Aircraft Electrical and Electronic Systems: Principles, operation and maintenance

nureddin isgenderov

Aircraft Electrical and Electronic Systems: Principles, operation and maintenance

Abstract
AI

Aircraft Electrical and Electronic Systems is designed for aircraft engineering students and professionals, providing essential knowledge on electrical, electronic, and digital theory. Covering EASA Part 66 Modules 11 and 13, this text serves as a reference for certifying mechanics, technicians, and engineers in maintenance activities for commercial and general aviation. Emphasizing foundational and advanced topics in aircraft electrical systems, it addresses the evolution of avionics and the transition towards all-electric aircraft, while maintaining a balance between technicality and accessibility for readers with varying levels of prior knowledge.

Figures (494)

Figure 1.2 A material with a loosely bound electron in its outer shell Figure 1.1 The Bohr model of the atom comb may result in a difference in charge between your hair and the rest of your body, resulting in your hair standing on end when your hand or some other differently charged body is brought close to it.

Figure 1.3. Free electrons and the application of an external force electric current is passed through a resistive heating element. Light is produced when an electric current flows through the thin filament wire in the evacuated bulb of an electric lamp. All electrons and protons carry an electrostatic

Figure 1.5 Electric field between isolated unlike charges

Figure 1.8 Electric field strength between two charged conducting surfaces Figure 1.7 Electric field between the two charged parallel metal plates of a capacitor surfaces (see Fig. 1.8). The electric field strength is given by:

Figure 1.6 Electric field between isolated like charges field defines the direction and magnitude of a force on a charged object. The field itself is invisible to the human eye but can be drawn by constructing lines which indicate the motion of a free positive charge within the field; the number of field lines in a particu- lar region being used to indicate the relative strength of the field at the point in question. human eye but can be drawn by constructing lines

Figure 1.9 A selection of capacitors with values ranging from 12pF to 1000,.F and working voltages ranging from 25V to 450V

reaction takes place. When several cells are connected together they form a battery. There are two types of cell: primary and second- ary. Primary cells produce electrical energy at the expense of the chemicals from which they are made and once these chemicals are used up, no more elec- tricity can be obtained from the cell. In secondary cells, the chemical action is reversible. This means that the chemical energy is converted into electrical energy when the cell is discharged whereas electrical energy is converted into chemical energy when the cell is being charged. You will find more information on aircraft batteries in Chapter 5.

Figure 1.11 A simple DC circuit consisting of a battery (Source) and resistor (load)

Figure 1.13 A selection of resistors with values ranging from 0.12 to 10MQ and power ratings from 0.1W to 15W

Figure 1.12 Relationship between V, /and R

Figure 1.19 Using Kirchhoff’s law to find the voltage that appears between terminals A and B— see Example 1.11 Figure 1.18 Using Ohm’s law to find the voltage dropped across the 9? resistor — see Example 1.11 across R (i.e. the potential drop between terminals A and B):

Figure 1.21 See Test your understanding 1.6 Figure 1.20 Using Ohm's law to find the value of R — see Example 1.11 Finally, since we now know the voltage, V, and cur- rent, J, that flows in R, we can apply Ohm’s law again in order to determine the value of R:

Figure 1.22 Relationship between P, / and V The formula may be arranged to make P, J or V the subject, as follows:

Figure 1.24 Field around a current-carrying conductor Figure 1.23 Field and flux directions for a bar magnet

Figure 1.26 Magnetic field around a single turn loop length of the conductor in the field, and the strength of the magnetic flux (expressed in terms of its flux density). The size of the force will be given by the expression:

Figure 1.25 A current-carrying conductor in a magnetic field pole and enter at a south pole; in other words, they travel from north to south, as indicated by the direc- tion arrows. The net effect of the coming together of these two magnetic force fields is that at position A, they both travel in the same direction and reinforce one another. While at position B, they travel in the opposite direction and tend to cancel one another. So with a stronger force field at position A and a weaker force at position B the conductor is forced upwards out of the magnetic field. If the direction of the cur- rent was reversed, i.e. if it was to travel towards us out of the page, then the direction of the magnetic field in the current-carrying conductor would be reversed and therefore so would the direction of motion of the conductor.

Figure 1.28 Demonstration of electromag- netic induction

Figure 1.29 Cutting lines of flux and the e.m.f. generated: (a) cutting lines of flux at 90°, e = Bly, (b) cutting lines of flux at 0, e = Bivsin 6

This effect is called self-inductance (or just induct- ance) which has the symbol L. Self-inductance is measured in henries (H) and is calculated from:

Figure 1.33 A selection of inductors with values ranging from 100nH to 4H and current ratings ranging from 0.1 Ato 10A

Fig. 1.37). The relationship between periodic time and frequency is thus:

Figure 1.34 Comparison of direct and alternating current

Figure 1.38 Average, r.m.s., peak and peak— peak values of a sine wave Figure 1.38 Average, r.m.s., peak and peak— peak values of a sine wave

Figure 1.39 Waveforms for a three-phase AC supply

Figure 1.40 Voltage and current waveforms for a pure capacitor (the current leads the voltage by 90°)

Figure 1.41 Voltage and current waveforms for a pure capacitor (the voltage leads the current by 90°)

Figure 1.42 Using CIVIL to determine phase relationships in circuits containing capacitance and inductance

Figure 1.43 Variation of inductive reactance, X_, with frequency, f

When alternating voltages are applied to capaci- tors or inductors the magnitude of the current flow- ing will depend upon the value of capacitance or inductance and on the frequency of the voltage. In effect, capacitors and inductors oppose the flow of current in much the same way as a resistor. The important difference being that the effective resist- ance (or reactance) of the component varies with frequency (unlike the case of a conventional resis- tor where the magnitude of the current does not change with frequency).

Figure 1.45 The impedance triangle where Z is the impedance in ohms (2), V is the alter- nating potential difference in volts (V) and J is the alternating current in amps (A).

Figure 1.46 The principle of the transformer

Here we will use Ns = turns-per-volt X Vg.

Figure 1.49 A selection of transformers for operation at frequencies ranging from 50Hz to 50 kHz

Figure 1.48 Hysteresis curves and energy loss from which

Figure 2.1 Regular lattice structure of a pure semiconductor material

Figure 2.2 Effect of introducing a pentavalent impurity

Figure 2.3 Effect of introducing a trivalent impurity

Figure 2.8 A forward-biased P-N junction diode

Figure 2.9 A reverse-biased P-N junction diode

Figure 2.10 Typical //V characteristics for germanium and silicon diodes

Figure 2.11 Various diodes (including rectifiers, signal diodes, Zener diodes, light- emitting diodes and silicon-controlled rectifiers)

ure 2.12 Symbols used for various common types of diode 6V). A similar effect (avalanche) occurs in less heav- ily doped diodes. These avalanche diodes also exhibit a rapid breakdown with negligible current flowing below the avalanche voltage and a relatively large current flowing once the avalanche voltage has been reached. For avalanche diodes, this breakdown volt- age usually occurs at voltages above 6V. In practice, however, both types of diode are commonly referred to as Zener diodes. The symbol for a Zener diode was shown earlier in Fig. 2.12, whilst typical Zener diode characteristics are shown in Fig. 2.13.

Figure 2.13 Typical Zener diode characteristic

Figure 2.14 A simple Zener diode voltage regulator

Figure 2.15 Triggering a_ silicon-controlled rectifier 2.2.3 Silicon-controlled rectifiers Silicon-controlled rectifiers (or thyristors) are three- terminal devices which can be used for switching and AC power control. Silicon-controlled rectifiers can switch very rapidly from a conducting to a non-con- ducting state. In the off state, the silicon-controlled rectifier exhibits negligible leakage current, while in the on state the device exhibits very low resistance. This results in very little power loss within the silicon- controlled rectifier even when appreciable power levels are being controlled.

Figure 2.17 Switching action of the diode in the half-wave rectifier: (a) D1 forward biased, (b) D1 reverse biased

Figure 2.18 Waveforms of voltages in the simple half-wave power supply The switching action of D1 results in a pulsat- ing output voltage, which is developed across the load resistor (R,). Since the supply is at 400 Hz, the pulses of voltage developed across R, will also be at 400 Hz even if only half the AC cycle is present. During the positive half-cycle, the diode will drop the 0.6V forward threshold voltage normally asso- ciated with silicon diodes. However, during the negative half-cycle the peak AC voltage will be dropped across D1 when it is reverse biased. This is an important consideration when selecting a diode for a particular application. Assuming that the secondary of Tl provides 28.75V RMS, the peak voltage out- put from the transformer’s secondary winding will be given by:

Figure 2.19 Adding a reservoir capacitor to a half-wave rectifier to make a simple DC power supply

Figure 2.20 Effect of adding a reservoir capacitor on the output of the half-wave rectifier

Figure 2.22 Switching action of the diodes in the bi-phase rectifier: (a) D1 forward biased and D2 reverse biased, (b) D1 reverse biased and D2 forward biased

two basic forms of full-wave rectifier: the bi-phase type and the bridge rectifier type.

Figure 2.23 Bi-phase power supply with reservoir capacitor

Figure 2.24 Waveforms in the bi-phase power supply with and without a reservoir capacitor

Figure 2.25 Full-wave power supply using a bridge rectifier

Figure 2.26 Switching action of the diodes in the full-wave bridge: (a) D1 and D2 forward biased whilst D3 and D4 are reverse biased, (b) D1 and D2 reverse biased whilst D3 and D4 are reverse biased

Figure 2.27 Full-wave bridge power supply with reservoir capacitor

Figure 2.29 Construction of a typical NPN transistor

Figure 2.30 Construction of a typical PNP transistor

Figure 2.32 Bias voltages and current flow in NPN and PNP bipolar junction transistors Figure 2.31 NPN and PNP BJT symbols and simplified junction models

Figure 2.33 Input characteristic (/3/Vge) for an NPN bipolar junction transistor

Figure 2.34 Output characteristic (/¢/Vce) for an NPN bipolar junction transistor

Figure 2.35 Transfer characteristic (/¢/lg) for an NPN bipolar junction transistor

Figure 2.36 Various transistors (including low-frequency, high-frequency, high-voltage, small-signal and power types)

A requirement of most amplifiers is that the out- put signal should be a faithful copy of the input sig- nal or be somewhat larger in amplitude. Other types of amplifier are ‘non-linear’, in which case their input and output waveforms will not necessarily be similar. In practice, the degree of linearity provided by an amplifier can be affected by a number of fac- tors including the amount of bias applied and the amplitude of the input signal. It is also worth noting that a linear amplifier will become non-linear when the applied input signal exceeds a threshold value. Beyond this value the amplifier is said to be over- driven and the output will become increasingly dis- torted if the input signal is further increased.

Figure 2.40 Symbol for an _ operational amplifier

Figure 2.41 Various integrated circuits (including logic gates, operational amplifiers and memories)

Figure 3.1 MIL/ANSI symbols for standard logic gates together with truth tables

Figure 3.5 Landing gear warning logic Figure 3.4 The majority vote logic

system is used to drive six warning indicators. Four of these are located on the overhead display panel and show which door (or doors) are left open whilst an indicator located on the pilot’s instrument panel pro- vides a master landing gear door warning. A switch is also provided in order to enable or disable the five door warning indicators.

Figure 3.8 Logic circuit based on transistor-transistor logic (TTL) integrated circuits

Figure 3.7 See Test your understanding 3.2

Figure 3.6 See Test your understanding 3.2

Figure 3.11 Generator control logic APU START button for a further attempt at starting!

Figure 3.12 Simple R-S bistables based on: (a) NAND gates and (b) NOR gates

Figure 3.14 Timing diagram for the D-type bistable memory) and as binary dividers. The simple circuit arrangement in Fig. 3.13 together with the timing diagram shown in Fig. 3.14 illustrate the operation of D-type bistables.

Figure 3.16 Truth tables for the J-K bistable

Figure 3.21 An eight to three line priority encoder and its corresponding truth table. Note that the enable (EN) input is active-low 3.7 Multiplexers

than one input becomes active, the input with the high- est priority will be encoded and its binary coded value Figure 3.21 shows an eight to three line priority encoder together with its corresponding truth table. Note that this device has active-low inputs as well as active-low outputs.

Figure 3.23 Switch equivalent circuits for some common types of multiplexer or data selector

Figure 3.24 A basic two to one multiplexer arrangement with its corresponding truth table

Figure 3.26 A four to one multiplexer. The logic state of the A and B inputs determines which of the four logic inputs (Xo to X3) appears at the output

Figure 3.25 Logic circuit arrangement for the basic two to one multiplexer

Figure 3.27 Logic gate arrangement for the four to one multiplexer

‘igure 3.29 Bus systems implemented on a modern passenger aircraft

gure 3.28 Unidirectional and bidirectional serial and parallel data

Figure 3.30 A simple system for serial data transfer between two avionic systems

Figure 3.31 A practical aircraft data bus ARINC 429 employs a unidirectional data bus standard known as Mark 33 Digital Information Transfer System (DITS). Messages are transmitted in packets of 32-bits at a bit rate of either 12.5 or 100 kilobits per second (referred to as low and high bit rate respectively). Because the bus is unidirectional, separate ports, couplers and cables will be required when an LRU wishes to be able to both transmit and receive data. Note that a large number of bus connec- tions may be required on an aircraft that uses sophis- ticated avionic systems.

Figure 3.32 A basic bus interface of receivers connected to the bus. With ARINC 429, no more than 20 receivers should be connected to a single bus. Since each bus is unidirectional, a system needs to have its own transmit bus if it is required to respond or to send messages. Hence, to achieve bidi- rectional data transfer it is necessary to have two sep- arate bus connections.

Figure 3.33 Basic components of a computer system 3.9 Computers

Figure 3.34 A typical aircraft computer system 3.9.1 Memories and data storage consists of a high-frequency square wave pulse train derived from an accurate quartz crystal controlled oscillator.

Figure 4.1 A conductor moving inside a magnetic field

Figure 4.2 A loop rotating within a magnetic field

Since E = Bly sin 0, the e.m.f. generated by the arrangement shown in Fig. 4.4 will take a sinusoidal form, as shown in Fig. 4.5. Note that the maximum values of e.m.f. occur at 90° and 270° and that the generated voltage is zero at 0°, 180° and 360°.

Figure 4.5 Sinusoidal voltage produced by the rotating loop e.m.f. generated by the loop is reversed after rotating through 180°. The generated e.m.f. for this arrange- ment is shown in Fig. 4.7. It’s worth comparing this waveform with that shown in Fig. 4.5—you should be able to spot the difference immediately! This can be achieved by modifying the arrangement shown in Fig. 4.3, replacing the brushes and slip- rings with a commutator arrangement, as shown in Fig. 4.6. The commutator arrangement functions as a rotating reversing switch which ensures that the

Figure 4.7 E.m.f. generated (compare with Fig. 4.5)

A simple DC motor consists of a very similar arrange- ment to that of the DC generator that we met earlier.

Figure 4.10 Torque on a current-carrying loop suspended within a permanent magnetic field a shaft in a permanent magnetic field, as shown in Fig. 4.11.

Figure 4.9 E.m.f. generated (compare with Fig. 4.7)

Figure 4.12 Magnetic field produced by a solenoid

Figure 4.18 Typical torque and speed characteristic for a shunt-wound DC motor

Figure 4.20 Typical torque and speed characteristic for a compound-wound DC motor Figure 4.19 Compound-wound DC motor

Figure 4.16 Typical torque and speed characteristic for a series-wound DC motor

Figure 4.22 Simplified construction of a single-phase AC generator By adding more pairs of poles to the arrangement cycles of output voltage for one single revolution of

Figure 4.23 Output voltage produced by the single-phase AC generator shown in Fig. 4.22 Example 4.1 slots in the high-permeability laminated core. These

Figure 4.24 Simplified construction of a two- phase AC generator

Figure 4.25 Output voltage produced by the two-phase AC generator shown in Fig. 4.24

Figure 4.27 Output voltage produced by the three-phase AC generator shown in Fig. 4.26

Figure 4.31 Phasor diagram for the three- phase star-connected system Figure 4.30 A complete star-connected three-phase distribution system

In a delta-connected three-phase system the line cur- rent is 6A. Determine the phase current. Example 4.3

Figure 4.32 A complete delta-connected three-phase distribution system

Figure 4.33 Arrangement of the field wind- ings of a three-phase AC motor

Figure 4.34 AC motor as a delta-connected load

‘igure 4.35 AC waveforms and magnetic field direction

Figure 4.36 Squirrel cage rotor construction

Figure 4.39 Relationship between torque and slip

Figure 4.38 Force on the rotor of an induction motor

Figure 4.40 Using an inverter to produce a variable output speed from an AC induction motor For values of torque within the working range of the motor (i.e. over the linear range of the graph shown in Fig. 4.39), the slip directly proportional to the torque and the per-unit slip is given by:

Figure 4.41 Capacitor starting arrangement Once the rotor is started rotating, however, it will continue to rotate and come up to speed. A field is set up in the rotating rotor that is 90° out of phase with the stator field. These two fields together pro- duce a rotating field that keeps the rotor in motion.

In some types of motor a more complicated arrangement is used with more than one capaci- tor switched into the auxiliary circuit. For example, a large value of capacitor could be used in order to ensure sufficient torque for starting a heavy load and then, once the motor has reached its operating speed, the capacitor value can be reduced in order to reduce the current in the auxiliary winding. A motor that employs such an arrangement, where two dif- ferent capacitors are used (one for starting and one for running) is often referred to as capacitor-start, capacitor-run induction motor. Finally, note that, since phase shift can also be produced by an inductor, it is possible to use an inductor instead of a capaci- tor. Capacitors tend to be less expensive and more compact than comparable inductors and therefore are more frequently used.

Figure 4.43 Main three-phase AC generator mounted in the engine compartment of a large passenger aircraft drive (in smaller aircraft), or engine/APU accessory gearbox in larger aircraft. Generators will have suffi- cient output to supply all specified loads and charge the battery(s). Most avionic equipment requires a regulated and stable power supply depending on its function, e.g. in the case of lighting, it would be incon- venient if the intensity of lighting varied with engine speed. Generator output is affected by internal heat and this has to be dissipated. Cooling methods can include natural radiation from the casing, however this is inadequate for high-output devices where ram-air is directed from a scoop and directed into the generator’s brush-gear and commutator. In some installations, e.g. helicopters, a fan is installed to provide cooling when the aircraft is hovering. We know from basic theory that a generator’s output will vary depending on the input shaft speed. A means of regulating the gen- erator’s output to the bus is required as is a means of overload protection. Test your understanding 4.4

Figure 4.44 Tail-mounted auxiliary power unit (APU) with three-phase AC generator (top of picture) brush erosion is accelerated. Additives can be incor- porated into the brushes that deposit a lubricating film on the commutator; this needs time to build up a sufficient protection; brushes need to be run in for several hours before the protective layer forms (this is often mistaken for contamination). The alternative is an in-built lubrication that is consumed as part of the natural brush wear, i.e. no film is deposited.

Figure 4.45 Overhead electrical panel fitted with main and APU generator indicators and displays. This panel provides the crew with an instant overview of the status of the electrical generating system and the AC and DC distribution buses)

jure 4.46 A practical brushless AC generator arrangement

Figure 4.47 Layout of a typical brushless AC generator

e 4.48 Typical main generator showing three-phase output cabling

Figure 4.49 Typical generator control unit (GCU)

electrodes (the anode and cathode) and electrolyte contained within a casing. Cell materials vary depend- ing on the type of battery performance required for a given cost. The simple primary cell (Fig. 5.1) causes an electron flow from the cathode (negative) through the external load to the anode (positive). The materi- als used refer to the two types of battery cell in wide- spread use on aircraft for the primary source of power: lead-acid or nickel-cadmium. These are maintained on the aircraft and treated as line-replaceable units; a full description of these two battery types is provided in this chapter. Cells used within other aircraft equip- ment or systems are typically made from lithium or nickel-metal hydride materials. These are not main- tained as individual items on the aircraft, they are installed/removed as part of the equipment that they are fitted into; in this case only a brief description is provided.

Figure 5.2 Cells and batteries: (a) connection of cells to form a battery; (b) symbols for cells and a battery; (c) typical battery casing

reactions take place on only one side; for this reason, there are always an even number of positive plates sandwiched between an odd number of negative plates. All positive plates are connected together as are all the negatives. The plates are interlaced and separated by a porous separator that allows free circulation of the electrolyte at the plate surfaces; the plates are all stacked within the cell container. The electrolyte is sul- phuric acid diluted with distilled (pure) water (H,SO,).

Figure 5.4 Lead-acid cell discharge characteristics

Figure 5.5 Nickel-cadmium cell discharge characteristics

Figure 5.8 Metal-hydride cell discharge characteristics

Figure 5.7 Lithium-ion cell discharge characteristics

Figure 5.9 Typical battery locations: (a) battery compartment (GA aircraft); (b) wing leading edge (Beech King Air); (c) nose equipment bay (medium helicopter); (d) externally mounted (small helicopter)

Figure 6.1 Vibrating contact regulator schematic This device comprises voltage and current regulators as shown in Fig. 6.1. They are used on small general

Figure 6.2 Vibrating contact regulator overview

inches in length) with cooling fins. Functions of each component are as follows:

There are many types and configurations of electronic voltage regulators. A representative type is illustrated in Fig. 6.5. The alternator master switch used in AC systems energizes the field relay and applies current to the base of TR, and the resistor network of Ry, Ro, Ry ;. This network, together with the Zener diode (Z) is used to establish the nominal operating voltage. Current flows through the alternator’s field coil via transistors TR, and TR3;, allowing the generator’s out- put to increase. When the output reaches its specified value (14 or 28V DC depending on the installation) Zener diode Z conducts which turns on transistor TR, shorting out transistor TR, and TR3. The genera- tor voltage falls and Zener diode Z stops conducting, thereby turning of transistor TR;. This turns transis- tors TR, and TR; back on, allowing the generator output to increase again. This operation is repeated

Figure 6.8 Static inverter installation Key point

This particular inverter has the following features: 6.5 Transformers

The APU itself is started from the main aircraft bat- tery. In some aircraft, the APU can also provide elec- trical power in the air in the event of main generator failure. The Boeing 787 aircraft has more electrical systems and less pneumatic systems than aircraft it is replacing. In this case the APU delivers only electri- cal power. Some APUs on larger four-engined aircraft are not certified for use while the aircraft is in flight.

Figure 7.1 Typical aircraft wires and cables: (a) single-walled, (b) twin-walled weight, but offer less voltage drop for a given length and lower heating effect due to ’R losses.

Figure 7.7 Fuse clipped directly into a terminal board operated at the minimum fuse rating for prolonged periods of time. Ambient temperature affects the cur- rent rating and response time of a fuse. They must be located close to power source to minimize the length of unprotected wire; at the same time they have to be accessible for replacement. Spare fuses must be carried on the aircraft and be accessible to the flight crew. Typical requirements are to carry 50% of each rating as spares, e.g. if the aircraft is fitted with four 10A and five 15A fuses, then two 10A and three 15A fuses should be carried as spares.

Figure 7.9 Aircraft circuit-breakers (thermal type): (a) internal schematic (closed and tripped), (b) external features

7.5.3 Limiting resistors These are used to limit current surges, primarily in DC circuits, where the initial current surge is large. When these circuits are switched they cre- ate large current flows that can be harmful to other components and reduce the power supply voltage for a period of time (determined by the time constants of the circuit). Limiting resistors are connected in series with such circuits and then automatically shorted out once the circuit current has stabilized. Typical applications of limiting resistors are found in engine starting circuits and voltage regulators. Limiting resistors are also used in fire extinguishing systems (see also Chapter 16). Fire extinguishers are activated by applying direct current (DC) through current-limiting resistors to the associated squib; this ruptures a disc that allows extinguishing agent to be expelled under pressure. The limiting resistors pre- vent inadvertent operation of the squib. In electron- ics, limiting resistors are used to protect devices such as diodes.

Figure 8.1 Battery busbar schematic The word ‘bus’ (as used in electrical systems) is derived from the Latin word omnibus meaning ‘for all’. Busbars are often formed from thick strips of copper, 8.1 Single engine/general aviation aircraft

principles of the current limiter, an electromechanical device is described.

Figure 8.7 External ground power from a battery pack

Figure 8.6 External power supply schematic 8.1.3 External power The aircraft battery provides an autonomous means of starting the engine. Certain types of operation, e.g. cold weather and repeated starts, could lead to exces- sive demands, resulting in a battery that is not fully charged. An external power supply system schematic (as illustrated in Fig. 8.6) provides power to the air- craft, even when the battery is flat, or not installed.

Figure 8.8 External ground power: (a) external ground power 3-pin connector, (b) external ground power connected engine in the hangar). External power in the form of another battery, a diesel-powered generator or adapted mains supply provides a source of energy. With external power available and switched on, power is now available at the battery relay output terminal via the external power relay. Power is also supplied to the closing circuit (the resistor and diode reduces the

yure 8.9 External ground power 3-pin connector schematic now different since the left generator is supplying less current. To illustrate this effect, if the left generator is supplying 40A and the right generator is supplying 50A, the voltage drops across the respective resistors are: switching loads off the bus is called load-shedding; this can be achieved by automatic or manual control. These loads are connected onto specific busbars that fulfil a specific function. These can be categorized into a hierarchy as illustrated in Fig. 8.11. Connections between busbars are via heavy-duty contactors, or breakers. Aircraft types vary, however the following categories are typical for many installations. supplying 40A and the right generator is supplying 50A,

igure 8.10 Simple twin engine equalizing system schematic This is a completely isolated twin generation system, sometimes called a non-parallel system used on twin-engine aircraft, see Fig. 8.12. Primary power is based on two main AC integrated drive generators (typically 40kVA on each engine). An APU genera- tor (40 kVA) is used as back-up in the event of a main integrated drive generator (IDG) failure. Note that the APU is normally a constant speed device in its own right; therefore an IDG is not required. The advantage of a split-bus system is that the generators do not need power is derived from step-down transformers to

Figure 8.13 Typical electrical power control panel features Key maintenance point

Figure 8.17 Differential current protection fault develops in the feeder line, current CT, flows (but

Figure 8.19 Electrical power breaker schematic In small general aviation (GA) aircraft, pilots need to be familiar with the electrical systems of their air- craft. Most light GA aircraft are only concerned about an alternator or generator failure. This type of aircraft The breaker is closed by an external control switch via contacts A; the coil remains energized via con- tacts B to ground. With the coil energized, the main and auxiliary contacts are closed and the spring is compressed. Contacts A latch the breaker closed, assisted by the permanent magnet. When a trip signal is applied (either by a fault condition or manual selec- tion) current flows to ground in the opposite direc- tion. The spring assists the reversed electromagnetic field and this breaks the permanent magnet latch. A Zener diode suppresses arcing of coil current across

Figure 8.20 Electrical power breaker installation

Figure 9.4 Micro-switch schematic: (a) internal schematic, (b) external features

Figure 9.3 Combined switch/light: (a) typical installation, (b) electrical schematic, (c) external features

of various items, e.g. flaps, gear, doors, etc and com- municates this information to other systems including:

Figure 9.8 Reed relay schematic (a) external schematic, (b) internal schematic between two pole faces formed by the frame of high These are used in control circuits with very low volt- ages or currents; the relay is extremely sensitive and can respond to levels in the order of mA/mV. This low level is often not suitable for the conventional spring- loaded armature device since (at very low pull-in/ drop-out voltages) the contacts would chatter, leading to spark erosion. Polarized relays use magnetic forces to attract and repel the armature instead of a spring force. The armature is a permanent magnet, pivoted between two pole faces formed by the frame of high

Figure 9.10 Relay installation (GA aircraft) Figure 9.9 Relay configurations

Figure9.11 Variable resistors: potentiometers, preset or rheostats 9.4.1 Solenoids

As the applied pressure from the fluid system increases, the tube will tend to straighten out, while 9.5.1 Bourdon tube

with the higher coefficient of expansion is on the outer side of the curve whilst the element is heated and on the inner side when cooled. The element can be formed as a switch contact, called a thermostat, to open or close a circuit when temperature limits are exceeded. Alternatively the element is formed in a spiral shape so that temperature changes cause a shaft to rotate; the typical application for this is an outside air temperature indicator, see Fig. 9.20 (a). The sensing portion of the indicator projects through windscreen (Fig. 9.20(b)) into the ambient air.

legs of a bridge circuit, one leg of which includes the unknown value.

Figure 9.25 Thermocouple principles: (a) the Seebeck effect, (b) simple thermocouple circuit, (c) practical thermocouple circuit

Key maintenance point Test your understanding 9.3

Figure 9.31 AC torque synchro system operation

Figure 10.2 Twin-piston engine aircraft starting the fuel/air mixture between the electrodes. Once the voltage exceeds the dielectric strength of the fuel/air mixture, it becomes ionized. The ionized fuel/air mix- ture then becomes a conductor and allows electrons to flow across the gap. When the electrons flow across the gap, this raises the local temperature to approxi- mately 60,000K. The electrical energy is discharged as heat and light across the air gap, thereby appear- ing as a spark with an audible ‘clicking’ sound. This energy ignites the fuel mixture in the cylinder. Brushes and commutators form the low-tension

via a gearbox drive. The dual-purpose machine is compound-wound and the field is connected via a changeover relay. During engine start it acts like a conventional motor until the engine is up to speed. The changeover relay then automatically connects the field windings to a voltage regulator and it becomes a conventional generator. The starter-generator system has reduced weight and component parts compared with having a separate starter motor and generator, thereby reducing overall operating costs.

conditions. Electrical energy is stored in the HEIU and then dissipated across the igniter plug.

Figure 10.10 Auxiliary power unit (APU) start and ignition slied via the closed 110% switch contacts; the starter

Referring to Fig. 10.9, the end of the igniter plug is formed with the outer casing insulated from the cen- tre electrode. The end of the igniter is coated with a semi-conductive ceramic material. The output from the HEIU heats the surface of the ceramic material and lowers its resistance. This creates a high-intensity flashover from the centre electrode to the outer and lowers its resistance. This creates a high-intensity

The cause of a wet start is most likely to be a defec- tive HEIU and/or igniter plug. The net result is no ignition in the combustion chamber and the accumula- tion of fuel. If compressor outlet air gets hot enough, it could ignite the fuel, causing a rapid expansion of the fuel/air mix (effectively an explosion) that could lead damage of the turbine section and eject flames from the engine exhaust. The procedure is to shut off the

Figure 10.15 Tachometer system installation

The AC generator tachometer consists of a permanent

Explain the terms N1, N2 and N3. How do these relate to the HP, IP and LP shafts?

Small gas turbine engines are normally fitted with several thermocouples to provide an average temperature and some redundancy in case of failure. Larger engines can be fitted with up to 21 thermo- couples; these are connected in parallel to provide an average reading of the gas temperature in the exhaust zone. The interconnecting cables between the thermocouple(s) and indicator have to be the same material throughout the system otherwise addi- tion junctions will be formed, thereby generating unwanted voltages. A typical thermocouple cable installation is shown in Fig. 10.22. an average reading of the gas temperature in the Key point

Figure 10.22 Thermocouple/cable installation Figure 10.21 Average temperature measure- ment

turbine blades to rotate against the force of a spring. The speed of turbine rotation is therefore proportional to fuel flow. Denser fuel imparts more energy into the turbine; the transducer is therefore measuring mass flow. The turbine is connected to the indicator via a synchro system; typical units of mass flow are given in pounds or kilograms of fuel per hour.

Figure 10.25 Fuel flow indicating system

LOTQUE Call disO DE MlCdsurcead Dy CHOCACINE Stra gauges into the shaft to measure the deformation (strain) of the shaft; these can either be metallic foil

Figure 10.29 Shaft torque sensor phase differences (a) no load conditions, (b) engine lightly loaded, (c) engine heavily loaded.

Figure 10.27 Typical helicopter torque indicator (twin engine) temperatures, its viscosity reduces and its lubrication

Figure 11.3 Fuel quantity float valve fuel quantity. One disadvantage of this method is that it is not very stable during aircraft manoeuvres.

A safer alternative to the drip stick method uti- lizes a transparent plastic rod. This method uses the principle of light refraction, and is based on fuel and air having different refractive indexes. When the tip of the rod is moved above/below the surface of the fuel, the light intensity emerging from the viewing end of the rod changes. A reading is then taken from graduations on the rod. Key point

Figure 11.9 Capacitance fuel quantity system — multiple tank units

Figure 11.10 Simple fuel pump system activated (see fire protection, Chapter 16). EE The delivery output from each pump feeds into the system via a non-return valve (NRV). Under normal operating conditions, each pump feeds own engine via a motor driven low-pressure (LP) cock. If a cen- tre tank is fitted as part of a three-tank instalation, this can feed either engine by a fuel transfer system as shown in Fig. 11.13. On some aircraft, tank pumps are located in a dry area of the wing root; on other aircraft the pumps are actually inside the fuel tank. In the lat- ter case, the crew have to maintain certain minimum fuel levels. Control switches for all pumps, cocks and valves together with warning indications are located on the overhead panel or the flight engineer’s station. LP cocks are automatically closed if the fire handle is activated (see fire protection, Chapter 16).

Figure 12.2 Strobe/xenon gas tube product

Key maintenance point Do not handle strobe tubes with bare hands; mois- ture causes local hot-spots that can lead to prema- ture failure.

Figure 12.5 Transistor lighting control circuit typical circuit is for flight instruments, engine instru- ments and switch panels. The increased quantity of instruments requires elec- tronic control in place of rheostat due to the higher loads. A transistor circuit provides electronic control as illustrated in Fig. 12.5; variable resistor Ry, varies the (relatively low) base current into the base of a PNP transistor; this controls the (relatively high) current through the collector and lamp. A typical transistor controlled lighting system is illustrated Fig. 12.6. The relatively low base currents in the respective transistors can now control a variety of lighting circuits:

Figure 12.6 Transistor lighting control system

Figure 12.9 Simplified master warning circuit

Figure 12.8 Master warning and caution lights

igure 12.14 Retractable landing/taxi light circui

Figure 12.17 Exterior lights (beacons/strobes)

The traditional location of the white light is on the tail cone or fin tip; some aircraft have the rear facing light on the trailing edge of each wing tip. The wing lamps are 20 W filament lamps, the tail lamp is 10 W. Coloured filters produce the specific colours; these filters must not shrink, fade or become opaque. A typical navigation light circuit is shown in Fig. 12.19. Note that the lights are controlled by single switch and protection device. The traditional location of the white light is on the

Figure 12.20 Position lights: (a) LED type; (b) wing tip location

Figure 12.21 Helicopter lights: Twin landing lights on nose; Navigation light below door; Additional landing light on rear skid

Figure 12.23 Rotating beacon (top of fin) with a clear glass front cover that is permanently

Figure 12.22 Anti-collision strobe lights

Figure 13.2 Public address (PA) selector panel

Figure 13.5 Comparison of satellite orbits

possible outcomes, or selectable switch positions.

Fibre optic cable bends need to have a sufficiently large radius to minimize signal losses and physical damage.

The light signal passes through holes in the end of the ceramic contacts when they are in direct physical contact with each other. The coupling nut on the plug barrel has a yellow band whilst the receptacle barrel has a red and a yellow band. A correct connection is made when the red band on the receptacle is at least 50 per cent covered by the coupling nut. This position indicates an effective connection in which the opti- cal fibres in the plug are aligned end-to-end with the cal fibres in the plug are aligned end-to-end with the

Figure 13.10 Air conditioning: passenger cabin zones

Figure 13.12 Air conditioning system control (right/centre of overhead panel) 13.7.2 Ventilation 13.7.3 Equipment cooling A re-circulation fan is used to re-circulate filtered cabin air back into the cabin to reduce bleed air requirements. Up to 50% of the cabin air can re- circulated for passenger comfort. The fan will switch off if either A/C pack is in high flow, giving a net reduction in the ventilation rate. Optimum cooling and reduced bleed air demand (hence reduced fuel consumption) is achieved with a combination of air re-circulation and automatic operation of the A/C packs. The ventilation rate is increased on larger aircraft with additional re-circulation fans for comfort levels. Although not part of the cabin system for the comfort

Figure 14.3 Trailing edge flaps control and indication Selecting ‘flaps up’ applies power to the motor through the closed contacts of the up-limit switch; the motor operates and the flaps start to move. The down- limit switch changes over as shown by dotted lines, and the flaps travel to the fully retracted position; note that both the limit switches are closed during

Figure 14.6 Control surface position indication

Figure 14.4 Flap position selection (top right of photo) Different aircraft types have a variety of control sur- face position indicators; the various flying control

ECAM comprises a series of integrated systems that display information to the crew in an efficient way. Referring to Fig. 14.8, aircraft sensors are cat- egorized into key monitoring functions; these sensors transmit data into two system data acquisition concen- trators (SDAC). The data is processed and supplied into flight warning computers (FWC). The FWCs are programmed to identify any inconsistencies in the data, and then output the data through three display management computers (DMC).

Level 2 failures — amber cautions: these are situations have no immediate or direct impact on flight

Figure 15.1 Aerofoil terminology Various systems are installed on aircraft to protect them from a variety of hazards including: stalling, ice, rain, unsafe configuration during takeoff and skidding. Stall protection systems provide the crew with a clear and distinctive warning before the stall is reached. The primary sensor required for this protection system is the aircraft’s angle of attack. Flying in ice and/or rain conditions poses a number of threats to the safe operation of the aircraft. Ice formation can affect the aerodynamics and/or trim of the aircraft. The anti-skid system (also called an anti-lock braking system: ABS) is designed to prevent the main landing gear wheels from locking up during landing, particularly on wet or icy runway surfaces. The configuration warning system (also known as a take-off warning system) provides a warning if the pilot attempts to take-off with spe- cific controls not selected in the correct position, i.e. an unsafe configuration. This chapter describes a variety of protection systems found on a range of aircraft. Basic aerodynamic principles state that increasing the angle of attack increases the lift produced by the wing; when the AoA reaches a certain angle, the air- flow over the wing becomes turbulent and the lift is dramatically decreased. When this occurs, the aero- foil is in a stall condition as illustrated in Fig. 15.3. It can be seen that the streamlined airflow is breaking up and becoming turbulent.

Figure 15.4 Angle of attack and attitude

Key point Figure 15.5 Boundary layer separation

Key point An aerofoil stalls at a certain angle; the speed at which the aircraft is flying makes little difference to the stall angle.

Figure 15.8 Stall warning: vane sensor installation Figure 15.9 Stall warning: vane sensor principles

Figure 15.14 Ice detector: ultrasonic sensor

gure 15.17 Pneumatic ice protection: (a) wing leading edge, (b) electrical connections De-icing allows ice to form and then be removed on a periodic basis.

Larger aircraft use engine bleed air for thermal anti- icing. Fig. 15.19 illustrates how ducts are located along the leading edge of the wing. Spray holes direct the air inside the leading edge before it is ejected overboard. Similar installations are used for tail plane and engine nacelles.

Figure 15.20 Windscreen de-icing circuit Nozzles on the fuselage (forward of the windscreen) direct the repellent spray onto the windscreens. failure. The motors are set by the control switch; typi- cal wiper speeds are:

Figure 15.22 Rain repellent system overview

Figure 15.23 Rain repellent system schematic

Figure 15.25 Take-off warning system: (a) functional diagram, (b) logic diagram

exists. (TCAS is an automatic surveillance system Figure 15.26 Configuration warning system

Figure 16.1 Overheat switch: (a) external features, (b) internal features, (c) electrical schematic

linear fire detectors are used on aircraft; these are based on averaging or discrete detectors. The alarm temperature of averaging linear detectors depends on the length of element heated. Discrete linear detec- tor systems are essentially independent of the length of element heated to provide an alarm; these eutectic salt. The response time for any fire detector is criti- cal to the safety of the aircraft; the standard test for a linear fire detector is for it to respond to a 1100°C flame with five seconds. Actual response time will depend on many factors including the convection of heat energy; this is why it is so important to route the fire detector in the critical areas of the engine nacelle. Linear fire detectors are manufactured in lengths of up to three metres; some types can be joined in series to so that all these areas can be covered in conven- ient lengths. Installing the detector in manageable lengths facilitates installation and removal during maintenance.

Figure 16.2 Overheat switches connected in parallel Test your understanding 16.1

Figure 16.4 Linear fire detector principles: (a) external features (b) detector loop, (c) resistance/temperature characteristics

Figure 16.8 Pneumatic detector characteristics rapidly heated; this now provides sufficient gas pres- sure to operate the alarm switch. When the sensor cools, the hydrogen is re-absorbed back into the cen- tre core, pressure reduces and the alarm switch opens. Should a leak develop in the sensor or responder, the integrity switch opens. The control unit detects an open circuit condition from the change of resistance and this is signalled as a fault condition. The char- acteristics of a pneumatic detector are shown in Fig. 16.8. The entire assembly is hermitically sealed; the responder can be located in the engine fire zone as illustrated in Fig. 16.9.

Figure 16.9 Responder located in the engine fire zone Figure 16.9 Responder located in the engine fire zone

Figure 16.10 Optical fire detection: (a) hydrocarbon fire and 4.4 micron CO, ‘spike’, (b) optical fire detector viewing angle materials; the detection method in widespread use is the smoke detector. The primary design con- sideration for this application is ensuring that the smoke reaches the detector in a timely way. Open area smoke detectors rely on the transfer of particu- late matter from the source of fire to the detector by convection. Smoke from burning materials in cargo, baggage areas often stratifies; in the early stages of a cargo/baggage fire there is insufficient heat energy to cause the smoke to rise through convection. Smoke will eventually accumulate in the roof area of the cargo/baggage area and so this is where the detec- tors are located. In larger cargo/baggage areas several detectors will be installed to speed up the detection time, which can be up to 60 seconds. In some aircraft, air samples are taken from numerous locations and supplied to a centralized ducted smoke detector. The average response time of aircraft engine fire detectors is less than five seconds; this is in response to a standard 1100°C, six-inch-diameter flame. The average response time of aircraft engine fire

Figure 16.11 Optical smoke detector principles detector uses a very small amount of radioactive iso- tope that is a source of alpha radiation. The radia- tion passes through an air-filled chamber between two electrodes; this radiation allows a low constant current to flow between the electrodes. Smoke enter- ing the chamber absorbs the alpha particles, thereby reducing the ionization and interrupting the current; this interruption is used to trigger a warning circuit. Ionization smoke detectors can be manufactured at a lower cost compared with optical detectors; however they have several disadvantages: detector uses a very small amount of radioactive iso-

basis when smoke enters through any of the sampling orifices.

Figure 16.13 Main deck/freighter smoke detectors

through the aperture left by the ruptured diaphragm. When the agent has been expelled, a discharge plug moves to the end of the discharge head; this causes an indicator pin to become visible. On the aircraft fuselage, a polythene plug is forced into the position shown in Fig. 16.16; this provides additional confir- mation that the extinguisher has been fired. If the fire extinguisher is exposed to high temperatures outside of its specified range, a pressure relief valve is opened and the contents discharged overboard.

Cargo bay fire extinguishers are operated from a 16.4.4 Fire extinguisher maintenance

Figure 16.17 Two-shot fire extinguisher system essential that any loss of gas pressure and/or extin- guishing agent be established. One way of ascertain- ing this is to weight the extinguisher on a periodic basis; the disadvantage of this method is the cost of removal/installation together with the logistic costs of transporting the item back to a workshop, together with the additional spares requirements. Most of these fire extinguishers are fitted with a means of indicat- ing the pressure of the contents. This can either be a direct reading gauge, typically using a Bourdon tube, or a pressure switch that is linked via an indicating system to provide a warning to the crew that they have a extinguisher with low pressure; either from a leak or when the extinguisher has been discharged.

Figure 16.18 Cargo fire extinguisher system

During the 1970s, studies were carried out by accident investigators and regulatory authorities to examine one of the most significant causes of aircraft accidents of the time: controlled flight into terrain (CFIT). This can be defined as an accident where a service- able aircraft, under the control of a qualified pilot, inadvertently flies into terrain, an obstacle or water. CFIT accidents usually occur during poor visual con- ditions, often influenced by other factors, e.g. flight crew distraction, malfunctioning equipment or air traffic control (ATC) miscommunication. With CFIT, the pilots are generally unaware of this situation until it is too late. The outcome of these investigations was that many CFIT accidents could be avoided with a ground proximity warning system (GPWS). A system was developed in 1967 to alert pilots that their aircraft was in immediate danger of CFIT. This system was further developed into the enhanced ground proxim- ity warning system (EGPWS) by adding a forward- looking terrain avoidance (FLTA) feature, made pos- sible via global positioning system technology. This chapter describes the generic name given to this type of protection: terrain awareness warning system (TAWS). Fig. 17.1. The combined systems provide a significant These early systems warned the crew of excessive descent rates or if they were flying towards rising terrain. TAWS can be considered as a ground prox- imity warning system (GPWS) combined with forward-looking terrain avoidance (FLTA), see Bio 171 The combined cyvcteme nrovide a cionificant

Figure 17.2 GPS principles: (a) Single satellite describes a circle on the earth’s surface, (b) Two satellites define two unique positions

capability; it can look ahead of and below the aircraft flight path. This can be depicted on a display that is colour-coded (Fig. 17.3) to show the required obsta- cle clearance: capability; it can look ahead of and below the aircraft

Figure 17.5 Low-range radio altimeter: (a) principle of operation, (b) analogue indication, (c) electronic indication

Figure 17.9 GPWS Mode 2 profiles: (a) excessive terrain closure rate, (b) excessive terrain closure rate, flaps and gear down message is produced together with a red text message consistent with the aural message. Mode 2 is inhib- ited below a minimum terrain clearance, typically 30 feet. The computer derives excessive descent rate alert and warning from the following primary inputs:

17.4.6 Mode 6 Altitude callouts is caused by unsafe terrain clearance after take-off or go around; the will produce a ‘TOO LOW TERRAIN’ caution. The various envelopes associated with Mode 4 are illustrated in Fig. 17.14.

Figure 17.12 GPWS Mode 3: (a) altitude loss after take-off profile, (b) altitude loss after take-off envelope

Figure 17.13 GPWS Mode 4 profiles: (a) flight into terrain when not in landing configuration — gear up, (b) flight into terrain when not in landing configuration — gear down/flaps up, (c) flight into terrain when not in landing configuration (take-off or go-around)

Figure 17.15 GPWS Mode 5: (a) excessive downward deviation from an ILS glide slope profile, (b) excessive downward deviation from an ILS glide slope envelope

Figure 17.16 GPWS Mode 6 altitude callouts Key point

Referring to Fig. 17.21, alert areas are projected along the flight path in three levels:

-igure 17.20 GPWS Mode 7: (a) wind sheer detection profile, (b) wind sheer detection envelope

Doppler frequency shift between A and B indicates a wind shear condition

Figure 17.21 Rotorcraft autorotation: (a) normal powered flight, (b) autorotation

the writing head. As each section of wire or tape passes

recorder is switched on, the wire is transferred onto

Figure 18.3 Underwater location transmitter

Figure 18.5 Crash-proof recorder tape drive The tape is transferred from one reel to another by a motor-driven belt, and transported over the write/read heads. The motor is a reversible stepper motor, located outside of the crash-survivable enclosure rotating at a nominal rate of 700 steps per second. Odd-numbered tracks are used when the tape is travelling in the nom- inal forward direction; even-numbered tracks are used in the reverse direction, see Fig. 18.5. When an entire track has been used, the direction of travel is switched and the next track utilized. At the end of recording on track eight, the recording will be switched back to track one, and old data is thereby overwritten. The tape travels at speeds up to five inches per second on an incremental basis when data is being written, or five inches per second continuously when data is being read. Beginning of tape (BOT) and end of tape (EOT) sensors are used to reverse the direction of travel and switch to the next track. These are specific requirements for the crash-survivable

There are many aircraft, however, that have lim- ited or no digital equipment. Depending on the size and category of aircraft, e.g. if it is used for public The acquisition units sample the data to be recorded and converts them into digital form suitable for record- ing. The logic unit also inhibits the recording (REC) function until the aircraft is capable of moving under its own power; this switch is controlled by one of many possible options, e.g. when ground power is removed, when an alternator comes on line or when ground speed increases over a specified value. The TEST switch allows the system to be tested on the ground by overriding the REC logic.

In the Harvard bi-phase format, each bit changes state at its trailing edge; either from high to zero or In the Harvard bi-phase format, each bit changes

Figure 18.8 Digital data formed into frames

Figure 18.9 Digital bits of logic one and logic zero

usually operates at around 65 kHz and an internal sig- nal (at around 600Hz) is provided for test purposes. Bulk erase can be performed by means of an erase switch; this is interlocked so that erasure can only be performed when the aircraft is on the ground and the parking brake is set. The erase current source is usually derived directly from the aircraft’s 115V AC 400 Hz supply. The magnetic tape (a continuous loop) is typically 308 feet in length and % in. wide.

between one electrical circuit to another, i.e. when a signal transmitted in one circuit creates an undesired effect in another circuit. A twisted pair cable consists of two independently insulated wires twisted around one another.

In computers, digital bandwidth refers to the rate at which data is transmitted/received; this is measured in bits per second (BPS). A digital communication network has a given bandwidth in terms of its over- all channel capacity and throughput (consumption). Channel capacity (in BPS) is proportional to the ana- logue bandwidth in hertz (Hz); this is the maximum amount of error-free digital data that can be transmit- ted via a communication link with a specified band- width in the presence of EMI. The deeper theory associated with this subject is beyond the scope of this book; students wishing to research this subject in more detail should refer to the Shannon—Hartley theorem. In computers, digital bandwidth refers to the rate at

Figure 19.10 Lightning strike; dissipation of energy through the aircraft

Electrostatic sensitive devices (including printed

Figure 19.15 ESD Typical wrist strap grounding point

Figure 20.3 Bulkhead wire looms: (a) sealant/potting, (b) connectors

Figure 20.1 Examples of installed wire looms

size and configuration of the aircraft will determine where connectors are needed and located. Other con- siderations are when aircraft sections (wings, fuselage, etc.) are manufactured at various geographical loca- tions and come together at a different place for final assembly. Finally, the need for inspection, removal and installation of equipment needs to be consid- ered; quick-release connectors are used for most line- replaceable units (LRU), see Fig. 20.5. Key maintenance point

Figure 20.6 Terminal blocks with connections must be withdrawn and the strands straightened. Excessive twisting will increase the diameter of the conductor, making it difficult to insert into the bar- rel. Once formed, the tag can be attached to terminal blocks with connections made to respective circuits. Some terminations benefit from a heat-shrink sleeve; this provides extra mechanical protection and support. Heat-shrink material is polythene based and reduces to a pre-determined diameter (but not length) when heated. The advantages of crimping are:

Figure 20.5 Quick-release connectors (LRU) an aircraft to ‘tap’ into a signal line. General rules for the use of splices will be given in the relevant aircraft documents; these will address the:

The bodies and shells of quick-release connectors are made from light alloy, stainless steel or cadmium- plated alloys. Keyways ensure the correct alignment and prevent any twisting movement. The pins slide- fit into the respective socket; the pins are gripped by springs inside the socket to provide the electrical con- tact surface. A number identifies each pin/socket; see Fig. 20.9 for examples.

Figure 20.7 General aviation connectors (D-shaped)

Connectors are released by screws or by twisting the quick-release mechanism. Pins and sockets are removed and inserted using special tools; they are crimped onto the wires in a similar way to crimping of tags and splices. the quick-release mechanism. Pins and sockets are

Figure 20.11 Direct bonding on composite structure

Figure 20.12 Indirect bonding on composite structure

Figure 20.14 Airframe ‘earth return’ circuit

yure 20.15 Earth/ground loops: (a) common return paths, (b) separate return paths

Figure 20.17 Multimeter; analogue or digital Multimeters combine several functions into a single

Figure 20.18 Avometer® (registered trade- mark of Megger Group Limited)

Decibels (dB) are a convenient means of comparing power, voltage and current in electrical and electronic circuits. They are also commonly used for express- ing gain (amplification) and loss (attenuation) in electronic circuits. They are used as a relative meas- ure (i.e. comparing one voltage with another, one current with another, or one power with another). In conjunction with other units, decibels are sometimes also used as an absolute measure. Hence dB V are decibels relative to 1 V, dB m are decibels relative to 1 mW, etc. The decibel is one-tenth of a bel which, in turn, is defined as the logarithm (base 10) of the ratio of power, P, to reference power, Pp. Provided that the resistances/impedances are identical, voltage and cur- rent ratios may be similarly defined, i.e. one-twentieth of the logarithm (base 10) of the ratio of voltage, V, to the reference voltage, Vp, or one-twentieth of the logarithm (base 10) of the ratio of current, /, to the reference current, Jp.

Table A7.1 Copper wire, American Wire Gauge (AWG)

References (1)

Air Conditioning 21-00-00 General 21-10-00 Compression 21-20-00 Distribution 21-30-00 Pressurization control 21-40-00 Heating 21-50-00 Cooling 21-60-00 Temperature control 21-70-00 Moisture/air contamination Chapter 22 Autofl ight 22-00-00 General 22-10-00 Autopilot 22-20-00 Speed-attitude correction 22-30-00 Autothrottle 22-40-00 System monitor 22-50-00 Aerodynamic load alleviating Chapter 23 Communications 23-00-00 General 23-10-00 Speech communication 23-20-00 Data transmission, auto. calling 23-30-00 Passenger address and entertainment 23-40-00 Interphone 23-50-00 Audio integrating 23-60-00 Static discharging 23-70-00 Audio & video monitoring 23-80-00 Integrated automatic tuning Chapter 24 Electrical Power 24-00-00 General 24-10-00 Generator drive 24-20-00 AC generation 24-30-00 DC generation 24-40-00 External power 24-50-00 AC electrical load distribution 24-60-00 DC electrical load distribution Chapter 25 Equipment & Furnishings 25-00-00 General 25-10-00 Flight compartment 25-20-00 Passenger compartment 25-30-00 Buffet/galley 25-40-00 Lavatories 25-50-00 Cargo compartments 25-60-00 Emergency 25-70-00 Accessory compartments 25-80-00 Insulation Chapter 26 Fire Protection 26-00-00 General 26-10-00 Detection 26-20-00 Extinguishing 26-30-00 Explosion suppression Chapter 27 Flight Controls 27-00-00 General 27-10-00 Aileron & tab 27-20-00 Rudder & tab 27-30-00 Elevator & tab ATA chapter/subsystem list 8 Appendix 27-40-00 Horizontal stabilizer 27-50-00 Flaps 27-60-00 Spoiler, drag devices, fairings 27-70-00 Gust lock & damper 27-80-00 Lift augmenting Chapter 28 Fuel 28-00-00 General 28-10-00 Storage 28-20-00 Distribution 28-30-00 Dump 28-40-00 Indicating Chapter 29 Hydraulic Power 29-00-00 General 29-10-00 Main 29-20-00 Auxiliary 29-30-00 Indicating Chapter 30 Ice & Rain Protection 30-00-00 General 30-10-00 Airfoil 30-20-00 Air intakes 30-30-00 Pitot and static 30-40-00 Windows, windshields & doors 30-50-00 Antennas & radomes 30-60-00 Propellers & rotors 30-70-00 Water lines 30-80-00 Detection Chapter 31 Indicating & Recording Systems 31-00-00 General 31-10-00 Instrument & control panels 31-20-00 Independent instruments 31-30-00 Recorders 31-40-00 Central computers 31-50-00 Central warning systems 31-60-00 Central display systems 31-70-00 Automatic data reporting systems Chapter 32 Landing Gear 32-00-00 General 32-10-00 Main gear & doors 32-20-00 Nose gear & doors 32-21-01 Actuators 32-30-00 Extension & retraction 32-40-00 Wheels & brakes 32-50-00 Steering 32-60-00 Position and warning 32-70-00 Supplementary gear Chapter 33 Lights 33-00-00 General 33-10-00 General compartment 33-20-00 Passenger compartments 33-30-00 Cargo and service compartments 33-40-00 Exterior 33-50-00 Emergency lighting Chapter 34 Navigation 34-00-00 General 34-10-00 Flight environment data 34-20-00 Attitude & direction 34-30-00 Landing & taxiing aids 34-40-00 Independent position determining 34-50-00 Dependent position determining 34-60-00 Flight management computing 34-70-00 ATC Transponder Chapter 35 Oxygen 35-00-00 General 35-10-00 Crew 35-20-00 Passenger 35-30-00 Portable Chapter 38 Water/waste 38-00-00 General 38-10-00 Potable 38-20-00 Wash 38-30-00 Waste disposal 38-40-00 Air supply Chapter 51 Structure 51-00-00 General Chapter 52 Doors 52-00-00 General 52-10-00 Passenger/crew

Aircraft Electrical and Electronic Systems: Principles, operation and maintenance

Sign up for access to the world's latest research

AbstractAI

Related papers

Related topics

Related papers

References (1)

Abstract
AI