B2-13f Electrical Student Resources

Student Resource

Subject B2-13f Electrical

Copyright © 2008 Aviation Australia All rights reserved. No part of this document may be reproduced, transferred, sold, or otherwise disposed of, without the written permission of Aviation Australia.

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Part 66 Subject

B2-13f Electrical

CONTENTS Topic Definitions

ii

Study Resources

iii

Introduction

v

Electrical Power (ATA 24)

13.5

Equipment and Furnishings (ATA 25)

13.6

Lights (ATA 33)

13.9

Fire Protection (ATA-26)

13.12

Ice and Rain Protection (ATA-30)

13.14

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DEFINITIONS Define 

To describe the nature or basic qualities of.



To state the precise meaning of (a word or sense of a word).

State 

Specify in words or writing.



To set forth in words; declare.

Identify 

To establish the identity of.

List 

Itemise.

Describe 

Represent in words enabling hearer or reader to form an idea of an object or process.



To tell the facts, details, or particulars of something verbally or in writing.

Explain 

Make known in detail.



Offer reason for cause and effect.


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STUDY RESOURCES JEPPESEN Sanderson Training Products: A&P Technical General Textbook - 2000 A&P Technical Airframe Textbook - 2001 A&P Technical Powerplant Textbook – 1997 AC 43.13-1B Aircraft Inspection and Repair – 1998 Aircraft Instruments & Integrated Systems – EHJ Pallett – 2000 Aircraft Maintenance Text 3 – Practical Science – Part 2 Basic Electricity – 1989 Aircraft Maintenance Text 4 – Basic functional Devices and Systems – 1989 National Training Materials for the Aerospace Industry – Module NAA02 (NAC) US Navy Electricity and Electronics Training Series – September 1998 Student Handout – Subject B1-11f


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INTRODUCTION The purpose of this subject is to allow you to gain knowledge of basic aircraft electrical systems. On completion of the following topics you will be able to:

Topic 13.5

Electrical Power

Explain the operation of an aircraft battery system and describe battery installation procedures Identify components of DC power generation systems and explain system operation. Identify components of AC power generation systems and explain system operation. Identify components of emergency power generation systems and explain system operation. Identify components of voltage regulation systems and explain system operation. Identify components of power distribution systems and explain system operation. Identify inverters, transformers and rectifiers and explain their operation. Identify circuit protection methods and explain their operation. Identify components of external/ground power systems and explain system operation. Describe safety precautions associated with electrical power.

Topic 13.6

Equipment and Furnishing

Identify Electronic Emergency Equipment and describe their requirements and operation Identify and describe the operation of Cabin Entertainment Equipment

Topic 13.9

Lights

Identify the following aircraft lights, state their purpose and explain their operation:  External o Navigation o Anti-collision o Landing o Taxiing o Ice  Internal o Cabin o Cockpit o Cargo  Emergency Describe safety precautions to be observed when performing maintenance on strobe lights.


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Topic 13.12.1 Fire Protection (ATA 26) Identify components of the following systems and explain system operation:  Fire detection  Smoke detection  Fire protection warning  Fire extinguishing Explain the testing of aircraft fire protection systems. Illustrate safety precautions to be observed when performing maintenance on fire protection systems.

Topic 13.12.2 Portable Fire Extinguishers Identify the various types of portable fire extinguishers and explain their application for different onboard classes of fire.

Topic 11.12

Ice and Rain Protection Explain types of ice formation. Identify ice detection components and explain their operation. Identify components of the following anti-icing systems and explain system operation:  Electrical;  Hot air; and  Chemical. Identify components of the following de-icing systems and explain system operation:  Electrical;  Hot Air; and  Chemical. Explain the operation of rain repellent and rain removal systems. List methods used for probe and drain heating and explain their operation.


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TOPIC 13.5: ELECTRICAL POWER (ATA-24) Aircraft Battery Systems All aircraft, almost without exception, use a negative-ground, single wire system. This means that the negative terminal of the battery is connected directly to the metallic portion of the aircraft structure, which places the structure at a negative potential. This allows electrical components to be connected to the structure for a return path to the battery to complete a closed circuit. This is commonly referred to as grounding. Benefits for grounding in this manner include the reduced chance of developing radio frequency interference (RFI) by having fewer wires that can emit RFI energy, and also reduced weight. The electrical components need only a single wire from the positive voltage source, and a negative connection to the aircraft structure. The exception to this design occurs when the aircraft is built using composite materials in place of metallic structures. As most composite materials are non-conductive, a two-wire electrical system must be used.

The illustrated circuit shows a typical battery circuit for a light aircraft. The positive terminal of the battery connects to the battery solenoid, which is a normally-open heavy-duty switch. Some manufacturers refer to a solenoid as a contactor, but both terms are used to describe components that essentially perform the same task; to remotely control a large current source with a small conductor and switch. When the master switch is closed, it completes a ground path to energise the battery solenoid coil. Once the coil is energised, the main power connection is closed in the solenoid to complete the circuit from the battery to the electrical distribution bus.

Freewheeling Diode A freewheeling diode is installed across the coil of the solenoid to eliminate spikes when the master switch is opened and the magnetic field from the coil collapses. Without a freewheeling diode, the voltage spikes could cause damage to sensitive electronic components. The freewheeling diode allows the induced emf (from a collapsing coil) to feed back into itself rather than back into the distribution system.

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Battery Temperature Monitor System Many aircraft that use Ni-cad batteries have a battery temperature monitor system. Typically, this system provides a continuous temperature indication of the battery or batteries. They may also warn of high temperature conditions. A high temperature indication enables action to be taken to isolate the battery and prevent possible damage arising from excessive battery temperature.

Illustrated is a typical battery temperature monitor system. There are two independent temperature sensors (thermistors) in each battery. Each thermistor is mounted on an intercell connector link of the battery. One sensor provides the temperature input to the temperature indication circuit, and the other (overheat temperature sensor) to the overheat warning circuit in the monitor.


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Typical Battery System

Typically, a battery may be used for starting an aircraft. Once the generator is on line, the generator will supply power to applicable systems. It will also charge the battery. Refer to aircraft maintenance manual regarding battery operation on ground etc.

In most aircraft battery installations, the battery will be connected hot to its relevant battery bus. It will only be connected to other DC buses when the battery is turned on. The battery switch will control the battery contactor. Normally, the battery is switched on when the aircraft is operating thus allowing the battery to charge. Most aircraft isolate the battery if external DC power is applied to the aircraft. This stops the build-up of hydrogen which is released from batteries during charging – Hydrogen is highly explosive.


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Ventilation requirements Lead acid batteries when fully charged, give off large quantities of hydrogen and oxygen. Adequate ventilation is therefore required to safely remove these fumes and dump them outside the aircraft. The aircraft’s slipstream can be used to create a positive airflow ventilation system for the battery compartment or sealed case, by creating a pick up and outflow pipe through the aircraft’s skin. Pressurised aircraft can also use the internal pressurised air, which is passed through a non-return valve and finally dumped overboard.

System A utilises a sump jar with a neutralising agent in it, so that the fumes pass over the felt pad, neutralising any acid or alkaline fumes before they exit the aircraft. This prevents damage occurring to aircraft skin. System B is a typical light aircraft system - does not incorporate sump jar. System C is a pressurised aircraft system which passes pressurisation system air through a non-return valve and into the battery case or compartment. It may or may not utilise a sump jar. Ni-Cad batteries in general do not produce large quantities of gas, therefore they do not usually require ventilation. To assist in maintaining more even temperatures in the battery case or compartment, some form of ventilation is usually included. If the battery is of the sealed case type the ventilation system will be connected to the case to maintain some airflow through the case.


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Battery connection Light Aircraft batteries have bolt on connections similar to auto batteries. Smear lugs with petroleum jelly (Vaseline) before connection to inhibit corrosion.

Larger fully enclosed batteries, where external connections cannot protrude though the case safely, will usually have a quick release type connector, as shown, fitted to the case. This type of connector requires some form of positive locking device to ensure that the connector does not unscrew during flight. If there is no inbuilt locking device, the connectors should be lockwired closed with 26 AWG copper wire (safety lockwire).

With all of these quick release type connections, it is important that the terminals be lubricated with white petroleum jelly to prevent any corrosion forming on the contacts which could affect serviceability or worse still, make the battery very difficult to remove. Remember, once a battery is fitted to an aircraft, it may stay there for many months before it is touched again, so if corrosion forms on the contacts you may not see it until it is too late.


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Battery Installation All aircraft batteries need to be securely mounted in the aircraft. Whether it has its own isolated compartment, or shares space with other equipment, it must be firmly bolted into a mounting cradle.

Large metal batteries will have lugs built into the cover through which the hold down bolts can pass. Plastic cased batteries will generally have a metal clamp over the top. Whether wing nuts, hexagonal nuts or quick release clamps are used for securing purposes, all should be lockwired for security purposes.

Battery Installation Procedures Before installing any battery in an aircraft, ensure the battery is the correct type. Refer to aircraft maintenance manual. Before installing, check that its mounting tray, box or compartment and the surrounding area are clean and free of corrosion or contamination. The only servicing permitted on lead acid batteries fitted to an aircraft, is the check and adjustment of electrolyte levels and cleaning of terminals. If any electrolyte is spilt, always neutralise it, before washing it away. Lead acid batteries require a solution of water and bicarbonate of soda, and Ni-Cad batteries require a boric acid solution or vinegar as a neutralising agent. Check that the vent and drain lines are in good condition and not blocked. Check the neutralising jar is serviceable.

Battery servicing Most aircraft have the negative terminal of the battery connected to the aircraft structure. When installing a battery, always connect the positive lead first. This way, if contact is inadvertently made between the battery and the aircraft with a wrench, it will not cause a spark. When removing a battery, always disconnect the negative lead first. If high or difficult lifting is involved don’t try to be a hero and do it all yourself. Get help or use some mechanical lifting device. Slide belt buckle around to the side if there is a possibility of contacting and shorting the terminals with it. Directly after removing a battery from an aircraft, fit insulator plug/cap into terminal connector. Remember: safe work practices must be applied to at all times.


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Generators – General Following on from earlier subjects, it is already known that electromagnetic generators, be they AC or DC, AC is always induced in the output windings. With alternators or AC generators, the generated current is passed to the output terminals as AC. With DC generators, the generated current is rectified to DC either mechanically by a commutator and brushes, or electrically by diodes as in DC alternators.

DC Alternators DC alternators produce relatively small amounts of current and, therefore, are typically found on light aircraft. Some DC alternators found on light aircraft are very similar to the alternators found on motor vehicles.

DC alternators do the same thing as DC generators; They produce AC that is then converted to DC before it enters an aircraft electrical system. The difference, however, is that in an alternator the magnetic poles rotate and induce voltage into a fixed, or stationary, winding. Furthermore, the AC produced is rectified by diodes instead of a commutator. All alternators are constructed in basically the same way. The primary components of an alternator include: •

The rotor.

•

The stator.

•

The rectifier.

•

The brush assembly.


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Rotors An alternator rotor consists of a wire coil wound on an iron spool between 2 heavy iron segments with interlacing fingers. Some rotors have 4 fingers while others have as many as 7. Each finger forms one pole of the rotating magnetic field. The 2 coil leads pass through 1 segment and each lead attaches to an insulated slip ring.

The slip rings, segments, and coil spool are all pressed onto a hardened steel rotor shaft which is either splined or has a key slot. In an assembled alternator, this shaft is driven by an engine accessory pad or fitted with a pulley and driven by a belt. 2 carbon brushes ride on the slip rings to bring a varying current into the rotating field.

Diode Rectifier Bridge The Diode Rectifier Bridge is responsible for the conversion or rectification of AC voltage to DC voltage. Typically, six or eight diodes are used to rectify the AC stator voltage to DC voltage. Half of these diodes are use on the positive side and the other half are on the negative side. Diodes are used as one-way electrical check valves. Passing current in only one direction, never in reverse. Diodes are mounted in a heat sink to dissipate the heat generated by the diodes. Diodes redirect the AC voltage into DC voltage so the battery receives the correct polarity


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Stator As the rotor turns, load current is induced into stationary stator coils. The coils making up the stator are wound in slots around the inside periphery of the stator frame, which is made of thin laminations of soft iron. Most alternators are 3 phase alternators. This means that the stator has 3 separate coils that are 120º apart. To do this, one end of each coil is brought together to form a common junction of a Y-connection.

With the stator wound in a 3-phase configuration, the output current peaks in each set of windings every 120º of rotation. However, after the output is rectified, the DC output becomes much smoother. Because an alternator has several field poles and a large number of stator windings, most alternators produce their rated output at a relatively low rpm. This differs from a DC generator which must rotate at a fairly high speed to produce its rated output.


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Wye Wound Stator Wye wound stators have three windings with a common neutral junction. They can be identified because they have 4 stator lead ends. Wye wound stators are used in alternators that require high voltage output at low alternator speeds. Two windings are in series at any one time during charge output.

Delta Wound Stator Delta wound stators can be identified because they have only three stator lead ends. Delta stators allow for higher current flow being delivered at low RPM. The windings are in parallel rather than series as like the Wye design.

Brush Assembly Two stationary carbon brushes ride on two rotating slip rings. Bushes are either soldered or bolted. Two slip rings are located on one end of the rotor assembly. Each end of the rotor field winding is attached to a slip ring. Thereby, allowing current to flow through the field winding.


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Generator Voltage Regulation Efficient operation of electrical equipment in an aircraft depends on a voltage supply that varies with a system's load requirements. Among the factors which determine the voltage output of a generator, the strength of the field current is the only one that is conveniently controlled. One way to control the field current is to install a rheostat in the field coil circuit. When the rheostat is set to increase the resistance in the field circuit, less current flows through the field coils and the strength of the magnetic field decreases. Consequently, less voltage is induced into the armature and generator output decreases.

When the resistance in the field circuit is decreased with the rheostat, more current flows through the field coils, and the magnetic field becomes stronger. This allows more voltage to be induced into the armature which produces a greater output voltage. One thing to keep in mind is that, the weaker the magnetic field is, the easier it is to turn the armature. On the other hand, if the strength of the magnetic field is increased, more force is required to turn the armature. This means that, when the load on a generator increases, additional field current must be supplied to increase the voltage output as well as overcome the additional force required to turn the armature


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Vibrating-type Voltage Regulator This basic rheostat principle is further developed by the addition of a solenoid which electrically connects or removes the field rheostat from the circuit as the voltage varies. This type of setup is found in a vibrating-type voltage regulator.

When the output voltage rises above a specified critical value, the downward pull of the solenoid's coil exceeds the spring tension and contact B opens. This reinserts the field rheostat in the field circuit. The additional resistance reduces the field current and lowers output voltage. When the output voltage falls below a certain value, contact B closes, shorting the field rheostat and the terminal voltage starts to rise. Thus, an average voltage is maintained with or without load changes. The dashpot P provides smoother operation by acting as a dampener to prevent hunting, and capacitor C across contact B helps eliminate sparking. With a vibrating-type voltage regulator, contact B opens and closes several times per second to maintain the correct generator output. Based on this, if the solenoid should malfunction or the contacts stick closed, excess current would flow to the field and generator output would increase.


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Three-Unit Regulator Certain light aircraft employ a three-unit regulator for their generator systems. This type of regulator includes a current limiter, a reverse current cutout, and a voltage regulator. The action of the voltage regulator unit is similar to the vibrating-type regulator described earlier. The current limiter is the second of three units, and it limits the generator's output current. The third unit is a reverse-current cutout which disconnects the battery from the generator when the generator output is lower than the battery output. Since contacts have a tendency to pit or burn when large amounts of current flow through them, vibrating-type regulators and three-unit regulators cannot be used with generators that require a high field current.


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Current Limiter In some generator systems, a device is installed which will reduce the generator output voltage whenever the maximum safe load is exceeded. The current limiter operates by having a current coil connected in series with the load, as shown in Figure. When electrical load demands are heavy, the voltage output of the generator may not increase sufficiently to cause the voltage regulator to open its contacts. Consequently, as the current flow increases so does the magnetic flux around the current coil. When sufficient magnetic flux/field is built up to overcome the armature spring tension then the contacts open. Opening the contacts places a resistor in series with the generator field winding, thereby reducing the generator output voltage which consequently reduces the output current. Maximum current adjustment on the current limiter is achieved by adjusting the spring tension applied to the armature, until the maximum current rating is set.

Reverse Current Cutout Relay (RCCR) The RCCR is a combination of relays in one unit. The combination provides current and voltage sensing. These relays allow current in one direction as a normal relay, but open the circuit for a reverse current. Normal flow is from the generator (GEN) terminal to the battery (BAT) terminal of the relay. L5 – voltage (shunt) winding made of many turns of fine wire. L4 – current (series) winding made of a few turns of heavy wire. In normal operation (Gen output higher than battery voltage) both coils are wound to assist each other. Initially at generator start-up, only the voltage coil actually pulls the contacts close. At this point the current coil comes into play. Should the generator output voltage drop below the battery voltage for any reason, current direction will change through the current coil (as battery now attempts to supply generator) and the magnetic field produced will oppose that of the voltage coil and the overall magnetic pull will decrease and the contacts will open. If the battery were not disconnected, it would discharge through the generator armature when the generator voltage falls below that of the battery. When this occurs the battery attempts to drive the generator as a motor. This action is called ‘motoring’ the generator and, unless prevented, the battery discharges in a short time.


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Carbon Pile Voltage Regulator Since contacts have a tendency to pit or burn when large amounts of current flow through them, vibrating-type regulators and 3-unit regulators cannot be used with generators that require a high field current. Therefore, heavy-duty generator systems require a different type of regulator, such as the carbon-pile voltage regulator.

The carbon-pile voltage regulator relies on the resistance of carbon disks arranged in a pile or stack. The resistance of the carbon stack varies inversely with the pressure applied. For example, when the stack is compressed, less air exists between the carbon disks and the resistance decreases. However, when the pressure is reduced, more air is allowed between the disks causing the resistance to increase. Pressure on the carbon pile is created by two opposing forces: a spring and an electromagnet. The spring compresses the carbon pile, and the electro magnet exerts a pull on the spring which decreases the pressure. Whenever the generator voltage varies, the pull of the electromagnet varies. If the generator voltage rises above a specific amount, the pull of the electromagnet increases, thereby decreasing the pressure exerted on the carbon pile and increasing its resistance. Since this resistance is in series with the field, less current flows through the field winding and there is a corresponding decrease in field strength. This results in a drop in generator output. On the other hand, if the generator output drops below a specified value, the pull of the electromagnet decreases and the carbon pile places less resistance in the field winding circuit. This results in an increase in field strength and a corresponding increase in generator output. A small rheostat provides a means of adjusting the current flow through the electromagnet coil. Adjustment to core pressure is done via adjustment screws.


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Voltage Regulator Types

Two regulator designs can be used: The Grounded Regulator type – The regulator controls the amount of battery ground (negative) going to the field winding in the rotor. The Grounded Field type – The regulator controls the amount of Battery Positive (B+) going to the field winding in the rotor.

Transistorised Voltage Regulator – DC Alternator Transistorised voltage regulators are taking over from the other forms of voltage regulators in modern aircraft. The regulator discussed in this section is used on a light aircraft alternator rectifier unit.


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Operation Refer to Figure and assume the master switch is turned ON. The charge indicator light illuminates gaining earth via the regulator D+ terminal, the alternator Ind, terminal, the alternator rotor and transistor T1. This small current gives the field initial excitation. Once the engine is started, the alternator output builds up. It is half wave rectified by the field exciter diodes, and supplies the alternator rotor field gaining earth via transistor T1. When the alternator voltage approaches battery voltage, the charge indicator light has near equal potential on each side and will extinguish. Alternator output voltage is impressed across resistors R1, R2 and R3. When the set voltage value is reached, the zener diode ZD will conduct increasing current flow through R6 and biasing transistor T2 so it conducts. This flow of current through T2 & R4 causes a voltage drop across R4 which makes the base of transistor T1 less positive & T1 switches off. T1 switching off stops the rotor field current and the alternator voltage drops. This drop in voltage across R1, R2 and R3 causes the zener diode to stop conducting and transistor T2 switches off. T2 switching off lowers the voltage drop across R4 and the base of transistor T1 becomes more positive and switches on allowing field current to flow. This process is repeated at a rate required to maintain alternator output voltage at the set value with varying load and speed conditions. The surge quench diode in the regulator is connected across the alternator rotor field winding. When power is applied to the field by transistor T1, the surge quench diode is reverse biased and will not conduct. When transistor T1 switches off the field current, the magnetic field around the field coil collapses inducing a current in the field coil with opposite polarity to that of the applied voltage (inductor theory). The opposite polarity forward biases the diode which conducts and shorts out the current induced in the field coil, preventing transistor T1 from sustaining any damage if the induced voltage exceeds the peak inverse voltage of T1. Note Late model transistorised regulators are now mounted in the alternator rectifier unit and are sealed. They are encapsulated, and if they are unserviceable, repair can only be done by replacing them with a new item. Some light twin engine aircraft have one regulator controlling both alternators and one regulator as an alternate that can be selected by the pilot if the first regulator fails.

Voltage Regulators Regardless of whether it is an AC generator, DC generator or DC alternator, typically the voltage regulator senses the output and controls (regulates) the field current under varying load conditions. As load conditions change, the voltage regulator adjusts the field current automatically (typically) maintaining the generator output at the required output voltage. Typically, aircraft DC generators/alternators are regulated to 28 VDC (14 VDC for 12 volt light aircraft) and AC generators to 115 VAC (200 VAC 3 Ø).


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AC Alternators AC alternators can be single phase, two phase, 3 phase and sometimes even 6 phase or more. However, almost all aircraft AC electrical systems use 3 phase. There are 2 types of alternators/AC generators: •

Revolving armature type.

•

Revolving field type.

Typically, revolving-field type AC generators are used on aircraft.

Revolving-armature Alternator The revolving-armature alternator is similar in construction to the DC generator in that the armature rotates in a stationary magnetic field as shown in figure.

In the DC generator, the emf generated in the armature windings is converted from AC to DC by means of the commutator. In the alternator, the generated AC is brought to the load unchanged by means of slip rings. The rotating-armature is found only in alternators of low power rating and generally is not used to supply electric power in large quantities. Note the AC output is via the brushes and slip rings and therefore must carry the full load current. This is a disadvantage of revolving-armature alternators.


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Revolving-field Alternator The revolving field type alternator has a stationary armature winding (stator) and a rotating field winding (rotor).

The advantage of this configuration is that the armature is connected directly to the load without sliding contacts in the load circuit. Direct connection to the armature makes it possible to use large cross-section conductors. Typically, aircraft AC generators are of the rotating-field type.

Brushless Alternators The AC alternators used in large jet-powered aircraft are of the brushless type and are usually air cooled. Since the brushless alternators have no current flow between brushes or slip rings they are very efficient at high altitudes where brush arcing is often a problem. As discussed previously, alternator brushes are used to carry current to the rotating electromagnet. However, in a brushless alternator, current is induced into the field coil through an exciter.


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A brushless alternator consists of three separate fields, a permanent magnetic field, an exciter field, and a main output field. The permanent magnets furnish the magnetic flux to start the generator producing an output before field current flows. The magnetism produced by these magnets induces voltage into an armature that carries the current to a generator control unit, or GCU. Here, the AC is rectified and sent to the exciter field winding. The exciter field then induces voltage into the exciter output winding. The output from the exciter is rectified by six silicon diodes, and the resulting DC flows through the output field winding. From here, voltage is induced into the main output coils. The permanent magnet, exciter output winding, six diodes, and output field winding are all mounted on the generator shaft and rotate as a unit. The three-phase output stator windings are wound in slots that are in the laminated frame of the alternator housing. The main output stator winding ends of a brushless alternator are connected in the form of a Y, and in the case of the previous figure, the neutral winding is brought to the outside of the housing along with the three-phase windings. The figure below is from a 747-400 Brushless Alternator. The figure clearly shows the rotor and stator, as well as the Current Transformer Assembly on the feeder cables.


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Brushless Alternator GCU (Generator Control Unit) The GCU actually monitors and regulates the main generator's output by controlling the amount of current that flows into the exciter field. For example, if additional output is needed, the GCU increases the amount of current flowing to the exciter field winding which, in turn, increases the exciter output. A higher exciter output increases the current flowing through the main generator field winding thereby increasing alternator output. Typically, the separate generator control unit (GCU) contains the PMG rectifier, main field regulator and aircraft bus contactor controls. The protective functions are also built into this control unit such as differential/feeder faults or overvoltage etc..

Integrated Drive Generator – IDG Some modern jet aircraft produce AC with a generator called an Integrated Drive Generator or IDG. An IDG differs from a CSD in that both the Constant Speed Drive (CSD) and the AC generator are in the same housing.

This is an example of a modern integrated drive generator (IDG). The term generator is still used even though these items are actually brushless alternators.

Flashing the Field A generator may lose it’s residual magnetism if left standing for very long periods or for some other reason. ‘Flashing the field’ will enable the generator to start producing an output if there is no residual magnetism. This is usually performed by using battery power directly connected to the field. Some aircraft have the facility to ‘flash the field’ from within the cockpit. Always refer to the manufacturers directions if required to flash the field of a generator.


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Differential / Feeder Fault This system is installed to protect against the possibility of a generator feeder cable shorting to airframe. Feeder cables are the power cables that run between the generators and the power distribution buses. Differential fault protection is typically the only protection for the generator feeder cables.

The output of each generator is sensed at the generator and also at the load bus. This is achieved using toroidal current transformers. If a difference exists between the two, then there is a ground or short in the feeder lines between the generator and the load bus. This fault is called a differential fault or a feeder fault. It is serious enough that when it trips the field relay and generator breaker, the field relay is typically locked out and may not be reset without resorting to an abnormal procedure to defeat the lockout. If one of the current transformers (in a pair) was to go open circuit, the outputs of both would be different and therefore sensed as a differential fault and the generator would be tripped off line.

These toroidal current transformers are typical of those used for differential fault protection of a 3Ø AC generator.


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Emergency Power Generation Emergency power generation is installed on many aircraft to supply essential power in the event of losing all normal power generation. Such a situation could arise if the engine driven generators stopped outputting. This could occur if the fuel ran out or the engines flamed out (eg. volcanic ash ingestion). Emergency generators are driven either directly by a wind driven generator or driven from a hydraulic motor being supplied pressure from a wind driven hydraulic pump. Emergency generators are, normally without exception, permanent magnet alternators and are typically constructed with a permanent magnet rotor and do not require battery power to function.

Air Driven Emergency Generator The illustrated emergency power supply is a manually deployed, air driven, emergency electric supply for small aircraft. This is a 28 volt DC generator. The unit is capable of supplying electric power to drive a basic complement of flight critical instruments. If a primary electrical system failure should occur, the pilot deploys the miniature turboalternator. The flight can then be continued to the intended destination, or terminated at a convenient airport under more casual conditions. The turboalternator is a true stand-by power supply. In the event the aircraft loses its total electrical system the alternator will still function. The unit is self-exciting, thus it will work in the absence of the battery. It utilises a Permanent Magnet Generator.


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Ram Air Turbine – RAT The RAT is used in emergency situations to generate sufficient power to keep an aircraft flying long enough to land. The RAT typically consists of a propeller that is spun by the highspeed air flowing past the aircraft. The spinning propeller powers a turbine that provides the emergency power needed to keep critical systems running, such as hydraulics, flight controls, and key avionics. In normal flight, the entire assembly is folded up and stored in the aircraft fuselage or wing. Airbus A330 is equipped with a ram air turbine as evidenced by this illustration.


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System Description (Airbus A330) The AC emergency generation which consists of the Hydraulic Electrical Generating System (HEGS) enables to restore part of the distribution network in case of emergency configuration (loss of the two main busbars: AC 1 and AC 2). The HEGS is composed of the following items: the hydraulic Constant Speed Motor (CSM), the AC generator & the Generator Control Unit (GCU). The CSM & AC generator assembly is called the Constant Speed Motor/Generator (CSM/G) 8XE. In case of loss of the two main busbars AC 1 & AC 2, the CSM is supplied by the green hydraulic circuit which is powered either by the green hydraulic engine driven pump(s) (one pump driven by engine 1 and/or the other by engine 2) if one engine at least is operative, or by the Ram Air Turbine (RAT) pump in case of loss of both engines. In the case of the loss of the AC main busbars, the CSM supplied by the hydraulic fluid drives the AC generator. The AC generator is controlled by the CSM/G Generator Control Unit GCU (1XE) and provides the AC power supply (115 VAC/400 Hz, 3 phases). Constant Speed Motor/Generator (CSM/G) supplies the AC ESS buses through the EMER GLC, & the DC ESS buses through the ESS TR. The two main batteries (system 24-35-00) supply the DC ESS BUS and, via the static inverter, the AC ESS BUS during the transfer to emergency generator operation after the loss of the main busbars. The AC ESS SHED BUS and the DC ESS SHED BUS are shed respectively by the relay 16XH and the contactor 1PH when the CSM/G is supplied by the RAT hydraulic pump and, in batteries only configuration. AC ESS BUS SHED (4XP) Network The AC ESS BUS SHED (4XP) supplies 3ø 115 VAC/400 Hz power: through the relay 16XH, to the sub-busbar 401XP, NOTE : This sub-busbar is automatically shed in emergency configuration when the CSM/G is supplied by the RAT pump or when the static inverter operates. Configuration of the DC network: DC SHED ESS BUS 8PP and the associated sub-busbars) is shed through the contactor 1PH in emergency configuration when the network is supplied by the CSM/G powered by the RAT PUMP, or by the batteries only. In emergency configuration (CSM/G), sub-busbars 407PP and 805PP are recovered through the contactor 9PH as soon as the LAND RECOVERY pushbutton switch located on the EMER ELEC control panel (overhead panel) is pushed.


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Constant Speed Motor/Generator System description Constant Speed Motor The constant speed motor has a variable displacement axial piston hydraulic unit which is controlled by a servo valve. This servo valve is controlled by the GCU speed regulator and transforms the hydraulic power of the Green hydraulic system into constant speed for the generator. The CSM is equipped at the inlet with a solenoid control valve. The solenoid is supplied with 28 VDC from BAT 2 first, then from the PMG when the steady state operation is reached. The solenoid control valve is used to isolate the hydraulic power from the CSM/G when not used. It is also used to stop the CSM/G in the event of failures detected by the CSM/G GCU. AC generator The generator consists of three stages: a main generator, an exciter, & a permanent magnet generator. It provides a three phase 115 VAC/200 VAC voltage at a constant frequency of 400 Hz with neutral output. Main generator The main generator is composed of a laminated stator, with a 6-pole, three-phase winding and a salient pole laminated rotor. The rotor is supplied from the rotating member of a 3 phase exciter via diodes mounted on a rotating carrier. The diodes are configured in 3-phase full wave. Exciter The exciter consists of: a laminated rotor stack containing a 3-phase 10 pole winding & a 10 coil wound laminated stator. The exciter receives its field current from the PMG via the GCU. Permanent Magnet Generator (PMG) The PMG consists of a laminated stator stack containing a 18-pole 3-phase winding & a permanent magnet rotor. The PMG supplies the necessary power for the following: the voltage regulator, the exciter field winding, the speed regulator, the solenoid control valve and the GCU protection circuit. CSM/G main features Voltage : 115/200 VAC plus or minus 4 VAC at POR Frequency : 400 Hz plus or minus 4 Hz Power Factor : 0.75 to 1 lagging The generator stator is cooled by forced air provided by an integrated fan driven its shaft. The generator rotor is cooled by the hydraulic fluid.


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The B767 hydraulic motor generator (HMG) provides power to standby system, captains flight instruments, & selected navigation, communication, lighting & anti-icing loads when both main AC buses are unpowered in flight. The HMG shutoff valve controls hydraulic fluid flow to the hydraulic motor generator. The HMG is composed of three major components, the hydraulic motor, electrohydraulic servovalve & generator all assembled as one line replaceable unit. An electrohydraulic servovalve controls motor displacement to maintain a constant generator speed of 12,000 RPM. The servovalve is controlled by a feedback signal from the generator output. Speed control circuits are located in the generator control unit (GCU). A mechanical overspeed governor provides backup speed control. The brushless, self-excited generator is comprised of a main generator, an exciter generator & a permanent magnet generator (PMG). The main generator develops 5 kva, 3 ø, 120/208 VAC, 400 Hz & 50 ampere, & 28 VDC. The HMG Generator Control Unit (GCU) provides both main generator voltage regulation & control sensing for system operation. The primary operational functions of the unit are voltage regulation & control field excitation, undervoltage & underfrequency protection, application or removal & lockout of the power ready signal to external bus transfer relays, & an electrical speed control signal to the electrohydraulic servovalve in response to deviations in generator output frequency.


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HMG “ON” With Loss of AC Buses With the left & right 115 VAC buses powered, K858 & K859, the L & R AC BUS OFF relays energised, & after a 0.5 second time delay, the K860 HYD GEN CONT relay close coil is energised. The loss of voltage to both main AC buses in flight for at least 0.5 seconds (both K858 & K859 relax) energises the open coil of K860. HMG Hydraulic Motor Function As the motor rapidly accelerates, the PMG provides GCU operational power. The PMG frequency is also used as a feedback signal to the GCU speed control module. The frequency feedback signal is compared to a reference frequency & any error is transmitted to the electrohydraulic servovalve to modulate displacement. HMG Electrohydraulic Servovalve Function Electrohydraulic servovalve primary components are the torque motor, jet pipe, slide valve & feedback spring (not shown). A change in generator load causes a speed change. The speed control module supplies servo power & completes the circuit to either increase or decrease speed coil. Motor displacement increases or decreases causing an increase or decrease in motor speed. When the slide valve feedback spring torque equals the torque motor torque, the jet pipe returns to center position stopping flow to the slide valve. The electrohydraulic servovalve maintains main generator output frequency at 400 +/-2 Hz under steady-state conditions. In the event of feedback signal or electrohydraulic servovalve failure a mechanical overspeed governor maintains the frequency at 430 +/-10 Hz. HMG Generator Function The GENERATOR portion of the HMG is composed of a permanent magnet generator, exciter generator & main generator with rotors mounted on a common shaft. The permanent magnet generator supplies 3ø 800 Hz to the GCU for excitation, control & protection functions. A ten-pole AC exciter generator supplies a 3ø AC output to a half-wave rotating rectifier bridge for main generator field excitation. The GCU voltage regulator controls excitation current to maintain a nominal 120 volts AC line to neutral at the Point of Regulation (POR). The main generator has two output windings, AC & DC; one to produce 5 kva, 400 Hertz, 120/208 volt three phase AC power & the other winding output is full wave bridge rectified to supply 24 to 30 VDC up to 50 amps. HMG “OFF” With AC Buses Recovered If power to both main AC buses is recovered, the ADP ’ON’ signal is removed & the hydraulic shutoff valve closes automatically after a 0.5 sec’s. The HMG shuts down & the bus transfer relays are de-energised. Normal Operation The permanent magnet generator (PMG) power enters the GCU & is three phase full-wave rectified to produce 28 volt dc power. The power is used by the GCU voltage regulator & sensing circuits. The average of the point of regulation ac voltage is monitored by the voltage regulator for supplying exciter generator field excitation.


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Generator Power Ready Relay When hydraulic motor generator output voltage & frequency reach the power ready limits (100 volts ac & 380 Hz) & K873 GEN RESET relay is relaxed, enabling signals are applied to gate 1. The Generator Power Ready Relay closes & dc power is applied to K861, K862, K863, K864. Ground signal removal from the sensing level shift circuit enables the undervoltage & underfrequency trip levels at 104.5 +/-1.5 volts ac (three phase average) & 385 +/-5 Hz. The fault signal path from gate 2 to gate 3 is inhibited.

Bus Transfer Relays When energized K861 HYD GEN POWER relay connects dc generator output to the hot battery bus & disables K873 GEN RESET relay. Energizing K862 INSTR TRANSFER BUS, K863 & K864 L(R) TRANSFER BUS relays connects ac generator output to the captain’s flight instrument, left & right transfer buses. Also, both EICAS computers receive a dc voltage input (K862 energized) that activates the ’HYD GEN ON’status & maintenance message (NVM in air). A ground applied from K864 to K868 EMER LTS ALT TRANSFER relay switches the emergency light power source from right dc bus to battery bus. Grounding the L radio distance magnetic indicator (L RDMI) inertial reference system (IRS)source select input changes heading data source from right IRS to center IRS.


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Fault Sensing Operation An undervoltage or underfrequency fault condition that lasts longer than 9 +/-1 seconds trips the generator power ready relay, de-energizing the bus transfer relays. The generator still receives excitation current. A ground applied to the sensing level shift circuit enables power ready sensing. K873 applies a reset signal automatically for 1.5 seconds after the trip occurs. If the fault condition has not cleared, the generator power ready relay is inhibited from closing by a fault signal from gate 2 through relay contacts to gate 3. If the fault condition has cleared, the 1.5 second reset signal removes lockout & after K873 relaxes, the voltage regulator is enabled. The generator power ready relay closes after power ready limits are reached. Bus transfer relays reclose automatically to re-energize loads. System Reset: f power to both main ac buses is recovered, the hydraulic motor generator shuts down & deenergizes the bus transfer relays.

Purpose The RAT is an air turbine driven hydraulic pump that provides emergency hydraulic power for the flight controls in the event that power is lost on both engines. Description The RAT is stowed inside the right, aft body fairing. When deploying, the RAT pivots downwards to extend the turbine into the airstream. An electric motor drives the unit for both deployment and stowing.


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Hydraulic Pump The Abex variable displacement nine piston, pressure compensated pump has a rated output flow of 11.3 gpm at 2140 psi at a governed speed of 4165 ± 45 rpm at a minimum airspeed of 125 knots. Electric Motor Extends and retracts RAT by rotation of motor arm and actuation link. 15° Limit Switch Permits partial retraction of RAT so that blade can be locked & blade centered switch can be closed. Down/Up Limit Switches Provides extend/retract motor shut off at full down or full up travel limits. Automatic Deployment Automatic deployment of the RAT only occurs in the air mode when N2 speed on both engines falls below 50% and airspeed is 80 kts or greater. Manual Deployment WARNING: RAT extends rapidly. To avoid serious injury clear RAT area of all personnel and guard area when extending. The RAT may also be extended in either the air or ground mode by depressing the RAT manual override switch located on the P5 overhead panel. Retraction Retraction of the RAT is possible only on the ground and is controlled by a guarded switch near the checkout module on the keel beam in the right main gear wheel well. Ground Checkout (Backdrive) The RAT checkout module contains provisions for ground testing of the RAT. To accomplish ground checkout the center hydraulic system must be pressurized with the air driven pump (ADP), (or adequate external ground hydraulic power).


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RAM AIR TURBINE CONTROL LOGIC Automatic Deployment Automatic deployment of the RAT occurs when in the air mode, the calibrated air speed is greater than 80 knots and both engines have an engine speed of less than the low idle speed. The engine speed circuits are used to monitor N2 on both engines. Manual Deployment Manual deployment of the RAT can be made at any time, ground or air mode, by depressing the guarded two position RAT control & indication switch S7 on the P5 overhead panel. Power is from the hot battery bus. NOTE: Manual and automatic deploy circuits utilise separate relays and limit switches to

provide system redundancy. Hot battery bus only required for manual deploy and retract, battery bus and hot battery bus required for auto deploy.

Indications An amber RAT unlocked light in the RAT control and indication switch (S7) on the P5 panel will illuminate any time the RAT is not in the stowed position and an EICAS RAT UNLOCKED advisory message will be annunciated. A green pressure light in switch S7 will illuminate when the RAT is developing above approximately 1700 PSI.


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Retract Control Before the RAT can be retracted the RAT control and indication switch on the overhead P5 panel must be reset in the OFF position. The RAT retract switch is located above the RAT tachometer module in the right wheel well. The blade lock and centered limit switch (13.5°) and the strut angle switch (15°) operate in parallel in the RAT retract circuit. When the RAT is fully extended the blade lock and centered switch is open (Blade lock disengaged) and the strut angle switch on the RAT trunnion is closed. The RAT will retract the first 15° through the strut angle switch. Above 15° the strut angle switch will open, and the retract circuit can be completed only when the blade is centered and the locking pin is in place, permitting the blade locked and centered switch to close. The Boeing 757 and 767 have a RAT that only supplies hydraulic power. A variable pitch blade of 41" produces about 15 GPM based on airspeed, only to the primary flight controls. 767 does have a small hydraulic driven generator off the primary hydraulic system. The generator is not part of the RAT. The 777 has a RAT with both the 15 GPM hydraulic pump to flight controls (not high lift, plus there are no primary and secondary named flight controls, just flight controls on 777) plus a 5 KVa generator. It also has a variable pitch blade about 50". 757, 767, 777 have auto deploy based on different conditions. 757 and 767 are in air, >80 kias, and N2 on both engines < 50 %.

Ram Air Turbine (RAT) and Engine Start and Ignition Panel (Boeing 767) RAM AIR TURB Switch - Deploys RAT when pushed and held for at least two seconds. PRESS Light - Indicates RAT is supplying satisfactory Pressure to the center hydraulic system. UNLKD Light - Indicates RAT is not Locked in the retracted position. The RAT is primarily used on military fighters and transports while some advanced designs employ emergency power units (EPU) that burn hydrazine to generate emergency power. However, APUs and RATs are designed for extended use while the EPU only functions for a brief period that allows the pilot to safely eject.


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RAT – Precaution The extension time for a RAT is approximately 2 seconds. Ensure all personnel are clear prior to deployment as such deployment could result in serious or fatal injury.


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Inverter An inverter is used in some aircraft systems to convert a portion of the aircraft's DC power to AC. This AC is used mainly for instruments, radio, radar, lighting, and other accessories. These inverters are usually built to supply current at a frequency of 400 cps, but some are designed to provide more than one voltage; for example, 26 volt ac in one winding and 115 volts in another. There are two basic types of inverters: the rotary and the static. Either type can be single phase or multiphase. The multiphase inverter is lighter for the same power rating than the single phase, but there are complications in distributing multiphase power and in keeping the loads balanced.

Rotary Inverter A rotary inverter is basically a DC motor driving an AC generator. Typically, they consist of a 28 V DC compound wound motor driving a 115 V AC single phase or 115/200 V AC three phase alternator. Frequency control is achieved by regulating the motor field current. AC output voltage is controlled by regulating the alternator field current.

Rotary inverters are reliable and produce a pure sine-wave output. Disadvantages of a rotary inverter are lack of frequency control, low surge capability (50% above maximum rating) and lower efficiency (50% to 80%). Rotary inverters are not as common as electronic (static) inverters.


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Static inverters Convert DC to AC, electronically. Static inverters are a solid state device with no moving parts, are quieter, lighter and more reliable and require less maintenance than rotary inverters because they do not have: 

Brushes, slip rings and commutators.



Bearings.



Rotating windings.

They provide a backup AC power supply in the event the main AC supply is lost on aircraft fitted with alternators. They only would supply essential AC in this case. Cooling holes must be kept clear. The primary reason for failures is overheating. When operating, a “whistle or whine” noise is normal. Inverters are usually operated in pairs so that in the event of one failing, the other will supply the AC bus automatically.

Static Inverter Basic Operation

Unless you were to become involved in the repair and maintenance of these inverters, you only need to know the basic principles of their operation, so we will refer to the block diagram Figure and observe these points: 

The filter prevents fluctuations and spikes on the 28 V DC supply.



A square wave generator (oscillator) produces a 400 Hz square wave output.



Pulse shaper, power driver and output stage shapes and transforms square wave into sinusoidal wave at the required voltage.



The output filter removes the odd harmonics ensuring a sine wave output.



A voltage regulator circuit ensures the correct voltage is maintained.


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Transformers Transformers allow the values of AC voltage and current to be changed through the use of mutual inductance. A typical transformer consists of two coils of wire wound around a common core, but not connected electrically. When an alternating current flows in the primary coil, a voltage is induced into the secondary coil. The amount of voltage generated in the secondary coil is equal to the voltage in the primary times the turns ratio between the two coils.

For example, 100 turns in a primary coil and 1,000 turns in a secondary equates to a turns ratio of 1:10. Therefore, if 115 volts flows across the primary, 1,150 volts are induced across the secondary. Since a transformer does not generate any power, the product of the voltage and the current in the secondary coil must be the same as that in the primary coil (less the transformer losses). Therefore, whenever volts are increased in a transformer, amperes must decrease by the same ratio. In other words, if the voltage is increased by a ratio of 1:10, the current must decrease by a ratio of 10:1. A transformer can have its primary coil connected directly across an AC power line and, as long as there is an open circuit in the secondary coil, the back voltage produced in the primary coil blocks the source voltage so almost no current flows through the primary winding. However, when the circuit is complete in the secondary coil, secondary current flows producing lines of flux that oppose the back voltage and allow source current to flow in the primary coil.

Some transformers may have multiple secondaries or tappings at various voltages (both higher and lower than supply voltage). A tap is nothing more than a wire connection made at some point on a winding between the very ends. Not surprisingly, the winding turn/voltage magnitude relationship of a normal transformer holds true for all tapped segments of windings. Issue B: January 2008

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Autotransformers An autotransformer is a form of variable transformer. In an autotransformer there is only one coil that acts as both primary and secondary. One lead of the coil is in common with both, while the other secondary lead connects to a movable brush that makes contact with the coil. The position of this brush determines the amount of secondary voltage.

Transformer Rectifier Principles It is often necessary to convert alternating current into direct current to power various circuits in the aircraft or within electronic equipment. The conversion of alternating to direct current is accomplished by a circuit referred to as a rectifier. Rectifier circuits typically employ diodes that allow current flow in only one direction.

In the illustrated example, if it were to be an aircraft TRU, it would transform 115/200 Volts AC to around 30 Volts AC and then would be rectified for a 28 Volt DC output. Typically, there is more than one TRU connected to a DC system thus providing redundancy.


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Transformer Rectifier Unit – TRU TRUs are a combination of static transformers and rectifiers used as a secondary or main conversion unit in aircraft having rectified AC power systems. Aircraft TRUs are designed to operate on a regulated 3 phase input of 200 VAC 400 Hz to provide a continuous DC output of 28 VDC (approx.). They consist of a transformer and two, three phase bridge rectifier assemblies. The transformer has a star wound primary winding and secondary winding is star and delta. Each secondary connected to individual bridge rectifier made up of 6 silicon diodes and connected in parallel.

A shunt connected to O/P enables reading of amps at aux terminals. Cooling of TRU’s is typically by natural convection. A thermal switch provides warning through a overheat (light) (set at 150 deg c to 200 deg c).

400 Hz 3 phase AC, when full wave rectified, has a DC voltage with a ripple frequency of 2400 Hz. A star to star transformer ’s input and output are in phase. A star to delta transformer’s input and output are 30 degrees out of phase. The outputs of both secondary outputs are then full wave rectified. Therefore, the parallel rectified outputs of the two secondary winding has a DC ripple frequency of 4,800 Hz. With such a high frequency DC ripple the DC voltage is almost linear and the capacitive effect of the rectifier diodes smooth the ripple to acceptable limits.


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Circuit Protection Protective devices are installed in electrical circuits to prevent damage caused by overloading a circuit or a short in a circuit. Overloading a circuit results from connecting loads that are too large for the wiring. A short, on the other hand, occurs when part of a circuit comes in direct contact with the return side of the circuit. When a short occurs, a path for current flow with little or no resistance is established. This results in large amounts of current flow and conductor heating. 2 main types of Circuit Protection Devices: •

Fuses

•

Circuit Breakers

Fuses One of the cheapest and most simple forms of protection for electronic circuitry is the fuse. Although the outer appearance of fuses varies widely in electronic and electrical appliances, the basic components of the physical construction are the same. These components are the Body or Case, Mounting and Fuse Link. The Body or Case can be made of glass, ceramic, fibre or other non-conducting material. The body not only holds the fuse together but its design and construction material have important functions. For example, a glass body allows easy identification of a blown fuse. Some ceramic fuses are designed for use in areas where a spark from a blowing fuse could be a hazard, especially in combustible environments. The Mounting forms the electrical connection to the circuit. There are several types of mounting including knife blade, cartridge or ferrule, and lead connection. The Fuse Link connects to the mounting and is the part of the fuse that ‘breaks’ the circuit. The fuse link is designed to disconnect circuit current once predefined limits are exceeded. The limits are determined by the size and composition of a fusing link. Once the current limit of the fusing link is exceeded, the heat generated by the current flowing through the fusing link will melt (or blow) the fusing link. There are various types of fuse link including bead, element filament and others.

In the cartridge fuse, the fuse link is enclosed in a tube of insulating material with metal ferrules at each end (for contact with the fuse holder). Some common insulating materials are glass, bakelite, or a fiber tube filled with insulating powder. In the slow blow fuse illustration shown above, this is not a spring but actually a wire wound resistor to limit current flow and introduce a time factor.


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Current Limiter The current limiter is another kind of fuse commonly found in aircraft power distribution systems and other high current DC power systems. It's a device that still behaves like a fuse but with much longer time constants. It's designed to be a protection from hard faults as opposed to continuous but small overloads. This device will take over 6 times longer to open than a fast blow fuse. This may seem like a really big difference but in a hard fault of this circuit, it's not uncommon to see currents of 1000A or more . . . even the current limiter will open in a few tenths of one millisecond . . . it will all be over before you know it!

There are some noteworthy characteristics of typical current limiters. Note that they will carry nearly 2X their rated current indefinitely. This makes them behave more like fusible links. The long time constant and overload capability of current limiters are not subject to the nuisance tripping we encountered occasionally with much faster fuses. Current limiters are used in large aircraft to protect long power distribution leads that carry huge starter-generator currents and even feeds from batteries. This level of protection is dictated in large aircraft.

Circuit Breakers Circuit breakers overcome the one major disadvantage a fuse has when protecting a circuit. That is, a fuse destroys itself whilst a circuit breaker ‘trips’ and can be reset once the fault is rectified. Another advantage the circuit breaker has is that it gives a visual indication that circuit parameters have been exceeded and the circuit is no longer operating. We will now look at the basic operation of the two types of circuit breaker. These are the magnetic circuit breaker and the thermal circuit breaker.


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Magnetic Circuit Breaker Referring to Figure, inside the sensing coil is a non-magnetic delay tube which houses a spring-biased moving magnetic core (core). An armature links the contacts to the coil mechanism, which functions as an electro-magnet. When the circuit breaker is set (button actuator pushed in, rocker actuator to the correct position etc.), current flows through the sensing coil and a magnetic field is created around the coil.

As current flow increases above the circuit breakers ‘rated’ current, the: •

Strength of the magnetic field increases.

•

Core moves towards the pole piece.

•

Magnetic circuit efficiency is increased and electro-magnetic force increases.

•

Core reaches fully ‘in’ position.

•

Armature is attracted to the pole piece.

•

Trip mechanism activates (and the circuit breaker is ‘tripped’ or ‘popped’).

•

Contacts are opened.


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Thermal Circuit Breaker A thermal circuit breaker uses a bi-metallic strip to operate. The basic internal construction of a thermal circuit breaker is shown in Figure. The thermal circuit breaker consists of three major assemblies: •

Bi-metal thermal element.

•

Contact type switch unit.

•

Mechanical latching mechanism.

In normal operation, current flows through the main contacts and thermal element to the load. The heat produced in the thermal element is radiated away quickly. When current exceeds the circuit breakers rating, (as in a short circuit) the temperature of the thermal element begins to build up. Due to the different rates of expansion of the two metals in the thermal element the following actions occur: •

The thermal element distorts.

•

The latch mechanism releases.

•

The main switch contacts open.

•

The push button extends.

•

The circuit breaker has now tripped.

•

No current is supplied to the load.


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Circuit Breaker Precautions We have seen the operation of the two types of circuit breakers, now we will look at the precautions associated with them. Firstly, if a circuit breaker has tripped, DO NOT immediately reset it. First, isolate the faulty circuit or component and repair it, then reset the circuit breaker. Resetting it immediately could result in further damage to the circuit or equipment being protected. Secondly, when replacing a damaged or unserviceable circuit breaker only use the circuit breaker stated in the applicable publication. Failure to adhere to this could damage the circuit or equipment the circuit breaker is protecting.

•

Small white band shows when tripped.

•

Aids pilot to see which C/B has popped.

•

Protects circuit wiring.

•

Can be reset – no need to carry spares.

•

Magnetic or Thermal (Bimetal) types.

•

Either trip free or non trip free.

•

Trip free will still trip even if trying to reset.

•

Thermal trip free must cool down.

•

Non trip free can be manually held closed.

•

Most common are trip free thermal.

Note the symbols used in schematics / wiring diagrams. All circuit breaker panels are clearly labelled to show circuits they protect and must clearly show their current rating in AMPS. The primary reason why circuit breakers and fuses are so close to the power supply (busbar) is to protect the wiring run from short circuits, which would cause an overload and possible fire. The secondary feature is to protect the component at the end of the wiring


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Reverse Current Cut-Out Relay A reverse current cut-out relay is used principally in a DC generating system either as a separate unit or as part of a voltage regulator. The circuit arrangement, as applied to tin generating system typical of several types of small aircraft, is shown in the figure. The relay consists of two coils wound on a core and a spring-controlled armature and contact assembly. The shunt winding is made up of many turns of fine wire connected across the generator so that voltage is impressed on all times. The series winding, of a few turns of heavy wire, is in series with the main supply line and is designed to carry the entire line current. The winding is also connected to the contact assembly, which under static conditions is held in the open position by means of a spring.

When the generator starts operating and the voltage builds up to a value which exceeds that of the battery, the shunt winding of the relay produces sufficient magnetism in the core to attract the armature and so close the contacts. Thus the relay acts as an automatic switch to connect the generator to the busbar, and also to the battery so that it is supplied with charging current. The field produced by the series winding aids the shunt-winding field in keeping the contacts firmly closed. When the generator is being shut down or, say, a failure in its output occurs, then the output falls below the battery voltage and there is a momentary discharge of current from the battery; in other words, a condition of reverse current through the cut-out relay series winding is set up. As this also causes a reversal of its magnetic field, the shunt winding-field will be opposed, thereby reducing core magnetization until the armature spring opens the contacts. The generator is therefore switched to the “off-line” condition to protect it from damaging effects which would otherwise result from “motoring” current discharging from the battery.


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Switched Reverse Current Relay This relay is adopted in the DC. generator systems of some types of small aircraft, its purpose being to permit switching of a generator on to the main bus bar, and at the same time retain the disconnect function in the event of reverse current arrangement is shown in the figure. In addition to a current coil the relay has a voltage coil, and a pair of contacts actuated via a contactor coil. When the voltage output is at a regulated value the current through the voltage coil is sufficient to actuate its contacts which then connect the generator switch and contactor coil to ground. The contactor coil is thus energized from the A+ output of the generator and so the auxiliary and main contacts close to connect the generator output to the battery and main busbar. The magnetic effect of passing through the current coil assists that of the voltage coil in keeping the pilot contacts closed.

During engine shut-down, the generator output voltage decreases thereby initiating a reverse current condition, and because the magnetic effect of the current through the current coil now oppose of the voltage coil, the pilot contacts open to de-energize the contactor coil; thus, the main and auxiliary contacts are opened to disconnect the generator from the battery and main busbar.


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Reverse Current Circuit Breakers These circuit breakers are designed to protect power supply systems and associated circuits against fault currents of a magnitude greater than those at which cut-outs normally operate. Furthermore, they are designed to remain in a “locked-out” condition to ensure complete isolation of a circuit until a fault has been cleared. An example of a circuit breaker designed for use in a DC generating system is shown in the figure. It consists of a magnetic unit, the field strength and direction of which are controlled by a single-turn coil connected between the generator positive output and the busbar via a main contact assembly. An auxiliary contact assembly is also provided for connection in series with the shunt-field winding of the generator.

The opening of both contact assemblies is controlled by a latching mechanism actuated by the magnet unit under heavy reverse current conditions. In common with other circuit breakers, resetting after a tripping operation has to be done manually, and is accomplished by a lever which is also actuated by the latching mechanism. Visual indication of a tripped condition is provided by a coloured indicator flag which appears behind a window in the circuit breaker cover. Manual tripping of the unit is effected by a push-button adjacent to the resetting lever. The figure is based on the circuit arrangement of a DC generating system used in a particular type of aircraft, and is an example of the application of a reverse current circuit breaker in conjunction with a cut-out relay. Unlike the reverse current cutout relay circuit, the relay controls the operation of a line contactor connected in series with the coil of the reverse current circuit breaker. Under normal current flow conditions closing of the relay energizes the line contactor, the heavy- duty contacts of which connect the generator output to the busbar via the coil and main contacts of the normally closed reverse current circuit breaker. The magnetic field set up by the current flow assists that of the magnet unit, thus maintaining the breaker contacts in the closed position. The generator shunt field circuit is supplied via the auxiliary contacts.


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When the generator is being shut down, or a failure of its output occurs, the reverse current resulting from the drop in output to a value below that of the battery flows through the circuit as indicated, and the cut-out relay is operated to de-energize the line contactor which takes the generator “off line”. Under these conditions the reverse current circuit breaker will remain closed, since the current magnitude is much lower than that at which a specific type of breaker is normally rated (some typical ranges are (generator) 200—250 A and 850—950 A). Let us consider now what would happen in the event of either the cut-out relay or the line contactor failing to open under the above low magnitude reverse current conditions, e.g. contacts have welded due to wear and excessive arcing. The reverse current would feed back to the generator, and in addition to its motoring effect on the generator, it would also reverse the generator field polarity. The reverse current passing through the circuit breaker coil would continue to increase in trying to overcome mechanical loads due to the engine and generator coupling, and so the increasing reverse field reduces the strength of the magnet unit. When the reverse current reaches the pie-set trip value of the circuit breaker, the field of the magnet unit is neutralized and repelled, causing the latch mechanism to release the main and auxiliary contacts to completely isolate the generator from the busbar. The breaker must be reset after the circuit fault has been cleared.


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Overvoltage Protection Overvoltage is a condition which could arise in a generating system in the event of a fault in the field excitation circuit, e.g. internal grounding of the field windings or an open-circuit in the voltage regulator sensing lines. Devices are therefore necessary to protect consumer equipment against voltages higher than those at which they are normally designed to operate. The methods adopted vary between aircraft systems and also on whether they supply DC or AC. An example of an overvoltage relay method applied to one type of DC system is shown in the figure.

The relay consists of a number of contacts connected in all essential circuits of the generator system, and mechanically coupled to a latching mechanism. This mechanism is electromagnetically controlled by a sensing coil and armature assembly, the coil being connected in the generator shunt-field circuit and in series with a resistor, the resistance of which decreases as the current through it is increased. Under normal regulated voltage conditions, the sensing coil circuit resistance is high enough to prevent generator shunt-field current from releasing the relay latch mechanism, and so the contacts remain closed and the generator remains connected to the busbar. If, however, an open circuit occurs in the regulator voltage coil sensing line, shunt-field current increases and, because of the inverse characteristics of the relay sensing coil resistor, the electromagnetic field by the coil causes the latch mechanism to release all the relay contacts to the open position, thereby isolating the system from the busbar. After the fault has been cleared, the contacts are reset by depressing the push button.


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The figure illustrates a method employed in a frequency-wild AC generating system, the full of which is provided by magnetic amplifiers. The output of the overvoltage protection magnetic amplifier is fed to a bridge rectifier and to the coil of a relay, via a feedback winding. The main contacts of the relay are connected in the normal DC supply switching circuit to the line contactor.

Under normal voltage output conditions the impedance of the magnetic amplifier is such that its AC output, and the rectified AC through the relay coil, maintain the relay in the deenergized condition. When an overvoltage condition is produced the current through the relay coil increases to a pre-determined energizing value, and the opening of the relay contacts interrupts the DC supply to the line contactor, which then disconnects the generator from the busbar. At the same time, the main control unit interrupts the supply of self-excitation current to the generator, causing its AC output to collapse to zero. The relay resets itself and after the fault has been cleared the generator output may be restored and connected to the busbar by carrying out the normal starting cycle. An overvoltage protection system adopted in one example of a constant frequency (nonparalleled) AC generating system is shown in basic form in the figure.


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The detector utilizes solid-state circuit elements which sense all three phases of the generator output. and is set to operate at a level greater than 130 ± 3 volts. An overvoltage condition is an excitation-type fault probably resulting from loss of sensing to, or control of, the voltage regulator such that excessive field excitation of a generator is provided. The signal resulting from an overvoltage is supplied through an inverse time delay to two solid-state switches. When switch s1 is made it completes a circuit through the coil of the generator control relay, one contact of which opens to interrupt the generator excitation field circuit. The other contact closes and completes a circuit to the generator breaker trip relay, this in turn, de-energizing the generator breaker to disconnect the generator from the busbar. The making of solid-state switch S2 energizes the light relay causing it to illuminate the annunciator light which is a white one in the actual system on which the figure is based. The purpose of the inverse time delay is to prevent nuisance tripping under transient conditions.


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Differential Current Protection The purpose of a differential current protection system is to detect a short-circuited feeder line or generator busbar which would result in a very high current demand on a generator, and possibly result in an electrical fire. Under these conditions, the difference between the current leaving the generator and the current arriving at the busbar is called a differential fault or a feeder fault. In an a.c. system current comparisons are made phase for phase, by two three-phase current transformers, one on the ground or neutral side of the generator (ground DPCT) and the other (the load DPCT) on the down stream side of the busbar. The figure illustrates the arrangement and principle of a system as applied to a single-phase line.

If the current from the generator is I, and the fault current between the generator and busbar equals If, then the net current at the busbar will be equal to I - If. The fault current will flow through the aircraft structure and back to the generator through the ground DPCT. The remainder of the current I - If, will flow through the load DPCT, the loads, the aircraft structure, and then back to the generator via the ground DPCT. Thus, the ground DPCT will detect the generator’s total current (I — If) + (If.) which is equal to I, and the load DPCT will detect I — If. If the difference in current (i.e. the fault current) between the two current transformers on the phase line is sensed to be greater than the specified limit (20 or 30 amperes are typical values) a protector circuit within a generator control unit will trip the generator control relay.


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Merz —Price Protection System This system is applied to some AC generating systems to provide protection against faults between phases or between one of the phases and ground. The connections for one phase are shown in the figure; those for other phases (or other feeders in a single-phase system) being exactly the same. Two similar current transformers are connected to the line, one at each end, and their secondary windings are connected together via two relay coils. Since the windings are in opposition, and as long as the currents at each end of the line are equal, the induced e.m.f’s are in balance and no current flows through the relay coils. When a fault occurs, the fault current creates an unbalanced condition causing current to flow through the coils of the relays thereby energizing them so as to open the line at each end.

Electrical Bus Bars In most aircraft types, the output from generating sources is coupled to one or more low impedance conductors known as bus bars. Bus bars are typically copper or aluminium bar to which input and output supply connections can be made. Known as a bus or buses. Some organised form of distribution is necessary for load sharing purposes, and for prioritisation under load shedding conditions. The more important the system – the more possible supplies it may have. Distribution is designed for the arrangement of busbars and distribution circuits so that they may be fed from different power sources. It is usual to categorise services into their order of importance and, in general, they fall into 3 groups: Vital, essential and non-essential. Vital Services Vital services are those which would be required after an emergency wheels-up landing, eg. Emergency lighting and crash switch operation of fire extinguishers. These services are connected directly to the battery. Essential Services Essential Services are those required to ensure safe flight in an in-flight emergency situation. They are connected to DC and AC busbars, as appropriate, and in such a way that they can always be supplied. Non-essential Services Non-essential Services are those which can be isolated in an in-flight emergency for load shedding purposes.


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Bus Tie Breaker - BTB The BTBs used in these types of circuits are interchangeable three phase contactors (same part No) they will however, perform different functions within the circuit. They may also be used as Generator Circuit Breakers (GCB) controlled by the GCU, or they may be known as Bus Tie Contactors.

These contactors contain three heavy duty contacts which connect three phase power to the generator busbars where it is distributed to the aircraft AC loads. They also contain a number of auxiliary contacts that are used to connect various control and indication circuits within the AC distribution system. Power supplied to the close coil causes the main contacts to close and the auxiliary contacts change, so the normally open contacts close and become mechanically locked in this position. Power can now be removed from the close coil and the breaker remains in the closed position. Power applied to the trip coil releases the mechanical lock and spring-tension forces the main contacts to open and the auxiliary contacts to return to the normally closed position.

Generator Circuit Breaker – GCB The main functions of the GCB in a split bus system are: • To connect the AC generator’s output to its respective bus if the generator is serviceable. • To ensure that when the generator’s output is connected to the generator bus the bus tie breaker is tripped therefore, ensuring that the AC loads receive power from only power source at any one time. • To isolate the generator’s output from the aircraft’s loads if a fault is detected by the generator control unit.


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Basic DC Aircraft This is a simple representation of a small aircraft with one generator and how the power is distributed. The battery would be charged from the bus when generator was outputting at a higher voltage. The inverter provides the aircraft with AC voltage.

Tied DC Bus System Buses are separated from each other, particularly with respect to order of importance for flight safety and in relation to redundancy following power malfunctions. Each generator feeds its own bus, and the tied breaker would close if only one generator was available such as a situation of a generator failure or engine failure.


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Basic Twin Engine AC System In this example the generators (alternators) supply their individual buses. The AC bus then feeds the TRU which provides the DC requirements of the aircraft as well as charge the battery. The emergency AC bus is fed from the AC bus directly. Should this AC supply fail for any reason the static inverter would be commanded to start and provide the Emergency AC from the battery.


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Split System Breaker If the split system breaker (SSB) will not close, it would be caused by a situation where the buses’ properties would be outside limits of voltage, phase rotation sequence or frequency. A tied bus system means the AC generators outputs are connected together onto the same bus. Before the generators can be connected together or paralleled they must be matched so: • Their output wave forms are the same. • They must have the same phase rotation. • Their output voltages must be the same. • They must have the same frequency. • Their voltages are in phase.

The phase rotation will be correct if the generators are all connected to the lines in the same way. If this is not correct, a correct synchronisation indication will not be achievable. Their output voltages will be within acceptable limits if the voltage regulators are set correctly.


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This is a schematic representation of an aircraft AC/DC power distribution system.


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Split Bus Distribution System

When operated on the ground with external power connected & no engines running, all loads will be supplied by the external power source via the external power relay & the closed contacts of the BTB (note both GCBs will be tripped & the changeover relay will be energised closed). When one of the aircraft’s engines is started & its generator is operating within system limits its GCB will energise close. This will cause the external power relay to open removing external power from the system loads. The BTB however, will remain closed, connecting the three phase power from the operational generator to all of the aircraft’s loads. When the second aircraft engine is started & its generator is operating within system limits its GCB will close. As the power to the BTB’s close coil is now redirected through the auxiliary contacts of the GCB to the BTB’s trip coil the BTB will trip causing the operational generators to supply their individual busbars. Power from the generator busbars is then distributed to the AC non-essential loads & TRUs. The essential AC loads are supplied from the AC essential bus which under normal operating conditions (both AC generators operating within system tolerance) is connected via a changeover relay to the No. 1 generator bus. If one of the generators should become inoperative (fail or be switched OFF), its GCB will trip. Power is now supplied via the auxiliary contacts of the tripped GCB to the close coil of the BTB causing the BTB to close. When closed the BTB connects both generator buses together thereby enabling the output from the serviceable generator to supply power to all system loads. If, for any reason, both generators should become inoperative (fail or be switched OFF) then both GCBs will trip. Battery power will energise the BTB closed however as both generators are off line no power will be available to operate the non-essential loads. In this situation the changeover relay between No. 1 generator bus & the essential AC bus will automatically deenergise & connect the essential bus to the static inverter. The system’s essential AC loads will be supplied with 115 VAC from the static inverter while the non-essential AC loads will be effectively isolated. As we saw in DC systems if the system operates for a long period of time under these conditions the battery will become discharged. Therefore, before this occurs the circuit’s battery relay will be turned “OFF” isolating the aircraft’s AC & DC essential loads from power while allowing the aircraft to operate under emergency DC power.


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External / Ground Power Electrical power is required for the starting of engines, operation of certain services during ‘turn around’ servicings, eg. Lighting, and for testing electrical systems during routine maintenance.

DC Power Receptacle / Plug Application of external DC to aircraft is via this type of plug (nato type). It has 2 large pins for the DC supply and earth return, and a smaller pin used for interlock sense and polarity check. It also ensures plug will only fit one way in receptacle. The reason why it is shorter than the main pins is that in the event of the plug being withdrawn with the circuit “live”, the external power relay will de-energise before the main pins are disengaged from the socket. This ensures that breaking of the supply takes place at the contactor, thus preventing arcing at the pins of the external power plug/connector.


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DC External Power Circuit As can be seen from the diagram, the short pin is connected to the positive pin of the plug. This connection is internal in the plug. Should the Ground Power Cart lead/plug be accidentally wired incorrectly, the external power relay will not energise, due to the reverse polarity diode, and thus, incorrect polarity will not be supplied to the aircraft.

This is a schematic representation of a typical aircraft DC external power system.


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AC Power Receptacle / Plug 4 Large pins. These are one for each phase and one for the earth. The 2 small pins are the interlock sense which must be seen correct before the AC can be connected to modern aircraft. This also means as if the plug is accidentally pulled or someone trips over the lead dislodging it, the circuit between (E) and (F) is interrupted cutting power. This is done so there can be no arcing of the big current carrying pins and sockets.

Auxiliary Contacts – Bus Tie Breaker and Generator Circuit Breaker Interlocking circuitry through auxiliary contacts of relative contactors and breakers is used to prevent unsynchronised power sources being connected in parallel. Refer back to the breaker illustrated in Figure and notice the auxiliary contacts which have both normally closed and normally open contacts. The bottom set of contacts will trip all the other generator breakers so ground-power cannot be connected to any generator. Switching ON the on-coming source automatically switches OFF the present source. For example, ground power is supplying the loads and the APU generator is switched ON. The APU breaker closes and its auxiliary contacts de-energise the ground power contactor, disconnecting ground power from bus system. This is a schematic representation of a typical aircraft AC external power system.


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ELECTRICAL SAFETY PRECAUTIONS Degrees of Electric Shock The human body depends on two body functions to maintain life: •

Breathing or respiration.

•

Heartbeat or circulation.

To maintain these two functions the brain, via the body’s nervous system, sends tiny electrical pulses to the muscles that control breathing and circulation. Should there be a failure of these two functions, then life will be in danger. When the body is exposed to an electric shock, the shock will cause current to flow within the body that can interfere with the electrical impulses that control the respiration and circulation functions. The degree of interference is dependent upon:

•

Where the current flows within the body.

•

The amount of current flowing.

•

The duration of the electric shock.


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Reaction of Human Body to Electric Shock The human body is not designed to withstand a continuous application of electricity. Muscular contraction caused by an electric shock can lead to serious injury or death.

Current (milliamperes) Effect 

up to 2 mA

limit of perception



2 mA to 8 mA

sensation becomes obvious and more painful.



8 mA to 12 mA

muscle spasms and pain



12 mA to 20 mA

unable to control muscles, it becomes impossible to let go of a “live” conductor.



20 mA to 50 mA

if current flow is across the chest, interference with the heart and lung muscles may result in breathing stopping.



50 mA to 100 mA if current flow is across the chest, heart commences ventricular fibrillation.



100 mA to 200 mA heart stops beating



above 200 mA

severe burns.

Immediate Action If you find someone who is in contact with live electrical equipment/wiring you must isolate the victim from the current source before trying to give first aid. WARNING Do NOT touch the victim until the power source is removed. WARNING Do NOT cut the cable. The action to be followed is: 1.

Isolate the victim from the power by either:

Turning the power OFF at the isolation switch (preferred method), OR If this is not possible, use of an insulating material to remove the victim from the source is essential. An insulating material is anything which will not conduct electricity. (Insulating materials could be dry wood, blankets and hoses etc.) THEN 2.


Give first aid to victim.

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Precautions 

Only qualified personnel are to perform electrical maintenance.



Maintenance to be carried out in accordance with applicable maintenance manual.



When disconnecting/connecting any electrical equipment, ensure power is OFF.



Ensure ALL power is isolated before working inside panels.



Before any maintenance on electrical equipment is carried out, the equipment must be isolated from electrical power. WARNING

Electrical equipment is to be isolated from power prior to maintenance. 

Attach safety notices (tags) to switches and circuit breakers as required



Before commencing work on any circuit or component in aircraft: o

Isolate power to circuit – pull circuit breaker / fuse, and

o

Tag relevant circuit breaker / fuse

Electrical circuit isolation When you are working on and around aircraft, there are many people performing all sorts of maintenance, from cleaning the aircraft to fuel replenishment and fault rectification. During these maintenance checks and servicing there are many dangers that could happen to an unwary person. The electrical power available on and around aircraft ranges from 12 volts DC (Direct Current) to 415 volts AC (Alternating Current). When working on aircraft it is often necessary to isolate equipment from the power supply to enable work to be carried out. This can be done by:  Switching off the circuit breaker or master switch.  Removing the fuse. This in itself is not enough to protect you from some other person turning on the power while you are still working on the equipment. You must also put signs on the switches to alert others that work is being done on the system. This should make you the only person who has authority to turn these circuits back on.


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Warning These signs do not prevent others from turning the switch back on or replacing the fuse or resetting the circuit breaker. The best way to ensure that no electrical power is applied, is to isolate the aircraft power supplies totally, if this is possible. Remove all rings, bracelets and similar metal items.

Never assume that a circuit is OFF. Do not tamper with safety devices. Verify that all safety interlocks operate correctly. Keep all tools and equipment in good working order. Use the correct tool for the job. Note: Incident One incident with an airframe mechanic who, while working on some totally unrelated equipment, accidentally contacted a live 28 V terminal with his gold wedding ring. Before the power could be turned off his ring welded itself to the terminal and glowed red hot, burning his finger to the bone. It was later amputated at the knuckle. Remember! The smell of burning flesh is not at all pleasant, especially when it is your own! 

Verify that capacitors have been discharged.



Do not remove equipment earth connections.



Do not use adaptors that defeat earth connections.



Do not work when you are tired or taking medication that can cause drowsiness.



Avoid working in poor light.



Avoid working in damp areas.


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Isolation Safety Equipment Mains powered equipment (i.e. 240 V or 415 V 50 Hz) is not to be used on an aircraft unless it is operated through an Residual Current Device (RCD) or Isolation Transformer. An RCD is sometimes referred to as an “Earth Leakage Circuit Breaker (ELCB)”. Ensure RCD is installed as close as possible to power outlet (ie. Hangar wall). This way it protects the entire circuit, extension lead and appliance, not just appliance. Most RCDs operate with a discrepancy of 30mA. There are also 10mA RCDs, but these are usually used in hospitals, etc.

RCD A RCD (Residual Current Device) is a device which detects an imbalance between the current 'leaving' and 'returning' from and to the circuit. If more current leaves than returns then some is leaking away (to 'earth') due to a fault.

Electrical Power Leads The following procedures are to be observed when using electrical extension leads: 

If at all possible, a Residual Current Detector (RCD) is to be used (sometimes called an Earth Leakage Circuit Breaker). Do not use when working in damp conditions or outside in inclement weather.



Extension leads must be fully unwound before use.



When routed through doors, the lead is to be protected to prevent the door being closed and damaging the cable.



If leads are to be run across walkways or driveways, etc, then feed-through covers or ducts are to be used for protection.


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Battery Maintenance Safety Precautions 

Never short out the terminals of a battery.



Do NOT charge batteries while still fitted to aircraft (on ground) as this could create an explosive situation.



Batteries release hydrogen and oxygen during charging – Hydrogen is highly explosive. Ensure NO sources of ignition come anywhere in the vicinity of charging batteries.



Acids and alkalines neutralise each other. Acid is very harmful to ni-cad batteries. Use separate equipment (ALL) and maintenance facilities for lead-acid and alkaline battery maintenance. Do not store close to each other.



A lead acid battery under charge will release large quantities of hydrogen and oxygen gasses as it reaches the fully charged state. Under no circumstances must the charger be connected or disconnected with the power turned on. The resulting spark will ignite the gas and could cause a massive explosion. A 4% build up of hydrogen in the air will explode.



When using sulphuric acid or potassium hydroxide to mix electrolytes, always slowly add to water. Do not do in reverse order. This causes splattering.



Ensure access to neutralising agents in clearly marked containers.



Only charge in a well ventilated facility.



Only ever add water to a battery in service – Never acid or alkaline



Be familiar with batteries and electrolytes to cope with them safely



Ensure the correct PPE is worn when performing maintenance on batteries: o

Safety full face shield.

o

Acid proof gloves.

o

Acid proof apron.



Acid and alkaline electrolytes are dangerous in eyes or on skin. Flush with water and seek medical attention for eyes and, if necessary, for skin.



Beware of jewellery – watches, rings, bangles. Can short terminals.

REMEMBER: SAFETY FIRST


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TOPIC 13.6– EQUIPMENT AND FURNISHINGS Purpose of Equipment Emergency locator transmitters help rescue crews find airplanes which land away from an airport. The transmitters send a radio signal to satellites, other airplanes, and air traffic control facilities. Rescue crews use information from these sources to find the airplane. Frequency: 

121.5 MHz. Civil aircraft distress frequency.



243 MHz. Military aircraft distress frequency.



406.025 MHz. Satellite frequency.

Modulation: 

Data biphase L encoding, short message, Serialized Aviation User Protocol.



121.5/243 – 100% AM tone, swept downwards 1500 to 500 Hz @ 3 sweeps/second

RF Output Power: 

406.025 - 5 Watts.



121.5/243 MHz – 250mW

Battery Endurance: 

50 hours minimum.

Physical: 

Length: 68cm



Width: 8.5cm


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In-Flight Entertainment System

Uses: 

In-flight movies



Video Games



Duty-free shopping



Customer Surveys

PES Function Groups Head-End Digitises audio and modulates onto an RF carrier. Audio and Video RF signals are mixed prior to distribution throughout the aircraft. Area Distribution Intelligent data routing to and from the seat columns, RF splitting with level adjustment and column power control. Seat Group Passenger entertainment selection, channel tuning, audio/video presentation and gives passengers access to the service functions


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Head-End Equipment Primary Access Terminal Gives cabin crew full control over entertainment system video player control, processing of credit cards, print receipts, preview movies. Maintenance crew can run full diagnostics of the system components and print failure reports.



80486 based PC running Windows



10 inch touch screen LCD



500MB hard disk



Magnetic credit card reader

Cabin File Server 80486 DX 4 / 100 MHz based PC with 500MB hard drive, automatic operation - no user action required.


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Video Players “Teac” triple deck VEP, VCR, VCP, VTR, VTP… 8mm video

Video Modulator Modulates up to 8 video signals (PAL or NTSC) and combines them onto a single RF cable 

Two video modulators



can be cascaded to give



16 video channels

Later models as used in TES process 24 channels Passenger Entertainment System Controllers Adds audio signals to RF signal generated by video modulator PESC-V (video): 

adds soundtrack to films etc

PESC-A (audio): 

adds “radio” stations and other pre-recorded messages.


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Area Distribution Equipment Area Distribution Box 

Distributes RF video and audio throughout cabin

ADB is also responsible for: 

Control of 115V power to seat columns



Routing of serial communications between seats and controllers


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Seat Group Equipment Seat Electronics Box (Audio Video Unit) 

AVU



Location - Throughout the aircraft under the Seats



Multiple



Each AVU controls the channel selection and volume for 2 or 3 seats


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Passenger Control Unit 

UPCU



1 per seat



Allows the passenger to control the system. Includes channel selection, volume control, game control, telephone, reading light on/off, and Flight Attendant call button


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System Block Diagram


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TOPIC 13.9 – LIGHTS Exterior Lighting

The above illustration shows most of the exterior lighting used on a modern commercial aircraft. The main lighting to be identified is;

1. Navigation 2. Anti Collision 3. Landing and Taxi 4. Ice detection and 5. Strobe lights


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Navigation Lights May also be known as position lights on some aircraft. Positioned on each wing tip and the tail of an aircraft (some aircraft have a dual light system where a white tail light is located at the end of each wing tip). Coloured, so other aircraft can determine the direction of travel of the aircraft. The tail light is clear/white, the port light is red and the starboard light is green. Aviation authority regulations require the lights are visible at certain angles. These angles are shown below. The main fault in lighting systems is bulb failure. When checking light systems always check that all bulbs in a multi-bulb light are working. Always replace a bulb with one that has the part number listed in the aircraft maintenance manual. When checking lighting systems with an intensity control, check the system through the full intensity range.

Navigation Light Operation May be selected to: •

Steady or Flash (not common on many of today’s aircraft)

•

Bright or Dim (also not a common feature on modern aircraft).

Most commercial aircraft now have a redundancy system, where two lamps are placed in parallel. If one lamp fails the other can be used until the aircraft arrives at a maintenance facility.


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Cautions Do not handle bulbs with bare hands, oil from skin will cause local heat spots and bulbs may fail prematurely. These bulbs operate at very high temperatures.

Anti-Collision Beacons These lights are flashing red beacons producing between 40 and 100 flashes per minute. They should be visible from all directions within an angle of 30 degrees above and below the horizontal plane of the aircraft. An aircraft can have one unit on top of the vertical stabiliser or one on top of the fuselage and one on the bottom of the fuselage. There are three types of anti-collision beacons which are described below.

Rotary Reflector Type Light aircraft use a rotating reflector type which has a small motor and gear train driving a reflector around a bulb. The light has a red lens. Larger aircraft use a sealed beam shining onto a double reflector rotated by a small motor and gear train above the sealed beam.

Rotating Bulb Type Two bulbs with a reflective surface on approximately half the bulb are positioned with the reflective surfaces back to back and are rotated by a small motor and gear train.

Strobe Type The strobe light as an anti-collision light utilises a red lens. It operates in the same principle as any other strobe light, where by a high intensity light is generated from a power source located independent of the strobe unit. This power source supplies high voltage to the strobe unit at the required flash rate.

White Collision Avoidance Strobe Lights Most aircraft have two of these lights, usually mounted in each wing tip but there can be one underneath and one on top. Larger aircraft have a third one mounted in the tailplane. A few light aircraft have only one mounted on top of the vertical stabiliser. The strobe lights give an extremely intense flash of white light for a very short period at a frequency of 50 to 80 times a minute. The flash should be visual from any point around the aircraft in an area 30 degrees above and below the aircraft horizon. They denote that the aircraft is in flight. Due to their intensity they are switched on just before takeoff and switched off after exiting the runway on landing.


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Strobe Light Flash Tube The flash tube is filled with argon gas and has a cathode, anode and an external trigger electrode. The clear tube can be made from glass or quartz. Glass tubes should not be touched by hand, so care should be taken when changing the tube. If the tube is touched it should be cleaned with a soft dry cloth. Power packs can be manufactured to supply one, two or three lights. If a multi-light installation is installed with single power packs, the packs have to be interconnected so the lights are synchronised to flash at the same time. Power packs are located independent of the strobe unit, usually in a water proof section of the aircraft.

Precautions The output power from the power pack can be lethal. Always ensure that input power to the power pack is disconnected for five minutes before any maintenance is performed. Unless otherwise stated in Manual. The white flashes produced by a strobe light can cause eye damage if looked at from a close position. •

Do not look directly at an operating strobe.

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Do not operate strobes during refueling operations.

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Do not touch the tube assembly with bare hands.

Strobe Lights Remember the precautions detailed under this heading above. CASA has reported on five on-the-ground explosions and a fatal crash involving strobe lights and tip tanks. It is suspected that because of the high voltage, bad connections (including the airframe return path) can cause arcs to occur and ignite fuel vapour leading to an explosion and/or fire. Care must be taken to ensure that connections are clean and secure and that there is return path bonding so arcs will not occur. Care must also be taken to ensure there are no fuel leaks or vapour sources in the area of these lights.


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Landing Lights These lights are normally cooled by ram air so beware not to overheat on ground. These lights generate a lot of heat so testing on ground must be carried out quickly. The type of bulb used in landing lights can be a pre-focused bulb placed in a parabolic reflector or a sealed beam. Most aircraft use sealed beams. When replacing lamps, tighten the retaining device evenly so the glass sealed beam will not be placed under stress, which may cause it to crack. A crack in a sealed beam allows air to enter the bulb and the filament will burn out.

Quartz Halogen Bulb These bulbs are used in some aircraft taxi lights, landing lights and floodlights. Their envelopes are made of quartz which withstands much higher temperatures than glass. The hotter filament temperature produces more light for the same electrical power and these bulbs have a longer life than the ordinary incandescent bulb. Care must be taken not to touch the envelope of this type of bulb. If it is handled it must be cleaned with lead-free spirit. Two types of landing lights exist. Fixed and retractable.

Fixed Landing Lights Fixed landing lights can be set in the nose or leading edge of the main plane behind a Perspex cover. Their beam is fixed and can only be adjusted by removing the Perspex cover and adjusting the beam guiding screws.

Retractable Landing Lights These are mounted in the wing, flaps or under the fuselage. When retracted they will be flush with the surface of the aircraft. They can be extended out to one or more positions and are driven by a small split field rotary actuator and a gear train.

Extend to One Position Only With this type when the control switch is selected ON, the landing light extends to a preset position where a cam depresses a micro-switch and stops the actuator motor.

Extend to Two or More Pre-Set Positions This system has more than one position on the selection switch, for example, APPROACH and LAND. For each selected switch position, there is a micro-switch which the cam will strike, stopping the light in that position. Some aircraft use these types of landing lights as taxi lights and on selection of taxi the lamp is stopped in a position suitable for taxiing. This system is also used when the landing light is located in one of the flap fairings. At each of the flap settings, a micro switch and cam arrangement is used to extend or retract the landing light to ensure that it remains parallel with the aircrafts path. This occurs automatically when the landing light switch is selected to on and the flaps are cycled through their desired range.


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Manual Selection to any Position Operated by a switch spring-loaded to centre in the OFF position from the EXTEND and RETRACT positions, this system allows the crew to stop the light at any angle. Microswitches limit full extension and retraction to prevent damage to the motor and gear train.


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Taxi Lights Most taxi lights are fixed and can be mounted alongside fixed landing lights, but set on a different angle. Many aircraft have them attached to the nose landing gear. By attaching it to the nose landing gear of the aircraft, the beam will follow the intended direction of the aircraft. The electrical circuit is the same as for fixed landing lights. Basic ON-OFF switch control. Some aircraft operate through a relay in the landing gear limiting use to periods when the landing gear is extended.

Retractable landing light and taxi lamp circuit for Cessna 310

Ice Detection or Leading Edge Lights Mounted on the side of the fuselage under a Perspex cover, these lights shine a beam down the wing leading edge. This permits observation of the wing to visually detect ice formation on the wing. Usually a sealed beam lamp is used. On some aircraft this type of ice detection is used in conjunction with a black strip of paint, located on the wing leading edge. Basic ON-OFF switch control.


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Interior Lighting Cabin Lighting Cabin lights are used to illuminate the main cabin of the aircraft. They are normally located on the ceiling of the aircraft or near the overhead lockers of the aircraft. Most modern aircraft have fluorescent lamps which use indirect lighting techniques to disperse the light around the cabin. The controls are located at the flight attendant station of the aircraft. In larger commercial aircraft, several control units maybe used to control different sections of the aircraft. Most modern commercial aircraft have the cabin lighting operating off the ground service bus, as this allows for the operation of the aircrafts cabin lighting without the need for the aircraft AC power to be fully activated. Lamp replacement should be done in accordance with the aircrafts manual. Never use lamps or ballast units not stipulated in the aircrafts manuals. Most military aircraft, fitted with cabin lighting, use incandescent lamps in dome lights. These are dispersed around the cabin to provide lighting. Some aircraft (cargo transport) have multi coloured lamps for the purpose of night operations. These lamps are usually red or another colour suitable for night vision equipment.


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Cabin Window Lighting This type of cabin lighting is used on most commercial aircraft to improve lighting of the seating area. They are usually fitted in parallel along the side of the interior of the aircraft. Fluorescent tubes are commonly used and most aircraft use a ballast unit instead of a fluorescent starter. They are controlled from the flight attendant station, either at the front or the rear of the aircraft. In some larger aircraft the window lighting maybe divided into sections controlled from separate locations within the aircraft. The window lighting system is run independently of other cabin lighting in most aircraft.

Passenger Advisory Signs These signs indicate FASTEN SEAT BELTS, NO SMOKING, RETURN TO SEAT and TOILET OCCUPIED-VACANT. They are illuminated by electroluminescent lighting or small incandescent globes which are placed behind a transparent cover coated with an opaque covering. The lettering is cut out of the opaque cover and the light shows the sign.


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How the system operates Advisory signs, except for the toilet, are controlled by the flight crew with switches in the cockpit and are simple circuits. However some aircraft also have automatic control of these signs which will illuminate them in certain situations, for example: • When the undercarriage lever DOWN position is selected • If there is any extension of the flap • If a decompression occurs. When any of the signs except the toilet are illuminated, an audio tone sounds over the PA system.


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Lavatory Lighting The Toilet Occupied-Vacant signs are simply a circuit-breaker, switch and the lights for each of the individual toilets. The circuit also uses a switch located in the door lock. When the passenger locks the door a hammer or plunger strikes against the micro switch in the door surround, this closes the circuit and illuminates the occupied sign. When the door is unlocked the micro switch is released and the sign extinguishes.


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Passenger Reading Lights These lights provide a restricted light in the reading area of a passenger and will not annoy other passengers. Usually have a master switch located on the Flight Attendant panel. Each light is individually switched and the circuit is simply powered from a circuit-breaker to individual switches and lights. Some aircraft reading lights are controlled by electronic units located in or near the seats.


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Cockpit Lighting

Cockpit lighting has to be sufficient to illuminate all controls, switches, instruments and panel inscriptions with uniform intensity, but not shine at, or disturb the crew in any way. Flood or background Cockpit lighting in modern day civil aircraft is usually white with some older civil aircraft have red lighting fitted. Military aircraft use red lighting or NVG green which cannot be seen outside the aircraft. There are 5 main types of cockpit lighting •

Integral lighting

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Post or Pillar Lights

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Trans-illuminated panels

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Floodlights

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Utility or Wander lights


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Integral Lighting This form of lighting is used to illuminate cockpit instruments. In older aircraft it is used for most instrumentation illumination. In modern aircraft commonly used for ECAM or EICAS light intensity control and for the standby instrumentation. In multi-pilot aircraft, both pilots are able to control their own instrument lighting, independent of the other. The pilot in command will also control the centre instrument panel (usually engine and standby instruments). This form of lighting comes from globes inside the instrument. One form has several smaller globes, usually six, positioned around the dial in a plastic ring. This diffuses the light which shines on the dial. In some instruments six red and six white globes are fitted. The operator can select which colour will be used.

Post or Pillar Lights These lights, as illustrated, are used in areas where there is insufficient light from other sources. They are screwed to the panel and have a hood which directs the light down and across an arc of 90 degrees. Usually the hood can be turned through 360 degrees so it can be adjusted to ensure the light illuminates the correct area.


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Trans-Illuminated Panels Trans-illuminated panels, or indirect lighting, are made from transparent plastic or Perspex which is painted with a translucent white, then an opaque black or grey paint. The opaque material is engraved with inscriptions such as PILOT’S PANEL LIGHTS around switches and other controls. Light shines through the engraved inscriptions. This method of cockpit lighting is most commonly used to display system controls and switching (e.g. hydraulic panel, electrical panel etc.). Operates through a dimmer unit, controlled by switches in the cockpit. Have students explain reason for this type of lighting. Minimum number of lamps can illuminate large areas of paneling.


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Floodlights Floodlights can be mounted to shine on overhead, side and centre consoles, floors, circuitbreaker panels and main instrument panels. They are usually incandescent globes with a dimming facility, but can be fluorescent tubes. Usually controlled through a switch and a dimmer control unit.


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Utility or Wander Lights These lights can be used as floodlights, map reading lights and emergency lights to replace a failed light system and are often powered from a hot battery bus so they can be used as emergency lighting in the event of a total power failure. They have a flexible chord and are mounted in a ball clip socket. They can be removed by hand and placed in other ball clip sockets around the cockpit or hand held. They can also be pivoted up and down and left to right. The rheostat controls the switching and intensity of the light. The bezel on some of these lights is used to focus the beam. Some have two lenses – one white and the other red.

Warning and Caution Lighting These cockpits lights are used to alert the crew to system failures and/or abnormalities. Two main colours are used. Amber – Used to indicate a system abnormality or unsafe condition. It does not necessarily mean a dangerous condition. Red – These warnings are designed to alert the crew to an unsafe condition which may affect the safe flying of the aircraft. These warnings require immediate crew action.

Cargo Compartment Lights These lights are fitted in cargo compartments. They can be fluorescent or incandescent and are simple light systems. Most commercial aircraft cargo lighting systems operate through a micro-switch located in the door of the cargo compartment, so once the door is closed, regardless of the lights being selected on or off, they are switched off upon closing of the door. Once the door is cargo door is opened, they can usually be operated through normal switching controls, located inside the cargo compartment or near the cargo door.

Loading Lights These lights shine out from the aircraft to illuminate areas where cargo and baggage loading takes place and are simple light systems. The controls for these lights are usually located on the outside of the aircraft, and in some aircraft, are controlled by a micro-switch in the aircrafts landing gear, which turns on the lights when the weight of the aircraft is placed on the wheels.


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Emergency Lighting These lights have to mark all exits and illuminate the cabin with enough light for the crew to operate emergency equipment and guide passengers to safety, without outside lighting. The intensity and position requirements of these lights are detailed in relevant civil aviation regulations.

Emergency Exit Lights Fitted over or near each emergency exit, these lights indicate EXIT and have a window at the bottom using white light to illuminate the exit, therefore acting as emergency lighting at the same time. In older aircraft they were powered from hot battery buses and could be switched manually by the crew or automatically by a crash or inertia switch, or when the aircrafts power supplies are deactivated. Modern aircraft usually have an independent power source, a nickel cadmium battery. If aircraft power fails the battery is automatically switched to power the light. Under normal conditions the battery is charged from the aircraft power system, via the aircraft’s DC power system. Each emergency exit point has its own power source.


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Portable Emergency Lights As well as the fixed lights over the exits there are portable emergency exit lights near the exits which have an internal battery pack. The light can be detached and used as a torch and will turn on if the aircraft power fails or its switch is placed to the ON position. Emergency lights with battery packs that turn on if the aircraft power fails, will turn on if they are not disarmed before power is taken off the aircraft. This will discharge batteries and render the aircraft unserviceable until the batteries are replaced with charged batteries. Should power be removed from an aircraft before the emergency lights are disarmed, power has to be reconnected and the lights disarmed to correct the situation.


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TOPIC 13.12: FIRE PROTECTION (ATA-26) The Fire Triangle In order to understand how fire detection and extinguisher systems work, you first need to know a little bit about fire. Three (3) things must be present at the same time in order to produce fire: •

Enough oxygen to sustain combustion,

•

Enough heat to raise the combustible material to its ignition temperature, and

•

Some sort of fuel or combustible material.

The important thing to remember is: take any of these elements away, and you will not have a fire or the fire will be extinguished. Essentially, fire extinguishers put out fire by taking away one or more elements of the fire triangle. Fire safety, at its most basic, is based upon the principle of keeping fuel sources and ignition sources separate.

Aircraft Fire Protection

An aircraft carries large amounts of fuel and has several sources of ignition – electrical devices, hot engines and hot exhaust gases. Fire in flight is extremely dangerous. The most likely place an aircraft fire will occur is in an engine. If the fire is not detected early and extinguished, the consequences could be disastrous.


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Fire/Overheat Detection Systems Detection systems must be capable of providing rapid detection of a localised fire or overheat condition and indication of the area in which some corrective action is required. Detectors must not automatically operate fire extinguishing units, although they maybe used to shut down power or fuel to certain areas or components. Fire detection systems must: •

Be reliable.

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Give immediate indication of fire / overheat situation and also when that situation is extinguished.

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If the system fails – it is more likely that system will be inoperative than to cause erroneous readings.

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That the system can operate independently of other systems.

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That the system is able to operate when main power systems are no longer operational. Most systems operate of the aircrafts 28VDC or the 28VDC Hot battery bus.

Thermal switch system The thermal switch fire detector illustrated, uses a bimetal strip to close a set of contacts at a preset temperature. It operates off the 28VDC battery bus, which allows for operation during periods when the aircraft’s main AC power is not connected.

As the temperature increases outside of the thermal switch sensor, the bimetal strip expands. When the outside temperature reaches a preset level, the contact points in the bimetal strip will make contact. This completes the circuit and gives a warning to the flight crew of a fire or overheat situation. Some types are made so that their operating temperature can be adjusted. This is done by adjusting the position of the sliding piston. This increases or decreases the distance between the contact points of the bimetal strip. This change in distance is directly proportional to the heat required to bring the contact points together. Also known as Thermostat switch or spot detector.


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Thermal switch system - How the system operates.

All the switches are bolted to the aircraft frame which connects one of the contacts of the switch to earth. The illustrated circuit shows that the other contact of each of the switches is connected in parallel to the coil of relay RL1. Should any switch close, a current path is made for the coil of RL1. When energised the contacts of RL1 complete a current path for the bell and red light. When the crew has been alerted, they can stop the bell from ringing by pressing the BELL CANCEL button. This action energises RL2, breaking the bell’s current path and completing its own coil’s current path. The ‘bell cancel’ button can be released, but RL2 will remain energised until the fire warning disappears. This system normally operates of the 28VDC battery bus.

Testing The test button completes an earth connection to R1 via all the wiring connecting the thermal switches. This checks the wiring for continuity. The bell should ring and the red light illuminate. If the bell cancel button is pressed, the bell should stop, but the red light remains illuminated. Individual switches can be tested for correct operation (continuity) by using an Ohmmeter.

Warning Always install a thermal switch of the correct part number as detailed in the aircraft manual. Installation of a switch of a different operating temperature could cause a false warning or no warning in the event of a fire.


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Thermocouple Detection Sensors You learnt about thermocouples previously. Illustration is of a typical thermocouple fire detector. The hot junction is in the open and therefore is exposed to the air around it, although it does have a protective frame. The cold junction is in the body of the detector which insulates it to some extent from the surrounding air. When the hot junction is exposed to heat from an external source, an increase in current is obtained. The increase in current is directly proportional to the temperature at the hot junction. When the current reaches a preset level, warning indication is activated. This system is used in such areas as engines and in the aircraft pneumatic air bleed system. Again for the purpose of redundancy, this system operates off the aircrafts 28VDC system

Inspection and maintenance Apart from the normal electrical inspection, the thermal switches have to be tested with a temp-cal, or similar tester, at periods detailed in the aircraft or component manual. The switches are heated by the tester and the temperature is indicated. A test light is connected across the switch contacts and the temperature recorded when the light illuminates. The temperature at which a particular switch operates will be detailed in the aircraft manual

Thermocouple Detector System In a typical thermocouple system, it has one or more thermocouples, called active thermocouples. These are placed in fire zones around an engine while a separate thermocouple, called a reference thermocouple, is placed in a dead air space between two insulated blocks. Under normal circumstances, the temperature of the air surrounding the reference thermocouple and the active thermocouples are relatively even and no resultant current is produced. However, when a fire occurs, the difference in temperature produces a current in the circuit and activates a warning.


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System test Selection of the test function passes current through a heater around the test thermocouple. Its output will cause the system to operate. The test checks the continuity of the circuit and is usually the only test applied to these systems. Each of the individual thermocouples is not tested normally.

Maintenance procedures Because of the low voltage currents involved, all connections must be clean and tight. If a thermocouple is replaced ensure it is connected in the correct polarity.

Continuous loop fire element detector system These systems use different types of elements and control systems, manufactured by the Fenwal and Walter Kidde companies. The elements are also called fire wires. The Fenwal system uses a single conductor element and the Walter Kidde system uses a dual conductor element. Both are used primarily for engine fire detection and overheating detection of bleed air systems.

Single conductor element A thin wall tube of approximately 2mm diameter made from the alloy inconel is filled with a compound. Embedded in the centre of the filling compound is a single nickel conductor. The filling compound can be in different forms. Two examples are: • Ceramic heads or porous aluminium oxide ceramic tube impregnated with a eutectic salt: the salt’s melting point is low and its resistance lowers drastically when it melts, but will increase again when solidified • aluminium oxide suspended in fibrous glass: the resistance of this compound lowers with heat, also its dielectric qualities improve so the capacitance of the fire wire is higher as temperature increases.


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Two conductor element The tube is made from the same alloy as the single conductor element tube and is about the same size. Two nickel conductors are embedded in the filling compound which is referred to as a thermistor material, having a negative temperature coefficient of resistance. That is, an increase in temperature decreases resistance.

Fire elements come in different lengths and are identified by part number. There may be several sections joined together to make up the one element for an engine. These sections are joined to each other and the aircraft wiring by small plugs. The element is positioned around the engine or other compartment and attached by clamps, sleeves or grommets. When replacing fire wire, ensure path of wire going on match’s wire being replaced.

Testing Testing is usually carried out through the aircrafts fire test system, which is done through the control unit. Always use the correct test procedure (as per the aircraft manual) when testing continuous loop systems. Most test functions test the integrity of the circuit and the fire wires.

Installation When installing an element the following points must be complied with:  Bends can be no sharper than 1 inch radius.  The grommet must protect the element from the clamp. The clamp should not touch the wire.  The split in the grommet must be placed so that the element will not pull through.  The element must not rub or touch anything.  The element must be free of contamination, as oil etc can give erroneous readings.  The element must not be placed too near a hot area.


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Inspection and maintenance Damage to the element can cause failure of the system to operate or a false warning. Damage can be caused by poor installation, being stood on or having tools or other items placed or dropped on them. Ensure not to kink or over bend fire wires. This will cause internal breakage of wire, making the system inoperative. When inspecting an element, you must also ensure that:  The element is not broken, there are no cracks in the surface and no part of the surface is chaffed.  Indentations in the surface do not exceed parameters detailed in the aircraft manual  Grommets are in good condition  All plugs are tight and lock wired. If not, inspect for cleanliness and internal damage.


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Controllers There are several types of control units that can go with fire elements. Some work using resistance only, some use capacitance only and some use both.

Controller using resistance only The twin conductor element works on the resistance property of its thermistor filler material. At a preset level of temperature, the element passes enough current to bias the solid state controller to switch the warning devices.

Controller using capacitance only The capacitance element acts as a capacitor, storing the positive half of each cycle and releasing it during the negative half of the cycle. As the temperature rises, the capacitance of the element increases and at a preset temperature the energy released by the element will cause the solid state controller to conduct and energise the warning relay.

Testing Testing is carried out by pressing the test switch which completes a DC current path for the test relay coil. When energised, the test relay disconnects the centre conductor from the bridge rectifier and connects it to earth. This causes a warning to occur. Notice the test checks the continuity of the system. Should there be a break in the element, the test will not be successful. When the test relay is de-energised, it reconnects the central conductor to the bridge rectifier so that the power source is connected to both ends of the centre conductor. If a break in the element occurs, power will still be applied to the entire element, so a warning would be given in the event of a fire, but a test would not work.


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PNEUMATIC CONTINUOUS LOOP SYSTEMS Two types of pneumatic continuous-loop systems Lindberg System Systron-Donner System Lindberg System Consist of a stainless steel tube with an inert gas inside and a discrete material capable of absorbing a portion of the gas. The amount of gas the material can absorb varies with temperature, and when the discrete material is heated, it releases the absorbed gas, increasing the pressure in the stainless steel tube. One end of the tube is sealed and the other is connected to a responder which consists of a pressure sensing element and a set of electrical contacts. When the contacts are forced closed by the increased pressure in the stainless steel tube, a warning system is activated (lights and/or audible warning)

To test a Lindberg system, low voltage AC is sent through the stainless steel casing heating the sensing element until the gas is released from the discrete material eventually resulting in the warning system activating. outer casing. When the test switch is released, the sensing element cools and the discrete material re-absorbs the inert gas, decreasing the pressure and opening the switch contacts in the transponder. When tested, the entire system is functionally tested encompassing the control unit, wiring and the temperature sensor.


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Systron-Donner System Also a pneumatic continuous system incorporating a stainless steel tube filled with helium, but incorporates a titanium wire running through the centre. The titanium wire acts as the gas absorption material, because it contains a quantity of hydrogen. The Systron-Donner system provides a warning if an average OVERHEAT is detected, in addition to a FIRE warning. At normal temperatures the helium gas pressure has insufficient force to close the diaphragm switch. However, when the AVERAGE TEMPERATURE along the length of the tube reaches an OVERHEAT level the gas pressure increases enough to close the diaphragm switch contacts, activating the alarm.

The FIRE warning function is provided by the titanium wire. When the titanium wire is exposed to localised heating such as a fire or a bleed air leak, it releases hydrogen gas which again increases the overall pressure within the stainless steel tube, triggering the warning system. Typically this system will activate if exposed to 1100° Celsius (2000° Fahrenheit) for 5 seconds. After the fire is extinguished the hydrogen gas is reabsorbed by the titanium wire and the responder contacts break, resetting the alarm (switching it off).


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To check system integrity, the responder unit contains an integrity switch which has the contacts held closed by the normal gas pressure exerted by the helium. When the system is serviceable and helium pressure is satisfactory (no leaks) depressing the FIRE TEST switch results in the fire warning alarms activating. However, if the helium pressure is below its normal pressure, as a result of a leak in the system, depressing the FIRE TEST switch will have no result, because the test voltage will not be applied to the warning circuit, because there is an open circuit at the integrity switch contacts.

A typical system consists of two separate sensing loops for redundancy. Both loops are required to sense a fire or overheat before an alarm will sound, however if one loop fails (integrity switch opens) the system control box (typically computerised) will isolate the defective loop and reconfigure to a single loop operation using the good loop. In the case of this occurring, the system control box would also set a maintenance code indicating the fire loop failure to initiate maintenance at the next possible opportunity. It is also part of all preflight checklists for flight crew and/or maintenance crews to perform a fire detection system integrity test before each and every engine start to confirm system serviceability before operating the aircraft engines.


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Overheat Detection System Operation The above system is a typical duct and compartment overheat detection system. The top circuit is for a compartment area of the aircraft. It has several detector sensors arranged in a parallel circuit. This connected to a warning light and to the Master Warning System (MWS). No test function is supplied to this circuit. The reset temperature of this circuit is 100°C. The bottom circuit is for duct overheat detection. Two detectors are located in separate areas of the aircraft (typically the pneumatic ducting) and are tested from the one test switch. The test switch does not test the function of the detectors, just the circuit. Both detectors are again connected to a warning lamp and to the MWS.


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Smoke detectors As there is usually smoke before flames, smoke detectors are used to detect fire in equipment areas, toilets and the cargo / baggage compartments of commercial and some military aircraft. In these larger areas by the time enough heat was being generated to activate a heat detector, the aircraft would be consumed in fire. Two main types exist: •

Photoelectric (comparison and Refraction types) and

•

Ionisation type

Comparison Photoelectric System This system has two photoelectric cells. One is in a sealed compartment, the other in a compartment through which a sample of air passes. Both cells receive the same amount of light from one projector lamp. If smoke gets into the sampling compartment, the amount of light to that photoelectric cell will be less than the light to the one in the sealed compartment. This causes an imbalance in the photoelectric cells, which are connected in a bridge circuit. When the imbalance reaches a preset level, a warning light in the cockpit or other area of the aircraft is activated (i.e. Toilet warning lights are also usually located in the cabin area).

In most modern aircraft the output from the cell is passed through a control box, which monitors the output of the cell and provides a warning through a centralised system. Testing is usually achieved by providing a source of smoke, which is put near the detector.


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Refraction photoelectric system If you look at the above diagram, you will see this system has an air sample passing through a chamber. A photoelectric cell is placed at 90 degrees to a light beam. If smoke comes into the chamber, some of the light beam is refracted to the photoelectric cell (When light is reflected onto photo cell – circuit resistance decreases allowing output to fire/overheat warning systems). The output from the photoelectric cell is amplified and at a preset level, it switches a warning device such as a light in the cockpit. In most modern aircraft the output from the cell is passed through a control box, which monitors the output of the cell and provides a warning through a centralised warning system.

Testing A small test light is aimed at the photoelectric cell and when test is selected. It provides a direct light source on the cell, triggering the system.


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Ionisation type If you look at the figure above, you will see air pass through a chamber. On one side of the chamber is an electrode with a very small amount of radioactive material. There is an electrode opposite it. The radioactivity causes ionisation of the oxygen and nitrogen gases in the air, so current is able to pass across the electrodes via the ionised gases. If smoke comes into the chamber it affects the amount of ionisation and will reduce the current flow between the electrodes. When the current flow falls below a preset value, a solid state sensor turns on the warning device in the cockpit.

Again, testing is carried out through the use of a source of smoke generation.


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Smoke Detector Systems Smoke Detectors are most commonly used in aircraft cargo and toilet compartments. All run on the 28 VDC system to allow for protection during periods when aircraft AC power is not available.

Lavatory Systems Most airlines of the world now have smoke detectors in the toilets to catch cigarette smokers. Example below is that of a 737 aircraft.

They operate with an internal warning system which normally consists of a warning tone and a small LED. On larger aircraft, they are also connected to the master caution system, which alerts flight crew. Not classified as a master warning system input, so no warning bell is activated. In most cases a chime and the appropriate caution lamp will illuminate.


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Cargo Systems

Provides warning to crew of smoke in cargo compartments. Connected directly to a master warning system or through a fire detector module to the master caution system. Roof mounted in cargo hold as this avoids damage and also smoke rises. Usually are recessed and have a protective cover or grill to ensure cargo hold items do not damage the detector.

Modern Cargo Smoke Detection System Operation Smoke is detected by one or more of the smoke detectors located in the aircraft’s cargo hold. The signal from the smoke detector/s is sent to the local control unit which processes the signal. The processed signal is sent to the control display unit which sends the signal to the cockpit display. The crew take the required action to extinguish the fire in the hold. Usually one bottle per hold, ETOPS aircraft are required to be fitted with two.


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Fire Extinguisher Systems Aircraft fire extinguisher systems are used for the purpose of removing fire from the aircraft. Most aircraft extinguisher systems will remove either heat or oxygen from the fire triangle to extinguisher fire from the aircraft. Most aircraft have fire extinguisher systems for various areas of the aircraft. •

Engines

•

Auxiliary Power Unit (APU)

•

Cargo Holds

Example Twin Engine System

Look at a twin-engine aircraft system with one bottle in each engine nacelle, operated by a pull lever which also runs the motorised gate value low pressure cock closed. This system is illustrated above. Each bottle has two outlets. Assume the PORT engine is on fire; the bottle selector switch should be in the first shot-position. The pilot shuts down the engine and pulls the Tee handle. The Tee handle switch completes a current path to the close side of the low pressure cock and the first shot cartridge of the PORT fire bottle which discharges into the PORT engine. If the fire persists, the pilot selects second shot PORT with the handle still pulled out. A current path is made to the starboard fire bottle second shot cartridge. The STARBOARD bottle discharges into the PORT engine and the flap value prevents its contents going into the empty port bottle.


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Description and Maintenance The Fire Bottle is a circular bottle, usually brown in colour, containing the extinguishing agent. They are usually stored in a non-pressurized section of the aircraft. Bottles without a gauge have to be removed and weighed to check their contents. Fire bottles have to be serviced at regular time periods, which will be detailed in the aircraft manual. The gauge indicates the pressure in the bottle. Some systems are also fitted with a remote gauge. The safety plug will operate if the bottle pressure exceeds a preset level. It is designed to release the agent from the bottle in a safe manner. The output from the safety plug is connected to the discharge disks located on the aircraft. The outlet is connected to plumbing which directs the extinguishing agent into the fire zone. There are some bottles with two outlets. The cartridge is electrically detonated when the fire handle is activated. Cartridges have to be serviced at regular time periods, which will be detailed in the aircraft manual. Otherwise known as squibs.

The frangible seal is broken by a plunger driven by the force of the cartridge’s explosion. The strainer stops pieces of the seal from blocking the plumbing. Two indicating disks which are placed in an easily seen position on the outside skin of the aircraft. The red disk is connected by plumbing to the bottle relief valve. If the relief valve releases pressure, the red disk blows out. If this occurs, the bottle has to be serviced. A yellow disk is connected by plumbing and a restrictor to the bottle outlet. If the bottle discharges normally, the yellow disc blows out. If this occurs the bottle has to be changed.


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Extinguishing agent pipelines are identified by Brown coloured tape. This tape also has diamond- shaped symbols and the words “Fire Protection”. Fire extinguishing agent is distributed in the engine / APU sections by spray nozzles and perforated tubing. This is also marked with the brown fire tape.

Extinguishing Agent The most common agent is Halon. It is a CFC based chemical. •

Halon is a colourless, odourless, non-toxic, gas.

•

It is fast acting - it cools the surface and chemically interferes with the combustion process.

•

The gas is attracted to heat, moves around barriers and tracks to a fire. It blankets the fire area and reduces the risk of flashback.

•

Halon is clean and evaporates completely, it's non-staining, non-corrosive and nonconductive.

•

Halon will not damage any surface, including electronics, fabrics and painted surfaces.

•

Halon does cause damage to the atmosphere and is in the process of being replaced.

•

Halon can be used equally effectively on flammable liquid (Class B) and flammable gas (Class C) fires.

Fixed installation aircraft fire extinguishing systems These systems are usually installed to extinguish fire in the engine, APU and cargo holds. Once the fire handle has been activated, voltage is applied to the cartridge (cartridges should fire with only 3 volts DC applied with very little current) – the cartridge then fires. The frangible seal is then broken, allowing the release of the agent, under pressure, through the strainer into the piping where it is directed to the fire area by spray nozzles and perforated tubing.

Fire Panel Relevant handle lights may flash or illuminate steady, depending on aircraft design. Some aircraft handles flash for engine turbine area overheat and illuminate steady for turbine fire. All use red for fire warning and amber for overheat indication.


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Crew procedure for an engine fire In the event of an engine fire, the crew will activate the necessary fire handle. When the handle is pulled or activated it closes off the fuel, hydraulics, pneumatics, electrics and the thrust reverser systems. The next movement of the handles will cause the fire extinguishers to be activated.

System Basics Most aircraft have two fire bottles fitted, one bottle per engine with cross-feed plumbing, so the bottle can be discharged into the other engine as a second shot. The discharge of the first bottle is called first shot and the discharge of the next bottle is called second shot. The electrical switching to discharge the bottles always has protection against inadvertent firing. This can be: • A push button guarded by a flap. The switch is illuminated, indicating the button to be selected. Most common in Airbus aircraft. • A toggle switch guarded by a hood • A Tee handle which has to be turned 90 degrees before it can be pulled. These are usually transparent and contain a fire warning light (see above photo). This helps prevent wrong selection because it is the illuminated handle that has to be pulled. This is the most in Boeing aircraft.

Example Twin Engine System

The above diagram illustrates a twin-engine aircraft system with one bottle in each engine nacelle (each bottle has two outlets), operated by a pull lever which also runs the motorised gate value low pressure cock closed.


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System Operation Assume the STARBOARD engine is on fire; the crew receive a warning through the aircrafts fire detection system. The bottle selector switch should be in the first shot position. 1. The pilot pulls the Tee handle. 2. The Tee handle switch completes a current path and the first shot cartridge of the STARBOARD fire bottle which discharges into the STARBOARD engine. 3. Power is provided through the fire handle and relay to the closed side of the low pressure fuel cock motor. This closes off fuel to the starboard engine. 4. If the fire persists, the pilot selects the second shot PORT with the handle still pulled out. 5. A current path is made to the port fire bottle second shot cartridge. 6. The PORT bottle discharges into the STARBOARD engine and the flap value prevents its contents going into the empty starboard bottle.

Fire Extinguisher System Operation The diagram below illustrates a Boeing 727 fire extinguisher system.


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System Overview The system is operated off the aircraft’s 28 VDC bus. Three engine discharge switches are located in the cockpit and are used to operate the engine selection valves and fire bottles. The system has two fire bottles, which are connected through a shuttle valve to all three engines. The bottle discharge sequence is controlled by the selection of the bottle discharge switches, located in the cockpit. A yellow discharge disk is located in the supply line, which is used to indicate when a bottle has been discharged under normal use. Bottle discharge indication is provided in the cockpit by bottle discharge lighting. A cut away is also provided to show the operation of the engine selection valves.

Operation The flight crew receive a warning of a fire and this is displayed on the cockpit fire handles. In this system the fire handle will illuminate red when a fire has been detected. The flight crew will then pull the relevant fire handle, after selecting the fire bottle to be used. The bottle to be used is selected via the transfer switch located on the pilot light shield panel. Once the handle has been operated, a circuit is completed from the 28VDC bus to the fire bottle selected. 28 VDC is also placed on the engine selector valve. This operates the valve so as to allow for the fire agent to enter the engine. Once the bottle has been fired, the agent forces the shuttle valve against the opposite bottle, preventing agent into the other bottle. The yellow disk is also expelled under the pressure of the agent. Upon the bottle firing, the bottle discharge switch is activated and completes the circuit for the bottle discharge cockpit lighting. The agent is then released into the applicable aircraft engine.


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APU Fire System. Most modern day commercial aircraft have a separate APU fire detection and extinguisher system. This is due to the APU being required to operate during ground operations. The APU fire detection system is usually a single or double fire loop detection system, which only detects fires and not overheat situations. The loop is connected to an APU Fire Detection Module, located in the main electronics area of the aircraft. Fire warnings are provided both in the cockpit and in a remote location of the aircraft – usually on the nose landing gear or in the main wheel well (see above photo).

From both of these locations an aural and visual fire warning is displayed. In most aircraft the activation of the APU fire warning system will automatically shutdown the APU. The APU fire bottle can be discharged from either the cockpit or remote location fire panels.


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PRECAUTIONS Never use an ordinary ohmmeter for checking extinguisher circuits - It can fire the cartridge! Use only a safety ohmmeter. Ensure all associated circuit breakers are pulled before working on the aircraft fire protection system. Ensure no static charge is present on person or tools prior to touching fire bottle cartridges. Cartridges are to be shorted at all times when NOT fitted to bottles.

PPE Always wear approved: •

Breathing Mask or protective eyewear as required:

•

Gloves:

•

Special overalls or protective clothing and

•

Protective footwear.

ALWAYS refer to the MSDS for the correct and current handling precautions. The employer should have a copy of the MSDS (under current OH&S law) for each chemical in use by the organisation.

Extinguisher Agent Properties Chemical may be: •

Toxic

•

Carcinogenic

•

Corrosive

•

Lethal


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Classes of Fire A CLASS fires are those involving ordinary combustible solids, such as, paper, wood, cloth, rubber, and many plastics. A class fires need a high amount of heat to break down the burning material and turn it into vapours that are consumed in the combustion process. To extinguish these types of fire cooling is required. The best method of cooling is to apply water directly to the burning material to reduce its temperature, stop it giving off vapours and therefore extinguish the fire. The most suitable type of extinguisher for this material is WATER which are typically rated 2A or 3A, (the A being for A class fires), but any extinguisher displaying an A rating, or a green pictograph depicting a waste paper bin and wood on fire, may be used.

B CLASS fires are fires involving flammable liquids. such liquids as petrol and oils, even high flashpoint liquids such as diesel fuel. B class fires can involve many different types of flammable liquids. Some, such as petrol and shellite are easily ignited at room temperature and will rapidly consume the material. These fires cannot be adequately cooled to bring the material below the point where flammable vapours are being produced. Other liquids may need to be heated to produce adequate vapours to burn, whilst these materials could be cooled to extinguish the fire using common water, we find that applying water to very hot liquids can result in a violent reaction as the water turns to steam causing the burning liquid to froth and boil over spreading the fire in all directions. Extinguishers suitable for B class fires are designed to smother the fire or remove the oxygen from the fire. Some also interupt the chemical reaction which is the fire. Extinguishers suitable for B class fires will carry a rating such as 20B or up to 80B (the B being for B class fires). They may also have a green pictograph displaying a fuel can and liquid on fire.


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C CLASS fires are fires in flammable gases. there are a wide range of gases in common use that are flammable, most common would be LPG and acetylene. C class fires involve flammable gases, which are already a vapour and ready to burn when mixed with sufficient oxygen and ignited. Cooling the fire will not be enough to extinguish it. In fact extinguishing the fire may lead to a greater hazard of explosion, by allowing unburnt gas to escape. If extinguishing the fire is neccessary to to control the fire, and the gas can then be controlled, then an extinguisher suitable for C class fires must be used. The common extinguishers don't carry a C class rating where the hazard exists a DRY CHEMICAL type extinguisher will be installed. This is the only common type of extinguisher suitable for flammable gas fires.

D CLASS fires are metals fires. some metals are capable of sustaining combustion if adequately heated and ignited. most commonly occurs in workshops machining high magnesium content metals. Fires involving metals are rare due to the amount of heat required to commence the combustion process. Typical areas where a D class hazard exists is in machine shops where magnesium based metals are being machined and small pieces of metal accumulate, the aircraft industry can also have metals that may burn. Cooling of a D class fire is not readily possible as the extreme heat and chemical reaction may cause the fire to flare violently with the introduction of a cooling agent. Extinguishers suitable for D class fires are a specialised area, they generally involve the use of specialised dry chemical agents (powders) that are expelled slowly from a special extinguisher so as to encase a D class fire and form a crust sealing off the metal from the air.


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E CLASS fires are those involving live electrical hazards. the actual material burning may be of any other class, but if live electrical components are suspected to be involved then it is essential the extinguishing medium be suitable for E CLASS fires.

Electricity does not burn, however electricity can be a cause, or involved in many fires, and it

is extremely hazardous to us. E class extinguishers are those extinguishers which contain materials that will not conduct electricity. These extinguishers will be rated for another class of fire and also contain a rating (E), the E is in brackets to indicate that it is not the burning material, but it is an important subsidiary risk. Where you can reasonably suspect that an electrical hazard exists always use an extinguisher that carries a rating that includes (E), the extinguisher may also carry a green pictograph displaying a powerpoint and plug.

F CLASS is particularly for cooking fats and oils. F class fires are included as a class of their own to allow for the Wet Chemical extinguisher which is particularly effective against fires involving cooking fats and oils, but is not suitable for other flammable liquid type fires (B class). F class will not appear on the extinguishers rating as a single letter, rather most B class extinguishers may be used and those which are particularly suitable may carry the words 'suitable for fires involving cooking fats and oils' alongside the rating.


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Extinguisher Types 3 most common extinguisher types: Water Dry chemical powder (DC) Carbon dioxide (CO2). Other less common extinguisher types are foam and wet chemical

Water Extinguishers – Class A Water or APW (air pressurised water) extinguishers are large, red extinguishers, filled about two-thirds with ordinary tap water and then pressurised with air. This type of extinguisher most effectively removes the heat element of fire triangle. When the extinguisher is activated, a powerful stream of water is forced out the nozzle. In passenger commercial aircraft, this type of extinguisher may be mounted only in the passenger cabin. Never use water to extinguish an electrical fire. Water is a good conductor, and there is a high risk that you may receive a severe electric shock if water used on an electrical fire. Electrical equipment must be unplugged and/or deenergized before you use a water extinguisher on it. Never use water to extinguish flammable liquid fires. The burning liquid may float and spread out on the water surface, thus spreading the fire.

Carbon Dioxide (CO2) Extinguishers Filled with non-flammable carbon dioxide gas under extreme pressure. Pressure is so great that when used, bits of dry ice may shoot out nozzle. Extinguishes fire by displacing air – thus removing the oxygen. Also provide some cooling of the fire. Recognised by its black band on a red body and a large, wide nozzle shield on end of a flexible hose. These extinguishers are frequently used in laboratories and other areas where flammable liquids are stored or used.

Dry Chemical Power Extinguishers Filled with a fine powder and pressurised by nitrogen. Extinguishes fire by coating fuel with a thin layer of powder. This layer separates the fuel from the oxygen in the air. Powder also interrupts the chemical reaction of fire. Identified by a white band on a red body. They are extremely effective at extinguishing fires. Dry chemical extinguishers come in 2 types 1. ‘AB(E)’ – indicating they are designed to extinguish class A,B, and (E) fires 2. ‘B(E)’ – designed to extinguish class B and (E) fires, but NOT class A Either one may be used for a class C (flammable gas) fire after the source of gas is turned off, and for a class F (cooking oil or fat) fire.


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Foam Extinguishers Foam extinguishers are suitable for either Class A, (combustible materials) and class B, (flammable liquid) fires. Giving a good flame knock-down, the foam extinguisher retains the features to both extinguish the risk and secure the hazard.

Wet Chemical Extinguishers Wet Chemical Extinguishers are a special purpose designed extinguisher, specifically for use in kitchens and restaurants, where the major risk is from burning fats and oils (class F fires)

Extinguisher Use Even though extinguishers come in a number of shapes and sizes, they all operate in a similar manner. Here's an easy acronym for fire extinguisher use: P A S S - Pull, Aim, Squeeze, and Sweep


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TOPIC 13.14: ICE AND RAIN PROTECTION (ATA-30) Ice Formation The formation of ice on aircraft is caused by water droplets in the atmosphere which are at a temperature below zero degrees centigrade, but which have to lose their latent heat before they will freeze. As the droplets strike the metal surfaces of the aircraft, the metal conducts the heat away from the water and the droplets freeze on the metal. Types of Ice There are two types of ice which form on aircraft while in flight, they are: 

Rime ice



Glaze ice.

Rime Ice

This ice formation is caused when small droplets freeze immediately they come in contact with the aircraft. It builds up slowly, is opaque, and has a rough surface which increases drag and changes the aerofoil shape, but is easy to break off. It is rough because the temperature of the air is very low and freezes the water before it has time to spread. This type of ice formation is mainly found on the top surface as the water freezes upon contact with the aircraft.

Glaze Ice

This ice forms when large droplets of super-cooled water strike an aircraft’s surface which is below zero degrees Celsius. The water spreads over the surface and freezes. Glaze ice is the most dangerous: it is clear, adds large amounts of weight to the aircraft and is hard to break off.


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Dangers of ice forming on an aircraft Ice forming on an aircraft is dangerous and can cause an accident for the following reasons: • Ice formation on the wings and stabiliser increases the aircraft’s weight and disturbs the smooth aerofoil shape, reducing lift. • Ice formation in piston engine carburettors can block the carburettor air intake causing engine failure. • Ice formation on a propeller of either a piston or jet engine will lower the efficiency of the propeller, make it ineffective or affect its balance and cause vibration. • Ice formation on jet engines’ air intakes restrict the airflow, causing loss of power or overheating. Ice formed when breaking off can cause engine compressor blade damage. • Ice formation can block ram air intakes. • Ice formed on masts and probes such as pitot tubes, stall warning vanes and antennas can cause the associated system to fail, as well as increase drag. • Ice formed on windshields can obscure vision. • Rain on windshields can also obscure vision.


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Ice Detection is carried out by two methods: Visual- look and see with the aid of lighting and witness marks (ice build up on the wings, windscreen or wipers). Method usually used by smaller aircraft, where costs are a factor. Also used by some older aircraft designs. Electrical – by the means of an electro mechanical device. This is the most common method of ice detection in modern aircraft.

Visual Ice detection Visual detection is the main form of ice detection on smaller and older aircraft and the reason for the ice detection (leading edge) lights. Black witness marks are used on some aircraft as a visual aid of detecting ice build up. As ice builds upon the leading edge of the wing, the black witness marks becomes more reflective and harder to see when the light from the leading edge lights are shined on them. This indicates to the crew that icing conditions are present. Ice will usually build up on windscreen wipers (very visual) in the event of flying in icing conditions.

Electrical Ice Detection Electrical and electronic detection methods are also used, of which there are three types. •

Pressure

•

Ice shave

•

Frequency


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Pressure probe ice detection This is a probe positioned in the airflow which has a pressure produced in it, acting on a diaphragm. Should ice form, it will restrict the probe opening, reducing pressure on the diaphragm and allowing a set of contacts to close. The contacts will supply current to an ice warning light and a heater in the probe to melt the ice formed in the probe. The crew takes the necessary action while the heater in the probe melts the ice formed in the probe.

Once the probe heater melts the ice, pressure on the diaphragm increases, opening the contacts and switching off the detection probe heater and associated warning light. The warning is repeated if ice forms again.

Ice shave system Most common method of ice detection. A sample of the air passes over a rotor driven by an electric motor. Two methods are used supply a warning light signal 1. If enough ice forms on the rotor, a blade in close proximity will shave it off. The shaving requires more motor torque, causing the motor to twist slightly in its mount. This action activates a micro-switch to turn on a warning light. If icing ceases, the shaving stops and the motor torque decreases. This deactivates the micro-switch, turning off the warning light. 2. It can also be sensed by the current draw, as the drum slows it draws more current to maintain its RPM, this being sensed by a current metre. When the current draw level exceeds a preset point, a warning light is activated in the cockpit. This system can also be used to control a de-icing system.


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Frequency probe system As illustrated, the probe is mounted in a suitable place on the aircraft and is often duplicated. Operation The probe is designed to have a natural resonant frequency in the ultrasonic range. An oscillator induces this frequency in a coil around the probe. If ice forms on the probe, its resonant frequency reduces due to the extra weight of the ice. When this reduction reaches a preset level, electronic circuitry switches on a crew warning light and a probe heater. The probe heater is kept on by the 5 second time delay and should melt all ice formed then switch off in the 5 seconds. If ice re-forms within the remaining 55 seconds that the warning light remains on, the process will be repeated. The warning light delay will continue to keep the light on while ice continues to form. For example we have a probe which is supplied by a microprocessor with 40 Hz causing it to vibrate. As ice builds up on the probe in the airflow it causes the frequency to slow which is sensed by the microprocessor, which in turn will give the crew a visual warn light to say they are entering icing conditions.

Types of ice control systems Ice control systems can be categorised as: • Anti-ice systems which prevent the formation of ice, and • de-ice systems which remove ice after it forms.


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De-icer Boots This system is in common use on propeller aircraft. If you look at the illustration, you will see a rubber boot fixed to the leading edge of an aerofoil. For many years aircraft have used deicing systems consisting of inflatable boots on leading edges and stabilisers. The inflatable boots are usually constructed with several separate air passages or chambers enabling some to be inflated while the others are deflated. These chambers or tubes in the boot are attached to plumbing from an air control valve. The valve will apply low pressure or vacuum to all the tubes when the de-icing system is not in use, so the boot is held hard against the aerofoil, presenting a smooth surface to the airflow. This vacuum pressure can be achieved by either engine driven pumps or by a simple ejector operating on the venturi principle. When in use, tubes will be inflated with high pressure air (approximately 18 psi) for a specified time (usually 6 seconds) then reconnected to the low pressure line. The tubes are inflated for the same period and reconnected to the low pressure line. The inflation of alternate tubes makes them protrude out and break the ice, which is blown away by the airflow.


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Pneumatic De-Ice System Operation

The diagram above is a typical deicing system for a twin-engine airplane. This system has wet-type air pumps on both engines. The discharge air from these pumps flows through oil separators and check valves into the deicer control valve. When the system is turned OFF, the air discharges overboard, and when it is turned ON, the air flows to the distributor valve timer. An electric motor drives the distributor valve in a timed sequence to the center tube of the outboard boots, then to the outer tubes of the outboard boots. The boots on the empennage then inflate and deflate, then the center tubes, and finally the outer tubes on the inboard wing boots inflate and deflate. The boots actuate symmetrically to keep the airflow disturbances even on both sides of the aircraft. This minimizes any flight or control problems caused by these disturbances. When the system is turned OFF, the distributor valve connects the suction side of the air pumps to the boots to hold them tightly against the leading edges. A suction-relief valve installed between the check valves and the distributor valve regulates the amount of suction that is applied to the boots. Proper actuation of the deicer system may be determined by watching the pressure and suction gages. The pressure gage fluctuates as the timer sequences the different boots, but the suction gage remains steady since the vacuum side of the pump is not used during normal operation of the system.


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Turbine-engine aircraft have a ready source of warm compressor bleed air for anti-icing, and they normally use thermal ice control. Some of the smaller turbine engines do not have an adequate quantity of bleed air for thermal ice control, but do have enough for inflating pneumatic deicing boots. Systems that use compressor bleed air for this purpose have a pressure regulator that lowers the pressure to the correct value and a venturi downstream of the regulator that produces suction when the boots are not inflated. This suction holds the tubes deflated and tight against the leading edges.

Thermal pneumatic systems The jet engine bleed air thermal pneumatic system takes hot air from the engine compressor and directs it between the aerofoil leading edge outer skin and an inner skin before exhausting overboard. If you refer to illustration, you will see the ducting and the airflow between the two skins. Heated air for anti-icing is obtained by bleeding air from the engine compressor. The reason for the use of this type of system is that large amounts of very hot air can be tapped off the compressor, providing a source of anti-icing and de-icing heat. The hot bleed air is mixed with ambient air and at approximately 350°F, flows through passages next to the leading edge skin. Each of the shutoff valves is pneumatically actuated and electrically controlled. When the temperature in the leading edge reaches approximately 185°F a thermal switch connected to the control solenoid of the shutoff valve causes the valve to close and shut off the flow of bleed air. An operational system check can be carried out by using an external source of air. Most systems are designed with a test plug to permit ground checking of the system without operating the aircraft engines. When using an external air source make certain that the air pressure does not exceed the test pressures established for the system. The air used for airframe leading edge de-icing / anti-icing is vented back out, in most aircraft, via small holes located under the wing. For other areas of the aircraft where thermal pneumatic air is used, this is exhausted via the rear of the leading edge surface, into the non-pressurised area of the airframe (e.g. vertical stabiliser, horizontal stabiliser).


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Chemical de-icing Although it is not often used on modern aircraft, chemical de-icing can carry out all de-icing requirements on slower aircraft. The common de-icing fluid used is a mixture of isopropyl alcohol and ethylene glycol. These substances emulsify with water and lower its freezing temperature so the ice will melt. The de-ice fluid also makes the surfaces slick so ice will have trouble reforming on them. Therefore, the de-ice fluid performs both a de-ice and anti-ice function.

How the system works Each system being chemically de-iced has an electric motor driving a pump. The pump supplies the de-icing fluid from a storage tank through plumbing to the area to be de-iced as follows: • Windshield de-ice: the fluid is sprayed over the windshield • Carburettor de-ice: the fluid is sprayed into the carburettor air intake • Propeller de-ice: the fluid is sprayed out along the blades • Wing and empennage: de-icing is achieved by the de-ice fluid being slowly released through a porous boot fixed to the leading edges. The circuit is shown. A rheostat (variable resistor) is used to control the speed of each pump motor, which controls the fluid flow rate.


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Propeller De-icing – Fluid System A typical anti-icing system consists of a control unit, a tank (anti-icing fluid) a pump to deliver fluid to the propeller and nozzles. The control unit contains a rheostat which is adjusted to control pump output. Fluid is pumped from tank to a stationary nozzle installed just behind the propeller on the engine nose case. As fluid passes through the nozzle, it enters a circular U-shaped channel called a slinger ring. A typical slinger ring is designed with a delivery tube for each blade. Once fluid is in the slinger ring, centrifugal force sends the anti-icing fluid out through the delivery tube to each blade shank.

Propeller Ice Control - Electrical Electrical heating is the preferred method of ice control for propellers due to the available supply of electrical power during flight. Rubber boots with heater wires embedded in the rubber are bonded to the leading edges of the propeller blades. Electrical current is passed through these wires to heat the rubber and melt any ice that has formed, allowing centrifugal force and airflow to carry the ice away. Electric elements are set in insulating material on the propeller leading edges and in the propeller spinner for the purpose of de-icing. There is no air intake de-icing on piston engines and some do not have propeller de-icing.


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Timer Unit The timer controls the sequence of current to each of the heater mats. The sequence of heating is important to provide the best loosening of the ice so it can be carried away by the centrifugal force, it is also important that the same portion of each blade be heated at the same time. Also if heat is applied for to long duration to the heating mat, delamination could occur, causing damage to the system.

Load Meter The ammeter monitors the operation of the system and assures the pilot that each heater element is taking the required amount of current. This system above would typically use a 28VDC power supply. Two heater elements are fitted to each propeller, an outboard and inboard heater element. Each of the elements is connected to the power supply via slip rings and brush box assemblies. The timer unit is common to all propellers used on the aircraft. With the system selected on, power is supplied through the circuit breaker and ammeter to the timer unit. The timer then supplies the power to the heater in the cycle sequence (1) and through the brush block and slip rings; power is delivered to the heater element. After this the current is then passed back to the slip and through to earth. When an aircraft flies through icing conditions, ice can build up in the engine’s inlet duct and on its inlet guide vanes. This disrupts airflow and reduces efficiency. Furthermore, large pieces of ice could break off and enter the engine causing serious damage to compressor blades. To prevent ice formation and ingestion, turbine inlet ducts are typically equipped with some form of anti-icing system.

Engine Air Inlet Ice Control Systems

When the anti-icing system is switched on, a bleed air valve directs hot air to the inlet duct leading edge, nose dome, and inlet guide vanes to prevent the ice from forming. An indicator light illuminates (usually amber) in the cockpit to show anti-icing is on. Amber is used as an advisory display as this system is not normally an operating system. Once the air has been used, it is vented out onto the airflow by venting ports located in the engine cowl. A disadvantage with this type of system is that, whenever bleed air is taken from a turbine engine, the power output is decreased. This is not normally a problem on larger modern turbofans, but can be a factor on smaller turbine engined aircraft.


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When the anti-ice is selected to on, this drives the motor in the engine anti-icing valve to the open setting. With the valve open, pneumatic system air is allowed to pass through the air regulator. The air regulator is fitted with a bimetallic spring coil. This controls the amount of air flow to the engine anti-icing system by the temperature of the air being delivered. As the pneumatic system air flows over the bimetallic strip, it increases the temperature of the bimetallic strip. Once the temperature increases to a predetermined level, the bimetallic strip will expand and reduce the airflow to the de-icing system. This will then control the airflow and the temperature of the de-icing surfaces. Excessive temperature can cause damage to these areas. By using this type of air regulator airflow and temperature can be controlled from the one device. From this point it is divided up and distributed to the spinner and nose cowl anti-ice systems. Air is then removed from the engine via venting holes, located around the nose cowl and the spinner of the engine.


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B2-13f Electrical

Engine Air Inlet Ice Control Systems This engine power loss may be alleviated by the use of an electric engine air inlet anti-icing system. This system utilises electric heating elements surrounding the engine air intake.

This system is only used on turboprop aircraft. Not used by high bypass turbofans as current required would be to large. Usually has two selections, fast and slow (low / high). This provides operation in various icing conditions, as continuous use of the fast (high) setting in low to medium icing could cause damage to the air intake heaters.


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Windscreen Wipers Aircraft windshield wipers can be powered by a pneumatic motor, a hydraulic motor or electric motor, the latter can be a DC or AC motor. This is the most common form of rain removal from aircraft windscreens. Most aircraft wipers can be operated independently of each other.

Do not operate on dry windscreens and ensure clear of foreign matter, as this will cause damage to the windscreen. Usually four position selector switch is used (Off, Park, Slow and Fast). Park is a momentary switch, which is used to park the wiper out of view once the wipers are no longer required. Operates against spring tension and once released the switch goes back to the OFF position. Forces Cam to operate wiper till internal park switch is activated, stop motor and wiper is then in the park position. The slow and fast operation provides power to the wiper motor. During the slow operation a higher resistance value is used.


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Pneumatic Rain Removal System Windscreen wipers have problems with blade loading and speed. With the advent of turbinepowered aircraft, the pneumatic rain removal system became feasible. The air blast forms a barrier that prevents raindrops from striking the windshield surface. Air is bleed from the aircrafts engine bleed air system. This air is then forced over the windscreen by small ducts. Usually cannot be operated independently for each pilot. Most commonly used on military fighter aircraft.


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Rain repellent system

Rain repellent fluid reduces the rainwater surface tension so that it forms into globules and is blown away by the airflow. The rain repellent fluid is in a pressurised container connected to a reservoir. Plumbing connects the reservoir via a solenoid valve to a spray nozzle at the bottom of each windshield; there is a solenoid valve for each spray nozzle. When operated, the solenoid valve only opens for approximately 0.25 seconds so that approximately 5 cubic centimetres of fluid is metered out on the windshield. This is achieved by a time circuit, stopping power to the solenoid valve after 0.25 seconds. Do NOT apply repellent on dry windows. Undiluted repellent will restrict visibility. Do NOT operate windscreen wipers after having applied repellent. Smearing may occur and reduce visibility Most aircraft have this system deactivated due to environmental reasons as CFC was used as the propellant. New products are now available if operator wishes to use.

Electric windshield anti-icing Windscreens are heated to prevent the build up of ice and also to make the windscreen more pliable and therefore, more able to withstand the effects of a bird strike. The laminated construction also helps to prevent the windscreen from shattering into small sharp pieces when broken. Some light aircraft have windshields made with fine wire elements embedded in a layer of vinyl, sandwiched between two layers of glass. 28 V DC current is passed through the element and a thermistor, also embedded in the windshield, is used to control the DC current to maintain a temperature of approximately 45°C.


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Larger Aircraft The construction of large aircraft windshields is shown in illustration. The conductive film can be stannic oxide or gold, rolled so thin that it is transparent. The temperature is monitored by the temperature sensing element (shown in centre of photo) and the temperature control unit turns the power relay on and off to regulate the temperature. Do NOT operate on ground for extended periods – only quick test. ALWAYS test in accordance with the aircrafts maintenance manual to prevent possible damage to the aircraft’s windscreen.

Temperature Control

We will look at a system powered by a 3-phase 200 V AC source which can be frequencywild or have constant frequency. In the previous figure you can see the three elements, one for each phase, and the normal and overheat temperature sensing thermistors. You may see windshields with several thermistors. They are placed there during manufacture and the best two are selected for the job. Issue B: January 2008

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Operation The 3-phase 200 V AC power source goes to an auto transformer. The transformer has two output voltages; both are higher than the input voltage. The lower voltage is applied to the windshield element when low heat is selected, and the high voltage is applied when high heat is selected. Assume the windshield control switch is selected to LOW. If the windshield temperature is below 45°C, the temperature controller will complete a current path for the coil of relay RL1. When energised, RL1 connects the lower voltage from the autotransformer to the windshield elements. The windshield elements heat the windshield until it reaches 45°C. Then the normal heat thermistor will cause the temperature controller to break the current path to the coil of RL1, which de-energises, breaking the current path to the heating elements. As the windshield cools, the thermistor causes the controller to switch the heat back on. If ice still forms on the windshield when LOW heat is selected, HIGH will have to be selected. This will cause the temperature controller to operate relay RL2 instead of relay RL1. Relay RL2 will switch the higher voltage output from the autotransformer to the heating elements, so more current will flow through the elements producing more heat. The temperature controller will still control the windshield temperature, so that it does not exceed 45°C. If the temperature controller or the normal thermistor fails, the temperature of the windshield can rise above 45°C. Should the temperature reach 55°C, the overheat thermistor will cause the other windshield temperature controller to break the earth circuits of relays RL1 and RL2. This will stop current flow through the windshield elements. The windshield can continue to operate in this condition but must be fixed when the aircraft lands. A magnetic indicator for each windshield will show NORM when the system is operating normally, OH if it is operating in the overheat condition, or OFF when off. Jet aircraft windshield strength The windshield of a jet aircraft has to take all the forces of the air acting on it at speeds up to 1,000 kilometres per hour. It must also be able to withstand the impact of a two kilogram bird at the cruise speed of the aircraft. Leaving the windshield heat selected to LOW even when there is no icing conditions will have the following results: • increase in the windshield’s impact resistance • increase of the windshield’s life • better dissipation of static electricity.


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Probe heaters In larger aircraft which are equipped for flight in bad weather, heaters are provided to reduce ice formations which would affect the airflow about the tube, or completely seal off the opening, there by producing erroneous readings. Any aircraft which, in the course of its operation, can fly at or above icing levels must be fitted with heated pitot and static probes to prevent the build-up of ice. What exactly is the icing level? As you know, the temperature of the atmosphere decreases as you go higher up. This means that any moisture which lands on the probe in flight, will eventually freeze and block up the probe, destroying any readings (see photo).

To overcome this problem, heating elements are fitted into the forward parts of the probe which will heat up and melt any ice build-up in flight. Where more than one Pitot static system is fitted, heater operation is totally independent of each other. Do NOT touch probes after flight, as units are extremely hot. Allow cool down time before attempting maintenance. Unless otherwise stated in the appropriate maintenance manual, anti-icing systems using electrical heating are NOT TO BE OPERATED ON THE GROUND. Without the cooling effect of airflow over the heater, it is possible that the temperature of the heater can rise sufficiently to cause severe damage to the heated area.

Where heated probes are fitted, the covers should be manufactured from cotton canvas rather than vinyl or synthetic materials, so that in the event of the heater probes being turned on with the covers still fitted, they will burn away rather than stick to the probes, possibly plugging up the openings.


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Pitot-Static Probe Heaters Circuit Operation Each main pitot-static probe is anti-iced by two electric heaters. On Ground On the ground, with engines not operating, the Probe Heaters are not powered. With any engine operating, relays R7423 and R7425 are energized. • The strut of the probe is supplied with 115 volts AC power. • Reduced power is supplied to the head of the probe. In Flight In flight, relay R7425 is deenergized, relay R7423 remains energized through the engine speed sense card or relay R7334. • The head and strut of the probe are supplied with 115 volts AC power. Display The power supply to the heaters is sensed by two current sensors. • If power is not supplied through the heater, the current sensor provides this information to EICAS through the EFIS/EICAS interface units. Test The heater operation can be checked on the ground by the Central Maintenance Computer. During test, relays R7421 and R7423 are energized. The head and strut of the probe are supplied with 115 volts AC power. The test results are observed on the CDU.

Pitot-Static Probe Heaters Circuit


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Total Air Temperature Probe Heater Circuit Operation Each total air temperature probe is anti-iced by an electric heater. On Ground On the ground the probe heater is not powered. In Flight In flight, relay R8268 is energized through the air/ground relay R7334. The heater is supplied with 115 volts AC power. Display The power supply to the heater is sensed by a current sensor. • If power is not supplied through the heater, the current sensor provides this information to EICAS through the EFIS/EICAS interface units. Test The heater operation can be checked on the ground by the Central Maintenance Computer. During test, relays R7421 and R8268 are energized. The probe heater is supplied with 115 volts AC power. The test results are observed on the CDU.

Total Air Temperature Probe Heater Circuit


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Water and Toilet Drain Heaters Heaters are provided for toilet drain lines, water lines, drain masts, and waste water drains when they are located in an area that is subjected to freezing temperatures in flight. Typical heater types used are: Integrally heated hoses 

Ribbon

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Blanket

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Wrap around patch

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Gasket

These heaters are designed for continuous operation. They usually operate in two modes – Air and Ground. This is controlled by the air to ground sensor system of the aircraft. Most systems operate at 26 VAC on ground mode and 115 VAC in air mode. Anti-icing may also use bleed air. Small line from the aircraft bleed air manifold is directed onto the probe.


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Drain Mast Heater Circuit Ground Operation Ground Handling Bus Powered With the Ground Handling Bus powered, relay R8277 is relaxed, relay R731 is energized and the Heaters are supplied with 42.5 volts. AC comes from the Drain Mast Heater Transformer. No Ground Handling Bus Power When power is supplied to the Main Airplane Busses, relays R8277 and R731 are energized and the Heaters are supplied with 42.5 volts. AC comes from the Drain Mast Heater Transformer. Air Operation In flight, the Ground Handling Bus is not powered, relays R8277 and R731 are relaxed. The Heaters are supplied with 115 volts AC from the Ground Service Transfer Bus.

Drain Mast Heater Circuit


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