - No.
TMUNAV701 LEVEL 1
UFD0100
In back-up mode, an ILS can be tuned through: A - RMP 1 or 2. B - RMP 1 only. C - The onside RMP only.
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34 - NAVIGATION 34-00-00 SYSTEM CONTROLS PRESENTATION
TMUNA2101 LEVEL 1
UFD0100
CONTENTS: Multipurpose Control Display Unit (MCDU) ADIRS Control Display Unit (ADIRS CDU) Radio Management Panel (RMP) Audio Control Panel (ACP)
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SYSTEM CONTROLS PRESENTATION MULTIPURPOSE CONTROL DISPLAY UNIT (MCDU)
TMUNA2101-T01 LEVEL 1
UFD0100
The Multipurpose Control and Display Unit (MCDU) allows the crew: - To display the Radio Navigation frequencies (automatically or manually tuned) on a specific page called RAD/NAV. - To align the Inertial Reference systems from a specific page called INIT via the FMGC. - To initiate tests for all navigation systems and for troubleshooting via the CFDIU.
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SYSTEM CONTROLS PRESENTATION ADIRS CONTROL DISPLAY UNIT (ADIRS CDU) The ADIRS Control and Display Unit allows the following functions: - To switch on the ADR and IR by setting a single control to NAV. - To disconnect the ADR output bus by a specific pushbutton. - To check the ADIRU operation - To align the IR instead of using the MCDU.
TMUNA2101-T02 LEVEL 1
UFD0100
NOTE: when set to ATT, the systems are still energized but the IR is in downgraded operation mode.
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SYSTEM CONTROLS PRESENTATION RADIO MANAGEMENT PANEL (RMP)
TMUNA2101-T03 LEVEL 1
UFD0100
The main function of the Radio Management Panels (RMP) is to control all communication frequencies. However they are also used for standby selection of Radio/NAV frequencies. The standby operation is used in case of dual FMGC failure, provided the NAV pushbutton switch has been pressed.
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SYSTEM CONTROLS PRESENTATION AUDIO CONTROL PANEL (ACP) The Audio Control Panels (ACP) enable to control the reception of all audio signals identifying the various beacons and stations.
TMUNA2101-T04 LEVEL 1
UFD0100
NOTE: DME identification signals can be selected by using the knob of the colocated VOR or ILS.
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UFD0100
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34 - NAVIGATION 34-20-00 STANDBY INSTRUMENT PRESENTATION
TMUNAVE03 LEVEL 1
UFD0100
CONTENTS: Standby Compass Standby Horizon Standby Altimeter Standby Airspeed Indicator (ASI) Metric Altimeter (Option)
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STANDBY INSTRUMENT PRESENTATION STANDBY COMPASS
TMUNAVE03-T01 LEVEL 1
UFD0100
The standby compass is located on the top windshield center post. It is stowed in normal configuration. A correction card is glued to the side of the compass assembly. Non-magnetic lamp: - A non-magnetic lamp assembly lights the compass card. It is controlled by a STBY COMPASS switch located on the INT LT panel on the overhead panel. Graduated compass: - The graduated compass card is attached to a magnetic element. It is free to rotate inside the compass bowl and 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 by positioning two small magnetic bars called compensators.
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STANDBY INSTRUMENT PRESENTATION STANDBY HORIZON
TMUNAVE03-T02 LEVEL 1
UFD0100
The Standby horizon is located on the center instrument panel and comprises the following elements: Roll pointer: - The roll information is given by a pointer which moves in front of a graduated roll scale. Roll scale: - The roll scale is graduated in 10 degree increments between -30 and +30 degrees and 15 degree increments up to 60 degrees. Flag: - The flag comes into view if a failure is detected in the electrical power supply or if the gyro rotor speed drops below 18000 RPM. Pitch drum: - The pitch drum is divided into two zones separated by 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 caging knob (Also used for shipping to protect the gyro).
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STANDBY INSTRUMENT PRESENTATION
TMUNAVE03-T03 LEVEL 1
UFD0100
STANDBY ALTIMETER The Standby Altimeter is located on the center instrument panel and comprises the following elements: Adjustable bugs: - Four manually adjustable bugs are provided for reference altitude setting. Altitude counter: - A display counter made up of two drums displays the tens of thousands, and the thousands of feet. When the altitude is below 10000 feet, the left drum displays black and white stripes. In case of negative altitude the left drum displays orange and white stripes. Altitude pointer: - The pointer indicates the hundreds of feet with 20 feet increments. To prevent the pointer from sticking, an internal vibrator is installed. It is only supplied in flight. Altitude dial: - The altitude dial is calibrated from O to 1000 feet with 20 feet graduations. Baro correction counter: - The baro correction is displayed on a counter graduated in hecto Pascals. Adjustment baro setting knob: - The knob enables adjustment of the baro setting in the range of 750 to 1050 hecto Pascals.
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STANDBY INSTRUMENT PRESENTATION STANDBY AIRSPEED INDICATOR (ASI)
TMUNAVE03-T04 LEVEL 1
UFD0100
The Standby Airspeed Indicator (ASI) is located on the center instrument panel and comprises the following elements: Ajustable bugs: - Four manually adjustable bugs are provided for reference speed setting. Speed pointer: - The pointer moves on a graduated dial. Speed dial: - The dial is made of two linear scales: one from 60kt to 250kt with 5kt increments, the other from 250 to 450kt with 1Okt increments.
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STANDBY INSTRUMENT PRESENTATION METRIC ALTIMETER (OPTION)
TMUNAVE03-T05 LEVEL 1
UFD0100
The Metric Altimeter is located on the center instrument panel and comprises the following elements: A display counter, made up of two drums displays the tens of thousands and the thousands of meters. Altitude counter: - When the altitude is below 10,000 meters, the left drum displays black and white stripes. In case of negative altitude the left drum displays orange and white stripes. Altitude pointer: - The pointer indicates the hundreds of meters with 50 meter increments. Altitude dial: - The altitude dial is calibrated from O to 1000 meters with 50 meter graduations. Baro correction counter: - The baro correction is displayed on a the lower counter and is graduated in hecto Pascals. Adjustment baro setting knob: - The knob enables adjustment of the baro setting in the range of 870 to 1050 hecto Pascals.
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UFD0100
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34 NAVIGATION
NAVIGATION
34-00-00 RADIO MANAGEMENT PANEL (RMP) PRESENTATION
TMUNA2201 LEVEL 1
UFD0100
CONTENTS: General Standby Navigation Keys Rotating Knob Standby/Course (STBY/CRS) Window Active Window
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RADIO MANAGEMENT PANEL (RMP) PRESENTATION GENERAL The radio Nav frequency selection can only be performed when the guarded NAV key LED is on, after a dual FMGS loss The ON/OFF switch controls the power supply of the RMP. 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 the displays to the last RMP memorized values (frequency and course). At any time, communication frequencies are still selectable, simply by pressing the corresponding key. Beat Frequency Oscillator (BFO) is set on or off by pressing the key. NOTE: The Microwave Landing System (MLS) key is a provision.
The desired frequency or course is set in the STBY/CRS window. Frequency becomes active by pressing the transfer key. STANDBY/COURSE (STBY/CRS) WINDOW The STANDBY/COURSE window displays a standby frequency or a course. Both can be changed by rotating the knob, but only the standby frequency can be made active by pressing the transfer key. If a course is displayed, the associated frequency is displayed in the ACTIVE window. NOTE: If a course is displayed on the STBY/CRS window, pressing the transfer key will display the ACTIVE frequency in both windows. ACTIVE WINDOW The active window shows the frequency in use of the system identified by the green LED on the selected key.
TMUNA2201-T01 LEVEL 1
UFD0100
ROTATING KNOB Two concentric knobs allow preselection of frequency for radio communication and standby navigation systems and selection of the required course for VORs and ILSs: - the outer knob controls the most significant digits, - the inner knob controls the least significant digits.
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UFD0100
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34 - NAVIGATION 34-00-00 DDRMI PRESENTATION
TMUNAVF02 LEVEL 1
UFD0100
CONTENTS: General Normal Operation Failure and Non Computed Data (NCD) Self Examination
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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 capability.
are
not
equipped
with
the
ADF
NORMAL OPERATION
TMUNAVF02-T01 LEVEL 1
UFD0100
The DME 1 Distance is displayed in the left hand window. The DME 2 Distance is displayed in the right hand window. 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 pointer by a selector switch.
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DDRMI PRESENTATION FAILURE AND NON COMPUTED DATA (NCD)
TMUNAVF02-T02 LEVEL 1
UFD0100
When a failure is detected by the DME or RMI monitoring circuits, the corresponding DME display window is blanked. In case of Non Computed Data (NCD), for example: - out-of-range station, the window shows white horizontal dashed lines. Heading information normally comes from ADIRU 1. If it fails, the heading is provided by ADIRU 3 after pilot switching. In case of VOR or ADF 1 or 2 receiver failure, a red flag comes into view and the corresponding pointer is set to the 3 o’clock position. In case of Non Computed Data (NCD), no failure flag appears, but the corresponding pointer is set to the 3 o’clock position.
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SELF EXAMINATION the DDRMI receive information ? 2. 3. 3.
TMUNAVF02 LEVEL 1
UFD0100
From which ADIRU can A - ADIRU 1 or B - ADIRU 2 or C - ADIRU 1 or
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34 - NAVIGATION 34-10-00 ADIRS PRINCIPLE
TMUADI001 LEVEL 1
UFD0100
CONTENTS: General ADM Functional Description ADM Inputs ADM Output ADR Computation IR Strapdown Ring Laser Gyro Accelerometer IR Computation
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ADIRS PRINCIPLE GENERAL
TMUADI001-T01 LEVEL 1
UFD0100
The Air Data/Inertial Reference Unit (ADIRU) comprises an Air Data Reference Unit and an Inertial Reference Unit, both included in a single unit. Data from external sensors (Angle of Attack, Total Air Temperature, Air Data Module) are used by the ADIRU. The ADIRUs are interfaced with the ADIRS Display Control Unit (CDU) for control and status annunciation.
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ADIRS PRINCIPLE ADM FUNCTIONAL DESCRIPTION A microcomputer processes an ARINC signal according to the discrete inputs and to the digitized pressure. ADM INPUTS The ADM Inputs are one pressure input and several discrete inputs. The ADMs are identical. The discrete inputs determine the ADM location and the type of pressure data (Pitot or Static) provided to the ADR. ADM OUTPUT
TMUADI001-T02 LEVEL 1
UFD0100
The ADM output is an ARINC bus which provides digital pressure information, type of pressure, ADM identification and BITE status to the ADIRU.
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ADIRS PRINCIPLE ADR COMPUTATION
TMUADI001-T03 LEVEL 1
UFD0100
The ADR processes sensor and ADM inputs in order to provide air data to users.
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ADIRS PRINCIPLE IR STRAPDOWN
TMUADI001-T04 LEVEL 1
UFD0100
In a strapdown Inertial Reference System the gyros and the accelerometers are solidly attached to the aircraft structure. The strapdown laser gyro provides directly accelerations and angular speeds.
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ADIRS PRINCIPLE RING LASER GYRO The three ring laser gyros, one for each rotation axis, provide inertial rotation data and are composed of two opposite laser beams in a ring. At rest, the two beams arrive at the sensor with the same frequency. An aircraft rotation creates a difference of frequencies between the two beams. The frequency difference is measured by optical means providing a digital output which, after computation, will provide rotation information. Stimulated
Emission
of
TMUADI001-T05 LEVEL 1
UFD0100
NOTE: Light Amplification Radiation (LASER)
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ADIRS PRINCIPLE ACCELEROMETER Three accelerometers, one for each axis, provide linear accelerations. The acceleration signal is sent to a processor which uses this signal to compute the velocity and the position. IR COMPUTATION
TMUADI001-T06 LEVEL 1
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Each ADIRU computes the laser gyro and the accelerometer outputs to provide Inertial Reference data to users.
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34 - NAVIGATION 34-10-00 ADIRS PRESENTATION
TMUADIA01 LEVEL 1
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CONTENTS: General MCDU ADIRS CDU Probes FCU GPS DMC DMC/PFD & ND ADIRS Switching Users Self Examination
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ADIRS PRESENTATION GENERAL
FCU
The Air Data Inertial Reference System (ADIRS) is composed of three Air Data Inertial Reference Units (ADIRU), each having their own set of probes and sensors and a common Control Display Unit (CDU).
The ADIRUs receive, from the Flight Control Unit (FCU), the Baro correction set by the crew.
MCDU
The Global Positioning System (GPS) provides data to the ADIRS, mainly A/C position and speed. The ADIRS processes the GPS data and provides pure GPS data, pure IR data and hybrid GPS/ADIRS data to users.
The Multipurpose Control and Display Units (MCDUs) are normally used to align the Inertial References, to initiate the ADIRU tests and to display ADIRU information. ADIRS CDU The ADIRS Control Display Unit is used as a back-up for Inertial Reference alignment. It is also used for mode selection, information display and status indication.
GPS
DMC The Display Management Computers (DMCs) 1 and 2 receive their data from their related ADIRU and from ADIRU 3. The Display Management Computer 3 (DMC3) receives information from all three ADIRUs, to operate as a back-up in case of DMC1 or 2 failure. DMC/PFD & ND
TMUADIA01-T01 LEVEL 1
UFD0100
PROBES The Air Data input parameters, such as total and static pressures, Angle Of Attack (AOA) and Total Air Temperature (TAT) are sent, from the related probes and sensors, to the three ADIRUs.
ADIRU 1 and 2 display information via DMC 1 and 2, on the corresponding Primary Flight Display (PFD) and Navigation Display (ND). ADIRU 3 operates as a back-up in case of ADIRU 1 or 2 failure.
NOTE : static and total pressure are sent to the ADIRUs via the ADMs.
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ADIRS PRESENTATION ADIRS SWITCHING Basically, ADIRU 1 is associated to the captain instruments, ADIRU 2 to the first officer instruments and ADIRU 3 is in standby. In case of failure of the Air Data Reference (ADR) or Inertial Reference (IR) function of ADIRU 1 or 2, the affected instruments and displays may be manually switched independently to ADIRU 3 by means of selector switches. USERS
TMUADIA01-T01 LEVEL 1
UFD0100
The ADIRUs are directly connected to other user system.
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SELF EXAMINATION
TMUADIA01 LEVEL 1
UFD0100
Which ADIRU does not supply DMC 1? A - ADIRU 1. B - ADIRU 2. C - ADIRU 3.
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34 - NAVIGATION 34-13-00 AIR DATA PROBES PRESENTATION
TMUADIB01 LEVEL 1
UFD0100
CONTENTS: Pitot Probes Static Ports AOA Sensors TAT Sensors Water Drain Self Examination
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AIR DATA PROBES PRESENTATION PITOT PROBES
TAT SENSORS
The total pressure is sent from the Pitot Probes to the Air Data Modules which convert it into ARINC words used by the Air Data Inertial Reference Units. Three pitot probes provide total pressure to three Air Data Modules (ADM) which convert this pressure into digital format (ARINC 429). ARINC words are then sent to the corresponding Air Data Inertial Reference Unit (ADIRU). The standby pitot probe supplies the standby Airspeed Indicator (ASI) and ADR3 through its related ADM.
The three ADIRUs receive Total Air Temperature information from two Total Air Temperature sensors.
STATIC PORTS
NOTE: that ADIRU3 receives the Total Air Temperature (TAT) from the TAT 1 sensor which is composed of two elements. WATER DRAIN The probes are installed in such a way that their pressure lines do not require a water drain, except for that of the standby static ports.
Each Air Data Module transforms the static pressure coming from the static ports into ARINC words. Six static ports provide static pressure to five ADMs which convert this pressure into digital format (ARINC 429). The standby static ports provide an average pressure directly to the standby instruments, and to ADR3 through a single ADM.
TMUADIB01-T01 LEVEL 1
UFD0100
AOA SENSORS Each ADIRU receives angle of attack information from its corresponding Angle Of Attack (AOA) sensors. The Angle Of Attack sensors are also called Alpha probes.
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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.
TMUADIB01 LEVEL 1
UFD0100
Which pressure line(s) need(s) to be drained? A - All pitot lines. B - All static lines. C - Only the standby static line.
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34 - NAVIGATION 34-36-00 MMR SYSTEM PRESENTATION
TMUMMRA01 LEVEL 1
UFD0100
CONTENTS: General ILS Principle GPS Principle Components ILS Indicating GPS Indicating Self Examination
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MMR SYSTEM PRESENTATION GENERAL The Multi Mode Receiver (MMR) system is a Navigation Sensor with 2 internal receivers. MMR = ILS + GPS ILS PRINCIPLE
TMUMMRA01-T01 LEVEL 1
UFD0100
The function of the Instrument Landing System (ILS) is to provide the crew and airborne system users with signals transmitted by a ground station. A descent axis is determined by the intersection of a Localizer beam (LOC) and a Glide Slope beam (G/S) created by this ground station at known frequencies. The ILS allows measurement and display of angular deviations and receives the Morse audio signal which identifies the ILS ground station. ILS operational frequency range: - LOC: between 108.1 and 111.95 MHz, - G/S: between 329.15 and 335 MHz.
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MMR SYSTEM PRESENTATION GPS PRINCIPLE
TMUMMRA01-T02 LEVEL 1
UFD0100
The NAVigation System Time And Ranging Global Positioning System ( NAV.S.T.A.R. GPS ) is a worldwide navigation radio aid which uses satellite signals to provide accurate navigation information. The architecture of the system is composed of 3 parts called segments: - Spatial Segment - Control Segment - User Segment SPATIAL SEGMENT The spatial segment is composed of a constellation of 24 satellites. These satellites are arranged in six separate orbital planes of four satellites each on a circular orbit and have the following characteristics: - 55° inclination to the Equator, - an altitude of approx 20200 km with an orbital period of 12 sideral hours. These satellites give: - the satellite position (Ephemeris of the constellation), - the constellation data (Almanach). - the atmospheric corrections.
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MMR SYSTEM PRESENTATION GPS PRINCIPLE (Continued)
TMUMMRA01-T03 LEVEL 1
UFD0100
CONTROL SEGMENT The control segment is composed of four monitor stations and one master control station which track the satellites, compute the ephemeris, clock corrections and control the navigation parameters and transmit 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.
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MMR SYSTEM PRESENTATION GPS PRINCIPLE (Continued) USER SEGMENT The principle of GPS position computation is based on the measurement of transmission time of the GPS signals broadcast by at least 4 satellites. This segment is constitued by the GPS receiver and defined as follows: - signal acquisition, - distance calculation, - navigation computation (Satellite choice, positioning, propagation corrections), - detection and isolation of failed satellites (GPS PRIMARY).
TMUMMRA01-T04 LEVEL 1
UFD0100
NOTE: When GPS mode is active, no VOR/DME/ADF data is used for navigation.
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MMR SYSTEM PRESENTATION COMPONENTS
TMUMMRA01-T05 LEVEL 1
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The components are two ILS antennas, 2 GPS antennas and two MMR units. The MMR system is also connected to: - Primary Flight Display (PFD) and Navigation Display (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) for ILS manual tuning. - Captain and First Officer Radio Management Panels (RMP) for ILS back-up tuning. - Audio Control Panels (ACP) for ILS audio signal. - Air Data and Inertial Reference Unit for GPIR data.
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MMR SYSTEM PRESENTATION ILS INDICATING
TMUMMRA01-T06 LEVEL 1
UFD0100
The ILS data appears on the PFD as soon as the ILS pushbutton switch on the EFIS control panel has been pressed in and on the ND when ROSE/ILS mode has been selected. ILS information is displayed in magenta. The ILS1 information is displayed on PFD1 and ND2. The ILS2 information is displayed on PFD2 and ND1.
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MMR SYSTEM PRESENTATION GPS INDICATING
TMUMMRA01-T07 LEVEL 1
UFD0100
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.
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SELF EXAMINATION The satellite orbital planes have an inclination of: A - 60°. B - 55° to the Equator. C - 45° to the Equator.
TMUMMRA01 LEVEL 1
UFD0100
The control segment is composed of: A - 3 monitor stations. B - 5 monitor stations. C - 4 monitor stations and one master control station.
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34 - NAVIGATION 34-42-00 RADIO ALTIMETER SYSTEM PRESENTATION
TMURADG01 LEVEL 1
UFD0100
CONTENTS: Principle Components Indicating
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RADIO ALTIMETER SYSTEM PRESENTATION PRINCIPLE
TMURADG01-T01 LEVEL 1
UFD0100
The Radio Altimeter (RA) System determines the height of the aircraft above the terrain during initial climb, approach and landing phases. The RA can therefore operate over non-flat ground surface. The principle of the radio altimeter is to transmit a frequency modulated signal, from the aircraft to the ground, and to receive the ground reflected signal after a certain delay. The time between the transmission and the reception of the RA signal is proportional to the A/C height.
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RADIO ALTIMETER SYSTEM PRESENTATION COMPONENTS
TMURADG01-T02 LEVEL 1
UFD0100
The components are two transceivers, two fans, two transmission antennae and two reception antennae. The RA system is also connected to the DMCs for display on the Primary Flight Displays (PFDs).
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RADIO ALTIMETER SYSTEM PRESENTATION INDICATING
TMURADG01-T03 LEVEL 1
UFD0100
The aircraft height data is displayed on the Primary Flight Displays for heights less than or equal to 2500 ft. 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).
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34 - NAVIGATION 34-43-00 TCAS PRESENTATION
TMUTCAB05 LEVEL 3
UFD0100
CONTENTS: Principle Components Indicating Self Examination
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TCAS PRESENTATION PRINCIPLE
TMUTCAB05-T01 LEVEL 3
UFD0100
The Traffic Collision Avoidance System (TCAS) is a system whose function is to detect and display aircraft in the immediate vicinity and to provide the flight crew with indications to avoid these intruders. The TCAS indications for flight plan modifications are in the vertical plane only. The TCAS detects the air traffic control system or TCAS equipped aircraft and maintains surveillance within a range determined by its sensivity. To evaluate threat potential of other aircraft the system divides the space around aircraft into 4 volumes. - OTHER TRAFFIC VOLUME. The OTHER TRAFFIC VOLUME is the first volume providing the presence and the progress of on intruder. (No collision threat). - PROXIMATE TRAFFIC VOLUME. The proximate traffic volume is defined by a given volume around the TCAS equipped aircraft. (No collision threat, but in vicinity).
- TRAFFIC ADVISORY VOLUME (TA). When the intruder is relatively near but does not represent an immediate threat, the TCAS provides an aural and visual information known as Traffic Advisory (TA). The TCAS aural messages can be inhibited depending on higher priority aural messages. - RESOLUTION ADVISORY VOLUME (RA). When the intruder represents a collision threat, the TCAS triggers an aural and visual alarm known as Resolution Advisory (RA), which informs the crew about avoidance maneuvers.
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TCAS PRESENTATION COMPONENTS
TMUTCAB05-T02 LEVEL 3
UFD0100
The TCAS components are two antennae, one TCAS computer and one TCAS/ATC control panel. Note: The TCAS/ATC control panels shown here after are given as examples. They may differ according to the aircraft configuration.
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TCAS PRESENTATION INDICATING
TMUTCAB05-T03 LEVEL 3
UFD0100
The TCAS indications appear on the PFD and the ND. The visual resolution and traffic advisory indications are associated with aural indications such as "TRAFFIC, TRAFFIC", "CLIMB, CLIMB"... The TCAS displays only the most threatening intruders.
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SELF EXAMINATION
TMUTCAB05 LEVEL 3
UFD0100
Which aircraft are detected by the TCAS ? A - All. B - Only ATC Mode S equipped A/C. C - ATC equipped A/C.
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34 - NAVIGATION 34-48-00 ENHANCED GROUND PROXIMITY WARNING SYSTEM PRESENTATION
TMUEGPA01 LEVEL 1
UFD0100
CONTENTS: General Principle Components Indicating
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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
TMUEGPA01-T01 LEVEL 1
UFD0100
The purpose of the Enhanced Ground Proximity Warning System (EGPWS) is to help prevent accidents caused by Controlled Flight Into Terrain (CFIT). When boundaries of any alerting envelope are exceeded; aural alert messages, visual annunciations and displays are generated. The basic GPWS modes generate aural and visual warnings corresponding to an aircraft behaviour when the alert envelope is penetrated. The "ENHANCED" features complete the basic GPWS modes: - Terrain Clearance Floor (TCF): Increase the terrain clearance envelope around the airport runway. - Terrain Awareness alerting and Display (TAD): Incorporation of a terrain database to predict conflict between flight path and terrain and to display the conflicting terrain.
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ENHANCED GROUND PROXIMITY WARNING SYSTEM PRESENTATION - PRINCIPLE EFFECTIVITY EFFECTIVITY ALL
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ENHANCED GROUND PROXIMITY WARNING SYSTEM PRESENTATION COMPONENTS
TMUEGPA01-T02 LEVEL 1
UFD0100
The system comprises an EGPWC, a control panel, two warning lights and two TERRAIN ON ND mode pushbutton switches. The EGPWS is connected to various navigation systems (WR, RA, ADIRS, ILS...). It processes the navigation data and generates alarms.
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ENHANCED GROUND PROXIMITY WARNING SYSTEM PRESENTATION INDICATING
TMUEGPA01-T03 LEVEL 1
UFD0100
The basic GPWS modes generate visual warnings through associated lights and synthetic warnings through the loudspeakers. The "ENHANCED" GPWS functions allow the terrain hazards to be displayed on the Navigation Display (ND).
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34 - NAVIGATION 34-51-00 DME SYSTEM PRESENTATION
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CONTENTS: Principle Components Indicating
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DME SYSTEM PRESENTATION PRINCIPLE
TMUDMEH01-T01 LEVEL 1
UFD0100
The Distance Measuring Equipment (DME) provides digital readout of the aircraft slant range distance from a selected ground station. The system generates interrogation pulses from an onboard interrogator and sends them to a selected ground station. After a 50 micro seconds delay, the ground station replies. The interrogator determines the distance in Nautical Miles (NM) between the station and the aircraft. The interrogator detects the Morse audio signal which identifies the ground station.
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DME SYSTEM PRESENTATION COMPONENTS
TMUDMEH01-T02 LEVEL 1
UFD0100
The components are two antennae and two interrogators. The DME system is also connected to: - Primary Flight Displays (PFD), Navigation Displays (ND) and Digital Distance Radio Magnetic 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 audio signal.
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DME SYSTEM PRESENTATION INDICATING
TMUDMEH01-T03 LEVEL 1
UFD0100
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 of the Digital Distance Radio Magnetic Indicator (DDRMI).
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34 - NAVIGATION 34-52-00 ATC SYSTEM PRESENTATION
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ATC SYSTEM PRESENTATION PRINCIPLE
TMUATCF08-T01 LEVEL 1
UFD0100
The Air Traffic Control (ATC) transponder is an integral part of the Air Traffic Control Radar Beacon System (ATCRBS). The transponder is interrogated by radar pulses received 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) of the aircraft on the ground controller’s radar scope. These replies enable the controller to distinguish the aircraft and to maintain effective ground surveillance of the air traffic. The ATC transponder also responds to interrogations from aircraft equipped with a Traffic Collision Avoidance System (TCAS) (Mode S).
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ATC SYSTEM PRESENTATION COMPONENTS
TMUATCF08-T02 LEVEL 1
UFD0100
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 to the aircraft configuration.
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34 - NAVIGATION 34-53-00 ADF SYSTEM PRESENTATION
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ADF SYSTEM PRESENTATION PRINCIPLE
TMUADFF01-T01 LEVEL 1
UFD0100
The Automatic Direction Finder (ADF) is a radio navigation aid. The ADF system provides: - An identification of the relative bearing to a selected ground station called Non Directional Beacon (NDB). - Aural identification of the ground station. The relative bearing is the angle between the aircraft heading and the aircraft/ground station axis.
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ADF SYSTEM PRESENTATION
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The combination of signals, received from two loop antennae and from one omni-directional sense antenna, provides bearing information. The ground stations operate in a frequency range of 190 to 1750 Khz. An additional Morse signal is provided to identify the selected ground station.
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ADF SYSTEM PRESENTATION COMPONENTS The Automatic Direction Finder system is composed of two receivers and two antennae. The ADF system is also connected to: - Navigation Displays (ND) and Digital Distance Radio 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) for manual tuning. - Captain and First Officer Radio Management Panels (RMP) for back-up tuning. - Audio Control Panels (ACP) for ADF audio signal.
TMUADFF01-T03 LEVEL 1
UFD0100
Note: ADF 2 system is optional.
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ADF SYSTEM PRESENTATION INDICATING The Automatic Direction Finder (ADF) system information can be displayed on the Navigation Displays (ND) system and on the Digital Distance Radio Magnetic Indicator (DDRMI). On the NDs, depending on the position of the ADF selector 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 ADF selector switch: - ADF 1 is represented by a single pointer - ADF 2 is represented by a double pointer. are
not
equipped
with
the
ADF
TMUADFF01-T04 LEVEL 1
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Note: Some DDRMIs capability.
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34 - NAVIGATION 34-55-00 VOR/MARKER SYSTEMS PRESENTATION
TMUVORG02 LEVEL 1
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CONTENTS: VOR Principle MKR Principle Components VOR Indicating MKR Indicating
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VOR/MARKER SYSTEMS PRESENTATION VOR PRINCIPLE
TMUVORG02-T01 LEVEL 1
UFD0100
The Very high frequency Omni-directional Range(VOR) system is a medium-range radio navigation aid. The VOR system receives, decodes and processes bearing information from the omni-directional ground station (working frequency range: 108 to 117.95 Mhz) The ground VOR station generates a reference phase signal and a variable phase signal. The phase difference between these signals, called bearing, is function of the aircraft position with respect to the ground station. The bearing is the angle between the Magnetic North and the ground station/aircraft axis. Furthermore, the VOR station provides a Morse identification which identifies the station.
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VOR/MARKER SYSTEMS PRESENTATION MRK PRINCIPLE
TMUVORG02-T02 LEVEL 1
UFD0100
The MARKER (MKR) system is a radio navigation aid which indicates the distance between the aircraft and the runway threshold. The MARKER (MKR) system is normally used together with the ILS system during an ILS approach.
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VOR/MARKER SYSTEMS PRESENTATION COMPONENTS
TMUVORG02-T03 LEVEL 1
UFD0100
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 Flight Displays (PFD) and VOR/ADF/DME Radio Magnetic Indicator (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) for manual tuning. - Captain and First Officer Radio Management Panels (RMP) for back-up tuning. - Audio Control Panels (ACP) for VOR/MKR audio signal.
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VOR/MARKER SYSTEMS PRESENTATION VOR INDICATING
TMUVORG02-T04 LEVEL 1
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The indicators show that the aircraft is flying from the ground station and is on the right,crossing and then on the left hand side of the course selected by the pilot.
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TMUVORG02-T05 LEVEL 1
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When the aircraft overflies the Marker, the type of Marker is display on the PFDs in different colors, and is indicated by an aural identification.
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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.
SYSTEMS
ON BOARD MAINTENANCEPAGE 1 of 26
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TECHNICAL TRAINING DEPARTMENT
Figure 1 shows a simplified block diagram of the MFC system.
FAULT ACTIVATE
FAULT ACTIVATE
MFC 1
MFC 1A STATUS
MFC 1B STATUS
INPUTS
INPUTS
MFC 1A
MFC 1B
OUTPUTS
PRIMARY SECONDARY
OUTPUTS
ELECTICAL POWER
ELECTICAL POWER
PRIMARY SECONDARY
MFC Block Schematic Diagram Figure 1
SYSTEMS
ON BOARD MAINTENANCEPAGE 2 of 26
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TECHNICAL TRAINING DEPARTMENT
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 1A
1B
2A
2B
FAULT
FAULT
FAULT
FAULT
OFF
OFF
OFF
OFF
MFC CONTROL PANEL (OVERHEAD) BITE ADV DISPLAY
8
4
2
F F
MFC
1
DATA BUS
F F
BITE LOADED
NORM FLT WOW & L/G
ERS MFC
DOORS
3
BOOTS
PTA/ERS MISC
2
MAG IND TEST
NAV 1
BRK FLT CTL
MFC MAINTENANCE PANEL (OVERHEAD)
MFC Maintenance & Control Panels Figure 2
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ON BOARD MAINTENANCEPAGE 3 of 26
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TECHNICAL TRAINING DEPARTMENT
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.
SYSTEMS
ON BOARD MAINTENANCEPAGE 4 of 26
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TECHNICAL TRAINING DEPARTMENT
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 1 2 3 4 5 6 7 8 9 A B C D E
8
4
2 F F
F F F F F F F F F F F F
1 F F F
F F
F F
F F F F F F
F
DEFINITION Right Main Gear Prime DnLk Prox Switch Fail Nose Gear Prime DnLk Prox Switch Fail Left Main Gear Prime DnLk Prox Switch Fail Right Main Gear Sec DnLk Prox Switch Fail Nose Gear Sec DnLk Prox Switch Fail Left Main Gear Sec DnLk Prox Switch Fail Left Main Gear WOW 1 Prox Switch Fail Nose Gear WOW 1 Prox Switch Fail Right Main Gear WOW 1 Prox Switch Fail Left Main Gear WOW 2 Prox Switch Fail Nose Gear WOW 2 Prox Switch Fail Right Main Gear WOW 2 Prox Switch Fail
F F F
F
End of list for selected system
Failure Codes - De-icing Boots System Table 1
SYSTEMS
ON BOARD MAINTENANCEPAGE 5 of 26
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TECHNICAL TRAINING DEPARTMENT
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 timespent 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.
SYSTEMS
ON BOARD MAINTENANCEPAGE 6 of 26
BASIC COMPLEMENTARY COURSE FOR AF & PP ENGINEERS
TECHNICAL TRAINING DEPARTMENT
Figure 3 shows the BITE flow sequence.
POWER POWER UP UP RESET RESET
PROTECTION PROTECTION
INITIALIZE INITIALIZE
CONTROL CONTROL
INPUT INPUT
OUTPUT OUTPUT
OPERATIONAL OPERATIONAL BITE BITE
BITE Flow Diagram Figure 3
SYSTEMS
ON BOARD MAINTENANCEPAGE 7 of 26
BASIC COMPLEMENTARY COURSE FOR AF & PP ENGINEERS
TECHNICAL TRAINING DEPARTMENT
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.
SYSTEMS
ON BOARD MAINTENANCEPAGE 8 of 26
BASIC COMPLEMENTARY COURSE FOR AF & PP ENGINEERS
TECHNICAL TRAINING DEPARTMENT
Figure 4 shows the Boeing 737 FMS Bite System.
Boeing FMS BITE System Figure 4
SYSTEMS
ON BOARD MAINTENANCEPAGE 9 of 26
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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.
SYSTEMS
ON BOARD MAINTENANCEPAGE 10 of 26
BASIC COMPLEMENTARY COURSE FOR AF & PP ENGINEERS
TECHNICAL TRAINING DEPARTMENT
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
DISK STORAGE
429 BUS INTERFACE
DISK STORAGE
POWER PROG
CHNG
COMP
RDY
XFER
R/W
FAIL
SPARE FUSE
PROG CHNG COMP RDY XFER R/W FAIL
DATA TRANSFER IN PROGRESS DATA CHANGE IS REQUIRED DATA TRANSFER IS COMPLETE UNIT READY FOR OPERATION DATA TRANSFER FAILURE UNABLE TO ACCESS DISK DATA SYSTEM TEST FAILURE
LINE FUSE ON/OFF
Boeing 737 Data Loader Figure 5
SYSTEMS
ON BOARD MAINTENANCEPAGE 11 of 26
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TECHNICAL TRAINING DEPARTMENT
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.
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ON BOARD MAINTENANCEPAGE 12 of 26
BASIC COMPLEMENTARY COURSE FOR AF & PP ENGINEERS
TECHNICAL TRAINING DEPARTMENT
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.
MAT KEYBOARD
MAT KEYBOARD SLOT
MAINTENANCE ACCESS TERMINAL (MAT) FLIGHT COMPARTMENT REAR RIGHT SIDEWALL
Maintenance Access Terminal (MAT) Figure 9
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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. SYSTEMS
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Figure 10 shows the MAT and cursor control device.
MAT DUAL DISK DRIVE
MAT DISPLAY
MAT CURSOR CONTROL DEVICE
SELECTION KEYS (3) TRACK BALL
POWER SUPPLY MODULE
BRIGHTNESS CONTROL
CURSOR CONTROL DEVICE
MAT & Cursor Control Device Figure 10
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ON BOARD MAINTENANCEPAGE 15 of 26
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Figure 11 shows the MAT display showing FDE data.
LINE MAINTENACE
EXTENDED MAINTENANCE
OTHER FUNCTIONS
HELP
N77701 TBC1234 KBFI/KMWH LEG STATRT WAS 1753Z 07 JUL 00 THIS DATA IS FROM LEFT CMCF
INBOUND FLIGHT DECK EFFECTS Select text of Maintenance Message, then select the MAINTENANCE MESSAGE DATA button to get more data.
MAINTENANCE MESSAGE DATA
Flight Deck Effects recorded during the present leg
FDE: F/D ZONE TEMP CTRL
STATUS
Fault Code : 216 011 00
FDE: CAPT RA FLAG
Maintenance Message: 34-42011 Approach
NOT ACTIVE 1948z 07JUL00
PFD FLAG
Fault Code : 343 311 31
REPORT
ACTIVE 1948z 07JUL00
ACTIVE 1941z 07JUL00
Radio Altimeter Transceiver (left) has an internal fault.
GO BACK
ERASE FAULT
MAT Displayed Data Figure 11
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ON BOARD MAINTENANCEPAGE 16 of 26
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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.
PMAT
SELECTION SWITCHES
POWER SWITCH
CURSOR CONTROL
PMAT RECEPTACLE
LCD DISPLAY
KEYBOARD DISK DRIVE
Portable Maintenance Access terminal (PMAT) Figure 12 SYSTEMS
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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.
AIRPLANE CONDITION MONITORING SYSTEM (ACMS)
ACMS REPORTS ACMS REPORTS ACMS XXXX REPORTS XX X XX XXXXXXX XXXXX XXXX XX X XX XXXXXXX XXXXX XXXX XX XXXXXXX XXXXX XXXX XX XX X XX XXXXXXX XXXXX XXXX XX XXXXXXX XXXXX X X XXXXXXXXXXXXX XX XXXX XX XXXXXXX XXXXX XXXXXXX XXXXXXXXXX XXXXXXXXXXXXX XX X XXXXX XXXXXXXXXX X XXXXXXXX XXXXXXXXXXXXX XXXXXXXX XXXX XXXX XXXXXXXXXX XXXXXXX XXXXXXX XXXXXXXX XXXX XXXX XXX XX XXXXXXXXX XXXXXXX XXXXXXXX XXXX XXXXXXX XXX XX XXXXXXXXX XXXXXXX XXXXXXXXXX XX XXXXXXXXX XXXXXXXXXXXXXX X X X X XXXXXXX XXXXXXX X X X X XXXXXXXXXXXXXX XXXXXXXX X X X XXXXXXX XXXXXX XXXXXXXXXXXXXX XXXXXXX XXXXXX XXXXXXXXXXXXXX XXXXXXXXXXXX XXXXXXX XXXXXX XXXXXXXXXXXX XXXXXXX
ACMF
XXXXXXXXXXXX XXXXXXX
PDF CMCF
QAR
AIMS FMCF
DCMF
TMCF
FDCF
TA DA
DFDAF
Boeing 777 ACMS Figure 13 SYSTEMS
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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.
RH DISPLAY LH DISPLAY
QAR FLIGHT COMPARTMENT PRINTER
MAT
PMAT
A I R C R A F T
FLIGHT CONTROL ARINC 629 BUS (3)
SDU
VHF TX/RX
SYSTEMS ARINC 629 BUS (4) ARINC 429 ANALOG DISCRETES
LEFT HAND AIMS CABINET
ACMS (Boeing 777) Figure 14
SYSTEMS
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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. SYSTEMS
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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.
PRESS SPARE DISK OPTICAL QAR
POWER ON
DISPLAY DISPLAY
PENNY & GILES
FAIL
LOW CAPACITY
MAINTENANCE
EJECT
OPTICAL DISK CARTRIDGE
MADE IN U.K.
QUICK ACCESS RECORDER
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.
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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.
SYSTEMS
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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).
AIRCRAFT CONDITION MONITORING SYSTEM (ACMS)
FLIGHT DATA RECORDER SYSTEM (FDRS)
FLIGHT MANAGEMENT COMPUTING SYSTEM (FMCS)
PRIMARY DISPLAY SYSTEM (PDS)
CENTRAL MAINTENCE COMPUTING SYSTEM (CMCS)
AIMS LEFT-HAND CABINET AIMS RIGHT-HAND CABINET
THRUST MANAGEMENT COMPUTING SYSTEM (TMCS)
DATA COMMUNICATION MANAGEMENT SYSTEM (DCMS)
AIMS System Figure 16
SYSTEMS
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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. SYSTEMS
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Interconnection board
SYSTEMS
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Figure 17 shows the flight compartment printer.
FAIL
PAPER
CUT
SLEW
RESET
TEST
TOP VIEW
SIDE VIEW
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.
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23 COMMUNICATIONS
23 - COMMUNICATIONS 23-73-00 CABIN INTERCOMMUNICATION DATA SYSTEM DESIGN PHILOSOPHY
TMUCI2A02 LEVEL 1
UFD4200
CONTENTS: General Principle Passengers Functions Crew Functions Cabin Systems Functions Monitoring And Test Functions Aircraft Systems Functions Cockpit Controls And Indicating
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CABIN INTERCOMMUNICATION DATA SYSTEM DESIGN PHILOSOPHY GENERAL
PASSENGER FUNCTIONS
Changing market demands require flexibility in customized cabin layouts and optional cabin systems. With the Cabin Intercommunication Data System (CIDS), the operator is able to change the cabin layout without hardware changes (e.g. cabin loudspeakers, PAX-equipment ...). This can be simply done by entering, on board, new cabin parameters in the software. The CIDS is a microprocessor based system. It monitors, tests, operates and provides control and monitoring of the cabin functions.
-
PRINCIPLE
CABIN SYSTEMS FUNCTIONS
To manage various functions, the CIDS has a central unit, the CIDS DIRECTOR. It is linked to the Forward Attendant Panel (FAP) for control and monitoring of the cabin functions. The Director then communicates, through a bus system, with Decoder Encoder Units (DEUs). The DEUs send (and receive) information to (and from) the cabin, passenger and crew systems. The Director has interfaces to other aircraft systems. Through a Programming and Test Panel (PTP) the CIDS can be programmed to customer demand. The PTP is also used to test the entire CIDS.
-
EFFECTIVITY EFFECTIVITY ALL
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.
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.
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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
-
EFFECTIVITY EFFECTIVITY ALL
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23 - COMMUNICATION 23-73-00 CABIN INTERCOMMUNICATION DATA SYSTEM PRESENTATION
TMUCI2B02 LEVEL 1
UFD4200
CONTENTS: General Directors Type A Decoder Encoder Units Type B Decoder Encoder Units Forward Attendant Panel Programming and Test Panel
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CABIN INTERCOMMUNICATION DATA SYSTEM PRESENTATION GENERAL
TYPE A DECODER ENCODER UNITS
The CIDS consists of the following components: - the Directors including ON Board Replaceable Memories (OBRM), - the Programming and Test Panel (PTP) including Cabin 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.
The type A DEUs provide the interface between the directors and the passenger related systems.
TMUCI2B02-T01 LEVEL 1
UFD4200
DIRECTORS For redundancy, two directors are installed. In normal operation of the CIDS, director 2 is in hot stand-by. Both directors receive the same inputs and perform the same computations. The outputs of the director in hot stand-by are disabled. The directors are connected through two CIDS busses to the type A and type B DEUs to carry the various data to the cabin equipment. The FAP, PTP and other systems are connected directly to the directors basically for control, indication and test of the CIDS functions.
EFFECTIVITY EFFECTIVITY ALL
TYPE B DECODER ENCODER UNITS The type B DEUs provide the interface between the directors and the attendant and cabin related systems. FORWARD ATTENDANT PANEL The Forward Attendant Panel (FAP) is installed at the forward attendant station. From the FAP, the various cabin systems can be controlled and monitored. PROGRAMMING AND TEST PANEL The Programming and Test Panel (PTP) is installed at the forward attendant station next to the FAP. The PTP contains the Cabin Assignment Module (CAM) which is used to store all information for the actual cabin layout.
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THIS PAGE INTENTIONALLY LEFT BLANK
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23 - COMMUNICATIONS 23-73-00 CIDS - DIRECTOR/DEU ARCHITECTURE
TMUCI2C03 LEVEL 3
UFD4200
CONTENTS: DEU A DEU B Self Examination
EFFECTIVITY EFFECTIVITY ALL
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CIDS - DIRECTOR/DEU ARCHITECTURE DEU A Twenty six type A Decoder Encoder Units (DEUs) are installed above the windows in the cabin ceiling and close to the center ceiling for the DEUs in the entrance area. The type A DEUs are connected to the directors via a top-line data bus (i.e. : two wire twisted and shielded cable). A broken wire in one top-line bus will only affect the type A DEUs behind the crack on this bus. The type A DEUs of the other top-line bus will work without disturbance.
TMUCI2C03-T01 LEVEL 3
UFD4200
PASSENGER SIGNS The passenger signs include NO SMOKING or the optional NO ELECTRONIC DEVICE lights, FASTEN SEAT BELT lights, NON SMOKER ZONE lights and RETURN TO SEAT lights in the lavatories. Furthermore, for the PAX call system, the seat row lights are connected to the type A DEUs.
LOUDSPEAKERS The loudspeakers are installed in the Passenger Service Unit (PSU), in each lavatory and close to the attendant station. They are all identical and are used for: - Passenger address announcements, - Call chimes (optional). PASSENGER CALL Pushbuttons are fitted in the PSU above each seat row and in the lavatories. READING/LIGHT POWER UNIT One R/L power unit for three reading lights installed in each Passenger Service Unit (PSU).
is
CABIN LIGHTS The cabin lights include: - Entrance area lights, - Lavatory lights, - Attendant lights, - Reading lights, - Cabin fluorescent strip lights.
EFFECTIVITY EFFECTIVITY ALL
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CIDS - DIRECTOR/DEU ARCHITECTURE EFFECTIVITY EFFECTIVITY ALL
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CIDS - DIRECTOR/DEU ARCHITECTURE DEU B Basically 4 type B DEUs (max. 6 - Optional) are installed near the exit doors in the center ceiling. They are connected to the directors via a middle line data bus. There are two supplementary DEU B mounts installed as a provision. The fig. on the next page shows a typical Type B DEU interface. It may vary with different locations and with specific airline requirements. SLIDE PRESSURE SYSTEM (Optional) The directors receive signals from the bottle pressure sensors via type B DEUs. If the pressure is low, the CIDS CAUTION light on the FAP comes on.
EPSUs The Emergency Power Supply Units (EPSUs) are connected to type B DEUs for the emergency lighting system test. DRAIN MAST The directors receive signals from the drain mast control unit via type B DEUs. If the drain mast heater or the control unit fails the CIDS CAUTION light on the FAP comes on. ATTND AND PANEL One Attendant Indication Panel is installed near each attendant seat for message purposes. AREA CALL PANEL One basic and one optional ACP can be connected to each DEU B.
TMUCI2C03-T02 LEVEL 3
UFD4200
DOOR PRESSURE SYSTEM (Optional) The directors receive signals from the bottle pressure sensors via type B DEUs. If the pressure is low, the CIDS CAUTION light on the FAP comes ON. CREW INTERPHONE SYSTEM The crew interphone system enables communication between cockpit crew and cabin attendants and between each attendant station. NOTE: From each attendant station it is possible to communicate with personnel at the service interphone connections.
EFFECTIVITY EFFECTIVITY ALL
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CIDS - DIRECTOR/SYSTEM ARCHITECTURE EFFECTIVITY EFFECTIVITY ALL
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23 COMMUNICATIONS
SELF EXAMINATION A break A B C
in one top line data bus: - Disables all DEUs. - Affects only type B DEUs. - Only affects the type A DEUs behind the crack on this bus. Panel (ACP) is connected to: directors directly. type A DEUs. type B DEUs.
TMUCI2C03 LEVEL 3
UFD4200
The Area Call A - The B - The C - The
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23 - COMMUNICATIONS 23-30-00 PASSENGER ENTERTAINMENT SYSTEM PRESENTATION
TMU23EA01 LEVEL 3
UFD4200
CONTENTS: General PES PES Video PRAM
EFFECTIVITY EFFECTIVITY ALL
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PASSENGER ENTERTAINMENT SYSTEM PRESENTATION GENERAL The passenger address and entertainment comprises the following basic functions:
system
- Passenger Entertainment System (PES), - Passenger Entertainment System Video
(PES video), - Pre-Recorded Announcements and boarding Music system (PRAM). The PES comprises the PES music, the passenger address and the passenger service. PES The PES transmits pre-recorded music programs, passenger address information, video and video sounds to the passengers. The audio signals can be heard through headphones connected to the Passenger Control Units (PCU).
TMU23EA01-T01 LEVEL 3
UFD4200
The PCU allows several music channels and video audio channels to be selected and the volume to be adjusted. The PCU also allows the reading lights and passenger calls to be remotely controlled through the Passenger Service System (PSS). All pre-recorded announcements (video and sound) and the passenger address messages, heard in the headphones through the PCU, have priority over the music and video sound entertainment channels.
EFFECTIVITY EFFECTIVITY ALL
The anouncements and passenger address messages are also broadcast through the passenger address loudspeakers, via the CIDS. The PES audio reproducers supply music channels to the Main Multiplexer and boarding music channels to the CIDS director. The CIDS broadcasts the boarding music through the passenger address loudspeakers. Boarding Music (BGM) channel and volume control is performed on the Forward Attendant Panel (FAP). The Main Multiplexer is connected to the CFDIU to ensure the passenger entertainment BITE function. PES VIDEO The PES video shows pre-recorded video movies and video announcements through different display units in the passenger compartment. The video sound is transmitted to the Main Multiplexer and to the CIDS. Therefore video sounds can be heard from the headset through the PCU or from the cabin passenger address loudspeakers. The in-seat video display units are supplied through the Main Multiplexer.
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PASSENGER ENTERTAINMENT SYSTEM PRESENTATION - SCHEMATIC EFFECTIVITY EFFECTIVITY ALL
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PASSENGER ENTERTAINMENT SYSTEM PRESENTATION PRAM The PRAM is an audio tape reproducer which contains pre-recorded announcements and boarding music supplied to the CIDS director. The announcements are also sent to the Main Multiplexer.
TMU23EA01-T02 LEVEL 3
UFD4200
The PRAM is controlled from the FAP.
EFFECTIVITY EFFECTIVITY ALL
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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.
SYSTEMS
INFORMATION SYSTEM PAGE 1 of 9
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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.
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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.
A/C SYSTEMS
AIRLINE COMPUTER SYSTEM
MAINTENANCE OPERATIONS
ACARS
VHF 3
TRANSMISSION NETWORK
FLIGHT OPERATIONS
PASSENGER SERVICES VHF TRANSMITTER/RECEIVER
ACARS Network Figure 1
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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.
AC ARS Management Unit (MU).
2.
Mu lti-Purpose Interactive Display Unit (MPIDU).
3.
Ide nt plug.
4.
Pr ogram 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)
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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.
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Figure 2 show the display layout of the MIDU.
IN Collins D A T A
DFDAU FAIL
SEND
NUMERIC ENTRY 13 : 02 : 58 FLT : 0123 0008
L I N K
1
2
3
4
5
6
7
8
9
0 CLR
RET
DEL
Multipurpose Interactive Display Unit (MIDU) Figure 2
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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 groundstation. 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.
SELF TEST
PPR ADV
PWR ON
ALERT RESET
PTR BUSY
PUSHBUTTON CONTROLS
DOOR LOCKING SCREW PAPER LOADING DOOR
PAPER CUTTING EDGE
ACARS Thermal Printer Figure 3
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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.
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Figure 4 shows a block schematic of the ACARS.
VHF 3 ANTENNA
Collins IN
DFDAU FAIL
SEND
NUMERIC ENTRY 13 : 02 : 58 D A T A
FLT : 0123 0008
L I N K
1
2
3
4
5
6
7
8
9
0 CLR
RET
DEL
MULTIPURPOSE INTERACTIVE DISPLAY UNIT
MANAGEMENT UNIT
VHF 3 TX/RX
FLIGHT DATA ACQUISTION UNIT THERMAL PRINTER
AIRCRAFT SYSTEMS
ACARS Schematic Diagram Figure 4
SYSTEMS
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777 AIRCRAFT MAINTENANCE MANUAL ELECTRONIC FLIGHT BAG - INTRODUCTION General
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 key EGP 101-999 * MAU - microwave antenna unit EGP 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 * *
The Electronic Flight Bag (EFB) lets the flight crew access to the electronic flight operation data, general purpose computing and communications. Abbreviations and Acronyms AC - advisory circular (FAA) ACARS - aircraft communication addressing and reporting system * 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) *
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777 AIRCRAFT MAINTENANCE MANUAL ELECTRONIC FLIGHT BAG - INTRODUCTION 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 - power EGP 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 order EGP 101-999 * TWLU - terminal wireless LAN unit EGP 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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777 AIRCRAFT MAINTENANCE MANUAL ELECTRONIC FLIGHT BAG - INTRODUCTION
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777 AIRCRAFT MAINTENANCE MANUAL EFB - GENERAL DESCRIPTION General The electronic flight bag (EFB) has two display units (DU) and two supporting electronics units (EU). The captain’s EFB system is independent from the first officer’s EFB system. Each EFB system consists of a DU and an EU. Description The EFB provides the flight crew with a paperless flight deck environment and enhance the quality of information available to the 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 a touch-screen that is a feature of certain applications (example: electronic logbook). In addition, the flight crew can also make use of the cursor control 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 * *
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777 AIRCRAFT MAINTENANCE MANUAL EFB - GENERAL DESCRIPTION
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