EAA CPL 010 Aircraft Technical

Eltanin Aerospace Academy

EAA-CPL-MAN-010v1

COMMERCIAL PILOT LICENCE AND INSTRUMENT RATING EXAMINATION

AIRCRAFT, TECHNICAL & GENERAL BOOK 2 Version 1.01

PREFACE These Study Notes have been written for pilots preparing for the South African Commercial Pilot Licence and Instrument Rating technical examination. They cover the full, published syllabus as at April 1999. This book has been re-written and re-drawn in accordance with the latest CAA Syllabus by Cedric Mew, Gr II Instructor. Eltanin Aerospace Academy acknowledges his invaluable assistance.

The syllabus has been split into chapters for ease of use. Each chapter usually covers one section of Aircraft Technical. Students should complete these exercises and questions before moving on to the next chapter. In addition a book of “Typical Examination Questions” is provided in your set of Home Study Notes. PLEASE NOTE THAT ALL THE DIAGRAMS ARE FOR STUDY PURPOSES ONLY. DISCLAIMER While every reasonable care has been taken to ensure that these Notes are free from error, no responsibility whatever is accepted for any actions or claims resulting from the use of these Notes. By purchasing these notes you accept these conditions and agree to be bound by them.

Copyright © are jointly held by AeroNav Academy and Eltanin Aerospace Academy

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AIRCRAFT, TECHNICAL & GENERAL - BOOK TWO CONTENTS AIM AND OBJECTIVE

Page 4

AMENDMENT LIST RECORDS

Page 5

Chapter 1

Airframes

Page 9

Chapter 2

Aircraft electrical systems

Page 15

Chapter 3

Fuel systems & fuel types

Page 47

Chapter 4

Undercarriages

Page 53

Chapter 5

Oxygen systems

Page 59

Chapter 6

Pressurisation and Air conditioning

Page 63

Chapter 7

Anti-ice and de-icing systems

Page 83

Chapter 8

Hydraulics

Page 93

Chapter 9

Pneumatics

Page 103

Chapter 10

Power plants

Page 107

Chapter 11

Engine parts and power ratings

Page 109

Chapter 12

Piston engines

Page 119

Chapter 13

Detonation and pre-ignition

Page 131

Chapter 14

Lubrication and cooling systems

Page 135

Chapter 15

Carburetion and fuel injection systems

Page 141

Chapter 16

Supercharging and turbo-supercharging

Page 151

Chapter 17

The gas turbine engine

Page 157

Chapter 18

Aircraft Elements –Bearings, Valves, Pumps Filters

Page 175

Chapter 19

High Speed Flight

Page 181

Chapter 20

Hydroplaning

Page 201

Chapter 21

Fire Systems

Page 207

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AIRCRAFT, TECHNICAL & GENERAL - BOOK TWO AIM The aim of this course is to give the student Commercial Pilot a sound technical knowledge of the Principles of Flight and the Operation of Aircraft. OBJECTIVE After completing the ATG module the student will have a good understanding of the following: ! basic principles of aerodynamics ! a comprehensive knowledge of the aircraft airframe including wing, cockpit and cabin windows and landing gear etc. ! piston, turbo-propeller and jet aircraft engines ! the different aircraft systems, some of which are: " Electrical system " Air-conditioning system " Fuel system " Pressurization

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AMENDMENT LIST RECORDS

Version No.

1.01

1.0

Date

Nov 2006 Jan 2008

Feb 2008

Copyright © 2012 EAA

Details

Complete re-design of these study notes. Aviation Courseware replaced all drawings with graphics drawings. Mr I. Forsyth was responsible for managing the update of the Aeronav manuals. This included the drawings and improving the text. Incorporation of Eltanin Aerospace Academy Notes

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CHAPTER 1– AIRFRAMES & SYSTEMS OBJECTIVE To achieve a thorough understanding of the basic principles of aerodynamics. CONTENTS OF THIS CHAPTER: ! The airframe and systems: ! The fuselage " Types of construction " Structural components and materials ! Cockpit and cabin windows "

Construction (laminated glass)

"

Structural limitations

"

Window heating

! Wing structure "

General construction

"

Types of wings

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AIRFRAMES – FUSELAGE AND WING CONSTRUCTION INTRODUCTION In flight aircraft undergo stresses in normal and abnormal operations. The whole airframe structure is designed to handle these stresses. The basic stresses are tension, compression, bending, torsion and shear loads. Tension – is the stress acting against another force that is trying to pull something apart. For example, engine power and propeller are pulling the airplane forward while the fuselage, wings, and tail section resist that movement because of the airflow around them. The result is a stretching effect on the airframe. Bracing wires in an aircraft are usually in tension. Compression – is a squeezing or crushing force that tries to make parts smaller. Aircraft wings are subjected to compression stresses. The ability of a material to meet compression requirements is measured in pounds per square inch (psi).

Bending – is a combination of two forces, compression and tension. During bending stress, the material on the inside of the bend is compressed and the outside material is stretched in tension.

Torsion – is a twisting force. Because aluminium is used almost exclusively for the outside, and, to a large extent, inside fabrication of parts and covering, its tensile strength (capability of being stretched) under torsion is very important. While in flight, the engine power and propeller twist the forward fuselage and the airframe is therefore subjected to variable torsional stresses during turns and other manoeuvres. Shear – stress tends to slide one piece of material over another. The aluminium skin panels of the fuselage are riveted to one another. Under flight loads shear forces try to make the rivets fails so selection of rivets with adequate shear resistance is critical.

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FUSELAGE – DESIGN AND CONSTRUCTION The fuselage structures of aircraft today can usually be divided into ! The truss structure – the longerons and subsidiary members carry the load and transmit the various stresses incurred. Construction is of wood, steel tube, aluminium tube, or other cross sectional shapes which may be bolted, welded, bonded, pinned, or riveted into a rigid assembly (Fig 1-1). ! Monocoque – The covering or skin is an FIGURE 1-1 integral structural or load carrying member. Monocoque (single shell) structure is a thin walled tube or shell which may have rings, bulkheads or formers installed within. It can carry loads effectively, particularly when the tubes are of small diameter. The stresses in the monocoque fuselage are transmitted primarily by the strength of the skin (Fig 1-2). Stringer

Curved skin covers airframe structure

Frame

FIGURE 1-2

loading. The skin is quite strong in both tension and shear and, if stiffened by other members, may be made to carry some compressive load. Since the skin of the structure must carry much of the fuselage's strength, it will be thicker in some places than at other places. In other words, it will be thicker at those points where the stress on it is the greatest (Fig 1-3).

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! Semi-monocoque – The most popular type of structure used in aircraft today. Stressed skin construction has come to be the standard for most aircraft builders. Here, internal braces as well as the skin itself carry the stress. The internal braces include longitudinal (lengthwise) members called stringers and vertical bulkhead. The metal skin exterior is riveted, or bolted and riveted, to the finished fuselage frame, with the skin carrying some of the overall

FIGURE 1-3

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WING STRUCTURE INTRODUCTION The main purpose of a wing is to give lift. The structure of a wing must be strong enough, yet still be light enough, for an aircraft to be able to generate lift. Over the years wings have developed from an internal structure where most of the loading and a skin formed the aerofoil to the present day where every part of the structure has a structural purpose. The materials used to cover wings have also changed over the years – from wood and fabric to metals such as aluminium and more recently – composites. PROPERTIES OF A WING ! Light in weight – a wing structure has to be designed to cope with torsional and bending loads and at the same time be sufficiently light in weight and yet generate enough lift so the aircraft can take off. ! Longitudinal stiffness – able to cope with bending loads. A reaction force is created by the joint of the wing to the fuselage. At this point the wing is fixed and therefore the upwards force along the length of the wing tries to bend the wing upwards. Reaction forces at the joint, compression and tension resist this bending and these forces are transferred throughout the length of the wing to stop the whole structure from bending. ! Torsional stiffness - able to cope with loads which would cause twisting such as the forces applied when using flaps and ailerons. Consider the wing to be an ‘I’ beam (Fig 1-4). If a force is applied off-centre the wing would twist. With so many variables in the types of loading to be experienced by a wing, the wing itself could never be designed so the force always acted at the torsional centre. Therefore the wing structure must be designed so that it can cope with torsional loading. Fig. 1-4 “I” Beam

GENERAL CONSTRUCTION With a few exceptions most wings are constructed of a spar or spars of different shapes and construction, ribs to provide contour or shape and stringers to support the skin. ! The spar is used to attach the wing to the fuselage. A conventional structure consists of front and rear spars with the metal skin attached to the spar flanges forming a torsion box. ! The ribs maintain the aerodynamic shape of the wing, support the spars, stringers and skin against buckling and pass loads created by the engines, undercarriage and control surfaces into the skin and spars.

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! The stringers provided that they are continuous in the spanwise direction, react to bending moment in the same way as the skin. They stiffen the skin by increasing the stress at which the skin buckles. ! The skin usually of a light but strong aluminium metal is riveted to the stringers giving the framework extra strength and rigidity. TYPES OF WINGS Braced or External Strut wings – bracing struts run from the fuselage to the point halfway along the wing. They anchor the spars in torsion and relieve them of a lot of their vertical load. This type of wing is mainly used for small high wing aircraft.

Fig. 1-5a External Strut Wings

Cantilever wings – Mainly used on high performance aircraft. No external bracing struts are utilized. The wings are usually tapered from a thin tip to a thicker root where the stresses are greatest. Fig. 1-5b Cantilever Wing

D Spar construction – The front spar is placed as near as possible to the point of maximum thickness of the wing. The skin of the leading edge is rigidly attached to it to form a D-shaped tube relieving the torsional stresses of the wing. The rear spar forms a mounting for the aileron and flaps and is connected by the skin and a light rib structure to the D-spar.

Fig. 1-5c D Bar Support

Torsion Box construction – A torsion box runs along the length of the wing providing all the torsional and longitudinal stiffness required in the wing. The torsion box is formed by the upper and Fig. 1-5d Torsion Bar lower surfaces of the skin rigidly attached to the front and rear spars in the form of a box. The skin between the spars is corrugated in order to increase the load carrying capacity and this is then covered with a thin metal sheet.

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Stressed Wing Construction A light alloy skin is riveted to the framework and is designed to stiffen the wing by taking some of the load. It produces a relatively strong wing without too large a weight penalty.

Fig. 1-5e Stressed Skin Wing

COCKPIT AND CABIN WINDOWS INTRODUCTION Since the modernization of aircraft into the jet age, cabin and cockpit windows have had to be adapted to withstand high speed and pressurization which they were not required to do in the aeroplanes of yesteryear. Several design factors and materials had to be incorporated into the manufacturing of modern cockpit and cabin windows such as strength, heat resistance, icing, durability and visibility etc. Cockpit windows are subjected to numerous stresses such as bird strikes, loading imposed by flight manoeuvres and pressurization. With advancements in technology various types of glass and synthetic glass have been created: ! Monolithic acrylic ! Laminated acrylic ! Acrylic/glass laminates ! Glass/glass laminates ! Cold heating film for heating windows ! Wire gribs

Fig. 1-6 Typical Cabin Window Construction (Pressurised)

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CONSTRUCTION Depending on the type of aircraft the windows would be constructed from especially hardened polycarbonate, laminated glass and/or acrylic materials. Cabin windows are normally constructed of single layer acrylic while various types of construction processes are used for cockpit windows. Plastic interlayers generally made of Polyvinyl Butyral (PVB) are placed between sheets of tempered glass to form a laminate. A wafer thin coating of polyurethane is vapour deposited between the layers of acrylic and glass to raise the chemical and mechanical resistance of the materials. The acrylic layer serves to prevent the glass from shattering. HEATING Windshield heating is provided by supplying power to heat up a wafer thin film of metal oxide that is vapour deposited on the inside of the outer pane. The heat generated provides protection from misting and icing. Fine heating wires such as used in cars is also incorporated into the glass to provide protection. Cockpit windows are also vapour coated with a layer of gold or silver in order to reflect solar radiation and protect the cockpit from heating up. One additional advantage of heating is that the vinyl, incorporated into the windscreen, becomes palpable and is therefore able to withstand the impact of bird strikes. LIMITATIONS Due to the extreme conditions in which aircraft operate the construction of cockpit windows and cabin windows is such that they have to be strong enough to withstand: ! temperature differences, ! surviving the impact of weather such as rain, hail, ice ! pressurization differences ! bird strikes

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CHAPTER 2 – AIRCRAFT ELECTRICAL SYSTEMS Section 1.01 CONTENTS OF THIS CHAPTER: ! Current and voltage ! Electricity and magnetism ! Aircraft generating systems " Power Supplies " Generator systems " Alternators ! Comparison of alternators and generators ! Switches

! Ignition systems " " " " " " "

Magnetos – types Magneto faults/failures Battery ignition system Ignition timing Single and dual point ignition Impulse coupling Induction vibrator

! Electrical miscellaneous

! Voltage regulators

! Abnormal operation/emergency handling

! Constant speed drive unit

! AC power supply

! Vibrating voltage regulators

! Ammeter

! Monitoring devices

! Brushless generator

" Reverse current relays " Sensor and warning lights " Ammeter ! Electrical distribution " Busbars " Terminals wiring fuses and switches " External power supplies ! Batteries " Construction " Types of batteries " Hazards and safety precautions

! Busbar ! Cabling ! Electric motors ! Engine starter motors ! Fault protection ! Frequency controller ! Light aircraft electrical circuits ! Indictor lights ! Inverters ! Load sharing ! Monitoring ! Multiple AC generator operation ! Protection unit ! Static electricity ! St. Elmo’s fire ! thermocouples

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AIRCRAFT ELECTRICAL SYSTEMS INTRODUCTION Many older aircraft had no electrical systems except for the magneto, which supplied energy for the ignition. Most contemporary aircraft have a full electrical system as well as a separate self-contained magneto or ignition system, which operates independently from the main electrical system. This means that if the aircraft electrical system is switched off, the engine will still run using power from the magneto/s. Most light aircraft use a 12v or 24v DC system consisting of the following basic components: ! alternator or generator ! battery ! master switch or battery switch ! busbar, fuses, and circuit breakers ! voltage regulator ! ammeter ! starter motor ! associated electrical wiring ! accessories. A schematic of a typical system found in General Aviation aircraft is shown in Figure 2-1. While the engine is running, an engine-driven generator or alternator provides electrical power. A battery provides the power for engine starting. It can also give a limited electrical supply for emergency use if the alternator (or generator) fails. Some aircraft have receptacles for external units to provide power for starting. This is very useful, especially for cold-weather starts. When starting engines with auxiliary power units (when the battery is dead) care should be taken that electrical energy is not forced into the dead battery. This would overheat the battery, which might explode. DEFINITION - CURRENT AND VOLTAGE ! Current is the flow of electrical charges (usually electrons) in an electrical circuit. The same effect as water flowing down a pipe. ! There are two kinds of electrical current – Direct Current produced by a DC Generator and Alternating Current produced by an Alternator. It is represented by the letter ‘I’. " When you switch on a torch the current flows from the negative terminal to the positive terminal, always in the same direction. This is called DIRECT CURRENT. " Electrical current from a standard household changes direction 60 time each second. Because the current flows first in one direction and then the other this is called ALTERNATING CURRENT. AC is used by power companies because it can be transformed to higher and lower voltages through transformers allowing them to transmit and distribute power with lower losses. ! Voltage on the other hand is a measure of pressure – Like how many pounds per square inch of air are in your tyres! It is represented by the letter ‘V’. The basic unit of electrical pressure is called the volt. The definition of a volt is that 1 volt is the amount of pressure required to force 1 amp of current to flow through 1 ohm of resistance.

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ELECTRICITY AND MAGNETISM If you move a wire through a magnetic field a small current (electricity) is created in the wire. This is called electromagnetic induction. Conversely, if you put electricity down a wire (conductor), the wire will have a magnetic field around it. The more wires you use and/or the greater the strength of the magnetic field the greater the effect becomes. These two principles are the basis for electric motors, generators and alternators. If you have one item, either movement or electricity, you can convert it into the other. To create a stronger effect you can use more turns of wire or windings (the term used for the wire in an armature)

Fig. 2-1 Typical Electrical System Schematic Copyright © 2012 EAA

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AIRCRAFT GENERATING SYSTEMS AND ASSOCIATED EQUIPMENT INTRODUCTION Generators and Alternators are the two prime means of producing electrical energy on an aircraft. They provide power for example: ! lights ! radios and other services ! a very important function is to recharge the batteries. Each converts mechanical energy into electrical energy by moving a conductor through a magnetic field. A generator produces direct current (DC) An alternator produces alternating current (AC). AC must be rectified or converted into direct current before being supplied to equipment (such as the battery) which needs DC. The following factors are required in order to generate a current of electricity by induction: ! A magnetic field ! A closed conductor ! Movement of the conductor at right angles across the magnetic field The principle on which generators and alternators depend for their operation is shown in Figure 2-2a below. A conductor, which is moved through a magnetic field, will experience an electromotive force (EMF). The EMF will cause a current to flow if the circuit is closed as shown. Figure 2.2b illustrates the electrical rule used to find the direction of flow on induced current in a generator. The ‘Right Hand Generator Rule’ says “with the index finger of the right hand pointing in the direction of the magnetic field (north to south) and the thumb indicating the direction of the movement of the conductor across the magnetic field, the second finger will point in the direction of the flow induced current.

Fig. 2-2a and b Principle underlying operation of the alternator and generator

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POWER SUPPLIES THE GENERATOR The generator/alternator is the primary source of aircraft electrical power. The battery is the secondary source. During normal engine operation the engine-driven alternator (or generator) supplies power for all aircraft electrical equipment. The battery is kept fully charged for use as an auxiliary source when the generator is not operating ("off line") or the engine is not developing power. The number of alternators or generators in a system depends on power requirements. Single-engine aircraft usually have one alternator or generator. Multi-engine aircraft may have two or more, depending on requirements and the number of engines. If AC current is required (e.g. for AC driven instruments) an inverter can be incorporated. An inverter converts DC to AC. This method is used in the Beechcraft King Air. A generator (which doubles as a starter motor) is used to produce electrical energy. If the main power generator is AC (an alternator), a rectifier must be incorporated to convert AC to DC for battery charging and other DC supplies. If the alternator voltage is higher than the battery or electrical system voltage, a "stepdown transformer" must also be incorporated. The combination of transformer and rectifier is known as a transformer-rectifier unit (TRU). REMEMBER:

RAD = RECTIFIER CONVERTS AC TO DC IDA = INVERTS DC TO AC

REQUIREMENTS OF A GENERATOR SYSTEM ! A generator or alternator must be very reliable over a wide range of climatic conditions. ! It must be able to maintain a continuous power supply for all power requirements. ! The supply must kept be at a specific voltage (phase also, in the case of an alternator) regardless of engine speed and loads on the system. DC GENERATOR A direct current generator supplies current that flows in one direction – from positive to negative. The purpose of the generator is to change mechanical energy into electrical energy which will supply power to operate all the electrical devices and keep the battery fully charged. The mechanical energy required to rotate the generator, driving it by pulleys, belts or gears is supplied by the engine. The generator operates on the principle of Electromagnetic Induction – electrons are made to move by magnetism. (See figure 2.1) In a DC generator the three factors necessary to produce a current of electricity by induction are present: ! It has a closed conductor called an armature ! A magnetic field ! Movement of the conductor or armature at right angles across the magnetic field In a DC generator the induced voltage which induces the current in the armature coils, alternates. The armature coils are connected to the commutator which is actually just an

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extension of the ends of the armature winding. Carbon brushes in contact with the commutator change the side of the coil to which they are connected every half revolution. This keeps the current flowing in the same direction and so supplies DC to the external field. Figure 2-3a shows DC generation and the current output from a typical DC generator.

Fig. 2-3a Current Output from DC Generator

MAIN COMPONENTS OF A GENERATOR

Field Coils

Drive end frame Field Frame Armature windings

Commutator end frame

Brushes Through bolts

Commutator

Laminated armature core

Field Pole Shoes

Brush holders and springs

Drive pulley and fan Armature

Fig. 2-3b Components of the DC Generator

FIELD FRAME – The main body of the generator. It supports the drive end frame and the commutator end frame and retains the residual magnetism which supplies the original magnetic field for the armature to cut in order to generate a current of electricity. COMMUTATOR END FRAME – Supports the commutator end of the armature, and provides a position for fastening the generator to the engine DRIVE END FRAME – Supports the drive end of the armature, and provides the other point of fastening to the engine

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DRIVE PULLEY AND FAN – Used to drive and cool the generator ARMATURE – is made up of armature windings, laminated iron core, commutator, armature shaft and insulators. The armature shaft carries and drives the armature parts and the insulators insulate the armature windings from the laminated iron core of the armature. COMMUTATOR – An electromagnetic device rectifies the AC within the generator into DC BRUSHES – usually made of carbon which is a good conductor with low friction necessary when coming into contact with the commutator. It transfers current to and from the commutator and assists the commutator to change the alternating current to direct current. NOISE SUPPRESSOR – The name describes its function. In order to prevent interference to radio/radar equipment the suppressor must be situated as close as possible to the generator.

Fig. 2-3c Noise Suppressor

ALTERNATORS Unlike a generator an alternator has no moving parts and as a result it is not only very reliable but also comparatively inexpensive to build and repair. An alternator can be considered as an AC generator. It produces alternating current (AC). AC current changes direction (alternates) in a regular wave-like manner. In AC two current reversals compromise a cycle. The frequency is the number of cycles per second. The principle of AC current and a schematic layout of an alternator are shown in Figures 2-4 and 2-5 below.

Fig. 2-4 Current produced by a simple alternator

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Fig. 2-5 Schematic of a typical AC alternator

! On an alternator a magnetic field is spun inside of windings of wire called a stator in order to generate electricity. ! Because the wires are fixed they can easily be connected directly to their outputs without the need for sliding contacts to carry the relatively high output current. ! The magnetic field is still generated via electro magnets mounted on a rotor and the relatively small field current that powers them is supplied to the rotor by two small brushes that each ride on a separate and continuous slip ring. ! Because these slip rings and the fact that the relatively heavy windings are fixed instead of rotating allows the alternator to be spun to much higher speeds. As a result maximum output is reached much sooner even at engine idle speeds. THE STATOR Three separate windings of wire in the stator are all set so that the AC current generated is slightly out of phase in each one. The stationary stator of the alternator takes the place of the rotating armature used in the DC generator. It has a laminated iron core and the frame is laminated to prevent eddy currents which could create heat which could burn the insulation off the stator windings. A DIODE ! In order to rectify the AC current into DC current in an alternator a diode is built into the alternator.

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! A diode is the simplest possible semiconductor device. It is a ‘solid state’ device that allows current to flow in one direction but not the other in much the same way as a turnstile at a stadium lets people go through in only one direction. ! It has no moveable parts. ! It relies on the different electrical properties of the materials it is made of to act as a one-way valve for current. ! A smooth and stable DC output is obtained by arranging diodes so that current from each of the three stator wires is only allowed to pass in one direction and by connecting the three outputs together. COMPARISON OF ALTERNATORS AND GENERATORS Several fundamental differences exist between generators and alternators. ! Unlike alternators generators can seldom produce sufficient electrical current at low engine rpm to operate the entire electrical system. This means that when engine rpm are low, power must be drawn from the battery – which has a limited life. This is particularly important during prolonged night ground operations. However, the alternator is capable of producing sufficient power for the increases in power requirements even when idling. ! Another advantage of the alternator is that the output is constant at most engine speeds. ! Alternators are lighter, cheaper to maintain and less prone to overload when loads are heavy. SWITCHES ! A master switch controls electric power for all aircraft equipment. ! The master switch activates all electrical circuits except the ignition system. ! Some aircraft have a battery switch, which is used in a similar fashion to the master switch. Often an alternator switch is fitted. This can isolate the alternator from the main circuit in the event of alternator failure. When the alternator switch is off, the battery is the only source of power. VOLTAGE REGULATORS Voltage regulators continually adjust the field current so that generator output voltage remains constant under all loads. Ohm's law states that the current in a closed circuit is directly proportional to the circuit voltage and is inversely proportional to the resistance of the circuit (V = IR).

Fig. 2-6 Voltage Regulator

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Generator output is directly proportional to generator voltage and inversely proportional to the circuit resistance. VOLTAGE OUTPUT The voltage output of a generator depends on three factors: ! the number of conductors, or armature windings ! the speed (rate) at which the magnetic lines of force are cut ! the strength of the magnetic field. Output voltage is regulated by increasing or decreasing the number of lines of force. This is accomplished by varying current strength in the generator field windings by changing the number of ampere turns, so altering the resistance. The current strength of the field windings can altered in several ways: ! connecting a variable resistor (rheostat) in series with the shunt circuit of a shuntwound generator. (In a series-wound generator the rheostat is connected in parallel with the field) ! full distortion, usually called third-brush regulation ! installing vibrating voltage regulators ! incorporating a carbon-pile voltage regulator. CONSTANT SPEED DRIVE UNIT (CSD or CSU or CSDU) ! Many aircraft services depend on stabilised frequency. ! This is only possible when the alternator rotates at a constant speed. ! Most aircraft have a CSD fitted, which regulates alternator rpm through a variable ration hydraulic drive mechanism. ! The CSD keeps the generator rpm constant independently of engine rpm and so ensures a constant supply frequency. ! A CSD is not required on aircraft fitted with constant speed engines. VIBRATING VOLTAGE REGULATORS ! Vibrating voltage regulators are used in light aircraft in conjunction with current regulators. ! The current and voltage from DC generators are regulated through interconnected coils. ! These are connected, so that only one of the regulators can operate at a time. MONITORING DEVICES REVERSE CURRENT RELAYS (RCR's) ! An RCR is an integral part of the voltage regulation system and prevents a storage battery from "motorising" a generator by reversing the current flow. ! When battery output is greater than generator output (especially at low generator operating speeds – low engine rpm), the RCR points open, thereby breaking the circuit.

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! This stops battery current flowing in the reverse direction to the generator and turning it into a motor. ! If the RCR fails, battery output will travel back into the generator field. ! This type of malfunction can be corrected by "flashing" the generator field. "Flashing" the field is passing a very brief current from the positive terminal to the negative terminal. SENSOR AND WARNING LIGHT ! In the event of over-voltage condition occurs, the over-voltage sensor will automatically remove alternator field current and shut down the alternator. ! A red warning light will then come on indicating that the alternator is not operating and that the battery is supplying all electrical power. AMMETER An ammeter is an instrument used to monitor the performance of the aircraft electrical system. It indicates the flow of current, in amperes, from the alternator to the battery or from the battery to the aircraft electrical system. When the engine is operating and the master switch is turned on the ammeter indicates the charging rate applied to the battery. In the event the alternator is not functioning or the electrical load exceeds the output of the alternator, the ammeter indicates battery discharge rate. ELECTRICAL DISTRIBUTION INTRODUCTION A busbar is the central distribution point for the main electrical system. It routes power from the generating source (battery/alternator/generator) to the electrical equipment. It is the common point from which power can be distributed throughout the system. Electricity is distributed in the aircraft to various services from busbars. Each generator supplies one or more busbars, from where current is supplied to various aircraft services. Each busbar serves various circuits, depending on the type of circuit and the number of generators. The busbar in DC circuits is usually a copper strip; in AC circuits it is simply three wires, one for each phase. Busbars are classified according to their importance. A typical aircraft system might have the following busbars: ! ! ! !

vital services busbar essential services busbar main busbar synchronising busbar (AC only).

BUSBARS VITAL SERVICE BUSBARS The vital services busbar is a DC busbar. It is supplied with power at all times direct from the batteries or via the secondary power distribution network. All vital services (those concerned with aircraft safety) are connected to this busbar. It is normally isolated from the main distribution system. Vital services include:

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fire extinguisher circuits crash relay switches and battery isolation engine relighting emergency power supplies (RAT/APUs etc) RAT and APU control (AC only) 28v DC control feathering controls powered flying controls (if no manual or hydraulic reversion)

ESSENTIAL SERVICES BUSBAR The essential services busbar supplies power to all services, which enable the aircraft to keep flying after an emergency (as well as under normal ops). These services may include: ! ! ! ! !

power flying control units (fly by wire – if reversion option available) radios and intercom. flight instruments fire detection, warning and extinguisher operation heaters and wipers.

Different manufacturers use different names for busbars. Where no provision is made for both essential and vital services busbars, the term "essential services" means "vital services". This busbar is normally DC and is supplied either from the battery through manual switching or by emergency power plants through appropriate transformers and rectifiers. Some aircraft have AC and DC essential services busbars. MAIN BUSBAR The main busbar is also called a load busbar or non-essential busbar. It may be either an AC or DC busbar. The generator supply is fed to the main busbar which is part of the main electrical system. Any secondary system is supplied from the main busbar via transformers and inverters etc. SYNCHRONISING BUSBAR The synchronising busbar is an AC busbar. It is used on aircraft fitted with alternators. It can be considered as the AC equivalent of the DC essential services busbar. It is always "on line". When a generator is on line to the synchronising busbar power is supplied to the essential services busbar via transformer rectifier units. In the event of total alternator failure, the RAT or APU can supply the synchronising busbar, after appropriate manual or automatic selection. CIRCUIT DIAGRAM Aircraft with complex electrical systems have circuit diagrams etched on the electrical control and indicator panels. This enables the Flight Engineer or pilot to monitor the operation of the various circuits, in particular when any fault-finding is required. Magnetic indicators show whenever a particular circuit has been made or broken and voltmeter/ammeters and frequency meters are included in the diagram. These are sometimes called mimic diagrams.

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TERMINAL, WIRING, FUSES AND SWITCHES TERMINALS Terminals facilitate the connection of electrical wire to junction boxes, terminal strips, or items of equipment. The tensile strength of the wire and the resistance of the terminal joint should be negligible compared with the resistance of the "cable run". Terminals can be soldered or solderless. WIRING – ELECTRICAL CABLE AND FLAMMABLE FLUID LINES When electric cables and flammable-fluid lines are installed along the same route, or run, they must be kept separate and the cable/s must always be above the fluid line/s. The following terminology is often used in conjunction with aircraft wiring systems: ! OPEN WIRING - Electric wire is often installed in aircraft without being carried in a conduit to facilitate maintenance and reduce weight. ! "SPAGHETTI" - Spaghetti is a soft slender tube of insulating material around wire or cabling. SWITCHES ELECTRICAL SWITCHES Switches used in aircraft are: ! snap-action switches ! toggle switches ! relay switches ! rotary selector switches ! push-button switches ! knife switches. Snap-action switches are best for rapid opening and closing of contacts, irrespective of the speed of the operating toggle or plunger. This minimises arcing. Switches are designated by the number of poles, throws and positions: ! The pole of a switch is its movable blade or contactor and the number of poles is the same as the number of terminals through which current can enter or leave the switch. ! The throw of a switch indicates the number of circuits which each pole can complete through the switch. ! The number of positions a switch has is the number of places at which the operating device (toggle, plunger, etc) will come to rest. Single-pole, single-throw (SPST) Switches are switches through which only one circuit can run. single-pole, double-throw (SPDT) switches are single-pole switches through which two circuits can be completed

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double-pole, single-throw (DPST) switches are Switches with two poles through each of which one circuit can be completed. double-pole, double-throw (DPDT) switches are those with two poles through each of which two circuits can be completed. A toggle switch that will come to rest at either of two positions (to open the circuit in one position and complete it in another) is a two-position switch. A toggle switch that is springloaded to the "off" position and must be held in the "on" position to complete the circuit is a single-position switch. Micro switches are switches that can open or close a circuit with a very small movement of the tripping device. Micro switches are usually the push-button type and are used as limit switches for automatic control of landing gears, actuator motors, etc.

FUSES AND CIRCUIT BREAKERS A circuit breaker is a protective device which can close or open a circuit. Circuit breakers are often used in place of fuses to protect a circuit as, unlike fuses, they can be reset. Some circuit breakers must be reset by hand; others are reset automatically. C/Bs are rated in amps and can be used as, or incorporated in, switches (Fig 2-7).

Fig. 2-7

MAGNETIC CIRCUIT BREAKER When a current flows in a circuit at a higher rate than desired, a magnetic circuit breaker generates an electromagnetic force through induction. This actuates a small solenoid which opens the circuit breaker. The circuit is interrupted and protected. Thus the circuit is being protected against the restraining force of a sprung detent. CIRCUIT PROTECTORS (THERMAL OVERLOAD SWITCH) The circuit protector is an automatic protective device that opens a circuit whenever the temperature of the protected unit rises above a certain value. It has two positions, automatic "ON" and automatic "OFF". The switch is often found in electric motor circuits. A bimetallic strip opens the switch when the temperature rises and re-makes the switch when it cools. FUSE TYPES CARTRIDGE FUSES The cartridge type of fuse is a wire strand passing between two terminals enclosed in a glass tube. The fuse forms part of the circuit and the wire burns out if the circuit overloads. This fuse usually fits into a cap that is screwed into the fuse mounting. Most fuse strips are made of an alloy of tin and bismuth (Fig2-8).

Fig. 2-8

CURRENT LIMITERS Current limiters normally have a copper fuse element which can withstand a higher than normal temperature for a short period. It melts and opens the current when a sustained overcurrent condition (as opposed to a transient overcurrent) occurs. The melting point is much higher than regular fuses. They can withstand a considerable overload for a short period before breaking the circuit.

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EXTERNAL POWER SUPPLIES RAM AIR TURBINE (RAT) The RAT is a temporary power supply which gives the crew time to rectify the generator faults. If this is not possible, an alternative source of power is used such as the APU (or AAPP). The RAT is an AC generator driven by ram air. Pulling a handle releases a turbine (fan) into the air stream, which drives it by ram air. It can supply a limited 200v, 400 Hz current in the event of total generator failure in flight. The RAT output is fed direct to the synchronising busbar (Essential AC busbar). The advantage of the RAT is that it can supply full output within two seconds. Usually the RAT will automatically shed the non-essential electrical loads when selected. A high IAS is required to maintain full output. This makes the RAT unsuitable for use at low airspeeds or during an emergency landing. AUXILIARY POWER UNIT (APU) The APU is a gas turbine, which drives an AC generator. It provides the synchronising busbar with a 200v, 400 Hz AC supply for use in emergency or on the ground when no external power unit is available. Gas turbines are difficult to start at high altitude so the APU is normally lit at lower altitudes, while utilising the supply from the RAT. GROUND POWER UNIT (GPU) The GPU provides power for aircraft servicing, for aircraft systems and engine starting. Protection circuits ensure that the GPU is not paralleled with either the APU or the generators. As each generator is brought on line, the GPU supply is automatically isolated from that generator busbar.

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ENERGY STORAGE - BATTERIES INTRODUCTION During normal aircraft operation, electrical power is supplied by the alternator or generator. However when power is required to: ! start the engines or ! cope with the failure of the generating system ! act as a stabilizer in the charging circuit some stored form of energy is required. The most common storage system is to use a storage battery. The most popular types in use are the lead-acid and the Nickel Cadmium batteries. The light aircraft electrical system normally uses a 12 volt lead-acid battery and large or modern aircraft use 24 volt Nickel Cadmium (Nicad) batteries. These two types of batteries are very different and cannot be interchanged. The tools, storage facilities and maintenance methods in each case are totally different. The battery is continually charged by the generator/alternator. The charging rate is controlled by a voltage regulator, which stabilises output from the generator/alternator. The generator/alternator voltage output is usually slightly higher than the battery voltage. For example, a 12v-battery system would be fed by a generator/alternator system of approximately 14v. This voltage difference keeps the battery charged. LEAD-ACID BATTERIES Most light aircraft have a Lead-acid battery. This battery creates an electrical current (amps) by a chemical reaction between lead plates immersed in weak sulphuric acid that acts as an electrolyte. CONSTRUCTION OF THE BATTERY The lead-acid battery consists of several separate cells, with each cell containing a number of lead plates. ! The negative plate is made of pure lead (Pb) (in a sponge form) and ! The positive plate is made of lead peroxide (PbO2). ! The plates are separated by porous insulators which keep the plates from touching one another while allowing free circulation of the electrolyte. ! The number of negative plates in a cell is always uneven. This ensures that there is always a negative plate either side of a positive plate.

Separators

Positive Plate

Negative Plate Fig. 2-9 A layout of the cells

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! All similar plates are connected to a common terminal. ! The plates and separators are immersed in an electrolytic solution of dilute sulphuric acid (H2SO4). The battery is housed in its own case in order to prevent corrosion from any Lead connector spillage of the acid. This case is made of hard rubber and an element is placed in each cell. Hard rubber covers are used on each cell and lead cell connectors connect the elements in series. A twelvevolt battery has six separate cells – two Six Cells volts per cell. A bituminous sealing Cell compound is poured around the covers and over the connectors after the cell Fig. 2-10 Battery Construction covers and cell connectors are in place in order to prevent the electrolyte from leaking. When the battery is charging a vent cap on each cell cover allows the hydrogen and oxygen to escape. DC current charges lead-acid batteries. The current causes chemical changes in the plates and electrical energy is converted into chemical energy. During the process, lead sulphate is removed from both the plates and the sulphuric acid content of the electrolyte is again increased. Vent Cap

Hard covering

When the battery is in use, the stored chemical energy is converted to electrical energy and current flows when the circuit is made. The rated capacity of a battery varies with the number, area, and thickness of the plates in each cell. Batteries are rated in ampere-hours, which indicate the number of amps the battery can deliver over a period of hours. ! A 45 ampere-hour battery for example, can deliver 1 amp for 45 hours or ! 5 amps for 9 hours. ! The voltage of each cell is 2v (2.2v when charged). A 12v battery has 6 cells and a ! 24v battery has 12 cells. INDICATIONS OF BATTERY STATE ! Excessive discharging of a lead acid battery is detrimental as it leads to heavy sulphation (lead sulphate forms on the outside of the plates as a result of the discharge chemical reaction). ! This deposit raises the internal resistance so that excessive heat is generated during rapid charging or further discharge. ! This in turn causes shedding of active material, which can short circuit the plates and result in internal discharge. ! In addition, the loss of active material reduces the capacity of the battery. The drain on a battery when starting large engines is very heavy. It is better to use a ground power unit. ! Batteries lose water through evaporation and in the charging process when water is split into its hydrogen and oxygen components. As a result, distilled water must be Copyright © 2012 EAA

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added to the electrolyte (battery fluid) to keep the dilute sulphuric acid at the correct concentration. Batteries should be kept in a warm place if temperatures are low. A discharged battery will freeze at -15°C, but will not freeze until around -65°C when fully charged. ! The gases liberated when a battery is being charged (or is discharging) require adequate ventilation to reduce the fire hazard and to ensure that harmful fumes will not infiltrate the aircraft. ! The electrolyte level should be checked periodically to ensure that the plates are covered. If necessary, the battery should be topped up with distilled water. A battery with inadequate electrolyte will not hold its normal capacity and the ammeter will indicate an abnormally high charge rate in flight. Never use non-distilled water as impurities in the water can react with both the electrolyte and the plates which will reduce the battery capacity. Cell Cover

Terminal Post Vent Cap Cell Connector

Hard Rubber Case

Separator

Sediment Chamber Plate Fig. 2-11 A typical lead acid

MEASUREMENT OF SPECIFIC GRAVITY Battery charge condition can be determined by a hydrometer. This measures the specific gravity (SG) of the electrolyte. Specific gravity of a fluid is the ratio of the density of a fluid relative to that of water ! Water is assumed to have a density of 1 kg per litre. ! The specific gravity of a fully charged battery is around 1.3. ! A flat battery will have an SG of between 1.2 and 1.24. Testing the specific gravity of the electrolyte in the battery cells tells the state of charge of the battery. ! The more sulphuric acid there is in the electrolyte the higher the charge in the battery. ! The less sulphuric acid there is in the electrolyte solution the lower is the charge in the battery.

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In order to take a specific gravity test hold the hydrometer steady in a vertical position so that the float in the hydrometer will move freely and not touch the sides. Draw enough electrolyte into the hydrometer to float the float. Do not allow the top of the float to touch the top of the float chamber as this will give an inaccurate reading. The reading should be taken at eye level.

Less buoyant, less charge

More buoyant, more charge

Fig. 2-12 Hydrometer

In a fully charged condition the positive plate is lead peroxide, the negative plate active material is sponge lead and the electrolyte is a solution of sulphuric acid and water with a specific gravity of 1.260 to 1.280. The electrolyte solution should be heavy, dense and buoyant and it should float the hydrometer float high, showing a specific gravity of 1.280. In a fully discharged battery the float in the hydrometer sinks deep into the electrolyte solution indicating a specific gravity of approx. 1.150. (As the battery discharges the sulphate in the electrolyte combines with the active materials in the positive and negative plates and the active materials in both plates gradually change to lead sulphate (see above).The sulphate then becomes less dense and consequently less buoyant. Continued use of the battery will result in almost all of the sulphate leaving the electrolyte and combining with the active materials in the plates. This reduces the electrolyte solution to water which is not dense and buoyant enough to support the hydrometer float at a high level.) The following table indicates the specific gravity of the solution at its ability to crank the engine at 80º(F) From 1.260 to 1.280 From 1.230 to 1.250 From 1.200 to 1220 From 1.170 to 1.190 From 1.140 to 1.160 From 1.110 to 1.130

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100% charged 75% charged 50% charged 25% charged Very little useful capacity Discharged

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NICKEL CADMIUM (NICAD) BATTERIES

EAA-CPL-MAN-010v1

Each cell produces around 1 Volt

A nickel cadmium battery has positive plates made of nickel hydroxide deposited on a fine nickel mesh screen. Negative plates are made of cadmium hydroxide and the electrolyte is dilute potassium hydroxide. All positive plates are connected together and all negative plates are connected together. Cells can be The specific gravity remains between 1.24 and replaced Metal individually 1.3 at room temperature. The plates are housed in a nickel-plated steel Fig. 2-13 Construction of a NiCad Battery container. A vent is incorporated in the case so gases can escape and distilled water can be added as required. Each cell produces a potential of about 1.9v when fully charged which falls to 0.9v when the cell is fully discharged. When the battery is charged, the negative plates lose oxygen leaving pure cadmium while the positive plates become more oxidised. ADVANTAGES OF THE NICAD BATTERY The NiCad battery has several advantages over the lead-acid battery: ! low maintenance, each cell is an individually unit ! the battery can remain in a low charge state without damage ! the freezing point is below operating temperatures ! it can tolerate a very high charge and discharge rate ! the battery is stronger and reliability is good ! the electrolyte undergoes little change ! it does not produce harmful fumes. BATTERY HAZARDS AND SAFETY PRECAUTIONS WATER During charging the water in the battery decomposes and leaves the cells as hydrogen and oxygen gases. The water must be replaced with distilled water. GASSES Charging batteries emit a mixture of hydrogen and oxygen. They are highly inflammable and will explode if a flame is brought to close to them. Lighted matches and torches should be kept away from charging batteries. TERMINALS Make sure that battery terminals are switched to the ‘OFF’ position before removing them. Removing terminals with the heater switch ‘ON’ will cause a spark at the battery terminal which could ignite the hydrogen-oxygen gas mixture in the cells and thus cause an explosion. 34

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REMOVAL OF BATTERY Remove the ground cable first and connect it last when removing or installing a battery as the use of battery pliers and wrenches may cause short circuits resulting in flashes which could cause an explosion. BATTERY ACID Battery acid is highly corrosive and great care must be taken to avoid getting it on the face or in the eyes. In the event of an accident wash the affected areas with large quantities of cold water. The acid will also eat holes in clothing so when carrying the battery use a battery strap and keep it level and well away from clothing.

IGNITION SYSTEMS INTRODUCTION - IGNITION Reciprocating engines require a very high voltage spark inside each cylinder for correct and complete combustion. The ignition system provides this high voltage spark to the spark plug. The high voltage is usually provided by a magneto. Two types of magneto ignition systems are in use :! high tension (HT) and ! low-tension (LT). Modern aircraft have double magneto ignition systems for safety and better engine performance. The double system may be two separate magnetos or a single "dual magneto" (two magnetos in one housing).

Fig. 2-14 Typical Magneto and Spark Plug Wiring Installation

Aircraft engines have two spark plugs for each cylinder. Battery ignition systems with coils (similar to motor car systems) are sometimes found on older aircraft and are reappearing in newly certified engines. A magneto system is superior to a battery system because the spark remains at a consistently high voltage.

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MAGNETOS A magneto is a combination of AC generator (permanent magnet type) and autotransformer. It uses a permanent magnet as the source of energy. Through gearing, the magneto rotates as the engine turns and a current is produced by induction. The resulting voltage is high but it is further boosted by a capacitor which is connected to the breaker points. The voltage is increased through various wiring and switching circuits and is fed to a distributor (which is often an integral part of the magneto system). The distributor feeds the high voltage current to the spark plugs in sequence. ! C – Coil Shaft ! E – Pole Shoe ! D – Magnets Magnetos developing less than 100v are classified as low tension (LT), those above 100v as high tension (HT). As a general rule, light aircraft use HT systems and larger aircraft use LT systems. Magnetos produce ionised gas as a byproduct. Holes allow cooling, ventilation and the products of ionisation to vent to atmosphere.

Fig. 2-15 Schematic Detail of Magneto

! The H.T. magneto system is divided into three circuits, the magnetic circuit, the primary electrical circuit and the secondary electrical circuit. ! The L.T. magnetos give less trouble and are more efficient than the H.T. magnetos. Because of the lower voltage there is less electrical leakage and they are lighter due to the use of lightweight leads from the magneto to the plugs. ! HT MAGNETO AND LT MAGNETO SYSTEMS The L.T. magneto system operates on the same principle as the H.T. magneto as far as the principle current flow is induced. The principal difference between an HT system and an LT system is the manner in which current from the magneto is distributed and raised to a higher voltage. HT ignition systems are often used on light aircraft. The system is composed of two magnetos and two distributors. The distributor is usually an integral part of the magneto. The switching sequence from the distributor current sequence is predetermined according to the firing order. The distributor consists of the rotor, and the distributor block. Embedded in the distributor block are terminals made from a non- conducting material. Each terminal is connected to the relevant spark plug. Long wires carry the HT electrical impulses in sequence to the spark plugs. A hot spark is produced when the plugs receive a high voltage impulse from the magnetos, hence the term "impulse-type magnetos".

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EAA-CPL-MAN-010v1 to

Secondary coil

Primary Coil

Distributor

Fig. 2-16 High Tension Ignition System

High altitude aircraft often have an LT system. In an LT system, the magneto does not have a "secondary" coil. The output voltage is kept low until reaching a secondary coil near the plug. This means that only a very short HT lead runs from the secondary coil to the plug. The system greatly reduces insulation losses and breakdowns, as well as flashovers which are the greatest source of malfunctions found when using HT systems above 25 000 ft.

Low voltage to the distributor

Distributor

Transformer – voltage stepped up

Fig. 2-17 Low Tension Ignition System

MAGNETO FAULTS/FAILURE Even today the vast majority of aircraft with piston engines use magnetos as their sole ignition source. In spite of over 100 years of experience with magnetos, extensive certifications of aircraft magnetos and quality control requirements by the aviation authorities, they still fail or require maintenance, more often than any other part of an aircraft engine. Copyright © 2012 EAA

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The keys to improved flight efficiency below all reduce the reliability of the magneto and especially its distributor. ! higher compression ratios or turbo chargers, ! reduced pumping losses, ! flight at high altitude, ! larger spark plug gaps and ! higher ignition voltages ! High Tension Lead Failure A common fault of magnetos is the break down of the high tension leads resulting in the failure of that particular lead to the spark plug. The high tension lead could also chafe against an engine part resulting in it shorting out. This will result in a loss of power. Primary Coil Failure The magnetos rely on a current passing through the primary coil to be activated. Any breakdown or fault in the primary coil will result in the magneto failing. Condenser failure The condenser is an integral part of the circuitry and should the condenser fail or the points either stick or not close will also result in the failure of the magnetos. BATTERY IGNITION SYSTEM The battery ignition system uses a battery instead of a magneto as the source of electrical energy. The ignition system is in fact part of the ordinary electrical system. The voltage is raised by a conventional ignition (induction) coil. The HT current then goes to a distributor for proper timing and distribution to the engine spark plugs.

Distributor

Ignition Coil

Primary Battery

Coil

Secondary Engine

Coil Mechanical Linkage

Breaker contact Points (Closed)

Cam

Breaker

contact

Points (Open) Fig. 2-18 Battery Ignition System

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IGNITION TIMING Ignition timing is important for efficient engine operation. If the spark is not produced at the optimum moment, incomplete combustion will take place. The time lapse between ignition and full combustion is affected by several factors : ! the engine (crankshaft) rpm ! the turbulence of the mixture in the combustion chamber ! the physical location of the spark plug ! the fuel quality ! the fuel/air ratio. Of these, the fuel/air ratio and rpm are the most important factors. SINGLE POINT AND DUAL POINT IGNITION Most aircraft engines have duplicated ignition systems (dual ignition). Dual ignition affords added safety and improved combustion.

Fig. 2-19 A Dual Point Ignition System

Dual ignition systems operate either with synchronised or staggered timing patterns. ! Staggered ignition timing is when the exhaust-side spark plug fires 4° – 8° before the intake-side spark plug. ! Synchronised timing is when both spark plugs fire at the same time. During normal combustion, the fuel-air mixture nearest the exhaust valve becomes contaminated with exhaust-gas residues. It burns more slowly than the mixture on the intake-valve side of the chamber.

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In staggered timing, the exhaust-side spark plug is timed to fire a few thousandths of a second before the other spark plug. This earlier start to combustion on the exhaust side compensates for the slower burning rate of the contaminated mixture. Faster burning of the "clean" mixture (intake side) approaches the ideal burning pattern. Both flame fronts meet in the centre of the combustion chamber. This type of combustion gives the even expansion which pushes the piston back with great power. IMPULSE COUPLING A magneto needs to run at a minimum of 300 rpm in order to function. This is known as the "coming-in speed". During engine start, the magneto speed is usually less than the coming-in speed. A booster magneto, or induction vibrator coil is incorporated which receives low voltage from the battery and steps up the voltage sufficiently to start the engine by boosting the spark produced by the magneto. Some magnetos incorporate an impulse coupling for engine starting. The impulse coupling does the same job as the booster magneto, or induction vibrator coil. When the magneto reaches a predetermined speed the impulse coupling cuts out and the magneto continues to provide the necessary current. INDUCTION VIBRATOR An induction vibrator supplies an interrupted low voltage to the magneto primary coil. This induces a high voltage in the secondary coil of the magneto. The high voltage produced in the magneto secondary coil is distributed to the spark plugs in the same way as the magneto spark.

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ELECTRICAL – MISCELLANEOUS ABNORMAL OPERATION – EMERGENCY HANDLING Emergency drills for generator malfunctions are detailed in the appropriate Aircrew Manual and Flight Reference Cards. On aircraft with only one generator, the drill is: ! to reduce electrical load to a minimum to conserve battery life. ! The generator should be brought back on line if possible. ! Load shedding should be carried out according to laid down procedures. ! Battery life will be limited and a safe landing should be made as soon as possible. ! Load shedding may include switching fuel pumps off so engine limitations must be observed. On aircraft with more than one generator the emergency drill will depend largely on whether alternative/emergency power is available. ! Where such an alternative power source is provided, this will be used and will support essential electrical loads if speeds and heights are flown as recommended. ! Where no emergency power supply is provided, the drill will be as for a single generator aircraft. AC POWER SUPPLY A rectifier is incorporated in the AC generator (alternator) to convert AC to DC for battery charging and other DC needs. When the alternator voltage is greater than battery or electrical system voltage, a step-down transformer must also be incorporated. The combination of transformer and rectifier is called a TR Unit or a Transformer-Rectifier Unit (TRU). AMMETER Electric current is measured in amperes (amps) by an ammeter. The ammeter displays the number of amps flowing in the circuit and also indicates if the battery is receiving a charge. Most ammeters have a zero datum in the upper centre of the dial with a positive indication to the right, and negative to the left. A positive value indicates that the battery is being charged. After drawing power from the battery for starting, the ammeter needle will indicate a positive value for approximately 30 minutes while the battery is re-charged. This gives a good indication of the drain on the battery during start. If the needle indicates a negative value, it shows that the output of the generator/alternator is inadequate; that is, current is being drawn from the battery to supply the electrical system. This could be caused by a defective alternator/generator or by an overload in the system, or both. Full scale discharge or rapid fluctuation of the needle usually means a serious generator/alternator malfunction. When this occurs the generator/alternator should be tripped out of the system and battery power conserved by reducing electrical loads. Not all aircraft have ammeters. Some have a generator discharge light. When this is lit, it indicates a system discharge because of generator/alternator malfunction. Copyright © 2012 EAA

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BRUSHLESS GENERATOR DC current starts the motor of the brushless generator and excites the armature winding. This excites the 3-phase field winding. AC is produced as the armature rotates. The effect is similar to rotating a small permanent magnet inside a fixed field winding and producing AC. The windings are constructed to give 3-phase power. BUSBAR Unlike the DC busbar (which is a copper strip) the AC busbar is composed of three separate wires. Each carries one of three phases. By convention these phases are labelled red, yellow and blue. The yellow phase is the reference phase for voltage and frequency. CABLING Moisture from the heating system in a pressurised cabin condenses on the cabin walls and can accumulate underneath the cabin floor. If it condenses on electrical connections and terminals or in cable connectors etc it will rust and cause high resistance. Aircraft wiring is run in spaced flat cables along the skin to allow better heat dissipation. Sometimes these cables are hard to check but fault-locating indicators (lights, circuit breakers etc) built into the electrical systems serve as guides to trouble areas. ELECTRIC MOTORS An electric motor is an arrangement of coils and magnets, which converts electric current into mechanical power. Electrical energy turns the motor armature. This rotation produces mechanical energy. Electrical power is used for engine start, landing gear, flaps, retractable landing gear, lights, electric propellers etc. Electric motors operate on the reverse principle of generators. Generators are rotated by mechanical force and produce electrical energy. Motors produce mechanical energy from electrical power. Electric power is supplied to the motor terminals, which in turn supply the commutator and a magnetic field is produced which makes the armature rotate. Gears to the service needing it transmit the physical energy produced. ENGINE STARTER MOTORS The series-wound DC electric motor can produce a high starting torque and good acceleration at relatively low rpm and high amperage. The motor speed varies according to the load need. For example, with a heavy load it will run at low speed. Conversely with a light load it will run at high speed. This enables it to be started on full load, making it very effective for engine starting. The starter motor is generally referred to as a series-wound motor. As the motor rpm increase, the engine momentum increases and the need for high torque decreases.

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FAULT PROTECTION Generator circuits need to be protected against certain faults: •

LINE-TO-EARTH OR LINE-TO-LINE FAULTS

Short circuits rapidly lead to generator overload and bring a strong fire risk. Fuses, circuit breakers, busbar distribution and transient protection units provide protection. •

OVER-VOLTING AND UNDER-VOLTING

The voltage regulator corrects under and over-volting provided that rpm are within the normal operating range. •

REVERSE CURRENTS

The reverse current cutout prevents reverse currents. FREQUENCY CONTROLLER The frequency controller provides small trimming signals to the CSDU (constant speed drive unit) so that the output frequency is maintained at 400 Hz. LIGHT AIRCRAFT ELECTRICAL CIRCUITS Light aircraft normally have a DC single wire, negative earth (ground) electrical system. The circuit runs through a single wire carrying positive current and is completed by earthing each electrical service to the airframe. In the wiring diagram, the earth connection is shown by the symbol. INDICATOR LIGHTS Indicator lights show if circuits or equipment are working, off or on standby. Indicator lights may be coded for colour, sequence, or position. They may also control intensity, and stationary or rotating. INVERTERS If AC is required in a DC circuit, an inverter can provide it. An inverter is a motor generator (rotary converter) which has separate AC rotor and DC armature windings which share the same field system. The inverter input is usually 28v, but can be 14v. Inverters usually supply 115v three-phase current. LOAD SHARING In a split busbar system with each generator supplying its own busbar, the total electrical requirements of the aircraft are divided approximately evenly between the generators. Where two or more generators are operating in parallel, the load must be evenly shared. The voltage regulator achieves this.

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LOAD SHARING BETWEEN ALTERNATORS Load sharing between alternators is similar to that for DC generators. An additional problem with alternators is the difference between real and reactive electrical loads. The real load sharing is controlled by the frequency controller, which senses the loads on each alternator and adjusts the torque on the alternator CSDs as necessary. The reactive load sharing is carried out by voltage regulation.

Power load shared equally by generators in parallel

Each generator supplies a bus with current.

Fig. 2-20 Types of Load Sharing

MONITORING Alternator output is measured on large aircraft by voltmeters, ammeters and frequency meters. Magnetic indicators also show when the alternator is supplying power to a busbar and (with warning lights) shows a supply failure. MULTIPLE AC GENERATOR OPERATION On multi-engine aircraft each engine normally drives one alternator. The three main methods of using the alternator outputs are summarised below. •

CONSTANT FREQUENCY PARALLEL BUSBAR

In this arrangement each alternator pair feeds a split busbar, which has a facility for interconnection through a circuit breaker. The two alternators can only supply the same busbar when frequencies and phases are accurately matched. A phase discriminator ensures that the alternators are synchronised before operating in parallel. In the same way, the two busbars can only be connected in parallel by the split busbar CB when voltage and frequencies are synchronised by a further phase discriminator. •

CONSTANT FREQUENCY – SYNCHRONISING BUSBAR

In this arrangement each alternator supplies its own load busbar. In addition, one alternator is always connected to the synchronising busbar. Each alternator supply is controlled by two circuit breakers. With the "A" (Alternator) breaker closed, the alternator feeds its own busbar. When the "S" (Synchronising) breaker is closed, the alternator supplies its own busbar and the synchronising busbar. •

PARALLEL-WOUND MOTOR

A parallel-wound motor produces high rpm with relatively low torque. This type of motor is used for small devices such as fans, centrifugal pumps and motor-generator units.

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PROTECTION UNIT A constant frequency AC generating system must include certain protection circuits to monitor safe system performance. The protection unit contains circuits to guard against faults such as under-volting, over-volting, incorrect phase rotation and line-to-line and lineto-earth faults. The protection unit controls the field relay and generator control breaker (GCB) and opens or closes them depending on system integrity. STATIC ELECTRICITY Static electricity is a build-up of electrons (or reduction of electrons) in a body. Two clouds may be charged with static electricity, but at different levels. When this occurs to a sufficient extent, an equalising discharge (lightning) may occur between them. In the same way an aircraft may build up a charge and "invite" a lighting strike. Static charge interferes with the radio magnetic field around the aircraft and may prevent use of radio equipment. Dischargers are fitted on various parts of the aircraft (behind the wings and empennage) to rid the aircraft of any static build-up. All parts of the aircraft are bonded to the main frame with two bonding "jumpers" (connectors) to form one complete circuit. All of the bonding connections are as short as possible to keep electrical resistance low. The path of the static electricity should be the least resistance possible so that static can travel to an earth. ST ELMO'S FIRE St Elmo's fire is a corona discharge of static electricity from an object in a highly charged electric field. It can be compared to lightning but the intensity is considerably lower. St Elmo's fire is impressive but quite harmless. THERMOCOUPLES A thermocouple (electric thermometer) is a self-energised heat sensing and measuring device. The thermocouple leads run from the engine to a cockpit indicator. A thermocouple is made from a copper gasket with two dissimilar metals attached (bi-metallic strip). The copper gasket is fitted at a suitable place in the engine. (This might well be the hottest running cylinder head). When the gasket is heated it expands and a voltage proportional to the temperature rise is induced where the metals join. This micro-voltage is proportional to temperature and is indicated on a sensitive voltmeter which is calibrated in °F or °C. The cylinder head temperature (of the hottest running cylinder) is thus indicated. Thermocouples are also used in fire sensing units in fire-detector systems. High temperature from flame or fire in a detector area will generate enough current (say 2.6 milliamps) to activate a sensitive relay which then illuminates a fire warning light and/or rings a bell. Thermocouple circuits in the exhaust cone of turbine engines indicate exhaust-gas temperature (EGT) on the instrument panel. Thermocouples are independent of the aircraft electrical system as the very small current is generated by exhaust gas heat.

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The term "Jet Pipe Temperature" (JPT) is sometimes used in place of EGT. The difference lies in where the thermocouple circuit is placed in the exhaust gas flow (jet pipe). Thermocouples and loop fire detectors are very efficient sensing devices in fire detection systems. The systems include detection circuits, warning circuits, test circuits and a fire bell.

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CHAPTER THREE – THE FUEL SYSTEMS

THE CONTENTS OF THIS CHAPTER: ! The Fuel System: ! Fuel Tanks " Structural components and types " Location of tanks on single and multi-engine aircraft " Sequence and type of refuelling " Unusable fuel ! Fuel Feed " Gravity and pressure feed " Cross feed " Schematic construction ! Fuel dumping system " Fuel system monitoring " Operation, indicators, warning systems " Fuel management (fuel tank switching) " Dipstick

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FUEL SYSTEMS INTRODUCTION Fuel systems ensure an adequate supply of fuel at the right pressure to the engine, irrespective of changes in aircraft attitude, altitude or speed and independent of environmental conditions. FUEL TANKS Regardless of the location of the aircraft fuel tanks, the basic components of a fuel system are the same as is shown in Figure 3-1. Fuel tanks are constructed with baffle plates to reduce sloshing of fuel within the tanks. This can be a contributory factor leading to vapour lock. Fuel tanks are vented to atmosphere to allow atmospheric pressure to be maintained in the tanks at all altitudes. The vents allow air to enter the tanks to replace fuel used by the engine. If the tanks were not vented, the pressure in the tanks would reduce as the fuel was consumed and the fuel flow to the engine would reduce. Also, the reduction in pressure in the tanks might make the tanks collapse inwards. The vent also provides for thermal expansion and contraction of the fuel. If an aircraft is refuelled to capacity and is allowed to stand in the sun with a blocked vent then the expansion of the fuel and the consequent increase in tank pressure can be sufficient to debond integral fuel tanks.

Fig. 3-1 Fuel Tank Installations in a High and Low Wing Aircraft

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Tanks are usually one of three types: ! The bladder type fuel tank is made of fibre-reinforced rubber or nylon, which is less expensive. The tank is usually flexible and is held in place with clips. This ensures that it conforms to the shape of the interior of the wing. Fig. 3-2a Bladder Type Fuel Tank

! The second main type of tank is the integral fuel tank. It is usually built as a part of the wing load carrying structure. It is termed a "wet wing".

Fig. 3-2b Integral Fuel Tank

! Rigid tanks are usually constructed from fibreglass, reinforced plastic or a light metal alloy. They are usually used as auxiliary tanks or drop tanks.

Drop Tank

Fig. 3-2c Rigid Type Fuel Tank

FUEL QUANTITY AND FLOW INDICATORS Fuel quantity indicators come in two basic types on modern general aviation aircraft. They are either mechanical or electrical (electronic). Indicating fuel accurately is difficult. It is made complex by the fact that the tank shape has to conform to that of the wing which is usually slender. Aircraft attitude varies widely with different manoeuvres which further complicates the task.

FILTERS, STRAINERS AND GASCOLATORS Fuel strainers are often installed in tank outlets. They may also be integral with the fuel boost pump when fitted. Fuel sump strainers are located at the lowest point in the fuel system, between the fuel tank and the engine, and usually have a very fine mesh. The gascolator is a centrifugal type strainer in which fuel is forced to circulate. This throws any denser compounds (water and dirt) to the outside of the device where they are collected and can be drained off. Copyright © 2012 EAA

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FUEL PUMPS Fuel pumps are either mechanically driven (via gears from the engine) or electrically driven. All aircraft engines have a mechanical pump, while only those requiring an auxiliary or backup pump have an electrical fuel pump. FUEL Aircraft engines (particularly piston engines) are highly sensitive to fuel grade. It is the responsibility of the pilot-in- command to ensure that the aircraft is serviced with the correct fuel. AVGAS Piston engine fuel (AVGAS) comes in various grades, which are colour coded. The standard colour codes are : ! 100 LL ("low lead") which is coloured blue. It is not available in the RSA. ! 100/130 which is coloured green. The octane rating of the fuel – e.g. 100 as in 100LL – is a measure of the antidetonation (anti-knock) properties of the fuel. The higher the octane rating of the fuel, the greater compression and heat a fuelair mixture can withstand without detonation occurring. Where two octane numbers are quoted – as in the case of 100/130 octane – the higher number represents the anti-detonation properties of the fuel when a full rich mixture is used, while the lower number represents the octane rating when a lean mixture is being used. ! Using fuel with an octane rating which is too low can lead to detonation. This is especially so at high power settings. ! Using fuel with a higher than specified octane rating can foul the spark plugs and damage the valves and valve seats. ! Using turbine fuel in a piston engine will make it stop. TURBINE FUEL Turbine engines are also sensitive to the type of fuel used. However, these engines are much more tolerant than piston engines. Typically, an engine may be approved for operation with a range of jet fuels such as Jet A-a, Jet B, JP-4, JP-5, JP-8, AVTUR and – for a limited period – with AVGAS (Typically 50 – 150 hours in any period between overhauls).

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AVIATION FUEL ADDITIVES Aviation fuel additives are compounds added to the fuel in very small quantities, usually measurable only in parts per million, to provide special or improved qualities. The quantity to be added and approval for its use in various grades of fuel is strictly controlled by the appropriate specifications. A few additives in common use: Anti-knock additives reduce the tendency of gasoline to detonate. Tetra-ethyl lead (TEL) is the only approved anti-knock additive for aviation use and has been used in motor and aviation gasoline’s since the early 1930s. Anti-oxidants prevent the formation of gum deposits on fuel system components caused by oxidation of the fuel in storage and also inhibit the formation of peroxide compounds in certain jet fuels. Only certain anti-oxidants are allowed in aviation fuels. Static dissipater additives reduce the hazardous effects of static electricity generated by movement of fuel through modern high flow-rate fuel transfer systems. Static dissipater additives do not reduce the need for 'bonding' to ensure electrical continuity between metal components (e.g. aircraft and fuelling equipment) nor do they influence hazards from lightning strikes. Corrosion inhibitors protect ferrous metals in fuel handling systems, such as pipelines and fuel storage tanks, from corrosion. Some corrosion inhibitors also improve the lubricating properties of certain jet fuels. Fuel System Icing Inhibitors (Anti-icing additives) reduce the freezing point of water precipitated from jet fuels due to cooling at high altitudes and prevent the formation of ice crystals which restrict the flow of fuel to the engine. This type of additive does not affect the freezing point of the fuel itself. Anti-icing additives can also provide some protection against microbiological growth in jet fuel. Metal de-activators suppress the catalytic effect which some metals, particularly copper, have on fuel oxidation. Only specific metal de-activators are allowed in aviation fuels. Biocide additives are sometimes used to combat microbiological growths in jet fuel, often by direct addition to aircraft tanks.

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CHAPTER FOUR – UNDERCARRIAGES

CONTENTS OF THIS CHAPTER: ! Gear Retraction and Extension Systems ! Gear emergencies ! Braking systems ! Suspension Systems

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UNDERCARRIAGES INTRODUCTION The undercarriage of an aircraft includes: ! the main undercarriage leg components as well as ! the wheels, ! tyres and ! brakes. It performs the essential function of absorbing shocks during landing, take off and to support the aircraft while on the ground. Landing speeds, weight, length of landing run and cross wind landing/take off capabilities all have an influence upon undercarriage design. The design of an undercarriage system must take into account: ! the necessity to provide minimum drag, ! must be strong enough to withstand the effects of braking, ! capable of withstanding drag and side loads imposed upon the system during landing with crosswind components and ! must enable the aircraft to be operated over different types of surfaces – grass, gravel etc. GEAR RETRACTION AND EXTENSION SYSTEMS Landing gear retraction mechanisms are classified according to the power source used for actuation. The most common systems are the electric, hydraulic and electro-hydraulic.

Fig. 4-1 Typical Electro Hydraulic Gear Actuation System

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In the electric system, a reversible electric motor is used to drive the landing gear extension/retraction system via a system of gears and levers. The hydraulic system uses an engine driven hydraulic pump to generate the hydraulic pressure in the system to transmit the energy to hydraulic cylinders for extending and retracting the undercarriage. In the electro-hydraulic system, the hydraulic pump is driven by a reversible electric motor which reduces the requirement for expensive valves which are needed to accomplish the reversible sequencing action. Pressure from the electric hydraulic pump extends the gear in one direction and retracts it in the other through actuators. Each gear is held up by hydraulic pressure but the main gear also has an hydraulically operated up-lock. The main gear is held in the locked down position by over centre travel of a spring-locked side brace. The nose gear is locked down by the over centre travel of a draglink and a hydraulically actuated down-lock. A pressure dump valve is provided for emergency landing gear extension. Most gear systems incorporate a method to prevent inadvertent ground retraction of the landing gear. This is usually a "squat switch" – a switch which breaks the circuit to the retraction motor when the landing gear strut is compressed. In addition to this device, most aircraft with retractable gear have a system designed to warn the pilot when the aircraft is in the approach configuration and the gear is not locked down. The system may even extend the gear under these circumstances. The system usually senses manifold pressure and airspeed and compares them with preset values to see whether the combination is in the typical range encountered during approach for landing. GEAR EMERGENCIES Gear emergencies – situations where the landing gear will not extend properly – are provided for in most undercarriage systems. Electrically driven systems usually incorporate a hand crank which can be used to extend the gear via the usual mechanism of gears and levers. It thus caters primarily for failure of the electric motor but cannot compensate for other problems such as a stripped gear set. Hydraulic and electro-hydraulic systems typically incorporate one of three systems. They can either be pumped down by hand, forced down by supplying a large quantity of pressurised gas to pressurise the system and extend the gear or they can be of the free-fall type. The ‘free fall’ system is preferable as it is cheap and simple to install and operate. It is usually used in a system where hydraulic pressure is used to hold the gear in the retracted position. This does mean that if the system should spring a leak, the gear will extend, reducing cruise speed and range. The hand pump arrangement is simple but has the disadvantage that it will be unable to produce enough system pressure to extend the gear in the event of a leak. This disadvantage is overcome with the use of a bottle of high-pressure carbon dioxide to pressurise the system to extend the gear. This system is usually expensive and once the bottle is discharged it cannot be used for a second attempt. Copyright © 2012 EAA

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BRAKING SYSTEMS The effectiveness of a braking system is based on the principle of rubbing two different surfaces together which then converts Kinetic energy to Heat energy. Braking systems on most general aviation aircraft are hydraulically actuated disc brakes acting on the main wheels. Most systems have an independent toe activated braking system for each wheel, which allows for differential braking. The hand brake is usually a system which either applies the brakes by mechanical means: ! a cable system, or ! a system that locks hydraulic pressure in the braking system once it has been applied by the toe brakes. Antiskid systems perform an important function in normal skid control and locked wheel skid control. The skid control box recognizes a skid, locked wheels and application of brakes. Skid control valves then remove some of the hydraulic pressure and permit the wheel to rotate and so maintain maximum braking energy.

Fig. 4-2 Typical Brake System

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SUSPENSION SYSTEMS (SHOCK ABSORPTION) Aircraft suspension systems fulfil a very important role even though an aircraft may only spend a small portion of its time moving on the ground. It is the most complex component of the undercarriage. As well as the requirement to provide a stable, comfortable ride, the system must dissipate any energy associated with a bounced landing. The various systems include: ! simple spring steel, ! rubber, or other polymers between metal layers, ! The Oleo System (Oil, air, piston and cylinder systems) ! The spring steel system is certainly the cheapest but the system offers very little energy dissipation capability. Shock absorption is accomplished by means of the cantilever blades of spring steel.

Spring Steel

Fig. 4-3a Spring Steel System

The rubber system is also cheap and is somewhat better in its ability to absorb energy. Rubber blocks are inserted within a cylinder and sliding tube and become the absorbent material under compression. Internally sprung wheels use the same system as the rubber system except that the absorbent material is a high tension spring.

Rubber blocks

Fig. 4-3b Rubber System

The oleo system, (Shock Struts) although by far the most complex and expensive to build and maintain, it is the most effective at absorbing and dissipating the energy of a bounced landing. This system is made up of two telescoping cylinders with both ends closed. The two cylinders, known as the piston and cylinder, form two chambers. The lower chamber is filled with hydraulic fluid and the upper chamber is charged with air. The top of the piston has a damping orifice through which a metering pin projects. Copyright © 2012 EAA

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On landing, compression of the cylinders forces fluid from the lower chamber, through the damping orifice, and compresses the air in the upper chamber. The flow of fluid through the orifice is controlled by the metering pin. During take-off the compressed air forces the oil back below the piston through the small damping orifice. A piston recoil valve or spring is fitted within the cylinders to prevent damage to the strut from the sharp impact at the end of the extension stroke. In order to maintain correct wheel alignment most shock struts are equipped with torque arms. Those without torque arms have splined piston heads which align with the splines in the cylinder ensuring correct wheel alignment.

Side struts/bracing struts – They may be found on aeroplanes whose main undercarriage retracts sideways (eg. Baron) to carry sideways loads. They would also be intended to support the force in a crosswind and during turning.

Fig. 4-3c Operation of Oleo Suspension System

Bungee chords – a simple and effective shock absorber system. One end of the triangular leg assembly is attached to the airframe and is free to pivot. A ‘bungee’ chord is attached over a rigid member of the airframe and around the free end of the leg assembly.

Hinged to airframe

Landing aircraft

Bungee extended Rigid frame

Bungee chord Fig. 4-3d Operation of Oleo Bungee Chord System

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CHAPTER FIVE – OXYGEN

CONTENTS OF THIS CHAPTER: ! Requirements for Oxygen ! Pressurisation and Oxygen ! Oxygen Systems ! Handling Precautions

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OXYGEN REQUIREMENTS FOR OXYGEN Flying at high altitudes has many advantages. High altitude flight gives better performance on most flights. The ability to fly above the weather is also very useful. Unfortunately, flying high imposes physiological limitations. The main problem with high altitude flight is the need for more oxygen and, eventually, oxygen at higher pressure as the altitude increases above 12000 ft. An idea of the importance of this limitation can be seen from the table below which shows the time of useful consciousness at various altitudes. 30 000 ft 28 000 ft 25 000 ft 22 000 ft 12 – 18 000 ft

1 – 2 min 2 ½ – 3 Min 3 – 5 Min 5 – 10 Min 30 Min +

+ (very variable)

Fig. 5-1 Typical Times Useful Consciousness

PRESSURISATION AND OXYGEN One way to cope with this limitation is to provide concentrated oxygen for the aircraft’s occupants to breathe individually. The alternative is to pressurise the aircraft cabin. In the latter case it is still necessary to provide a supply of concentrated oxygen for emergency use; such as a cabin depressurisation at altitude. The oxygen system may be permanently installed as shown in Figure 5-2 below or may be one of the many portable types available.

Fig. 5-2 Typical Oxygen System Installation

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OXYGEN SYSTEMS: Oxygen is supplied to the user in ways, which vary from model to model, and between manufacturers. The most common system is the "continuous flow" or "constant pressure system" which supplies oxygen at a constant rate for a given altitude.

Fig. 5-3 Components of a Typical Continuous Flow Installation

HANDLING PRECAUTIONS Oxygen is dangerous under certain conditions. The following cautions should be observed when using and handling oxygen : ! Do not expose oxygen cylinders to high temperatures. This will cause a rapid rise in internal pressure. The cylinder may explode or the safety valve (if fitted) may rupture. ! Never use oil or grease when using or handling oxygen. This restriction also means that oxygen system fittings must never be lubricated. Oxygen will react with these substances and can cause spontaneous combustion. ! Never use medical oxygen in place of aviator's breathing oxygen. The former has a higher moisture content which could freeze in the supply lines. ! Never use oxygen which has any form of odour. Good oxygen does not smell. ! Replenish oxygen in a well ventilated area – not inside a hanger where leaking oxygen could accumulate and create a fire hazard. Remember oxygen is denser than air and will sink to ground level. ! Never replenish oxygen during refuelling or other maintenance operations where there is a supply of material which could combine explosively with the oxygen. ! Never smoke in an area where oxygen is being used or bottles are being re-charged.

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CHAPTER SIX – CABIN PRESSURISATION AND AIR CONDITIONING

CONTENTS OF THIS CHAPTER: ! Cabin pressurization " Definitions and Terminology " Ram air system " Engine bleed system " Compressor / blower system " Cabin pressure differential " Pressure differential control " Cabin pressurization controller " Rate selector " Safety devices ! Air conditioning " Air conditioning requirements " Cooling methods " Air cycle cooling " Vapour cycle cooling " Conditioned air distribution " Typical air conditioning system " Gasper air system " Ventilation " Humidity control

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CABIN PRESSURISATION AND AIR CONDITIONING SYSTEMS INTRODUCTION As aircraft altitude increases, the amount of oxygen available in the atmosphere to sustain life decreases. The adverse physiological effects of altitude, rapid changes in altitude, composition of the atmosphere and extremes of temperature and humidity necessitate that the environment in an aircraft cabin or cockpit be carefully controlled within specified limits. The limiting height is usually around 10 000 feet, although periods may be spent at higher altitudes without ill effects. If an additional supply of oxygen is not available, air under pressure must be supplied to all on board so that the efficiency, safety and comfort of crew and passengers are maintained whatever the ambient conditions encountered. The effective cabin altitude of an aircraft must be limited to between 6,000 and 8,000 feet. When air is supplied under pressure, it has the effect of reducing the "cabin altitude" to a level well below the actual ambient altitude (that is, air in the cabin is at a greater pressure than the air outside the cabin – the ambient atmosphere). The pressurised air is supplied and controlled through a cabin pressurisation and air conditioning system (known as air conditioning "packs"). DEFINITIONS AND TERMINOLOGY CABIN PRESSURISATION "Cabin pressurisation" implies sealing the cabin and providing an air supply from an external source together with pressure control by manipulating the air flow rate through the cabin.

Fig. 6-1 Typical Piston Engine Aircraft Pressurisation System

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Cabin pressurisation enables an aircraft to fly at high altitudes while keeping the passengers and crew comfortable (and alive). The concept of the system is quite simple. ! High-pressure air is continuously pumped into the cabin through the pressurisation inlet ports of the air-conditioning system. ! Incoming air increases the air pressure in the cabin and raises it to that corresponding to a lower altitude. ! Cabin air pressure is regulated by the rate at which the air is allowed to flow out. The rate of flow is controlled by an "outflow valve" (or valves) which is/are in turn controlled by the cabin altitude selected by the pilot. ! AIR CONDITIONING Air Conditioning can be defined as the simultaneous regulation and control of factors, which affect the atmosphere within the aircraft. These factors are: ! temperature ! humidity ! ventilation. The temperature of air entering the cabin is controlled by air conditioning "packs". After leaving the engine, pressurised air passes through a heat exchanger in order to reduce the temperature. This is necessary because the air temperature is raised significantly in the compression process. The heat exchanger is usually an air-to-air device which uses atmospheric air as the coolant. Incoming air is mixed with air drawn from the cabin or with air-conditioned air before entering the cabin through the usual air vents. These vents are built into the aircraft and should not be confused with the air "louvres" which merely circulate air through the cabin. At high altitude in cold atmospheric conditions, the cool air supply to the air-to-air heat exchangers can be reduced or cut off to keep the cabin warm. A considerable part of the heat required is obtained from the occupants of the cabin themselves. TERMINOLOGY PRESSURE VESSEL refers to that portion of the aircraft designed to withstand the pressure difference between air inside the cabin and outside. CABIN ALTITUDE refers to the pressure altitude of the air within the cabin. The cabin pressure altitude is always lower or equal to the pressure altitude of the air outside the cabin (the aircraft pressure altitude). Cabin altitude is usually expressed in thousands of feet in the standard atmosphere (flight level). CABIN PRESSURE is the actual pressure which exists within the pressure vessel and is usually measured in psi. OUTSIDE AIR PRESSURE OR AMBIENT PRESSURE is the actual pressure of the ambient air and is usually measured in psi. PACK/S is a term used for the air conditioning system

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GASPER/S is a term used for the re-circulatory fans. These merely move the air in the same way as a desk fan in an office. PRESSURISATION RELIEF VALVE is a valve designed to reduce the pressure in the pressure vessel by releasing air to the atmosphere if the preset pressure differential pressure is exceeded. VACUUM RELIEF VALVE prevents ambient pressure from exceeding the cabin pressure – cabin pressure is increased to prevent fuselage from collapsing. DUMP VALVE cabin air is ‘dumped’ to atmosphere by means of a lever or switch in the cockpit which activates the dump valve. CABIN CONTROLLER is the selector, usually a rotary switch, on which the desired cabin altitude is set. OUTFLOW VALVE is a valve, usually at the rear of the pressure vessel, which can be manipulated, either manually or automatically, to control the pressure inside the cabin. PRESSURE DIFFERENTIAL is the difference between cabin pressure and ambient pressure. It is usually expressed in psi. Referring to Figure 6-2, it can be seen that for a constant pressure differential, the difference between cabin altitude and aircraft altitude increases with increasing altitude. This is because the rate of pressure change with altitude decreases as altitude increases, being roughly 30 hPa per 1 000 ft at sea level and 20 hPa per 1 000 ft at 20 000 ft.

Figure 6 – 2 Cabin altitude variation with differential pressure and altitude

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CABIN PRESSURISATION Three types of pressurisation systems are in use: ! ram air system ! engine bleed system ! compressor or blower system. ! RAM AIR SYSTEM The ram air system is a simple system used in smaller aircraft. (Figure 6-3 below) Ram air is usually taken in at the nose of the aircraft or at the base of the vertical stabilizer. After circulating through the cabin, the air is then discharged to atmosphere via a spill vent.

Figure 6 – 3 Typical ram air system

ENGINE BLEED SYSTEM This type of system is widely used on turbine powered aircraft where a ready supply of pressurised air is available from the compressor section of the engine. (Figure 6-4 below.) ! Air is bled from the high-pressure section of the compressor. ! Because the air is compressed, it will be hot and must be cooled before it can be piped into the cabin. ! It then passes through one of two heat exchangers where it is cooled. This first heat exchanger is normally referred to as the primary heat exchanger. ! The air is then compressed further as shown to raise the pressure and temperature. This enables the process to be repeated. ! The re-heated air then passes through a secondary heat exchanger for additional cooling. ! This air is then expanded (by reducing the pressure) through a turbine before it is supplied to the cabin. ! The energy released by the air expanding through the turbine is used to drive the compressor. Copyright © 2012 EAA

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This type of system where the compressor/turbine combination is self-sustaining is known as a bootstrap system.

Figure 6-4 Typical Turbine aircraft pressurization system

The bootstrap system requires a supply of high-pressure air to operate. On the ground with the engines stopped, this high-pressure air is normally supplied by the APU (auxiliary power unit). The various combinations of switching valves, which allow different air sources to supply compressed air, are shown in Figure 6-5.

Figure 6 – 5 Bleed air or bootstrap air conditioning system

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Figure 6 – 6 Typical heat exchanger arrangement

Figure 6 – 7 Valve switching combinations for supply of compressed air

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THE COMPRESSOR SYSTEM The compressor system is similar in principle to the bleed air system. The only major difference is that high-pressure air is supplied by a compressor. The compressor is driven by the engine through an accessory drive. The compressor is usually the lobed or roots blower type as shown in Figure 6-8. Older jet transports may still be found with these compressors (sometimes known as "turbo-compressors").

Figure 6 – 8 Compressor or blower air conditioning system

CABIN PRESSURE CABIN PRESSURE DIFFERENTIAL The choice of cabin pressure for any particular aircraft system will be influenced by factors such as: ! structural strength and the resultant weight penalty ! the possibility of explosive decompression caused by hijacking etc ! the aircraft type – high or low altitude transport, fighter/bomber etc ! the rate of change of absolute pressure which can be tolerated by the occupants. This will depend on whether they are fully trained aircrew, or passengers (who will not necessarily be of the same standard of physical fitness as aircrew) ! the availability of the required pressure. In practice, the maintenance of sea level conditions at extreme altitudes would impose an unacceptable weight penalty. At 15 000 ft a cabin altitude equivalent to that at sea level can be maintained with a 6½ psi differential, but as a general rule in transport aircraft, pressurisation commences from ground level and continues up to cruise altitude. At cruising altitude, the cabin altitude usually selected is the equivalent of that at about 8 000 ft. 70

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PRESSURE DIFFERENTIAL CONTROL Cabin blowers are rated to deliver the correct amount or air at altitude and engine speed recommended for normal cruising conditions. Below that altitude or at higher engine rpm the blowers deliver a supply greater than that required. Therefore a method is required so that surplus air can be vented to atmosphere. The usual system is to incorporate a mass flow controller in the pressurisation system. The controller actuates spill (dump) valves, or flow control valves, so that a constant air mass flow passes into the system, irrespective of variations in air density or in the output from the supply source. As a rule, flow controllers are based on the use of pressure-sensitive diaphragms, which operate electrical contacts. These, in turn, make or break an electrical circuit to relay switches. The switches control the operation of spill valve actuators, thereby regulating the rate of flow of air to the cabin. Although it is usual to exercise control over the air supply to the cabin, particularly during the climb or descent, cabin pressure is usually independent of the rate of air input. It is usually regulated by varying the rate of discharge from the cabin. In many pressurisation systems a proportion of the air supply is allowed to leak out through a venturi nozzle. The venturi is designed to pass a maximum flow of air which is determined by the throat area, thereby reducing the amount to be handled by automatic valves and sensing units, which usually comprise a pressure controller and a cabin discharge valve.

Figure 6 – 9 Cabin controller – Auto/Manual Control

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Another system varies the opening in an outflow valve, which increases or decreases the flow of air to atmosphere. This outflow valve is usually situated at the rear of the cabin so that there is a positive flow of air throughout the cabin. CABIN PRESSURISATION CONTROLLER Most cabin altitude controls simply involve setting the desired cabin altitude on the control panel (the cabin controller). The system will maintain this altitude up to the aircraft pressure altitude where differential pressure is the maximum designed. When maximum differential is reached, the cabin altitude will then climb but at a slower rate than the aircraft rate of climb.

FLIGHT ALTITUDE SCALE

As the cabin altitude selector is rotated, this scale displays the flight altitude, which corresponds to a 3.6 psi differential

CABIN ALTITUDE RATE CONTROL Set cabin altitude rate of change. Range: 50 to 2000 feet per min. Index Mark: 300 feet per min (approx)

BAROMETRIC PRESSURE SETTING CONTROL

Sets altimeter setting into pressure controller

CABIN ALTITUDE SELECTOR

Selects desired cabin altitude. RANGE: -1 000 to +10 000

Figure 6 – 10 Cabin controller – Altitude Selector

RATE SELECTOR The rate of change of cabin pressure is important, chiefly because of its effect on the ventilation of the inner ear. This ventilation controls the pressure differential across the ear drum, the equalisation of pressure being accomplished through the Eustachian tube (Valsalva's manoeuvre). Too high a rate will mean that passengers may well have difficulty in accommodating pressure changes. Too low a rate can impose operational limitations and reduce the aircraft rate of descent, which may pose air traffic problems. A maximum rate of change of pressure allowable is usually taken as 500 feet per minute. The figure of 300 feet per minute is usually set on the cabin controller, as most passengers can easily cope with this rate.

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The cabin altitude should not normally be permitted to exceed the equivalent of ± 8 000 ft except in emergency. This latter requirement is determined by the maximum altitude which passengers may be allowed to experience without discomfort. SAFETY DEVICES Cabin pressure and its rate of change are controlled by the regulated release of air to atmosphere through a discharge valve in the aircraft skin. Although the system can normally be manually selected on or off, pressurisation parameters are preset and automatically monitored and controlled by the control system. In the event of a system malfunction or cabin differential limits being exceeded, the control system is over-ridden by a safety valve which senses an unacceptable cabin differential. The valve automatically opens to dump air, or to allow air to enter the cabin in the event of too rapid a descent. AIR CONDITIONING AIR CONDITIONING REQUIREMENTS For the best crew performance and to ensure passenger comfort, the cabin temperature should be maintained between 18 and 24°C. This range needs to be provided in outside temperatures, which can vary, from 40 °C on the ground to -76 °C at altitude. The need for temperature control in aircraft air conditioning systems arises from four main variables: ! ambient conditions outside the aircraft ! the rate of heat release from equipment inside the cabin ! the rate of heat release from the occupants of the cabin ! the heat input into the aircraft structure from kinetic heating. On the ground ambient conditions can vary from tropical summer heat at midday in calm air to an Arctic winter gale. In flight, the worst heating condition will be at high altitude in dull weather, while the worst cooling condition will be during high-speed flight at low altitude in bright sunlight. Standard design conditions are laid down to provide for these extremes. The amount of heat transmitted into or out of the cabin by radiation or convection depends on such factors as the use of light or dark paints, the "wetted" surface area of the cabin, the type and thickness of insulation used, the number of windows and the velocity of the airflow over the cabin surface.

Figure 6-11 Typical Cold Air Unit Schematic

Depending on altitude and the four factors listed above, the problem of maintaining a heat balance within the pressure cabin will depend on whether there is a net gain of heat into, or net loss of heat from, the cabin. As a rule, the air conditioning system incorporates Copyright © 2012 EAA

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refrigeration equipment, which is used whenever cooling is required and is by-passed when it is necessary to provide heating. A diagram of a typical cold air unit as used in a Boeing aircraft is shown in Figure 6-11. The normal practice in this case is to cool the charge air in a ram-air-cooled heat exchanger (pre-cooled) then use mechanical refrigeration to obtain sub-cooling. Mechanical refrigeration is obtained by means of an air turbine across which the charge air is expanded. The turbine is coupled to a suitable brake, such as a fan or compressor, which absorbs the energy given up during expansion of the charge air, which subsequently enters the pressure cabin at a temperature of ± 1-10° C. Temperature control is usually affected by by-passing the cooling turbine and, under extreme conditions, the heat exchanger. By mixing hot and cold air in varying proportions the cabin temperature can be adjusted to suit the requirements of passengers and crew. Figure 6-12 illustrates a typical temperature control installation. In all temperature control systems one of the chief problems is that of thermal inertia. Depending on the length of ducting and the bulk and heat absorption properties of the associated equipment and aircraft structure, there will tend to be a delay between the signal received by the temperature control valves and the response at the place where a change in temperature is required. As far as possible temperature control systems are designed to anticipate significant temperature changes and thereby reduce the delay, which would otherwise occur. Pneumatic and electrically-actuated systems are widely employed, any deviation from the datum setting resulting in signals being transmitted to the appropriate control valves which then adjust the proportions of hot and cold air fed to the cabin.

Figure 6 – 12 Temperature control schematic

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Maintenance of an even temperature throughout the cabin increases with the size of aircraft and the number of occupants. Localised temperatures in the vicinity of radio and electronic equipment and galleys may vary considerably from the mean datum. Similarly, the flight deck temperature often tends to be higher than that in the main cabin. It is because factors such as cabin heat load, humidity, thermal inertia and temperature distribution need to be related to a common standard that the term "equivalent temperature" has been devised. Another measure of temperature effectiveness is termed "effective temperature". This requires knowledge of air temperature, humidity and air velocity. To take account of radiant heat a further term has been devised known as "corrected effective temperature". The use of these terms indicates that control of temperature when discussing air conditioning involves several factors. The term’s "equivalent temperature" “effective temperature” etc is used when providing recommendations for operating procedures.

Figure 6 – 13 Effective temperature – "Comfort zone"

Humidity is very important. Figure 6-13 shows a series of temperature and humidity conditions within which the same degree of comfort will be experienced. The area ABCD, within which these conditions apply, is termed a comfort zone. Note the dry bulb air temperature limits are 18-29°C and the humidity range is 30-70 % RH. While most cabin air conditioning systems can provide heating and cooling, the main air conditioning requirement is to cool pressurised air. This is because air temperature is raised significantly in the compression process. After leaving the engine, or compressor, the compressed air passes through a heat exchanger, which cools it. Various types of heat exchangers may be found. They may use atmospheric air or a refrigerant such as Freon to cool the incoming pressurised air.

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COOLING METHODS All methods for providing cooling rely on a basic cycle: ! compression of a gas or vapour ! heat rejection at the higher temperature produced by the compression process ! expansion or pressure reduction either through a pressure reducing valve or through a device such as a turbine. ! heat absorption at the lower temperature either via an evaporator or by direct mixing with the cooled gases. AIR CYCLE COOLING There are two variations on this cycle, ! the basic air cycle cooling system and ! the bootstrap system. In the basic air cycle system, air is first compressed until the pressure is substantially higher than that in the area to be cooled. The heat from compression is then dispersed in a heat exchanger. The air is then expanded through a turbine where a lot of the work of compression is extracted thus lowering the temperature further. The air is then supplied to the cabin. ! This system has the advantage of being relatively light and simple and relatively cheap. ! The main disadvantage is the low efficiency of the system and inability to provide significant ground cooling. Compressor

Heat exchanger

Expansion turbine

Ducting

Fig. 6 - 14- Basic Air Cycle System

These problems are as a result of the relatively small pressure rise available in single stage or multiple stage compressors without any intercooling. Thus the heat rejection at high pressure is relatively small. This effectively means that the volumetric flow rate required to provide adequate cooling must be very high which in turn raises the engine fuel consumption. In the bootstrap system, air is cooled, usually between two stages of compression. This allows the air to be raised to a higher pressure with less work input. In addition, the energy released when the air is expanded through the turbine is usually used to drive the second stage of compression further increasing the system efficiency. (Figure 6-15.)

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Figure 6-15 Air cycle cooling (boot strap)

VAPOUR CYCLE COOLING Vapour cycle cooling systems follow the basic cycle, but enjoy several very distinct advantages. ! they consume far less power than an air cycle system for a given amount of cooling power ! they allow the cooling action to be provided precisely where required ! they allow for effective humidity reduction on the ground. The system has the advantage that the working fluid or refrigerant is separated from cabin air and that it operates in two phases, liquid and gaseous.

Figure 6 – 16 Vapour cycle cooling

This enables large quantities of heat to be dispersed in a very small space when the refrigerant is forced to condense. Large quantities of heat are absorbed from the air supplied to the cabin when the refrigerant is evaporated. The coolant is compressed by a compressor then it passes to the condenser. The refrigerant is cooled in the condenser. Because of the high pressure, the air condenses give off large amounts of heat. From the condenser the refrigerant passes to the evaporator via an expansion valve. The expansion valve lowers the pressure so that the fluid will evaporate when brought into thermal contact with the air to be supplied to the cabin. Copyright © 2012 EAA

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CONDITIONED AIR DISTRIBUTION Conditioned air leaving the mixing chambers is collected in the main distribution supply duct. The temperature of the air will be directly related to the setting of the cockpit and passenger temperature selectors. Overheat protection is provided by temperature sensors located in the supply duct. An overheat condition in the supply duct will cause the appropriate mix-valves to drive to the fully cold position and the "DUCT OVERHEAT" light will illuminate. A temperature higher than the duct overheat will close the appropriate pack valve and the "PACK TRIP OFF" light will illuminate. A TYPICAL AIR CONDITIONING SYSTEM A typical cabin air conditioning system for a large aircraft is shown schematically in Figure 6-16. The sequence of the system is: ! Refrigerant, usually Freon or an equivalent in gaseous form is compressed by a compressor, which is driven by an electric motor. It can also be directly driven by the engine through a belt or gear drive system. ! The high-pressure refrigerant gas is then passed to a condenser. This is nothing more than a heat exchanger where the gas gives off the heat of compression and as a result condenses to the liquid state. ! From the condenser, liquid refrigerant with a temperature close to ambient passes to the evaporator via a restricting valve that lowers the pressure. ! As a result of the lower pressure the refrigerant evaporates, drawing the latent heat of vaporization from the cooling coils of the evaporator, which in turn cools the cabin air passing over them. "Bleed" air from the engines is normally supplied to the air conditioning and pressurisation system. As an alternative, the APU (auxiliary power unit) or, on the ground, an external ground supply may be used as an alternate source of air. Engine bleed air is controlled by engine bleed valves. These allow air to flow to a heat exchanger, then to the main pneumatic duct. This duct is a wide bore piping system, which is pressurised by air and is the main source for supplying pressurised air to the cabin. The duct may also be used to supply air for starting the engines. Figure 6 – 17 Typical air-conditioning schematic

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GASPER AIR SYSTEM The gasper air distribution system provides air to individual crew and passenger positions. This air is colder than the air supplied by the main air conditioning system. A movable control nozzle at each crew and passenger outlet can change the direction and amount of airflow. When the gasper air supply is low (during taxi and single pack operation), the gasper fan may be switched on to increase airflow in the gasper air distribution system. VENTILATION Air in an aircraft needs to be changed continuously to ensure an adequate supply of oxygen and removal of the products of respiration. Typically regulations require the supply of 1 lb. of fresh air/minute/person under normal conditions and 0.5lb/minute/person in the event of an engine or other system failure. Most commercial aircraft meet these requirements easily. HUMIDITY CONTROL For passenger comfort the relative humidity needs to be around 60% and the cabin temperature at 18 °C. This requirement becomes very demanding as cruising altitude increases. The absolute humidity of the atmosphere decreases with height because cold air can hold very little water. Warming this air and supplying it to the cabin results in a very low relative humidity. The low relative humidity also reduces the skin temperature of the aircraft occupants because it causes rapid evaporation of skin moisture thus making them feel cold. Raising the relative humidity to the required level of around 60% would require carrying large amounts of water resulting in a considerable reduction of the useful load. In practice, a relative humidity of around 30% is maintained by re-introducing any water separated out in the water separator.

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TYPICAL EXAMINATION QUESTION 1. When engine air bleeds are used for air conditioning purposes, the surge control bleeds remain open:

2.

3.

4.

a)

at low power settings

c)

at all power settings.

6.

7.

8.

80

at high power settings

In a pressurised aircraft the differential pressure is: a)

aircraft altitude divided by cabin altitude and increases as the aircraft descends

b)

cabin pressure in relation to outside pressure and increases as the aircraft climbs

c)

cabin pressure in relation to outside pressure and decreases as the aircraft climbs up to the structural maximum.

Cabin altitude in a pressurised cabin is controlled by: a)

controlling the amount of inflowing air

b)

controlling both the inflow and outflow of air to the cabin

c)

controlling the amount of outflowing air.

In a pressurisation system, air entering the cabin is cooled and can be controlled by: a)

5.

b)

a water separator

b)

an outflow valve

c)

an air cycle machine.

If the maximum differential was reached during the climb in a pressurised aircraft and the system is under manual control, it will be necessary to: a)

increase the aircraft altitude

b)

increase the cabin altitude

c)

close the outflow valves.

In a pressurised aircraft differential pressure is produced when: a)

the cabin pressure is reduced (compared with ambient)

b)

the cabin pressure is increased (compared with ambient)

c)

ambient pressure is increased.

Safety devices are required in cabin pressurisation systems: a)

to prevent excessive internal pressures

b)

to prevent excessive external pressures

c)

to prevent both excessive and insufficient internal pressures.

If the fuselage of an aircraft has a slight leak, pressurisation is maintained by the action of the pressurisation controller which: a)

closes the outflow valve slightly

b)

opens the outflow valve slightly

c)

opens the inflow valve further.

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Cabin differential is:

10.

a)

the difference between cabin altitude and aircraft altitude and is usually negative

b)

the difference between cabin pressure and ambient pressure and is usually positive

c)

the difference between cabin pressure and ambient pressure and is negative.

In an aircraft with a pressurised cabin, differential pressure is obtained by: a)

increasing the pressure in the cabin

b)

decreasing ambient pressure

c)

increasing ambient pressure.

11. While flying at a constant altitude in a pressurised aircraft, the cabin altitude decreases slightly below that set on the controller. The action of the cabin pressure control system will be to: a)

open the outflow control valve slightly so as to increase the outflow of cabin air

b)

close the outflow control valve slightly so as to decrease the outflow of cabin air

c)

open the inflow control valve slightly.

ANSWERS

1A 2B 3C 4C 5B 6B 7C 8A 9B 10A 11A

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CHAPTER SEVEN- AIRCRAFT ICING

CONTENTS OF THIS CHAPTER:

Aircraft icing: !

The effect of ice on engine performance

!

Physical principle of ice formation

!

Prevention and removal of ice

!

Carburettor heat

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AIRCRAFT ICING, DE-ICING AND ANTI-ICING SYSTEMS INTRODUCTION Icing on aircraft is highly undesirable. It produces various problems including: ! drag increase from an alteration of airfoil shape resulting in loss of speed, range, endurance and altitude capability. ! reduction in the stalling angle of attack which increases stalling speed and hence approach speeds. ! engine and airframe vibration as a result of uneven accumulation of ice on propellers and unsteady airflow over accumulated ice. ! increase in aircraft mass with the associated decrease in aircraft performance. ! possible jamming of the control surfaces as a result of droplets running back and freezing in contact with the cold surface. ! reduction of radio transmission strength and reception ability as a result of ice accumulation on antennae. ! engine failure either as a result of ingesting large chunks of ice or as a result of inlet blockage.

Figure 7-1 The effect of ice accumulation on stalling angle

Icing can be treated in two ways: ! either it is prevented from forming or ! it is dispersed after forming. Various types of de-icing and anti-icing systems are found on aircraft designed to operate in "All Weather Operations". Several systems can provide anti and de-icing capability. Each system has advantages and disadvantages and operating peculiarities. A typical installation is shown in Figure 7-2 below.

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Figure 7-2 Typical icing protection installation

PNEUMATIC SYSTEMS Pneumatic systems are usually found on the leading edges of wings and control surfaces. They are used for providing a de-icing action and do not provide any anti-icing function. The typical mode of operation is shown in Figure 7-3 below. Compressed air inflates the "boots" from the pressure side of the vacuum pump or by bleed air from the compressor section of the compressor (in a gas turbine engine).

Figure 7-3 Typical pneumatic de-icing boot installation

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The system has the advantage of simplicity of installation and concept. However, it has certain disadvantages: ! Unless the system is automated, it requires a high degree of pilot skill for effective use. Ice is allowed to accumulate up to a certain thickness (usually between 0.25" & 0.5") before the boots are activated. If the boots are activated when the ice is too thin it will not crack properly and will not be blown away by the slipstream. If activated when the ice is too thick, the boots will be unable to crack it effectively – with the same result. ! The rubber boots require heavy maintenance. They must be kept clean so that ice will not stick to the surface. ! Treatment for ultra violet damage is also required. ! The situation is further complicated by the fact that they are given a conductive coating to protect against charge build up and lightning strikes. This precludes the use of the normal petroleum based cleaning fluids.

Figure 7-4 Action of pneumatically driven rubber boots

HEATED SURFACES Thermal systems usually use along the inside of the leading edge of the aerofoil and distributed around its inner surface. There are several methods used to provide heated air. These include: • Bleeding hot air from the engine turbine compressor. • Engine exhaust heat exchangers. • Ram air heated by a combustion heater. Contour-etched outer skin

De-icing air duct

Heated air ducted span wise Inner skin

Front spar

Leading edge diaphragm Inner skin

Figure 7-5 Example Of A Leading Edge Exhaust

Thermal System

Enlarged Cross section View of Leading Edge

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Heater mats differ in design and construction according to their purpose and environment. The latest mats have elements, which are made from a range of alloys woven in continuous filament glass yarn (A continuous strand of twisted threads of natural or synthetic material such as wool, cotton, flax or nylon used in weaving or knitting. Continuously heated elements

Intermittently heated elements

Glass cloth layers

Insulating material

Figure 7-6 Heater Mats found on the leading

Electrical element

Inside of leading edge

edge of an Engine Nacelle

Other elements are made from nickel chrome foil. The insulating mats are usually PTFE (A synthetically engineered compound: PolyTetra-Fluoro-Ethylene.), which gives a much higher limiting temperature than synthetic rubbers. Smaller surfaces such as propeller blades, stall warning devices and pitot tubes are usually heated electrically by using electrical heating elements as an anti-icing measure. Larger surfaces (such as leading edges and engine intakes) are usually heated by hot air, which may be extracted from the compressor stage of the engine; the exhaust gases or produced by a dedicated combustor. Figure 7-7 Heated Propeller

Each method provides an anti-icing action. None are suitable for de-icing since once water has frozen on the surface, the frozen layer acts as a good insulator preventing further melting. It can make the heater element overheat. WINDSCREEN DE-ICING Various systems are used for de-icing of windscreens: ! Alcohol – see ‘Fluid Systems’ below ! Electrical heating elements are built into the tempered glass of the windscreen. Some aircraft have thermal electrical switches which automatically turn on the heating Copyright © 2012 EAA

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element when temperatures become low enough for icing to occur. ! Hot air is directed via ducts to the windows. FLUID SYSTEMS Fluid systems simply distribute a layer of alcohol or glycol mix on the surface needing protection. Fluid is applied to aircraft windshields and may be applied to propellers by means of "slingers". A recent development is the certification of a porous leading edge of the propeller, which allows the fluid to ooze through providing a de-icing and antiicing action without disrupting the airflow.

Figure 7 – 8 Liquid de-icing arrangement for Propeller

INERTIAL SEPARATORS (ICE VANES) Inertial separators are often found on turboprop engines. Their operation relies on the fact that ice and rain, – being denser than air – find it more difficult to negotiate sharp bends. As shown in Figure 7-9, in normal operation, the inlet ducting provides a smooth inlet path for the air into the engine. This is necessary because sharp bends cause flow separation and pressure losses, which in turn cause power losses and reduced fuel efficiency. In icing conditions the inertial separators are deflected. This provides a more tortuous route for the incoming air. The result is that air can negotiate the path to the engine whereas ice and rain are unable to "turn the corner". They bypass the engine via a door, which opens in the inlet duct.

Figure 7-9 Turboprop inlet in normal operating position

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Figure 7-10 Inertial separators in the extended position

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CHAPTER EIGHT – HYDRAULICS CONTENTS OF THIS CHAPTER: The basic principles of hydromechanics: !

Hydraulic fluids

!

Schematic construction and functioning of hydraulic systems

!

Main, stand-by and emergency systems

!

Accumulators

!

Reservoirs

!

Operation, indicators, warning systems

!

Ancillary systems

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HYDRAULICS INTRODUCTION Hydraulic pumps deliver high-pressure fluid flow to the pump outlet. They are powered by mechanical energy sources to pressurize fluid. Hydraulic Power is the power source of choice, rather than manual and mechanical systems, when used in larger aircraft for operation of systems where larger control surfaces are to be deployed in stronger airflows such as flying controls, flaps, heavier and larger undercarriage systems and wheel brakes. THE ADVANTAGES OF HYDRAULIC POWER: ! it is simple and reliable ! it has good power to weight ratio ! it is capable of transmitting very high forces ! because the units are compact they may be placed at or in close proximity to the control they are to operate.

Fig. 8 – 1 – A Simple Hydraulic Pump

SYSTEMS OPERATED BY HYDRAULICS: ! Flight Controls " Flaps " Ailerons " Elevators " Rudders " Spoilers " Slats ! Nose wheel steering ! Brakes ! Reverse thrust systems ! Windshield wipers ! Undercarriage extension and retraction

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BASIC PRINCIPLES OF HYDROMECHANICS HYDRAULIC FLUIDS Hydraulic fluids are fluids housed in a unit and used to transmit and distribute forces to various units being actuated by the pilot. Properties Of Hydraulic Fluid Anti-foaming properties – necessary because foaming will affect the operation of the pump Flash Point – Flash point is the temperature at which a liquid gives off vapour in sufficient quantity to ignite momentarily or to flash, when a flame is applied.

Liquid gives off vapour,

when flame is applied …

it momentarily ignites …

and returns to a liquid state.

Fig. 8 – 2 – Flash Point

A high flash point is required in hydraulic fluids because it indicates a good resistance to combustion and a low degree of evaporation at normal temperatures Fire Point – Fire point is the temperature at which a substance gives off vapour in sufficient quantity to ignite and continue to burn, when exposed to a spark or flame. As with flash point, a hydraulic fluid needs to have a high fire point, considering the extreme temperatures in which it is required to function.

Liquid gives off vapour, …

when flame is applied …

the liquid ignites …

and continues to burn.

Fig. 8 – 3 –Fire Point

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Chemical stability – The ability to resist oxidization and deterioration over long periods is necessary. Viscosity – The liquid must be dense enough to give a good seal at pumps and pistons yet not be too thick so as to resist flow.

Fig. 8 – 4 –Example of Viscosity

Types Of Hydraulic Fluid Hydraulic fluids are manufactured using different base materials. A note of warning – attempting to mix the different types of fluid is dangerous – they do not mix. ! Vegetable base hydraulic fluid (Blue in colour) ! Mineral base hydraulic fluid (Red in colour) ! Phosphate ester hydraulic fluid (purple or green in colour) Vegetable base hydraulic fluid Primarily composed of castor oil and alcohol and is highly flammable. Natural rubber seals are used with this fluid. Mineral base hydraulic fluid This is a petroleum based fluid and synthetic rubber seals are used in this system. It is also highly flammable. Phosphate ester hydraulic fluid Non-petroleum based with a high flash point of 6000 degrees. Once the heat source is removed this fluid will stop burning.

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BASIC CONSTRUCTION OF A HYDRAULIC SYSTEM The basic system is composed of the following parts Reservoir

Hand pump

Power pump

Pressure gauge

Filter

Relief valve

Pressure regulator

Selector valve

Accumulator

Actuating unit

Check valves

Reservoir

Hydraulic fluid

Power pump

Hydraulic filter Pressure regulator Accumulator

Check valves Hand pump

Pressure gauge

Selector valve

Pressure relief valve Actuating unit

Fig. 8 – 5- Diagrammatic representation of a typical hydraulic system

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RESERVOIRS Hydraulic systems require a reservoir in which the fluid displaced when the servo jacks are retracted is stored until required again. It also performs a secondary function of cooling the fluid and allowing any air absorbed to separate out. TWO TYPES OF RESERVOIRS ARE AVAILABLE: ! The Integral Reservoir which is incorporated within some major component to hold a supply of fluid such as in the brake master cylinder. ! The In-Line Reservoir which is housed as a complete unit and is connected directly with other components in the system. CONSTRUCTION OF A RESERVOIR: Filler neck Connection for vent or pressurisation line Strainer Normal fluid level

Strainers are fitted into the filler neck, to prevent the entry of foreign objects during refilling.

Glass sight gauge

Fin

Fin

Baffles Connection for return line Baffles are installed in the reservoir to prevent swirling and surging of the fluid. If this were not the case then air would enter the fluid, and cause foaming.

Fin

Fin

Standpipes A standpipe is commonly fitted to the reservoirs, and allows for emergency use of the appropriate control in the event of a fluid loss due to leakage.

Connection for emergency system pump

Connection for main system pump

Fig. 8 – 6- Components of a reservoir Diagrammatic representation of a typical hydraulic fluid reservoir

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Vented systems where the flow of the fluid is due to atmospheric pressure and gravity. Pressurised systems used on some aircraft where atmospheric pressure becomes too low to supply the pump with adequate fluid. Different methods of pressurization can be utilised: ! Using an Aspirator or air pump ! Installing an additional hydraulic pump at the reservoir outlet to supply fluid under pressure to the main hydraulic pump. ! Air pressure from the cabin pressurization system ! Bleed air of the compressor stage of a turbine Baffles - In order to prevent surging and swirling which cause air to enter and cause foaming by mixing with the fluid, baffles are installed inside the reservoir. Fluid Filters/strainers - These are incorporated in the filler neck to prevent contamination by foreign material during refilling which could cause damage to the seals and pumps. ACCUMULATORS In order to absorb shocks and sudden changes in system pressure, Hydraulic systems include an accumulator. ! They dampen hydraulic surges and pressure ripples caused by pump operations and sudden changes in pressure caused by operation of components such as jacks and valves. ! They store power for limited operation – e.g. to maintain aircraft parking brake pressure for long periods. ! Supplement power pumps when a number of systems are activated at the same time. In the event of systems requiring only a limited duration of operation under emergency power such as wheel brakes and undercarriages, the stored energy of an accumulator can be used.

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PRINCIPLE OF OPERATION While the power pump is providing sufficient pressure within the system, the fluid below the diaphragm is providing pressure, which pushes the diaphragm toward the compressed nitrogen, compressing it further. The highly compressed nitrogen then acts as a charged “spring” waiting for the pressure to decrease Fig. 8 – 8 Principle of operation

When the pressure in the system decreases, either due to malfunction, or excess demand, the compressed nitrogen forces the fluid back into the system under pressure COMMON ACCUMULATORS: Bladder Accumulators: A one piece steel sphere with a metal diaphragm separating two chambers within the sphere.. At the bottom of the unit is a large screw plug through which the bladder is inserted. Diaphragm Accumulators: A two-piece steel sphere – consists of two metal half ball sections fastened together at the centreline. A synthetic rubber diaphragm separates the two chambers. Piston Accumulators: a piston, inserted in a sealed cylinder, is charged with air below while fluid pressure is maintained above the piston. PUMPS ‘Constant delivery’ type systems are so called because the pump is continually delivering fluid even when the systems are not operating. ‘Constant pressure /Life-Line’ type systems are so called because as the outlet pressure rises beyond the predetermined pressure the actual fluid delivery from the pump is reduced by a pressure operated device. Depending on the degree of pressure required of the pumping mechanism in hydraulic systems, different types of pumps are used. Gear Type Internal or External Gear – used for normal to medium pressure requirements (2100 – 3500 psi) Internal gear pump uses internal gears to pressurize fluid. The pump’s power source causes the internal gears to turn forcing fluid through the pump outlet. External gear pump uses external gears in the same manner as the pump above.

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Fig. 8.9 – Gear Type Pump

Vane Type Used for low pressure requirements (1-150 psi). The vane type pump is used to pressurize fluid. The pump’s power source causes the vane to rotate. As the vale rotates, blades on the vane push fluid out the pump’s outlet.

Fig. 8.10 – Vane Type Pump

Piston Type Pumps used when high pressure systems (3600-6000 psi) are required. Axial-piston type uses an axially-mounted piston to pressurize fluid. Mechanical motion from the pump’s power source moves the piston through a chamber pressurizing the fluid it comes in contact with. The Radial Piston Pump uses pistons mounted radially about a central axis to pressurize fluid.

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Figure 8-11 Piston Type Pump

VALVES Pressure Relief Valves are situated at critical points in the hydraulic system in order to prevent over pressurisation. In the event the pressure regulator becomes inoperative the system pressure will increase and exceed maximum design pressure. As a result the system will fail. In order to avoid this a pressure relief valve is incorporated to relieve the system. The relief valve is always set to operate at a pressure of 100 psi higher than the pressure regulator ‘cut out’ pressure. Pressure adjusting screw

Adjusting screw cap

Pressure adjustment

Compression ring

Compression spring Return port

Return port

Return port

Ball Ball seat sleeve

Pressure port

Pressure port

Ball stud Pressure port Fig. 8.12a – Typical two port pressure relief valve

Fig. 8.12b – Typical four port pressure relief valve

Control valves are usually the ‘on/off’ type rather than the variable throttle type and movement may be achieved by mechanical, hydraulic or electrical means, depending on the application.

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Non-return Valves prevent fluid returning along a pipe carrying fluid to various components.

Outlet port

Spring

Inlet port Ball

Flow direction marking Fig. 8.13 – Non Return Valve

MAIN, STAND-BY AND EMERGENCY SYSTEMS A simplified description of the operation of a hydraulic pump is described below: In a simple pump powered system a control valve transmits pressure from the pump to the hydraulic jack. When selecting ‘down’ at the valve the hydraulic pressure is fed to the ‘down’ side of the jack and the pump will work to maintain system pressure during and after jack travel. Fluid from the un-pressurised side of the jack will be pushed through the return part of the system circuit back into a reservoir. On selection of ‘up’ at the valve, hydraulic pressure is removed from the jack ‘down’ side and applied to the ‘up’ side. Fluid displaced by the retraction of the jack is returned to the reservoir. In the interests of safety and reliability of hydraulic systems, power sources for the primary flying control systems must have a back-up with the capacity to provide continued control for an indefinite period after failure of the primary system. Secondary systems such as undercarriages and brakes must also have back-up with capacity to operate them for one landing, for example: ! Powered flying control units must be able to revert to manual control ! More than one hydraulic source to be connected to the failed powered system – hydraulically controlled primary control systems are powered by at least two totally independent hydraulic systems so that the failure of one does not jeopardize operation of the other.. ! The ‘Primary System’ is usually dedicated to powering the flight control units. ! The ‘Secondary System’ is used to describe the system providing flight control backup and powering other services such as wheel brakes and undercarriages. ! A third system is known as an ‘Auxiliary system’ and

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! The fourth is known as the ‘Emergency’ or ‘Back-up System’. In order to ensure that system operation can be continued after failure of one hydraulic pump two other pump sources are provided: " A pump driven from the aircraft’s normal secondary power system. " One pump powered by an emergency source such as an emergency Power Unit or a Ram Air Turbine. Also in the interests of safety, a shear pin on the drive shaft from the electric motor to the hydraulic pump is incorporated in order to prevent mechanical damage in the event that the hydraulic system jams or that pressure rises too high. This weak link breaks before the other parts are damaged. WARNING SYSTEMS A warning light called an ‘Annunciator’ will flash in the event of hydraulic system failure.

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CHAPTER NINE – PNEUMATICS

CONTENTS OF THIS CHAPTER: !

Power sources

!

Components

!

Pressure systems

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PNEUMATICS INTRODUCTION Pneumatic or air pressure is used in modern aircraft to duplicate aircraft systems, usually for emergency purposes. ! Wheel brakes ! Undercarriage ! Flaps. ! Cabin and cockpit pressurization and heating ! Engine de-icing ! Augmentation of flying controls. The pressures required by pneumatic services are very high - 3 00 psi or higher. Great care must be taken when working with pneumatic hoses and accumulators. Nitrogen is usually used in place of air, as it is inert. ADVANTAGES AND DISADVANTAGES There are advantages and also significant disadvantages to using compressed air over hydraulic or electrical systems: ! Compressed Air Is Lighter ! No Fire Hazard ! Air Is Freely Available ! Air Is Much Cleaner ! Less Chance Of Contamination Or Corrosion However: ! Air compresses quickly and so does not allow the same precise control of power as hydraulics ! The components used in the system are not easily sealed ! Tracing leaks in the system is difficult POWER SOURCES Pneumatic power is generated and stored in a number of different ways, each relevant to the specific end use. The availability of high temperature, high pressure air as a by product of the propulsion system or even the forward motion of an aircraft provides a cost effective source of heat or pressure energy. COMPONENTS Pneumatic systems do not utilize the components familiar to you as incorporated in a hydraulic system such as reservoirs, accumulators, regulators or pumps.

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Control valves - When the services of a particular system are required the appropriate control lever activates the control valve and that system then operates. When not required the control valve remains in the shut off position. Relief valves – To prevent excessive pressure build up relief valves are incorporated in the system. Check valves – to allow air to flow in one direction. The valve closes if the air attempts to reverse direction. Filters – Foreign materials such as metal particles and oil are prevented from entering the system by means of filters. Moisture from the system is removed by means of a separator which contains a pressure sensing device and relief valve. This removes the moisture by blasting air at approximately 3000 psi each time the compressor shuts down. Restrictors – used to reduce the rate of airflow. A large inlet port and small outlet port allows the restrictor to control the rate of airflow and the speed of operation of an actuating unit. There is also a variable restrictor which operates in a similar manner but has an adjustable needle valve which can be turned into or out of a small opening thereby varying the airflow. PRESSURE SYSTEMS High pressure systems – Air, stored at pressures ranging from 1000 to 3000 psi, usually in metal bottles, is used mainly for emergency use such as the operation of landing gear and brakes. These metal bottles have two halves, one being a recharge valve and the other a control valve. The bottles are recharged after use by two or three stage piston type air compressors. Medium Pressure Systems – These systems do not usually include air bottles. Bleed air is taken from the compressor stage of a turbine engine and lead through tubing via pressure controlling units to the operating units. These systems operate at pressures between 100 and 150 psi. Low Pressure Systems – in reciprocating engines continuous pressures ranging from 1 psi to 10 psi are generated by means of a vane pump. It is used to drive some of the aircraft’s gyro instruments.

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NOTES

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CHAPTER TEN – POWER PLANTS

CONTENTS OF THIS CHAPTER: !

A short introduction to the different Engine Types

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POWER PLANTS DIFFERENT ENGINE TYPES Piston engines, turboprops, turbojets, turbofans and all the various types of rocket, all appear to work on different principles. However these propulsion mechanisms, despite their differences, all rely on the same basic principle for their operation: They all propel large masses in the direction opposite to their motion. In order to propel any mass in a direction, a force has to be applied to the mass and in accordance with Newton's third law: An equal and opposite force is then exerted on the object providing the propulsive force. The difference in the engines and hence their application is the relative values of mass accelerated and the velocity to which the mass is accelerated. As shown below, the propeller engine accelerates a large mass to a relatively low velocity while the jet engine accelerates a smaller mass to a much higher velocity.

Large Airflow

Moderate Acceleration

Small Airflow

High Acceleration

Fig. 10 – 1 Fundamental difference between jet and propeller propulsion

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CHAPTER ELEVEN – ENGINE PARTS AND POWER RATINGS

CONTENTS OF THIS CHAPTER: The principle parts of the piston engine: !

Major parts and assemblies " " " " " " " " " " " " "

Cylinder piston Connecting rod Crankshaft Main bearings Lubrication system Stroke Engine displacement Swept volume compression ratio Scavenging of cylinder head Brake horse power Horsepower Torque

!

Engine Power Ratings

!

Engine efficiency

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MAJOR ENGINE PARTS INTRODUCTION Conveniently, the principal parts, and the names for these parts, are essentially the same for all piston engines, regardless of type or configuration. Using an air-cooled, radial engine as a typical example, the major parts and their primary functions are described below. Combustion Chamber Piston Cylinder Spark Plug

Crankshaft Camshaft (2)

Exhaust Valve

Inlet Valve

Crank Case

Magneto Connecting Rod Carburettor

Oil Sump

Fig. 11.1 – General components of a four stroke piston engine

CYLINDER All engines have one or more cylinders. The cylinders are circular, airtight chambers in which a mixture of fuel and air is burned inside the cylinder to provide the pressure needed to force a piston down. The space in which the fuel-air mixture is burned is called the combustion chamber. The inside diameter of the cylinder is called the bore. The closed top of the cylinder is known as the cylinder head. PISTON The piston is essentially a plunger which moves back and forth (or up and down) in the cylinder. At the top of the piston is the piston head. Its sides are known as the skirt. To make the combustion chamber airtight, the piston is provided with steel or cast-iron rings, called piston rings, which fit in piston ring grooves machined around the skirt of the piston. These press against the cylinder wall, preventing oil and the combustion gases from escaping through the small space between the cylinder wall and the piston.

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Circlip

Compression Ring

Crown

Piston Pin

Oil/Scraper Ring Fig. 11.2 – Piston Components

The upper rings, called "compression rings", seal against the leakage of air and vaporised fuel during the compression stroke and the leakage of gases during the power stroke. The lower ring (or rings), called an oil or "scraper" ring, controls the amount of oil on the cylinder wall. The piston has strength holes through each side of the skirt, about equidistant from the top and bottom of the piston, into which fits a piston pin. The connecting rod is attached at the centre of the pin by a hole in the end of the rod through which the piston pin passes. CONNECTING ROD The connecting rod serves as the link between the piston and the crankpin on the crankshaft. The hole at the upper, small end of the connecting rod, which accommodates the piston pin, and the hole at the larger lower end of the rod that fits on the crankpin are each provided with some form of bearing. CRANKSHAFT The heart of a piston engine is the crankshaft. All the pistons are connected either directly or indirectly to this all-important part, which changes the reciprocating (alternately moving backwards and forwards) motion of the pistons to rotary motion. Rotary motion is necessary to drive the propeller. The crankshaft changes the action of the pistons to rotary motion because the crankpin is offset from the centre of rotation of the crankshaft. Copyright © 2012 EAA

Connecting Rod

Piston

Crank Shaft Fig. 11.3 – The piston and crankshaft

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THE CRANCKCASE The crankshaft is housed in the cylindrical shaped crankcase. The crankcase serves the following purposes: ! It houses the crankshaft ! It is a base mounting for the cylinders ! It is oil tight and serves as housing for internal lubricating oil ! It contains thrust bearings which transmit power to the propeller

Crankshaft

Crankcase Oil Sump

Fig. 11.4 – The Crankcase Assembly

MAIN BEARINGS The main bearings of a piston engine hold the crankshaft in the crankcase or cylinder block. VALVES Collets

Cup

Valve Spring

Groove in valve stem for collet

Valve Head

Exhaust valve is often sodium filled to assist with valve head cooling

Fig. 11.5 – The different components of a valve

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Two valves are to be found in each cylinder: ! The inlet valve ! The exhaust valve The inlet valve port allows fuel and air to enter the cylinder where it is burnt off and the exhaust gas then exits through the exhaust valve port. Because valves are required to retain their strength at very high temperatures they are usually made of tungsten or chromium steel. The Camshaft is responsible for the opening of the valves. LUBRICATION SYSTEM When two metal surfaces slide over one another (as in an engine bearing) without lubrication, excessive friction will soon produce enough heat to destroy both surfaces. To prevent this, all metal-to-metal working surfaces within an engine are kept coated with a film of lubricating oil. The film of oil assures a long operating life of the moving parts, because the oil prohibits metal-to-metal contact which, in turn, minimises friction. The oil, in direct contact with both surfaces of a bearing, slides with the surfaces. Friction then occurs only between the layers of the oil itself. The engine lubrication system provides a constant flow of oil in the correct amount and at the required pressure to provide adequate lubrication at all bearing surfaces within the engine. Oil both cools and lubricates the various engine parts. The principle internal parts of a piston engine lubrication system generally consist of: ! an oil sump at the bottom of the engine ! an oil pressure pump for supplying oil at high pressure through lines and passages to the bearings, gears, and other parts ! an oil filter to remove all foreign matter from the oil ! oil jets (orifices) to control the flow of oil under pressure at strategic points through the engine ! an oil scavenge pump (or pumps) to pick up the used oil as it collects in the sump and return the oil (usually through an oil cooler) to the main oil-storage tank. ! the tank (which is outside the engine) collects and stores the oil for reuse in the engine. DEFINITIONS AND POWER RATINGS STROKE The stroke is the distance that the piston travels within a cylinder.

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ENGINE DISPLACEMENT The total volume of the cylinders (full stroke) measured in cu ", cm3 or litres.

TDC BDC

Stroke

Fig. 11.6 – Engine Displacement

SWEPT VOLUME COMPRESSION RATIO The degree of compression given in the piston engine is referred to as the compression ratio, which in simple terms is: the volume in the cylinder with the piston at BDC (bottom dead centre) divided by the volume at TDC (top dead centre). If the compression ratio is quoted as 10 to 1, for example, this means that the original volume is reduced to one-tenth during compression. High compression ratios are easy enough to obtain provided care is taken to allow adequate clearance between the piston crown and the valves, which may be off their seats a little way as the exhaust stroke finishes and the induction stroke begins.

1 10

This definition of compression ratio is perfectly Fig. 11.7 – Compression ratio of 10:1 clear when used in connection with piston engines as being a ratio of two volumes, but unfortunately the same expression is sometimes used in gas turbine terminology for a ratio of two pressures. For anyone dealing with one engine type only, the different meanings of compression ratio are perhaps not important, but when comparisons are made between two engines as different as the turbine and the piston engine care must be taken to make the meaning clear. The difficulty arises because the pressure brought about by reducing the volume is not in simple proportion to the change in volume. SCAVENGING OF CYLINDER HEAD The piston does not completely sweep the cylinder. With the piston at TDC there is still the clearance volume unswept, and therefore unscavenged. If the exhaust valve closes at TDC, as it does in the theoretical cycle, the charge introduced into the cylinder during the next stroke would be contaminated by this volume of burnt gases still left in the cylinder.

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POWER James Watt, the inventor of the steam engine, found that an English work horse could perform at a rate of 33000 foot pounds per minute. Thus, the horse could lift a 1000 pound weight 33 feet in one minute, hence the term horsepower. As power is the amount of work accomplished during a measured interval, power combines force, distance, and time. Thus, the equation for power is: P = Work = F x D where : P = power, F = force, D = distance, T = time. T T

HORSEPOWER Definition – the rate of doing work. HP =

FXD

T X 33 000 Where HP = horsepower, F = force (lbs), D = distance in feet, T = time (minutes). NOTE : IF THE TIME IS IN SECONDS, THE CONSTANT IN THE DENOMINATOR BECOMES 550 INSTEAD OF 33 000. BRAKE (SHAFT) HORSEPOWER Brake horsepower (BHP), or shaft horsepower (SHP), as it is often called, is the useful horsepower delivered outside the engine which is available through a rotating shaft to do practical work, such as drive a propeller. If the effect of friction on the various engine parts and the power required to do such things as open the valves, drive the engine magnetos, fuel pump, oil pump, supercharger and other components, were neglected, brake horsepower and indicated horsepower would be the same. INDICATED HORSEPOWER The theoretical power developed in the combustion chamber of a frictionless engine. TORQUE For an aircraft piston engine, the force part of the equation is the force that the engine is exerting internally to make the engine drive shaft or the propeller shaft turn. This turning force, which comes from within the engine, is called torque.

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The usual unit for measuring torque is the pound-foot (lb-ft), a pound-foot being a force of one pound acting through a moment arm, or radius, of one foot. (force x distance) Pivot point

Applied Force

Torque arm (r) (distance from the pivot point to the point at which the force is applied) Fig. 11.8 – Definition of Torque

ENGINE POWER RATINGS All aircraft engines have what are called "ratings". These ratings are usually the measure of the guaranteed output of the engine under certain specified operating conditions. The ratings, in terms of horsepower, rpm and sometimes manifold pressure, are normally stipulated in the engine specification data sheet published for each type, series and model engine by the engine manufacturer. The ratings are defined by a military specification (Mil Spec) for military engines and by the Federal Aviation Administration (FAA) for commercial engines. In brief, an engine rating represents the maximum power setting which can be used when operating a given type, series and model engine under the specific conditions stated for each of the ratings. Each engine model must pass a qualification test to demonstrate its ability to meet the requirements of each rating, whether military or commercial. As indicated by the following definitions of the standard engine ratings for aircraft piston engines, the ratings for military and commercial engines are not quite the same : Take-Off Power Rating (Military And Commercial Engines) Take-off rated power is the maximum power which may be used for take-off. Use of the rating is time-limited to five minutes. Military Power Rating (Military Engines Only) Military rated power is the maximum power which may be used for a limited period when tactical requirements necessitate a high power setting at the expense of some reduction in the normal time between overhaul (TBO) periods. Use of the rating is time-limited to 30 minutes. Maximum Continuous Power Rating (Commercial Engines Only) Maximum continuous rated power is the maximum power, which may be used continuously for an unlimited time. In practice, the rating is used only in an emergency on large engines.

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On small engines, it may be used for normal operation. Over the past 25 years, the FAA has variously used the expressions "normal rated power", "maximum except take-off (METO) power" and "maximum continuous power". Whenever the take-off power rating and one additional rating are shown for an engine, no matter what the additional rating is named, the additional rating (which may be the same or lower than the take-off rating, but never higher) will be the maximum continuous power rating of the engine. Frequently, as on the typical engine specification data sheet, maximum continuous power will simply be called "rated horsepower". Normal Rated Power (Military Engines Only) Normal rated power for military engines is comparable to maximum continuous power for commercial engines. Specifically, normal rated power is the maximum power, which may be used continuously. MAXIMUM CLIMB AND MAXIMUM CRUISE Although maximum climb and maximum cruise are power settings established for each engine model by the engine maker, they are not considered "ratings" as such, because they need not be approved by the FAA. As their names indicate, these are the maximum power settings (in terms of rpm and manifold pressure) recommended for climb and cruise. Each power setting, from that used for the take-off rating to that for maximum cruise, has a set of operating limits which must not be exceeded when the engine is being operated at that particular setting. The limits are usually applied for engine cylinder-head temperature, oil inlet temperature, oil pressure, torque meter oil pressure (when applicable) and, frequently, fuel pressure. The purpose of the various engine ratings is to conserve the life of the engine and to make the time between engine overhauls predictable. If definite, uniform standards of operation were not provided, a careless or unwitting pilot might literally wear out an engine in a very few hours of flying. This could be dangerous. Working an engine too hard or too long at the higher power settings will cause abnormal deterioration of internal parts. Conversely, the more conservatively an engine is operated, the longer it will last. Strict adherence to the prescribed use of the engine ratings and never exceeding the operating limits will assure safe, dependable engine operation throughout the normal life of the engine between the prescribed engine overhaul periods. ENGINE EFFICIENCY The reason for using different propulsive devices is the differing speed range in which various aircraft are designed to operate. The efficiency of a propulsive device depends on the ratio of the speed of the propelled air and the speed of the aircraft as follows

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Efficiency = 2 (Ve + Va). Va It can be seen that the closer the speed of the propelled air to the speed of the aircraft the higher the efficiency of the system. This variation of efficiency for various propulsive devices is shown in Figure 11-9.

Figure 11 – 9 Variation of efficiency with speed

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CHAPTER TWELVE – PISTON ENGINES

CONTENTS OF THIS CHAPTER: !

Piston engine types

!

Four-stroke cycle " " " "

Induction stroke Compression stroke Power stroke Exhaust stroke

!

Valve timing

!

Valve lead

!

Valve lag

!

Valve overlap

!

Ignition timing

!

Combustion

!

Mixture ratios

!

Temperatures

!

Cylinder pressure in normal operation

!

Cold start

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PISTON ENGINES INTRODUCTION - PISTON ENGINE TYPES The aircraft piston engine comes in many shapes and forms as shown in Figure 12-1. ! The radial engine was introduced in order to reduce engine weight (the crankshaft is the heaviest component) and the radial engine enabled the use of an extremely short crankshaft with only one throw. As speeds increased however, the large frontal area imposed too large a drag penalty and the arrangement was phased out. ! The inverted V and the in-line piston engine have the advantage of a very low frontal area. However they presented problems with landing gear retraction. Also because the frontal area in a single engine aircraft is determined by the cabin cross section, this type of engine did not find great favour. ! The horizontally opposed engine is more popular nowadays. It is typically used with up to 6 cylinders. However, a greater number of cylinders mean that air cooling of the whole engine becomes a serious problem.

Figure 12-1 Various piston engine layouts.

Each of the configurations mentioned has various modes of operation: ! The Diesel or compression ignition engine operates by compressing intake air to a very high pressure. Fuel is then added by injector and ignites because of the high temperature generated by compression of the air. ! The two-stroke engine produces a very high power output for a given weight and would be ideal for aircraft use but for its unreliability and inefficient operation. This is primarily because the fuel and oil need to be mixed together and routed via the crankcase to the cylinder for combustion. ! The four-stroke spark-ignition engine is by far the most common.

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FOUR-STROKE CYCLE – THE IDEAL CYCLE The sequence of operations by which the engine converts heat energy into mechanical energy is known as the four-stroke cycle: ! A mixture of petrol and air is introduced into the cylinder during the induction stroke ! and compressed during the compression stroke. The fuel is ignited at this point and the heat generated causes a rapid increase in pressure. ! This rapid pressure increase drives the piston down on the power stroke. ! Finally, the waste products of combustion are ejected during the exhaust stroke. These four basic operations, which comprise the operating cycle of this type of engine are shown and described in detail below. INDUCTION STROKE ! During the induction stroke, the inlet valve is open and the piston descends in the cylinder. This action lowers the pressure so that the mixture of fuel, vapour and air is forced in by the pressure outside. ! Because of the inertia of the mixture and the flow resistance through the passages of the inlet manifold, the flow of gas tends to lag. ! It is thus not possible to fill the cylinder up to quite the same pressure inside as outside. ! This lower pressure reduces the power output of the engine because the power output depends on the mass of air, which can be drawn in with each stroke. ! The ratio of the weight of gas in the cylinder at the end of the induction stroke to that which would be there if there was no resistance or inertia is called the volumetric efficiency of the engine.

Figure 12-2 Induction stroke

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COMPRESSION STROKE ! As the piston moves upwards, the inlet valve closes and the gas is compressed. ! By squeezing the gas into a smaller space, the pressure it exerts when burnt is increased proportionately. ! The ratio of its volume at bottom dead centre (BDC), i.e. when the piston is at its lowest point of travel in the cylinder, to that at top dead centre (TDC) is known as the compression ratio, which in aero engines is usually in the region of 6.5:1. It should be noted that the compression ratio is a ratio of volumes and is not a measure of the increase of pressure in the cylinder. ! As the gas is compressed it becomes heated adiabatically – in the same way that a bicycle pump warms up in action – as well as by conduction from its hot surroundings. The pressure consequently rises to a higher value than that to be expected from the reduction in volume alone.

Figure 12-3 Compression stroke

POWER STROKE Just before the piston reaches TDC on the compression stroke, a spark ignites the gas. As the flame spreads through the combustion chamber, the intense heat raises the pressure rapidly to a peak value which is reached when the piston is about 10° past TDC. The gas continues to burn and its pressure falls as the piston is forced down until, towards the end of the power stroke, combustion is complete and the pressure on the piston is comparatively small.

Figure 12-4 Power stroke

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EXHAUST STROKE With the exhaust valve open, the piston ascends, forcing out the spent gases. Here again it is important that the flow should be as free as possible. An obstruction would not only exert a backpressure on the piston, but it would also result in an undesirable amount of burnt gas remaining in the cylinder. This would contaminate the fresh charge brought in during the next induction stroke. At the end of the exhaust stroke the exhaust valve closes, the inlet valve opens, and the cycle begins again.

Figure 12-5 Exhaust stroke

VALVE TIMING Volumetric efficiency depends to a large extent on valve timing. For best volumetric efficiency, valves are not timed to open and close at the beginning and end of the respective piston strokes. They open before the stroke begins and close after the stroke ends. EXAMPLE ! In a particular reciprocating engine the intake valve opens 19° ‘before top centre’ (BTC) of the intake stroke (the piston is actually at the end of the exhaust stroke) and closes 62° ‘after bottom centre’ (ABC) of the intake stroke (the piston is actually at the beginning of the compression stroke). ! Thus the intake valve is off its seat for a total of 261° of crankshaft travel (19°+180°+62°). ! The exhaust valve, however, opens 60° ‘before bottom centre’ (BBC) of the exhaust stroke (the piston is actually on the power stroke) and closes 20° ‘after top centre’ (ATC) of the exhaust stroke (the piston is actually on the next subsequent intake stroke). ! Thus the exhaust valve is off its seat for a

FIGURE 12-6 VALVE T IMING

total of 260° of crankshaft travel (60°+180°+20°). Copyright © 2012 EAA

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These figures are used in the following text to express other factors, such as valve lead and valve lag, and to find valve overlap and the total degrees of crankshaft travel the valves would be on their seats. VALVE LEAD The early opening (prior to the actual piston stroke) of a valve is known as valve lead. In the preceding valve-timing example the intake valve lead is 19° and the exhaust lead is 60°. VALVE LAG The late closing (after the end of the actual piston stroke) is known as valve lag. In the preceding valve-timing example the intake valve lag is 62° and the exhaust valve lag is 20°. VALVE OVERLAP ! In a four-stroke cycle there is a short period of crankshaft travel during which both valves are open (or off their seats) at the end of the exhaust stroke and at the beginning of the intake stroke. This is known as "valve overlap". ! In the preceding valve-timing example the intake valve opened 19° before TDC and the exhaust valve closed 20° after TDC. ! Valve overlap in this engine is 39° (19° + 20°). ! Valve overlap permits fresh fuel-air mixture to enter the cylinder early. This helps to scavenge (or push out) what is left of the exhaust gas from the preceding cycle. ! Valve overlap thus aids in increasing the volumetric efficiency of the engine cylinder as it ensures that fewer burnt gases are left in the cylinder from one cycle to the next. IGNITION TIMING Once the spark ignites the mixture, the flame travels steadily but rapidly through the mixture. The mixture does not burn all at once, but takes some time to burn through completely. The actual speed of combustion depends on : ! The compression ratio - a highly compressed mixture burns more rapidly than one less compressed. ! The quality of the mixture - in other words, how evenly the fuel is distributed in the air. This depends on how much swirling, generally known as gas turbulence occurs in the cylinder. It has only quite recently been discovered that if the compressed mixture was not swirling round in the cylinder, the high speeds now obtained would be impossible, so designers set about defining main induction passages, shapes of valves and compression spaces of such a shape that the mixture will go on swirling while being compressed. In this way very rapid burning of the whole volume is obtained.

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COMBUSTION FACTORS NECESSARY FOR EFFICIENT COMBUSTION Many of the required properties of fuels are more easily studied when considering their uses. Although piston engines and turbines burn similar fuels, different burning and operating conditions generally demand separate consideration: ! Liquid fuels must be volatile as conversion to the gaseous state must take place before burning. ! When carburettors are used to supply the fuel, it is widely believed that it is best to complete evaporation before the mixture enters the cylinders. Otherwise there is a risk that the cylinders will not receive an equal share of the fuel or additives. Additives improve certain aspects of performance. These additives will be discussed later but it is interesting at this stage to look at the problem of sharing out the minute quantities of these substances to the cylinders: ! One of the best known "dopes" is tetraethyl lead which increases resistance to detonation but which must be supplied along with ethylene debromide to counteract the ill effects of the lead. ! Ethylene debromide tends to evaporate with the lighter elements of the fuel leaving tetraethyl lead behind with the liquid fuel. When evaporation in the induction system is incomplete, distribution of the remaining liquid between the cylinders is seldom equal. This can result in some cylinders being deprived of tetraethyl lead and suffering from detonation whilst others are supplied with too much tetraethyl lead and insufficient ethylene debromide to safeguard the valves from lead deposits. If distribution is not right an excellent fuel may do more harm than good. So as well as deciding which fuel properties are desirable, it is also important to appreciate how the fuels should be used. MIXTURE RATIOS CHEMICALLY CORRECT MIXTURE Carburetion, as applied to the internal combustion engine, is the process by which air and fuel vapour are mixed in suitable proportions and the supply of this mixture regulated according to the requirements of any given operating condition. Although air and fuel vapour will burn when mixed in proportions ranging between roughly 8 to 1 (8 parts air and 1 part fuel) and 20 to 1 (by weight), complete combustion occurs only at an air/fuel ratio of about 15 to 1. With a chemically correct mixture, all the hydrogen and carbon in the fuel combine with all the oxygen in the air to form carbon dioxide and water vapour. The atmospheric nitrogen takes no active part in the combustion process beyond moderating the rate of burning. The chemically correct mixture, however, does not give the best results because the temperature of combustion is so high that power is lost through a phenomenon known as dissociation. Copyright © 2012 EAA

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Some degree of detonation also occurs at such temperatures and the loss of power is thereby aggravated. Dissociation is a momentary splitting-up of the products of combustion into their separate elements, during which heat is absorbed which would otherwise have helped to raise the combustion pressure. Although these elements re-combine later in the power stroke and the lost heat is regained; it is too late to be of much value. RICHER MIXTURE With a mixture about 10% richer than the chemically correct one, the excess petrol absorbs sufficient latent heat during vaporisation to obviate dissociation and detonation at moderate power outputs. But, as there is a tendency for some cylinders in an engine to receive a weaker mixture than others owing to the difference in inertia between air and petrol droplets, it is the practice to err on the safe side by supplying a mixture about 15% richer at normal cruising powers. Since engine power is a product of engine speed and the mean effective pressure, higher power outputs involve an increase in either or both of these factors. Such increase however, involves an increase in mixture temperature and therefore in the tendency to detonation. M E P = (the average pressure in the cylinders during one cycle) Therefore, when higher power is required, as in climbing, the mixture is further enriched to about 20% above chemically correct, to obtain the necessary cooling. For take-off, when maximum power is employed, the figure rises to 30% or higher. Apart from its cooling function, the fuel required at high power is wasted because there is no oxygen available to burn it. In practice, excess fuel vapour is not exhausted as such. The oxygen is shared out to some extent, so that carbon monoxide is produced as well as carbon dioxide. With very rich mixtures, however, some carbon particles fail to unite with oxygen at all and are exhausted as black smoke. WEAKER MIXTURE Cooler burning is also obtained with a mixture weaker than the chemically correct one, partly because less petrol is burnt per power stroke and also because the rate of burning slows down (owing to the greater proportion of inert gas in the cylinder). With extremely weak mixtures, the flame rate may be so slow that combustion is still taking place when the inlet valve next opens and "popping back" tends to occur in the inlet manifold. For these reasons, power tends to decrease as the mixture weakens from the chemically correct ratios. As engine speed is reduced, however, the exhaust gas velocity falls and more combustion products are left behind in the cylinder, whilst at still lower speeds there is even a tendency for gas to be sucked back by the descending piston through the exhaust valve before it closes. The consequent dilution of the free charge is such that, to maintain smooth running, a progressively richer mixture must be supplied as idling speed is approached.

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EFFICIENCY A correct mixture of fuel and air is one in which there is just the right amount of air to burn the fuel completely and leave no unconsumed oxygen in the products of combustion. When a correct mixture is completely burned the exhaust gases consist of steam (formed by the combination of hydrogen from the fuel with oxygen from the air), carbon dioxide (formed by the combination of carbon from the fuel with oxygen from the air) and nitrogen. A weak mixture is one in which there is more air than the minimum required for complete combustion. In a weak mixture it is usual for all trace of combustible matter from the fuel to burn up completely leaving some unconsumed oxygen in the exhaust. The weak mixture would appear to be just what is required for economy and, within limits, this is so. However, like many another good thing it can be overdone. Too weak a mixture carries through the engine a large excess of air. This air has to be compressed and pushed out again without contribution to the useful work. A weak mixture also burns slowly and allows the piston to pass TDC before combustion is complete, so that the maximum pressure is less than it would have been if all combustion had been completed before TDC. In weak-mixture running, the gases may still be burning when the exhaust valves open, so exposing the valves to much higher temperatures than would have prevailed with complete combustion before expansion. It is this slow burning which accounts for signs of overheating with weak mixtures, as a certain amount of the heat never has a chance to be converted to work by expansion and has to be carried away by the cooling system. A rich mixture is one in which there is not enough oxygen to unite with all the fuel and some combustible matter is carried away unconsumed in the exhaust system. Naturally a volatile fuel like petrol does not survive as a liquid in the furnace-like conditions of the combustion chambers. What usually happens is that there is insufficient oxygen for all the carbon to yield carbon dioxide and some carbon monoxide appears in the exhaust. If there had been plenty of oxygen the carbon monoxide would have united with it to form carbon dioxide and in the process it would have released quite a lot more heat. In the case of extremely rich mixtures carbon, in the form of soot or black smoke, can be seen coming from the exhaust. In spite of the apparent wastefulness of rich mixtures they are deliberately used for certain running conditions. For example: ! When full power is needed, it is usual to supply a slightly rich mixture to ensure that every trace of oxygen is used up and the maximum heat extracted from the mixture taken into the engine. Engine power really depends on the amount of air "swallowed" per minute. It is essential that all the air contributes to the power. In a subsequent section it will be shown that it is not an easy matter to increase the air consumption beyond a certain amount, hence the anxiety to make full use of all that does get in. As the evaporation of fuel takes up heat from its surroundings, the excess fuel does a useful job by helping to cool the inside of the cylinders at a time when the cooling system is on full stretch because the engine is developing maximum power.

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! At the other end of the power scale, when the engine is idling, a slightly rich mixture is again supplied. There may only be enough heat in this running condition to evaporate the lighter elements of the fuel. Without the excess fuel not enough vapour would be available at the right time. Depending on the length and shape of the manifolds, certain cylinders may be starved of fuel or only receive heavier un-evaporated elements unless the supply is generous enough to make sure that every cylinder gets sufficient vapour to start combustion. TEMPERATURES Inevitably some of the heat of combustion passes to components which are exposed to the flames. As gas temperatures are high enough to burn valves, valve seats, sparking plugs and pistons, arrangements must be made to remove heat from these parts and transfer it to the coolant and the cooling air. As the heat carried away by the cooling system has done no useful work, everything possible must be done to minimise it. A compact combustion chamber which has minimum metal surrounding the gases reduces the path through which heat may pass. The coolant flow should be no more than is absolutely necessary to keep the vulnerable parts at a safe temperature. Overcooling only increases heat loss and may also cause condensation inside the engine, contaminating the oil and increasing the rate of wear. As the spark which starts the burning is usually timed to take place about 20° before TDC, combustion has a good chance of being completed before the piston has moved very far along the power stroke. The heat energy in the gas is reduced during the power stroke by the amount of work done on the piston and it would seem good business to reduce the store of heat as much as possible in this way. However, in a conventional engine the use of a high expansion ratio to exploit the heat in the gases involves the use of an equally high compression ratio. For example, if the gases in the clearance volume at the beginning of the power stroke are expanded to 10x that volume in an effort to make the best use of the heat, the fresh charge will in due course have to be compressed to 1\10 the volume occupied at the beginning of the compression stroke. Such a restriction on the amount of fuel limits the power available and is a point which must be monitored during the design and development stages – the search for higher efficiencies might in certain circumstances result in lower power! CYLINDER PRESSURE IN NORMAL OPERATION MANIFOLDS, MANIFOLD PRESSURE AND MANIFOLD TEMPERATURE On most engines, a system of intake pipes designed to offer a minimum of internal air resistance conducts the fuel-air mixture from the carburettor or the supercharger to the intake ports of the cylinder. The configuration of the piping and the number of pipes depends upon the number of cylinders and general design of the engine. The intake pipes and the associated parts of the system comprise what is known as the intake manifold. On all but the smallest engines, the

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pressure in the intake manifold is measured in inches of mercury (" Hg) by an instrument called a manifold pressure gauge on the aircraft instrument panel. Manifold pressure is an important operating parameter because it is a primary indication of the weight of the fuel-air charge going into the cylinders. On un-supercharged engines, the pressure in the intake manifold is normally less than the pressure of the surrounding atmosphere. When the engine is supercharged, manifold pressure is generally higher than the atmospheric pressure at medium to high engine powers and less than atmospheric pressure at low engine powers. Manifold temperature also influences the weight of the air charge which enters the cylinders. On supercharged engines, the manifold temperature is normally lower than the surrounding atmospheric air temperature because of the drop in temperature which occurs when fuel is vaporised in the air passing through the carburettor. In supercharged engines, the manifold temperature is usually higher than atmospheric temperature because a temperature rise occurs when the supercharger compresses the air (or fuel-air mixture) entering the engine. Although manifold temperature determines, in part, the weight of the charge entering the cylinder, the temperature is not indicated in the aircraft because it is difficult to measure accurately. It also bears a relatively fixed relation to engine rpm and atmospheric air temperature for any given model of engine. So, although manifold temperature is of interest to an engine performance engineer, it is of no special value to pilots. COLD START Starting from cold, particularly extreme cold, may present difficulties with low volatile fuel unless steps are taken to either: ! heat the engine, ! prime the cylinders with highly volatile fuels, or ! dilute the lubricating oil so that the crankshaft may be spun more briskly to improve the prospects of getting a start. (A fast cranking speed helps both the carburettor and the ignition system to work properly and there is less time for the fuels to condense in the cylinders.) A high volatile fuel helps the starting process by vaporising readily. Unfortunately this same quality may result in vapour forming where it is not wanted in the pipelines and passages of the fuel supply system to such an extent that the flow of fuel is obstructed. SUMMARY Although fuel properties are vitally important, the main purpose of feeding fuel to the engine is to release heat. Naturally it is expected that every trace of combustible matter is used up and that the engine design is such that a worthwhile proportion of heat is converted to mechanical energy. The way in which the mixture of fuel and air burns in the cylinder depends to some extent on the proportions of fuel and air in the mixture and certain standard mixtures are recognised as a basis for discussions on performance. Copyright © 2012 EAA

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CHAPTER THIRTEEN – DETONATION AND PRE-IGNITION

CONTENTS OF THIS CHAPTER: !

Effects of detonation

!

Causes of detonation

!

Pre-ignition

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DETONATION AND PRE-IGNITION DETONATION After ignition, the flame normally travels smoothly through the combustion chamber until the charge is all burned. The rate of burning may be as high as 60 ft/sec. This may seem very fast in view of the size of the cylinder but nevertheless, it is a steady burning process. Combustion is relatively silent, with a regular pressure rise and a steady push on the piston. However, when detonation occurs, combustion begins normally, but at an early stage the temperature and pressure of the unburned portion of the mixture is raised so high that it ignites spontaneously. The flame velocity is approximately 1 000 ft/sec. The cylinder walls and piston receive a hammer-like blow (knocking) which gives rise to the characteristic "pinking" noise. Motorists will be familiar with this sound, although it will not be heard while airborne because of propeller and other noises. The rate of pressure rise is too great to be accommodated by movement of the piston. This means that much of the chemical energy released is wasted as heat, instead of being transformed into mechanical power. A schematic of the normal and detonating combustion processes is shown in Figure 13-1 below.

Figure 13-1 Normal combustion results in a smooth pressure rise

Figure 13 – 2 Detonation or explosive combustion results in shock

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EFFECTS OF DETONATION ! Loss of power which follows the reduced thermal efficiency is accompanied by a rise in cylinder temperatures. This could be excessive at certain positions. ! Burning of the piston crown and exhaust valves will result from prolonged or severe detonation. ! In addition, the high maximum combustion pressures can lead to the collapse of the piston crown and loosening of the valve seats. ! Also to be expected are carbonising of the oil in piston-ring grooves (with consequent burring of the piston walls) and ! Oil vaporisation because of the hot gases forcing their way past gummed-up rings into the crankcase. CAUSES OF DETONATION Detonation is caused by: ! excessive temperatures and pressures within the cylinder ! These excessive temperatures and pressures are in turn caused by: " high manifold pressures " high intake air temperatures " an overheated engine. ! using fuel with insufficient "knock resistance" (octane rating too low). ! Anything that raises the temperature or pressure of the mixture before it is burned, such as the use of warm air (carburettor heat) at high power, overheated cylinders or ‘hot spots’ inside them, and high manifold pressure (boost) with unusually low rpm, is likely to cause detonation. This last factor causes detonation because, at the reduced induction velocity, the resistance to flow is less and consequently the volume efficiency is greater. Therefore, at a given boost a greater weight of charge is admitted per stroke and higher cylinder pressures are obtained. In addition, because of the lower engine rpm, the fuel air mixture will have a longer period of contact with the cylinder walls and will be heated to higher temperatures increasing the chances of detonation occurring. Fuels differ considerably in their resistance to detonation, depending on the composition. Fuels with high octane ratings allow an increase in compression ratio (increasing thermal efficiency) with a resultant gain in economy and increased power. These fuels also permit an increase in permissible boost pressure and therefore greatly increased power. The power output of an engine is almost directly proportional to the weight of air consumed per unit time. Higher boost pressure increases this weight. Engines must be designed or modified to take advantage of the higher grade fuels. Using higher grade fuels in a low-performance engine will not produce more power or even greater economy. It may, on the other hand, cause fouling of the cylinders and eventual mechanical failure.

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Detonation results in engine vibration and excessive cylinder-head temperatures (CHT). If detonation is allowed to continue over an extended period of time, structural damage to the engine will result. As detonation can almost never be recognised in an aircraft, protection must be taken against its possible occurrence. Operational control of detonation can usually be achieved by selecting and using the correct approved grade of fuel and by observing recommended operating limits for manifold pressure, rpm, supercharging and CHTs. One of the most important variables in the control of detonation is the proportion of fuel added to a given quantity of air. An extra rich fuel-air ratio lowers the temperature of the charge in the cylinder and lowers the rate of combustion, keeping the unburned portion of the charge cooler. The fuel-air ratio necessary to protect an engine is determined by test. SUMMARY Detonation should be avoided. Certain, easily followed, techniques which can be used to control suspected detonation are: ! enrich the mixture – this reduces combustion and so engine temperatures ! reduce power (by reducing rpm and/or manifold pressure according to manufacturers instructions) ! ensure that the carburettor heat is in the COLD position ! open cowl flaps and the oil-cooler doors, if installed ! lower the nose and accelerate to a higher airspeed. This will increase the flow of cooling air over the engine and lower the engine temperatures. PRE-IGNITION Pre-ignition is (as the name implies) ignition of the combustible mixture prior to being ignited by the spark plug. It is caused by local "hot spots" within the combustion chamber which ignite the combustible gases. Typically, these hot spots would be red hot carbon deposits on the spark plugs, or anywhere else inside the cylinder. Hot spots result from using an incorrect mixture, incorrect ignition, incorrect timing and/or overheating. Pre-ignition reduces engine power output and causes engine roughness because the pressure in the cylinder rises as a result of the burning gases well before the cylinder reaches top dead centre. Pre-ignition differs from detonation as the flame front proceeds at the normal rate after ignition and does not proceed explosively as it does in the case of detonation.

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CHAPTER FOURTEEN – LUBRICATION AND COOLING

CONTENTS OF THIS CHAPTER: ! The Oil System " Oil system components " Turbine engine dry-sump lubrication " Turbine engine wet-sump lubrication " Lubrication methods " Pressure pumps " Scavenger pumps " Oil pressure relief valve " Oil pressure gauge " Viscosity " Characteristics of aircraft oil " Functions of engine oil

! The Cooling systems " Liquid cooling " Air cooling

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LUBRICATION AND COOLING INTRODUCTION When two surfaces are in contact and moving there is always friction because, even apparently smooth surfaces, have minute imperfections. In order to overcome the potential for development of intense heat, lubricating of the affected parts is essential. The primary method of cooling would be oil lubrication. There are two basic phases of lubrication – film lubrication where the surfaces are separated by a substantial quantity of oil and boundary lubrication where the oil film may be only a few molecules thick. In general, the parts to be lubricated and cooled include the main bearings and accessory drive gears and the propeller gearing in the turboprop. OIL SYSTEM COMPONENTS The oil system components used on gas turbine engines are: ! Tanks. ! Pressure pumps. ! Scavenger pumps. ! Filters. ! Oil coolers. ! Relief valves. ! Breathers and pressurizing components. ! Pressure and temperature gages lights. ! Temperature-regulating valves. ! Oil-jet nozzle. ! Fittings, valves, and plumbing. ! Chip detectors. Not all of the units will be found in the oil system of any one engine. But a majority of the parts listed will be found in most engines. The major difference between a wet-sump system and a dry-sump system is the location of the oil reservoir. TURBINE ENGINE DRY-SUMP LUBRICATION In a turbine dry-sump lubrication system, the oil supply is carried in a tank mounted externally on or near the engine. With this type of system a larger oil supply can be carried and the oil temperature can be controlled. An oil cooler is usually included in a dry-sump oil system. It is designed to furnish a constant supply of oil to the engine. This is done by a swivel outlet assembly mounted inside the tank, a horizontal baffle mounted in the centre of the tank, two flapper check valves mounted on the baffle, and a positive-vent system.

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All oil tanks have expansion space. This allows for oil expansion after heat is absorbed from the bearings and gears and after the oil foams after circulating through the system. Some tanks also incorporate a de-aerator tray. The tray separates air from the oil returned to the top of the tank by the scavenger system. Usually these de-aerators are the "can" type in which oil enters a tangent. The air released is carried out through the vent system in the top of the tank. In most oil tanks a pressure build-up is desired within the tank. This assures a positive flow of oil to the oil pump inlet. This pressure build-up is made possible by running the vent line through an adjustable check-relief valve. A dipstick is usually used to measure the quantity of oil or, in some aircraft; electric quantity indicators give direct readings on a gauge in the cockpit.

Feed Oil Return Oil

System Relief Valve

Oil Tank

Pump Relief Valve

Figure 14 – 1 Oil Tank

TURBINE ENGINE WET-SUMP LUBRICATION In some engines the lubrication system is the wet-sump type. Because only a few models of centrifugal-flow engines are in operation, there are few engines using a wet-sump type of oil system. The reservoir for the wet-sump oil system is the sump mounted on the bottom of the accessory case or lower crankcase. A bayonet-type gage indicates the oil level in the sump. Two or more finger strainers (filters) are inserted in the accessory case for straining pressure and scavenged oil before it leaves or enters the sump. These strainers aid the main oil strainer. A vent or breather equalizes pressure within the accessory casing.

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A magnetic drain plug may be provided to drain the oil and to trap any ferrous metal particles in the oil. This plug should always be examined closely during inspections. The presence of metal particles may indicate gear or bearing failure. A temperature bulb and an oil pressure fitting may be provided. LUBRICATION METHODS Pressure Lubrication ! In a pressure lubrication system, a mechanical pump supplies oil under pressure to the bearings ! Oil flows into the inlet of the pump through the pump and into an oil manifold which distributes it to the crankshaft bearings ! Although pressure lubrication is the principle method of lubrication on all aircraft engines, some engines use splash lubrication also Splash Lubrication and Combination Systems ! Splash lubrication is never used by itself. It is so called due to the moving parts picking up oil and ‘splashing’ it around. ! All lubrication systems are pressure systems or combination pressure/splash systems Mist or Spray Lubrication ! High pressure oil is fed to the main bearings by pipes and ducts and then into the hollow crankshaft via holes drilled into the journals. ! Oil passes through holes in the crankpins to the big end bearings and then escapes in the form of a mist or spray due to the centrifugal forces of the rotating crankshaft. PRESSURE PUMPS ! Oil pressure pumps may be the gear type or vane type ! The gear type pump is used in the majority of reciprocating engines and uses close fitting gears that rotate and push the oil through the system The gear-type pump consists of a driving and a driven gear. The engine-accessory section drives the rotation of the pump. Rotation causes the oil to pass around the outside of the gears in pockets formed by the gear teeth and the pump casing. The pressure developed is proportional to engine RPM up to the time the relief valve opens. After that any further increase in engine speed will not result in an oil pressure increase. The relief valve may be located in the pump housing or elsewhere in the pressure system for both types of pumps. SCAVENGER PUMPS These pumps are similar to the pressure pumps but have a much larger total capacity. An engine is generally provided with several scavenger pumps to drain oil from various parts of the engine. Often one or two of the scavenger elements are incorporated in the same housing as the pressure pump.

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OIL PRESSURE RELIEF VALVE To allow for oil flow in the event of filter blockage, all filters incorporate a bypass or relief valve as part of the filter or in the oil passages. When the pressure differential reaches a specified value (about 15 to 20 psi), the valve opens and allows oil to bypass the filter. Some filters incorporate a check valve. This prevents reverse flow or flow through the system when the engine is stopped OIL PRESSURE GAUGE ! An oil pressure gauge is an essential component of any engine oil system ! These gauges generally use a bourdon tube to measure the pressure ! They are designed to measure a wide range of pressures Oil pressure is registered from that point where oil enters the engine. Oil, under pressure, flowing into a semi bent tube, attempts to straighten the tube. This straightening force is amplified through a system of gears and levers and positions a needle on a calibrated dial in p.s.i. VISCOSITY ! Viscosity is technically defined as the fluid friction of an oil or: ! More simply, it is the resistance an oil offers to flowing ! Heavy-bodied oil is high in viscosity and pours or flows slowly ! At high temperatures oil becomes thin (low viscosity) CHARACTERISTICS OF AIRCRAFT OIL ! It should have the proper body (viscosity) ! High antifriction characteristics ! Maximum fluidity at low temperatures ! Minimum changes in viscosity with changes in temperature ! High anti-wear properties ! Maximum cooling abilities ! Maximum resistance to oxidation ! Non-corrosive FUNCTIONS OF ENGINE OIL ! Lubrication, thus reducing friction ! Cools various engine parts ! Seals the combustion chamber ! Cleans the engine ! Aids in preventing corrosion ! Serves as a cushion between impacting parts

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THE COOLING SYSTEM INTRODUCTION The job of the cooling system is to prevent damage to the engine parts which could result from high temperatures. The intense heat generated when fuel and air are burned mandates that some means of cooling be provided for all internal combustion engines. Reciprocating engines are cooled either by passing air over fins attached to the cylinders or by passing a liquid coolant through jackets surrounding the cylinders. LIQUID COOLING In liquid cooled engines the oil cooler is used to reduce oil temperature by transmitting heat from the oil to another fluid usually a mixture of water and ethylene glycol. Since the fluid flow through the cooler is much greater than the oil flow, the fuel is able to absorb a considerable amount of heat. This reduces the size and weight of the cooler. Thermostatic or pressure-sensitive valves control the oil temperature by determining whether the oil passes through or bypasses the cooler. AIR COOLING There are three main parts of the cooling system on any air-cooled aircraft: ! high pressure air, low pressure air, and airflow. The high pressure air is typically produced at the front of the engine by the propeller forcing air in the cowling air inlets. This forced air is at a higher pressure than the static air around the aircraft. There are two types of air cooling systems on air-cooled piston engine general aviation aircraft. ! One type is when air from a high pressure area is forced down between the cylinders to a low pressure area; this type is called "downdraft". ! The "updraft" is when the high pressure air below the engine is forced up between the cylinders to the low pressure area. Both of these systems function the same way and they both require baffling boxes in the high pressure area underneath the cowling. The baffle seals (usually a type of rubber) meet with the cowling to form a closed seal to prevent leaking. The aluminium supports box in the high pressure air and give strength and flexibility to the seals. Most cooling problems are due to a lack of high pressure from old worn baffle seals, but not always. Engine cooling is also dependent on the low pressure area - the low pressure area is due to gaps (intentional) at the bottom of the cowling. The faster moving air passing by the bottom of the cowling picks up the slower moving air from inside the cowling creating the low pressure area. There are a couple of keys to the low pressure area. The cooling air exit at the bottom of the cowl needs to be the same surface area as the air inlet in most cases. To generate enough low air pressure at slow airspeeds cowl flaps are usually added to the aircraft in some instances. If there is not enough low pressure air the aircraft will experience cooling problems. The combination of high air pressure and low air pressure creates air flow.

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CHAPTER FIFTEEN – CARBURETION AND FUEL INJECTION SYSTEMS

CONTENTS OF THIS CHAPTER: ! The carburettor " The float chamber " The main Jet " Diffuser " Accelerator pump " Idling jet ! Mixture Control ! Carburettor Icing ! Fuel Injection Systems " Advantages " Air throttle assembly " Fuel control assembly " Fuel injection pump " Fuel manifold valve " Fuel discharge nozzles

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CARBURETION AND FUEL INJECTION SYSTEMS INTRODUCTION Carburetion is the process of supplying vaporised air and fuel mixed in suitable proportions for combustion. Fuel/air vapour burns when mixed in proportions from 8:1 to 20:1 (air-to-fuel by weight). The most efficient combustion (complete combustion of all oxygen and fuel present) occurs at a ratio of 15:1. This is known as the "chemically correct mixture" or stochiometric ratio. However, in practice, it is not possible to operate at this value because of uneven mixing of the fuel/air mixture. A more common ratio is about 13:1. Engines either have a carburettor or some form of fuel-injection system. THE CARBURETTOR THE FLOAT TYPE CARBURETTOR A schematic diagram of a typical float type carburettor is shown in Figure 15-1.

Figure 15 – 1 Typical simple float type carburettor

The carburettor is attached to the engine inlet manifold. It supplies atomised air and fuel in the correct ratio to the engine. Air flows into the carburettor from the air inlet after passing through the air filter through the choke tube or venturi. In the choke tube the pressure of the flowing air is reduced according to Bernoulli's principle. As a result, fuel is forced into the air stream by the higher atmospheric pressure in the fuel line and float chamber. The fuel enters as a fine spray in order to assist in the evaporation process. The fuel/air mixture is then distributed to the cylinders via the inlet manifold during the induction stroke. The throttle or “butterfly” valve controls power output of the engine by restricting the rate of airflow through the carburettor and hence to the engine.

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Figure 15 – 2 Butterfly valve action under different operating conditions

THE FLOAT CHAMBER The float chamber acts as a fuel reservoir. The fuel flows through a filter and into the float chamber. The flow is controlled by a needle valve attached to the float. As the level of the fuel rises so does the float and in the process the valve closes. With the drop in fuel the float drops and the valve opens again allowing fuel to enter the chamber. A vent in the chamber allows the atmospheric pressure to push the fuel out through the diffuser which is located in an area of lower pressure. THE MAIN JET In order to limit the fuel flow once the throttle butterfly valve is fully open the fuel must pass through the main metering jet before reaching the discharge nozzle. The rate of fuel discharge depends on the size of the hole of the jet. A large jet will produce a rich mixture and therefore more fuel while a small jet will produce a lean mixture and less fuel. DIFFUSER / DISCHARGE NOZZLE The flow characteristics of fuel and air differ. As the throttle is opened and the air demand increases, the proportion of fuel to air rises. This tendency is corrected by air bleeding through a diffuser. The diffuser is located in the throat of the venturi. It is a narrow tube with very small holes in the sides which allow air to be drawn in and to be pre-mixed with the fuel before it is discharged from the end of the nozzle. When the engine is running the fuel level falls by an amount which increases as the velocity of the air in the choke tube increases. As the throttle is opened the holes in the diffuser passage are slowly uncovered and the depression over the fuel is reduced sufficiently to maintain a constant fuel/air ratio. Figure 15 – 3 Diffuser action Copyright © 2012 EAA

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In addition, the air entering through the diffuser holes mixes with the fuel and helps to atomise it. The action of the diffuser is shown in Figure 15-3. Advantages of a diffuser: ! It improves vaporization of the fuel ! it achieves an even mixture ratio over different throttle settings ! it reduces the required fuel metering force To supply the rich mixture required at high power, extra jets, controlled by cam operated valves are fitted. Their output supplements the main jet OUTPUT. ACCELERATOR PUMP The accelerator pump can give an added spurt of fuel into the choke tube. This is required to enable smooth and rapid response of the engine with rapid opening of the throttle. The system is necessary because fuel is denser than air and takes longer to respond to throttle movements. The extra fuel from the accelerator pump prevents weakening of the mixture through the intake of too high a proportion of air. SLOW RUNNING (IDLING SYSTEM) The idling jet supplies a mixture richer than the main jet. At low engine speeds, with the butterfly nearly closed, the volume of air passing into the engine is so small that the pressure drop in the choke is insufficient to draw fuel past the nozzle. In addition, the fuel which is drawn into the air stream does not mix properly with the air. An enriched mixture at idle speed is necessary to ensure that the engine is supplied with a combustible mixture. Figure 15 – 4 Operation of the idle jet

Above the butterfly valve considerable suction exists and permits a fuel supply through the slow running inlet. This inlet, with its own idling jet, allows fuel to be drawn into the engine for slow running. The size of the jet is such that the rich mixture necessary under these conditions is obtained. An air bleed opening into the choke below the butterfly valve assists atomisation. A cut-out valve (idle cut-off) is used to shut down the engine. It is particularly necessary when the engine is so hot that it would continue to fire even with ignition off. The valve leans the mixture by introducing atmospheric air. It leaves the carburettor and the fuel lines full, for ease of starting.

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MIXTURE CONTROL A mixture control system overcomes the problem of a rich mixture and consequently a decrease in power and economy at high altitudes. At high altitudes/high temperatures the air is less dense and as a result the volume of air passing through the carburettor weighs less. However, the density of the fuel does not change. This results in too much fuel for the amount of air. In order to correct this, the amount of fuel entering the carburettor venturi must be reduced by ‘leaning’ the mixture. A venturi type carburettor meters fuel in accordance with the volume of air passing through the choke, ensuring that the same amount of fuel is drawn through the discharge nozzle at all altitudes. Because it is not sensitive to changes of air density passing through it this characteristic produces a steady richening of the mixture as altitude is gained. The fuel input remains the same and therefore the air/fuel ratio becomes too rich. TYPES OF MANUAL MIXTURE CONTROL ! The Needle Type mixture control – operated by means of a lever in the cockpit. A needle valve found between the fuel chamber and the main jet is raised or lowered which increases or decreased the opening in the valve. " Full rich – valve wide open, leaner mixture – " Valve half closed and " Idle cut off – needle tightly closed. ! The back suction mixture control – an adjustable control valve which is located in the float chamber vent line. It regulates the pressure in the float chamber. Atmospheric pressure is drawn from the atmospheric vent and low pressure is drawn through a vent channel from the venturi. " Valve fully open – maximum atmospheric pressure, maximum pressure differential between bowl and diffuser – maximum fuel flow to the diffuser – full rich mixture setting. " Valve slowly closed – atmospheric pressure restricted – lower pressure gradient between bowl and diffuser – leaner mixture. CARBURETTOR ICING FORMATION Carburettor icing can occur under any condition of high humidity, when the outside air temperature is between -15ºC and +35º C. Icing in the choke tube is caused by a combination of two factors: ! EXPANSION As pressure decreases in the choke tube, an adiabatic expansion takes place, resulting in a decrease of temperature. ! LATENT HEAT When liquid fuel is drawn from the nozzle, it changes state to a vapour, absorbing latent heat resulting in a temperature decrease.

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Figure 15 – 6 Summary of atmospheric conditions conducive to carburettor icing

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SYMPTOMS OF CARBURETTOR ICING ! In an engine with a fixed pitch propeller, carburettor icing will produce a drop in rpm and rough running. ! In an aircraft with a constant speed propeller, carburettor icing will produce a drop in manifold pressure followed by a drop in rpm and rough running. AVOIDING CARBURETTOR ICING A number of systems are used to prevent the accretion of carburettor ice: ! Sometimes the carburettor is mounted within a warm oil jacket. ! On most light aircraft carburettor heating can be used. ! This supplies warm air from a jacket which surrounds the exhaust manifold into the choke tube. ! A change of flight altitude and/or flight conditions can reduce the possibility of carburettor icing. FUEL INJECTION SYSTEMS The modern tendency in aero engines is to use fuel injection instead of float-type carburettors. The fuel injection system sprays atomised fuel directly into the inlet manifold. In some cases it sprays the fuel directly into the cylinders. Advantages of the fuel injection system are: ! better mixture distribution as fuel is sprayed into the air just before it enters the cylinder. Each cylinder therefore receives the same amount of fuel regardless of the length of the induction pipes to that cylinder. ! fuel metering is more accurate therefore no fuel is wasted. ! there is less chance of icing (the possibility of evaporative icing in the inlet manifold close to the cylinder is very remote). ! the system is not sensitive to "g" forces as there is no float system relying on gravitational forces for its operation. ! volumetric efficiency is better ! the correct mixture is maintained at all times which ensures rapid throttle response and avoids problems at idle and very high power settings. ! Inverted flight becomes possible ! Cold starting is made easier ! Throttle response is more rapid and smoother Most fuel injection systems operate on the same basic principles. The schematic layout of a typical system is shown in Figure 15-7 below. An airflow metering device controls air flowing into the engine. Air is then distributed to the cylinders via the inlet manifold. Fuel is supplied to a flow divider under pressure from where it is distributed to the fuel injectors in the proportion necessary to match the airflow measured at the airflow metering device. Copyright © 2012 EAA

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Figure 15 – 7 Schematic of a typical fuel injection system.

The main functioning components of the fuel injection system are: ! The Air/Fuel Control Unit ! The Fuel Injection Pump ! The Fuel Control Assembly ! The Fuel Manifold Valve ! The Fuel Discharge Nozzles THE AIR THROTTLE ASSEMBLY This unit consists of three control elements: ! two in the fuel control assembly and ! one for air control which is situated in the air throttle assembly. The butterfly valve within the throttle assembly controls the amount of air to the cylinders via the air intake. The throttle assembly is similar to a normal carburettor but with no venturi or other restrictions. THE FUEL CONTROL ASSEMBLY Metered fuel is piped to the fuel distributor from the Fuel Control Unit. From here a separate fuel line carries fuel to the discharge nozzle in each cylinder head. The central bore of the fuel control assembly consists of a circular metering valve at one end and a circular mixture control valve at the other end. 148

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A metering plug is situated between these two valves and has two passages. The one joins with the fuel return port and the other passage connects the mixture chamber to the metering chamber. Excess fuel is allowed to flow through a machined groove in the mixture control valve to the metering valve and then to the return fuel outlet by the rotation of the mixture valve to the full rich position. The metering valve, which is linked to the throttle system, has a cam shaped face and as the throttle is opened, rotation of the cam allows for a fuel air ratio at all engine speeds.

Fig. 15 – 8 Shrouded Fuel Nozzle

THE FUEL INJECTION PUMP The shaft of the fuel injection pump is connected to the accessory drive system and is a positive displacement rotary vane type pump. Fuel enters the pump assembly via the swirl well and any vapour in the fuel is forced upwards and returned to the fuel tanks by the vapour ejector. Because the pump is designed to deliver more fuel than is required a return line is provided. FIG. 15 – 9 FUEL PUMP WITH ANEROID VALVE

A metering jet and pressure relief valve is built into the return line (or recirculation path) so the pump delivery pressure and metered fuel amount can be maintained throughout the range of the engine speeds. Fuel flow is directly proportional to the speed of the engine. Should the engine driven pump fail, boost pressure can be used from an auxiliary pump as a source of pressure. Copyright © 2012 EAA

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THE FUEL MANIFOLD VALVE Fuel enters the distributor under pressure from the fuel/air control unit. The distributor contains the following: ! An inlet ! A diaphragm chamber ! Outlets to the nozzles ! Valve assembly (plunger and spring) When the engine is running the pressure of fuel entering the diaphragm chamber raises the plunger off its seat and allows the fuel to pass through the outlets. A screen at the inlet traps any foreign particles. A calibrated gauge in lbs/hr or gall/hr is taken from this point. All the outlets to the nozzles are blocked when the engine is not running due to the action of the spring which holds the plunger on its seat. THE FUEL DISCHARGE NOZZLES ! Fuel discharge nozzles are mounted in the cylinders of the engine with the outlet into the intake port. ! They contain a central bore counter sunk at each end. ! One end contains a jet for calibrating the nozzles and the other end used for fuel/air mixing. Air inlets are situated on the outside of the nozzles. Screens are fitted over these air inlets to ! Prevent dirt and foreign matter from entering the nozzles.

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CHAPTER SIXTEEN – SUPERCHARGING AND TURBO SUPERCHARGING

CONTENTS OF THIS CHAPTER:

! Superchargers ! Turbo-supercharges ! Intercoolers ! Aftercoolers

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SUPERCHARGING AND TURBO-SUPERCHARGING INTRODUCTION - THE NEED FOR SUPERCHARGING It is of course well known that atmospheric pressure decreases as altitude increases, resulting in a decrease of air density (pounds per cubic foot). As engine power is proportional to the "weight-flow" of air into the cylinders the power falls off as the weight of the air per cubic foot decreases. (Although the engine actually takes in air by volume, or cubic feet per unit of time). If the aircraft throttle is left in a fixed position (such as full throttle) the engine will deliver less and less power to the propeller as altitude is gained. To enable the engine to deliver adequate power at high altitude, an air compressor called a supercharger or turbosupercharger may be provided to compress the air entering the engine before it is taken into the cylinders. Because the power developed by the engine depends entirely upon the weight of the air drawn into a cylinder, the power produced will be maintained at sea level quantity if the pressure of the entering air (and therefore its density) is increased to the pressure that would normally be encountered at or near sea level. ! Superchargers are driven by a shaft which takes power from the engine crankshaft ! Turbo-superchargers are powered by a small turbine placed in the gas stream of the engine exhaust. SUPERCHARGERS The supercharger is a special air pump whose primary function is to compress the incoming charge so that a greater weight of air may be handled by the fixed volume of the cylinders, since horsepower output is directly related to weight rather than volume of air. In the supercharger a system of gears enables the engine crankshaft speed to be increased many times to drive a centrifugal compressor called an impeller at a very high number of revolutions per minute. The impeller often rotates seven to twelve times as fast as the engine crankshaft. In the supercharger, the impeller compresses the incoming air by centrifugal force. The air at high pressure then passes through a diffuser section of the supercharger, where a series of vanes straightens the airflow, increases its static pressure (the pressure of a gas at rest) and decreases its velocity before the air goes to the cylinders. A supercharger compressor comprises one or two stages (or single-speed/two-speed) depending on the horsepower of the engine. A stage is the number of times compression occurs. In the case of a two-speed supercharger when operating at high altitudes a high speed clutch (a high blower) is engaged which drives the impeller at a ratio of 10:1. Supercharging is the method most commonly used to increase the air-pumping capacity of an engine, because it requires no increase in engine rpm or in the size or number of cylinders. Furthermore, supercharging adds relatively little to the overall weight of the power plant.

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The conventional supercharger is a centrifugal air compressor placed between the carburettor and the air intake pipes. On radial engines, the supercharger is usually housed between the engine power section and the accessory rear section. The principal components of a supercharger (see figure below) consist of three units: ! the impeller, ! the diffuser and ! the collector.

Figure 16 – 1 Typical single stage supercharger

TURBO-SUPERCHARGERS

installation.

A turbo-supercharger is basically a centrifugal compressor powered by a turbine driven by the engine exhaust gases after they are collected from the cylinders but before they are discharged to the outside air. Turbo-superchargers are usually mounted on the outside of the engine and are sometimes used in conjunction with a singlestage mechanically driven supercharger inside the engine. Figure 16 – 2a Operation and installation of a typical turbocharger system

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Figure 16 – 2b Operation and installation of a typical turbocharger system

COMPARISONS BETWEEN SUPERCHARGER & TURBOCHARGER Although a turbo-supercharger accomplishes the same result as a supercharger, it has one distinct advantage: ! A turbo-supercharger allows the engine to continue to produce sea level rated power all the way up to the maximum operating altitude of the engine, whereas the brake horsepower of the engine decreases as the additional stage of a two-stage supercharger cuts in. ! Power drops off with a two-stage supercharger because the two-stage supercharger, like the single-stage supercharger, is driven by the engine crankshaft. Every increase in the compression supplied by the supercharger requires more power from the engine to drive the supercharger and this, in turn, means that there is less power available to drive the propeller. The turbo-supercharger, on the other hand, accomplishes its task of compression by taking power from exhaust gases that are already superfluous to engine needs. When the exhaust gases are used to drive the supercharger turbine wheel, the major effect this has on the engine is to increase the exhaust back pressure slightly above the normal sea level value. The increased exhaust back pressure serves to produce an effect on engine operation that is similar to the effect of sea level, regardless of the altitude at which the engine is flying. The back pressure remains almost constant as altitude is gained, until the critical altitude of the turbo-supercharger is reached.

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This critical altitude is the altitude at which the supercharger drive turbine attains its maximum allowable speed. Until a turbo-supercharger reaches its maximum allowable rpm at high altitude, the degree to which a turbo-supercharger compresses the fuel-air charge (or air, in the case of fuelinjection engines) is determined by the waste gate. The waste gate is a pilot-operated valve in the engine exhaust manifold between the engine and the turbo-supercharger. The pilot can set the amount of compression provided by the supercharger to any value. ! Closing the waste gate forces all the engine exhaust gases through the supercharger drive turbine, thus increasing the speed of the turbine and the compression supplied by the supercharger. ! Opening the gate permits the exhaust gases to escape to the outside air without acting on the turbine. This allows the turbine to slow down, which reduces the amount of compression The engine above is typical of small aircraft, turbo-supercharged opposed engines. The total degree of supercharging achieved is somewhat less than that customarily obtained when turbo-superchargers are used on large radial engines because opposed engines are not normally also supplemented with mechanically driven internal superchargers. This is due, principally, to the lack of space. Since no intercooler is needed, the entire supercharging system for an opposed engine can be supplied in a single, externally mounted unit. INTERCOOLERS AND AFTERCOOLERS The compression process in a supercharger and turbo-supercharger, as well as the heating as a result of the exhaust gases in a turbo-supercharger, significantly raises the temperature of the air which is to be fed to the engine for combustion. This is undesirable because the raised temperatures reduce the air density which reduces the effectiveness of the supercharging or turbo-supercharging process. Intercooling and aftercooling are simply operations in which the intake air is passed through an air to air heat exchanger after compression or between two compression processes. ! The prime effect is to increase the air density by lowering the intake air temperature. ! A secondary effect is to reduce the combustion temperatures and hence the cylinder head temperatures.

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CHAPTER SEVENTEEN – THE GAS TURBINE ENGINE

CONTENTS OF THIS CHAPTER: !

Turbine types " Impulse Turbines " Reaction turbines

!

Turbine blade material requirements

!

Exhaust systems " Exhaust system sections " Exhaust system operating conditions " Special exhaust ducts

!

Advantages of turbine engines

!

Gas turbine terminology

!

Starting technique

!

Types of starters " Electric starters " Air starters

!

Engine starting in flight

!

Turbojets

!

Turbofans

!

Turboprops

!

Turbine engine ignition systems

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THE GAS TURBINE ENGINE and STARTER SYSTEMS INTRODUCTION The turbine must provide power to drive the compressor and accessories. In the case of turboprop and turbofan motors shaft power must be provided for a propeller or rotor. This is done by extracting energy from the hot gases released from the combustion system and expanding them to lower pressure and temperature. High stresses are involved in this process and for efficient operation, the turbine blade tips may rotate at speeds up to 1 300 ft/sec. The continuous flow of gas to which the turbine is exposed may have an entry temperature between 700 and 1 200°C and may reach a velocity of 2 000 ft/sec in some parts of the turbine. To produce the driving torque, the turbine may consist of several stages, each employing one row of stationary nozzle guide vanes and one row of moving blades. The number of stages depends on whether the engine has one shaft or two and on the relation between the power required from the gas flow, the rotational speed at which it must be produced and the diameter of turbine permitted. However, with the advent of higher compression ratios, the tendency in recent years has been to increase the number of stages. The number of shafts varies with the engine type. High compression ratio engines usually have two shafts, which drive high and low pressure compressors. On high by-pass ratio fan engines with an intermediate pressure system, another turbine is interposed between the high and low-pressure turbines, thus forming a "triple-spool" system. On some propeller or shaft-power engines, the driving torque is derived from a free-power turbine. The shaft driving the propeller or the output shaft to the rotor blades of a helicopter, through a reduction gear, may be mechanically independent of other turbine and compressor shafts. The mean blade speed of a turbine has a considerable effect on the maximum efficiency possible for a given stage output. This is because the gas pressure drop through the turbine is proportional to the square of the blade speed. For a given output the gas velocities, deflections and hence losses, are all reduced with higher mean blade speeds. Stress in the turbine disc also increases as the square of the speed unless the sectional thickness (hence the weight) is increased disproportionately. Because of this, the final design must be a compromise between efficiency and weight. The by-pass engine enables a smaller turbine to be used than in a pure jet engine for a given thrust output and it operates at a higher gas inlet temperature, thereby obtaining improved thermal efficiency and power/weight ratio. The turbine depends for its operation on the transfer of energy between the combustion gases and the turbine. This transfer is never 100% because of thermodynamic and mechanical losses, but the more thoroughly it is accomplished, the higher is the efficiency of the turbine. An approximate value is 90%. After the gas expands due to the combustion process, it forces its way into the discharge nozzles of the turbine where, because of their convergent shape, it is accelerated.

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It reaches the approximate speed of sound (at gas temperature ± 2 000 ft/sec). At the same time the gas flow is given a "spin" or "whirl" in the direction of rotation of the turbine blades. On impact with the blades and during the subsequent reaction through the blades, energy is absorbed, causing the turbine to rotate at high speed and so provide the power for driving the turbine shaft and compressor. The nozzles and blades of the turbine are "twisted", the blades having a stagger angle, which is greater at the tip than at the root. The reason for the twist is to make the gas flow from the combustion system perform equal work at all positions along the length of the blade and to ensure that the flow enters the exhaust system with a uniform axial velocity. This results in certain changes in velocity, pressure and temperature occurring through the turbine.

Figure 17 – 1 A typical turbine blade showing the twisted contour

The losses, which prevent the turbine from being 100% efficient, are caused by a number of reasons. ! A typical three-stage turbine would suffer a 3.5% loss because of aerodynamic losses in the turbine blades. ! A further 4.5% loss is incurred by aerodynamic losses in the nozzle guide vanes, gas leakage over the rotor blades and exhaust system losses. These losses are of approximately equal proportions. ! The total losses result in an overall efficiency of 92%. The flow characteristics of the turbine must be very carefully matched with those of the compressor to obtain the maximum efficiency and performance of the engine. If, for example, the nozzle guide vanes are allowed too low a maximum flow, then a back pressure would build up, causing the compressor to surge. Too high a flow would cause the compressor to choke. In either condition a loss of efficiency would occur very rapidly. TURBINE TYPES Turbines are classified into three basic types according to the designed mode of operation: ! impulse ! reaction ! impulse-reaction

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IMPULSE TURBINES In the impulse type turbine, there is no change in gas pressure between rotor inlet and rotor exit. Any change in pressure occurs in the nozzle guide vanes where pressure is reduced in order to increase the velocity of the gas flow. Nozzle Guide Vanes

Rotor Buckets Direction of Rotation

Gas Flow

Figure 17 – 2 Impulse type rotor blade arrangement

REACTION TURBINES In the reaction type turbine, no change in pressure occurs in the nozzle guide vanes, whose sole function is to alter the direction of gas flow to that required by the rotor blades. The pressure change required to affect an increase in velocity is accomplished during the passage of the gases through the rotor blades. Nozzle Guide Vanes

Rotor Direction of Rotation

Gas Flow

Figure 17 – 3 Reaction type rotor blade arrangement

Most turbines are the impulse-reaction type – a combination of the above two types designed to achieve the desired turbine diameter and the proper match with the compressor. TURBINE BLADE - MATERIAL REQUIREMENTS The turbine blades, while glowing red-hot, must be strong enough to carry the centrifugal loads due to rotation at high speed. A blade weighing only two ounces may exert a load of over two tons at top speed. It must withstand the high bending loads applied by the gas to produce the many thousands of turbine horsepower necessary to drive the compressor.

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Blades must also be resistant to fatigue and thermal shock, so that they will not fail under the influence of high frequency fluctuations in the gas conditions. They must also be resistant to corrosion and oxidization EXHAUST SYSTEMS The jet engine exhaust system directs the turbine discharge gas into the atmosphere to create thrust. The thrust is the result of the momentum increase of the gas passing through the engine and the pressure in the exhaust gas acting over the area of the exhaust nozzle. EXHAUST SYSTEM SECTIONS A simple exhaust system is shown in Figure 17 – 4. It consists of three sections:

Figure 17 – 4 A Basic Exhaust System

Exhaust Unit This is the section immediately aft of the turbine discharge. It consists of a cone and struts. The cone creates a diverging duct, to reduce the velocity of the gas stream to about M0,5. This reduces frictional losses. The cone and struts also reduce energy losses from gas flow across the face of the turbine and from swirl. Turbine Disk

Exhaust Strut

Figure 17 – 5The cone and strut assembly Tail Cone Copyright © 2012 EAA

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Jet Pipe This is simply a passage to convey air from the exhaust unit to the propelling nozzle. Optimally, it is as short as possible but design is constrained by the location of the jet engine in the fuselage or pod.

Insulated Jet Pipe

Figure 17 – 6 The jet pipe location in the Hawk

Propelling Nozzle The simplest propelling nozzle is a convergent duct. This raises the velocity of the gases. Under most operating conditions the velocity reaches Mach 1 at the discharge. It is not possible to raise the velocity any higher and the duct is said to be “choked”. Additional energy in the gas stream appears as pressure outside of the engine. This high pressure contributes partially to increased thrust but energy is wasted as expansion continues downstream of the engine, where it makes no contribution to thrust. This is called underexpansion.

Propelling Nozzle

Figure 17 – 7 Typical nozzle found on fighter type aircraft

To recover this energy, some engines use a convergent-divergent duct, shown in Figure 17-8. In the convergent-divergent duct, the gases reach Mach 1 in the throat, and then continue to increase in velocity in the diverging section. The balance to this increase in momentum of the gas stream is a force on the walls of the diverging section. A component of the force acts along the longitudinal axis of the engine, contributing thrust.

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Converging Nozzle Section

The performance of convergingdiverging ducts depends on the ratio of propelling nozzle area to throat area. Unfortunately, optimal performance is obtained only over a very narrow range of operating conditions (speed, temperature, altitude). To operate efficiently over a wide range of operating conditions the propelling nozzle area must be made adjustable.

Figure 17 – 8 gas flow through a convergent-divergent nozzle

The increased mass of the equipment to achieve variable geometry offsets the efficiency gain, so this is seldom done for non-afterburning engines. EXHAUST SYSTEM OPERATING CONDITIONS The temperature of the gases entering the exhaust system is typically 550 – 850 °C, with velocities ranging from 250 – 400 m/s. These conditions are severe, calling for nickel or titanium construction. The rest of the aeroplane must also be insulated from these high temperatures by a stream of ventilating air around the jet pipe and/or an insulating blanket of material with low thermal conductivity. Sound insulating material is sometimes also applied. As all of the components expand and contract during with temperature changes, the design must allow them to do so without distortion or build up of internal stresses. (Note that because of the high temperatures, the speed of sound in the exhaust system is well above that in the ambient air.)

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SPECIAL EXHAUST DUCTS Turbo-propeller engines extract most of the energy from the gas stream to drive the propeller. Nevertheless, purposeful exhaust duct design can still make a significant contribution to total thrust (e.g., on a King Air 200, about 10% of total power is from jet thrust). Controllable, directional ducts are to be used on V/STOL aeroplanes (e.g. Harrier). THE ADVANTAGES OF TURBINE ENGINES Large, high-speed airliners would not have been possible without the development of the turbine engine. Already by the end of the World War II, the limited potential for further increase in power obtainable from piston engines was apparent. The regular use of jet transport aeroplanes also saw a quantum improvement in flight safety. The table below shows the advantages of the jet engine over piston engines: JET

PISTON

Total power output

Almost unlimited – 10000 hp is normal

1000 hp is best achieved at high altitude and speed

Power to mass ratio

Worst is about 40N/kg

Best is about 20N/kg

Small frontal area

33 KN/m2

20 KN/m2

Economy

Good

specific

consumption and maintenance costs Most Efficiency

efficient

at

fuel

Lower specific fuel consumption

low

and high maintenance costs

high

speeds; by-pass increases efficiency at moderate speeds

Propeller compressibility losses make high-speed flight inefficient

Large engines are very difficult to

Cooling

Easily cooled by airflow

Reliability

At least four times more reliable than piston engines

Frequent overhauls required and

and no deterioration of reliability with size increase

reliability decreases increase in power output

cool effectively

with

GAS TURBINE TERMINOLOGY ESHP: Equivalent shaft horsepower. Under static conditions, approximately 2.6 pounds of thrust is equal to 1 horsepower. Thus: ESHP = JET THRUST ± 2.6.

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GROSS THRUST: Gross thrust is the thrust generated at the exhaust neglecting the momentum of the incoming air stream. GROSS THRUST = M AIR (V EXHAUST ) + A EXHAUST (P EXHAUST – P ATMOS). ITT: Inter-turbine temperature. The temperature of the engine gases between the high and low-pressure turbines. Ng: This is usually associated with turboprop engines. It is the speed of the gas generator (turbine or compressor). NF: This is usually associated with turboprop engines. It is the speed of the free (or power turbine in the case of twin spool engines). N1: The speed of the low-pressure compressor stage. N2: The speed of the high-pressure compressor stage. NET THRUST: Net thrust is the thrust which results from the change in momentum of the mass of the air and the mass of fuel which pass through the engine, plus an additional force at the exhaust represented by the difference in static pressure at the nozzle and the ambient static pressure. NET T HRUST = OUTGOING GAS MOMENTUM – INCOMING AIR MOMENTUM – INCOMING FUEL MOMENTUM NET THRUST = M AIR (V EXHAUST – V AIRCRAFT ) + M FUEL (V EXHAUST) + A EXHAUST (P EXHAUST – P ATMOSPHERE). This can be approximated as: NET T HRUST = M AIR (V EXHAUST – V AIRCRAFT) + A EXHAUST (P EXHAUST – P ATMOS). THRUST HORSEPOWER: Thrust horsepower is the useful horsepower developed by an engine when it is producing thrust. Obviously the engine only produces useful power when the aircraft is moving. As such: THRUST HORSEPOWER (THP) = T HRUST (LBS) X SPEED (MPH) 375 or, in metric units : THRUST POWER = T HRUST (N) X SPEED (MS-1) THRUST SPECIFIC FUEL CONSUMPTION (TSFC): This is the fuel flow of the turbojet engine per unit of thrust produced. TSFC = FUEL FLOW (LBS /HR) NET THRUST (LBS) TIT: Turbine inlet temperature. The temperature of the engine gases at the entrance to the turbine stage. STARTING TECHNIQUE ! The standard technique for starting a turbine engine requires it to be rotated to a suitable speed when the start sequence is initiated.

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! When the correct rpm are maintained, fuel is supplied and the fuel/air gas enters the combustion chambers. ! The ignition source is energised at a suitable stage and the fuel/air gas starts to burn. ! When all necessary engine parameters are stabilized and checked to be within prescribed limits, the igniters are de-energised and the engine is self-sustaining. ! The functioning and co-ordination of the two systems – rotation of the engine up to speed and ignition – is usually controlled automatically after initiation of the start cycle. A typical "log" of events during the starting cycle is shown below.

Figure 17 – 9 A typical starting sequence for a gas turbine engine

TYPES OF STARTERS The types and power sources for a turbine starter vary, but the basic requirement is basically the same for all commercial aircraft. ! The system must be economical and reliable. ! The system must transmit high torque to the engine assembly in order to accelerate it smoothly to a speed where the gas flow through the engine is sufficient for it to selfsustain and then accelerate further to operating speed.

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ELECTRIC STARTERS Electric starting is used on some turboprop and turbojet engines. The starter is usually a DC motor coupled to the engine through a reduction gearbox and a clutch or ratchet mechanism. This mechanism disengages after the engine reaches self-sustaining speed. A typical electric starting system is shown in figure 17 – 10.

Figure 17 – 10 A typical electric starting system

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AIR STARTERS The most commonly used system for starting the engines of a commercial aircraft is with an air starter. It is also used on some military aircraft. Among the advantages which air starting enjoys over other systems are: ! simplicity of operation ! economical to operate ! the equipment is light. Power is transmitted from a turbine rotor at the air intake, through reduction gearing and a clutch to the starter output shaft. The starter turbine can be supplied with air from: ! an external ground supply ("start cart") ! an air source in the aircraft – auxiliary power unit (APU) or, in some older aircraft types, an air bottle ! an engine which is running ("cross feeding") The actual air supply is controlled electrically through a pressure-reducing valve. This is opened when an engine start is selected and it should close automatically at a predetermined engine speed. The clutch is also disconnected automatically when the engine accelerates to normal idle rpm and the starter stops turning. A typical air starter system is shown in figure 17 – 11. Turbine Exhaust

Air Intake HP Air Supply

Auxiliary Power Unit Electric Starter Air Control Valve

LP Air Supply Exhaust Air Engine Air Starter Figure 17 – 11 A typical air Starter Turbine

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ENGINE STARTING IN FLIGHT Engines can be re-started in flight without difficulty. The necessary rotational speed is provided by adjusting airspeed and the start sequence is performed. It may be necessary to descend from very high altitudes to be able to accelerate the engine to the required speed detailed in the aircraft manual. The conditions required are frequently referred to as being the "envelope". Care has to be taken that the necessary parameters detailed in the manufacturer’s handbook are observed. Exceeding these parameters can damage the motor.

TURBOJETS The turbojet (as shown below in Figure 17 – 12) has the same basic components as the first jet engine invented by Sir Frank Whittle. During the normal turbojet working cycle, the intake air is compressed by the compressor and then directed to the combustion chambers where fuel is added and ignited. This action substantially increases the volume of the air by raising the temperature and hence decreasing the density. From the combustion chamber the air flows to the turbines where between 60% and 80% of the added energy is extracted to drive the compressor and keep the engine running.

Figure 17 – 12 Turbojet schematic

The energy remaining in the gases is then converted into kinetic energy and the gases exhaust from the tailpipe at high speed which produces the required thrust.

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TURBOFANS The turbofan is a variation on the turbojet engine but with better subsonic performance characteristics and lower noise levels. The fan attached to the front of the engine accelerates a large mass of air to a lower velocity than in the turbojet. The air which is accelerated by the fan is either directed outside the engine through ducting, exhausted or it is mixed with the air which has flowed through the core of the engine and then exhausted. The effect of adding the fan arrangement to the basic turbojet configuration is to improve the power to weight ratio of the engine considerably and to reduce the specific fuel consumption (sfc). This is the amount of fuel (mass) burnt per hour per pound of thrust produced. The term "bypass ratio" is associated with the turbofan. This refers to the ratio of the volume of air, which bypasses the combustor to that which passes through the combustor. Compressor Section

Fan/By-Pass Air

Main Jet Efflux

Inlet Air Fan Section

Figure 17 – 13a Low Bypass Turbofan (Military)

A typical high bypass ratio is one of 6:1. Fan Section

Drive Power Turbines

With this setup, about 70% of the trust is developed by the fan itself.

Figure 17 – 13b Turbofan engine layout

Like turbojet engines, turbo fans may be manufactured in single or multiple spool arrangements, depending on the engine design objectives. In multi-spool arrangements, the low-pressure turbine usually drives the fan.

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TURBOPROPS

Figure 17 – 14a Turboprop Applications

The turboprop engine can be considered an extreme development of the turbofan engine. The design is aimed at the low end of the speed regime and is intended to obtain the best of both worlds – that is the smoothness, high power/weight ratio and fuel flexibility of the turbine engine with the high efficiency at low speeds of the propeller. The propeller is driven via a reduction gearbox which reduces the rpm from around 30 000 rpm to about 1 900 rpm. This will depend on engine and propeller size. FIXED SPOOL and FREE SPOOL ARRANGEMENT The output shaft to the gearbox may be the same shaft which connects the turbine and compressor wheels. In this case, it is said to be of the "fixed spool" arrangement. This type of system necessitates the propeller running at full rpm whenever the engine is running. This, in turn, means that thrust can only be controlled during ground operations by varying the pitch. This arrangement has disadvantages: ! This configuration is very noisy indeed. ! Another feature is the need to stop the engines in the fully fine position to make starting easier.

Propeller

Compressor

Turbine

Common Drive Shaft Reduction Gearing

Figure 17 – 15a Fixed Spool Arrangement

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An alternative arrangement is to couple the output shaft to a separate shaft driven by the low pressure turbine. This type of arrangement is known as the “free spool” arrangement (Figure 17 – 15b). The high pressure (HP) turbine which drives the compressor is then known as the gas generator turbine, while the low pressure (LP) turbine which drives the propeller is known as the free turbine. This arrangement, despite the added complexity, has many advantages: ! Firstly, the free turbine need not be turned during starting operations. This reduces the load on the starter motor considerably as it does not have to turn the gears or the propeller. ! Secondly, the free turbine can be operated at lower rpm than the power turbine. This enables a simpler and lighter gearbox to be used. ! Finally, the free turbine can be slowed at any time while the power turbine continues to operate at full rpm. This simplifies and quietens ground operation greatly.

Free Power Turbine

Propeller Drive Shaft Reduction Gearing Figure 17 – 15b Free Spool Arrangement

TURBINE ENGINE IGNITION SYSTEMS A typical turbine-engine DC ignition system consists of: ! a relay, ! two DC ignition units and ! two igniter plugs. Switching is controlled from an ignition master switch, a throttle relay and a starter switch. The ignition system operates only during engine start. A turbine-engine ignition system supplies a very high-capacitance voltage to the igniter plugs. The igniter plugs (usually two) ignite the fuel in two burner cans (combustion cans). The other burner cans are ignited by flame-crossover (flame-travel) tubes. When the engine is stabilised, ignition is shut off and the engine continues to operate by self-ignition.

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If "light-up" is delayed, unburnt fuel accumulates quickly. When combustion finally takes place the excess fuel burns and can produce excessive turbine temperatures. This is known as a "hot start".

Figure 17 – 16 Different types of Igniters

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FILL IN THE GAPS Fill in the gaps (or circle the correct answer) in the following exercise referring to the text in the previous chapter. 1. Turbine engines cannot be started until the turbine has reached a ________ _____. 2. When the correct rpm are maintained, ____ is supplied and the fuel/air mixture is ignited in the __________ ________. 3. The initial sources of ignition are ________. 4. When the engine is stabilized and burning under normal conditions, the combustion process is _____-__________. 5. In civil aviation, engines are normally started with ________ or __________ ___ starters. 6. Engines can easily be re-stared in flight by adjusting ________ to a suitable value before supplying fuel and ________.

ANSWERS 1. suitable speed 4. self-sustaining.

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2. 5.

fuel/combustion chambers electric/compressed air

3. the igniters. 6. airspeed/ignition

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CHAPTER EIGHTEEN - AIRCRAFT ELEMENTS BEARINGS, VALVES, PUMPS & FILTERS

CONTENTS OF THIS CHAPTER:

!

Bearings " Bearing materials " Plain bearings " Ball and roller bearings

!

Valves

!

Types of valves " " " " "

!

Pumps " " " " " "

!

Thermal Relief Valves Pressure Release Valves Selector Valves Check Valves Restrictors

Gear type pumps Diaphragm type pumps Vane type pumps Piston type Self regulating pumps Centrifugal pumps

Filters " Strainers " Sediment traps " Paper filaments

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AIRCRAFT ELEMENTS BEARINGS, VALVES, PUMPS & FILTERS BEARINGS Bearings are needed to reduce friction between moving surfaces, to transfer loads and to allow various parts and components to rotate in a uniform circular motion. Bearings are mainly used in aero engines for: ! main crankshaft bearings ! connecting-rod (“con-rod") bearings ! thrust bearings. BEARING MATERIALS A good friction bearing is made from material strong enough to withstand the high pressures imposed on it while in motion. It must also permit adjacent surfaces to move or rotate with minimum friction and/or wear. The moving parts or units must be kept in position to very close tolerances to provide efficient, quiet operation. Bushes are usually made of copper/brass or nylon. PLAIN BEARINGS Plain bearings (called friction bearings) are sleeve-shaped. Each surface supports or is supported by another surface. These bearings are normally only subjected to radial loads, although they can be designed to accept thrust loads. ! Plain bearings are mainly used as crankshaft bearings. These bearings consist of a metal sleeve in a section or divided sections. ! They are also used on the camshaft for in-line and opposed engines and with the cam ring/cam plate in radial engines. ! Connecting-rod bearings are also plain bearings. ! A split bearing is a plain bearing made in equal halves for easy installation. ! A "bush" is also a type of plain bearing. BALL AND ROLLER BEARINGS Ball bearings consist of an inner race or track, one (or two) sets of steel balls and an outer race. The steel balls may be arranged in one or more rows. The inner race (and sometimes the outer race) has grooves which fit the radius of the steel balls. Roller/ball bearings (thrust bearings) are often used on the propeller shaft. In a roller bearing the roller itself rotates between an inner and an outer race (holder). Both races are made of case-hardened steel. When a roller is tapered, the inner and outer races are cone-shaped. This type of bearing can handle radial and thrust loads. Straight roller bearings are used where large radial loads act in the engine. Some roller bearings are specially designed to accept thrust loads.

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The advantage of the roller bearing is that it allows large contact surfaces to carry high thrust and radial loads. Ball bearings are designed to produce very little friction and give almost friction-free operation. VALVES Valves are used for various applications. Because they are required to retain their strength at very high temperatures they are manufactured of chromium or tungsten steel. A basic valve construction consists of: !

Valve seat which runs in bronze valve guides pressed into the cylinder head.

!

Camshaft which opens the valves, one for the intake valve and one for the exhaust valve.

!

Pushrods which are moved back and forth by the camshaft

!

Rocker arm assembly which pushes the valves open

!

Timing gears which ensure correct valve timing

!

Fig 18 – 1 Valve Assembly & Associated Operating Parts

Coiled spring which closes the valves.

Intake valves are usually larger than exhaust valves. Exhaust valves are filled with a solution of sodium or mercury which serves as an effective cooling agent. Clearance between the valve and the rocker-arm is essential for efficient engine operation. If the clearance is excessive it results in the valve opening late and closing early with loss of power. In order to obtain an efficient vibration free engine correct valve and ignition timing is essential. TYPES OF VALVES Thermal Relief Valves and Pressure release valves – hydraulic fluid is incompressible and mechanical damage can be caused to components if over pressurisation occurs. To avoid this relief valves are situated at critical points in the system. They operate by balancing system pressure against an internal reference spring. If system pressure rises above spring pressure the valve opens allowing fluid to escape into the system return pipes and so reduces pressure. The valve re-seals automatically once system pressure returns to below the reference level.

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Selector valves – also called non-return valves, control the direction of movement of fluid. They are similar in construction to relief valves. A valve poppet is held closed by a weak internal reference spring. Pressure of fluid flowing in the desired direction can overcome spring force, and fluid can therefore flow through the valve. However, if fluid pressure upstream of the valve is reduced the poppet snaps closed to prevent a fluid flow reversal. Check valves – are used when fluid is required to flow in one direction only. Fluid enters the inlet port and forces the valve to open against the pressure of the spring. Once the flow of fluid stops the spring returns the valve to its seat, blocking the reverse flow. Restrictors –are a type of control valve used in pneumatic systems where the outlet port is smaller than the inlet port thus restricting the amount of flow to the required system. PUMPS The majority of engine or motor driven pumps are positive displacement, rotary swash plate pumps. Two basic variations of this type of pump are used in aircraft systems: CONSTANT DISPLACEMENT PUMPS:

Fig. 18 – 2 Constant Displacement Pump and Control System

These pumps absorb constant driving power whatever the output demand – when pressure in the system reaches an upper limit cut-out valve allows fluid to bypass the pressure line and flow back to the reservoir. Because large volumes of high pressure hydraulic fluid are constantly being circulated, great attention must be paid to cooling the fluid to maintain it within temperature limitations. Four basic types of displacement pumps are described below: Gear type pump – draws oil from the tank. It consists of two gears with closely meshed teeth. Oil is picked up by the teeth and trapped between them and the gear housing, carried around and discharged through the outlet port into the oil screen passage. Excess oil is fed back to the pump inlet.

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Diaphragm type – Commonly used in the fuel systems of aircraft. Consists of a diaphragm which is connected via a rocker arm to a camshaft. The diaphragm is located within the pump chamber containing an inlet and outlet port with their respective check valves. The rocker arm transmits the back and forth movement to the diaphragm resulting in fuel being sucked into the inlet and forced out of the outlet port. Because it is a positive displacement pump it requires a pressure relief or by-pass valve.

Fig. 18 – 3 Diaphragm-Type Pump

Vane type – Employed in the vacuum and fuel systems. It consists of a number of vanes that freely slide in and out of a rotor which is housed slightly off centre in the pump casing. The vanes are flung outwards due to centrifugal force and are always in contact with the inner casing. Because the rotor is offset from centre there is always a large volume on the one side and a small volume on the opposite side.

Fig. 18 – 4 Vane-Type Fuel Pump

The expansion of the volume draws in the fluid through the intake port and discharges it under pressure through the outlet where the volume has decreased. Because it is a positive displacement pump it requires a pressure relief valve. Piston type – Mostly used in the hydraulic system due to the extreme pressures that may be generated by it. It consists of a piston within a cylinder with the respective inlet and outlet Copyright © 2012 EAA

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ports and their check valves. These pumps are also available as double action pumps and are employed where it has a dual function to perform. It consists of a single element double piston with the two respective outlets. It is a positive displacement pump Self Regulating Pumps These pumps are usually chosen for Primary and Secondary systems. As the name implies it automatically varies its output to meet system demands. Centrifugal pumps – In this type of pump, the fluid is fed to the centre of the rotating impeller (eye of the impeller) and is thrown outward by centrifugal action. As a result of high speed of rotation the liquid acquires a high kinetic energy and the pressure difference between the suction and delivery sides arises from the conversion of kinetic energy into pressure energy. The impeller consists of a series of curved vanes so shaped that the flow within the pump is as smooth as possible. The greater the number of vanes on the impeller, the greater is the control over the direction of motion of liquid and hence the smaller are the losses due to turbulence and circulation between the vanes.

Fig. 18 – 5 Centrifugal pump

The liquid enters the casing of the pump, normally in an axial direction, and is picked up by the vanes of the impeller. In the simple type of centrifugal pump, the liquid discharges into a volute, a chamber of gradually increasing cross-section with a tangential outlet. FILTERS Contamination of fluid with even minute particles will damage and degrade system performance. Adequate system filtration ensures that particles introduced into or generated by the system are removed by various types of filters. Strainers – made of fine wire gauze and shaped as a tubular screen. In the event of clogging and to maintain an uninterrupted flow of fluid, these strainers may be designed either to collapse or may employ relief valves. Sediment traps – as their name implies, sediment carried in fluid is deposited into traps which are always situated at the lowest part of a system. This material may then be removed from the system by means of drain valves. Filters – specially treated paper filaments which prevent micron particles from entering the engine. They are housed in replaceable filter cartridges or in holders where the filter only can be removed and replaced. Fluid passes through the inlet of the holder, forced through the filter element and passes out the outlet.

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CHAPTER NINETEEN – HIGH SPEED FLIGHT

CONTENTS OF THIS CHAPTER: The differences between High Speed and Low Speed Flight ! Bell’s Experiment ! Compressibility of Air ! Subsonic, transonic and supersonic speeds ! Free stream Mach number ! Local Mach number ! Critical Mach number ! The Mach Meter ! Wave propagation from a moving source ! Shock waves ! Expansion wave drag at supersonic speeds ! Swept wing planforms ! Delaying shock induced separation ! Vortex generators ! Stalling speed at height ! The effects of increasing Mach number

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HIGH SPEED FLIGHT INTRODUCTION Piston engined aircraft are unable to fly much faster than ± 500 mph. Jet engined and rocket engined aircraft are capable of much higher speeds. Within the earth's atmosphere, the "speed limit" is governed by such factors as the melting point of metals and the effectiveness of the refrigeration of the aircraft. Outside the atmosphere, the gravitational effect of the heavenly bodies comes into play. Supersonic speeds are well established with the Concorde aircraft, but supersonic flight is constrained by the damaging effect of the supersonic pressure wave – the "sonic boom" – and further development in the civil aviation field is doubtful at the time of writing. THE SPEED OF SOUND DEFINITION The speed of sound is defined as the speed at which a very small pressure disturbance is propagated in a fluid under specified conditions. BELL'S EXPERIMENT All sound waves travel at measurable speeds. These speeds depend on variables such as the density and the temperature of the media through which the waves travel. A well-known experiment first carried out by the physicist Graham Bell showed the characteristics of sound in different media. A bell is covered by an air-filled glass jar. When a current is connected to the bell, it can be heard ringing. The vibrations from the bell clapper striking the bell are transmitted to the wall of the jar by the air in the jar. As air is withdrawn from the jar (leaving a vacuum), the ringing becomes fainter. If all the air was exhausted from the jar, silence would result, even though the bell can be seen ringing.

Figure 19 – 1 Bell's experiment

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The speed of sound at sea level in the standard atmosphere is 661.5 kt. It can be calculated from the formula: LSS (local speed of sound) = k√ √T where k is a constant (usually taken to be 38.95) and T is the temperature in ° Kelvin. The speed of sound at a particular density is known as Mach 1. The name derives from an Austrian physicist, Dr Ernst Mach, who developed a theory relating to the measurement of airflow. COMPRESSIBILITY OF AIR At any speed, the air around a wing undergoes changes in pressure. ! At low speeds, the small pressure changes cause negligible variations in the density of the air and the flow around the wing is termed "incompressible". ! At high speed though, significant changes occur to the airflow. The pressure changes are fairly large and some significant changes in air density occur. While airflow at low speed is considered as incompressible, at high speed the reverse holds. Compressibility must be taken into account. LOW SPEED FLIGHT As a wing moves through air, local changes in the velocity of the air near the wing produce pressure disturbances in the airflow around the wing. These "disturbances" are propagated through the air at the speed of sound. At low speeds these "disturbances" move ahead of the aircraft. Because they are travelling faster than the aircraft – they are moving at the speed of sound. They also travel in other directions but we are concerned mainly with those moving directly ahead. The thickness of the pressure waves is a mere 0.0001" and they completely envelop the aircraft in subsonic flight. Figure 19-2 shows pressure build-up as an aircraft increases speed to that of sound.

Pressure disturbances

Less Than the Speed of Sound

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Sound Barrier (Mach 1) At the Speed of Sound

Shock Wave

Greater Than the Speed of Sound

Fig. 19 – 2 Pressure Build-Up With Increasing Air Speed

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HIGH SPEED FLIGHT As aircraft speed increases, it catches up with the pressure waves moving ahead of it (at the local speed of sound) until it is travelling at the same speed as the pressure disturbances themselves. This is when aircraft speed is M1.0. As the aircraft approaches M1.0, a compression wave forms at the leading edge of the wing and any changes in air velocity or pressure are quite sudden and sharp. The pressure wave ahead of the wing (which is already travelling at the speed of sound) will be "caught up" by the wing as it (the wing) reaches the speed of sound. Subsonic Acceleration Flow

Subsonic Deceleration Flow

Flow Direction Changes well Ahead of the leading edge

Fig 19 – 3 Typical Airflow Around An Aerofoil

When the wing is travelling faster than the speed of sound, the airflow ahead of the wing is not influenced by the pressure field of the wing since pressure disturbances cannot be propagated faster than the speed of sound. Figure 19-4 shows a typical supersonic flow pattern for a supersonic wing. Initial Expansion Wave

Flow direction does not change ahead of leading edge

Final Expansion Wave

Trailing Edge Oblique Shock Wave

Leading Edge Oblique Shock Wave Figure 19 – 4 Typical Supersonic Flow Pattern – Supersonic Wing

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SUBSONIC, TRANSONIC & SUPERSONIC SPEEDS AIRFLOW AROUND A WING Subsonic and supersonic airflow patterns around a wing are like surface waves on water. A surface wave is caused by a pressure disturbance. The bow of a ship moving slower than the wave is not in contact with the wave. This is similar to subsonic flight. As the ship approaches the speed of the wave ahead, a bow wave is formed. The bow wave grows stronger as the speed of the ship increases and its angle decreases. A supersonic aircraft would produce similar bow waves through air as shown in Figure 19-6. Figure 19 - 5 Bow waves of a boat

The shape of "bow waves" of supersonic aircraft are the same as shown in Figure 19 – 6.

Figure 19 - 6 Bow Waves Of A Faster Moving Boat

Compressibility effects depend more on Mach number (MN) – the relationship between airspeed and the speed of sound – rather than on airspeed alone. These effects are not just found at airspeeds near Mach 1. The airflow around certain aerodynamically shaped parts of an aircraft can reach very high speeds when the aircraft speed itself is well below Mach 1. This means that subsonic and supersonic flow can exist at the same time. Flight speeds are defined in certain regimes: Subsonic -

below M 0.75

Transonic -

M 0.75 to M 1.20

Supersonic -

M 1.20 – M 5.0

Hypersonic -

M 5.0 and above.

(These speeds are approximate). Various flow speeds exist in the different speed regimes. For instance, in the subsonic flow regime, subsonic airflow exists on all parts of the aircraft. In the transonic regime, the airflow is partly subsonic and partly supersonic. Similarly, in supersonic and hypersonic flight, some parts of the boundary layer are subsonic but the predominating airflow is supersonic.

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FREE STREAM MACH NUMBER The "Free Stream Mach Number" is the MN of the airflow sufficiently far away from the wing to be unaffected by the passage of the wing (the aircraft) through it. The formula is given by the relationship : M = V ÷ A or, MFS = TAS ÷ LSS (LSS = local speed of sound). MFS is sometimes known as the flight MN. It is the true MN as shown on the Machmeter (ignoring small errors). LOCAL MACH NUMBER (ML) At a particular MFS, the airflow is accelerated in some areas and decelerated in others. The local speed of sound will also change because of temperature changes around the aircraft produced by the movement of the aircraft through the air. ML can be defined as the ratio of the speed of the airflow around a point and the speed of sound at that point: ML = airflow speed at a point. LSS at that point ML may be the same, higher or lower than MFS. CRITICAL MACH NUMBER (Mcrit) As airspeed increases, the aircraft begins to overtake the outermost pressure waves formed ahead of it as it approaches LSS. At this airspeed, the airflow above the wing may already be sonic. Figure 19-7 illustrates the case of an aircraft approaching LSS and the behaviour of the airflow around the wing section.

0.80 0.75 M = 0.70

0.85

0.80

0.75

0.70

0.70

Subsonic Aerodynamic Centre of Balance Figure 19 – 7 Subsonic Flow Approaching LSS

The design of an aerofoil provides for the reduction of pressure above the wing compared with the pressure below the wing. At airspeeds of M0.77 to M0.85 (depending on the shape of the aerofoil), the pressure waves emitted by the top surface of the wing move out, ahead of the wing, at the speed of sound, in the normal fashion. However, as the airflow above the wing reaches LSS, at, say, M0.85 for a particular wing, a shock wave is formed above the wing. This shock wave is at right angles to the airflow. The importance of the shock wave is that immediately ahead of the shock wave the airflow is supersonic and subsonic immediately behind it. Copyright © 2012 EAA

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The shock wave is very thin, about 1/10 000" thick and the air within it is in a state of very high compression. Each molecule of air in a subsonic flow is affected to a greater or lesser degree by the motion of all other molecules of air in the whole airflow. When the shock wave occurs, the airflow is at different speeds in different parts. Figure 19-8 refers. Supersonic Region

Normal Shock (Weak)

1.00

Normal Pressure (Waves) Subsonic Flow 0.95 0.90

0.95

1.00

1.02

0.90

M = 0.85 0.85

0.85

Subsonic Aerodynamic Centre of Balance

Figure 19 – 8 Airflow at critical MN

At supersonic speeds, a moving air molecule can only influence another molecule in that part of the airflow contained in the "Mach cone" behind the molecule. No "warning signal" can be sent ahead of the aircraft (as in a "bow wave"), which would "warn" the airflow ahead of the approaching aircraft. The MFS when any ML reaches unity is termed the critical Mach number (Mcrit). Mcrit for an aircraft or a wing section varies with angle of attack. It also marks the lower limit of the speed band where ML may be either subsonic or supersonic. This "band" is known as the "transonic range". USING THE MACHMETER High-speed aircraft have a Machmeter fitted as well as an ASI. As LSS varies directly with temperature, the Mach meter is of much more use than an ASI. Cruise regimes at high altitude are best achieved by using the Machmeter as a speed reference rather than the ASI. All cruise tables are predicated on MN, not RAS/CAS. In the climb to high altitude, airspeed is used as a reference. A typical climb graph or table will show "280/0.80". Figure 19 – 9

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This indicates that the aircraft is flown at a steady RAS of 280 kt until M0.80 is reached. From then on, M0.80 is the target speed for the climb. WAVE PROPAGATION FROM A MOVING SOURCE As mentioned above, aircraft at speeds below M1.0 transmit pressure disturbances ("waves") in all directions. ! One obvious effect is that an approaching aircraft can be heard before it is seen when it is travelling below the LSS. ! An aircraft flying at or near M1.0 is seen before it is heard, in the same way as a gunshot may be seen before it is heard. ! Another effect of an aircraft travelling below LSS is that the airflow ahead of it is "warned" of its approach and has time to divide and allow the passage of the aircraft with minimum disturbance. If air was not compressible, any pressure waves transmitted by a moving aircraft would be transmitted at an infinitely high speed and the disturbance caused to the air by the progress of the aircraft through it would be felt everywhere instantaneously. The flow pattern around the aircraft would also be independent of the airspeed. However, since air is compressible, a different situation prevails. A change of air density will mean a change of pressure. The speed of propagation of a pressure wave has a finite value, which is equal to the LSS. The distance between an aircraft and the pressure waves ahead of it will decrease as the aircraft approaches LSS. The effect of the "advance wave" will also decrease and there will be a change in the airflow and pressure pattern around the aircraft and the wing as airspeed approaches LSS. SHOCK WAVES AT Mcrit The airflow above a wing in subsonic flight will cause local velocities on the surface, which are greater than the free stream velocity. This can produce compressibility effects at speeds below LSS. In the transonic regime, mixed subsonic and supersonic local speeds occur and the first significant effects of compressibility are experienced. Figure 19-10 illustrates subsonic flow patterns. Increased Speed of Airflow

Downwash

Decreased Speed of Airflow Upwash Figure 19 – 10 Subsonic Flow Patterns Copyright © 2012 EAA

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SHOCK WAVES If a blunt object is placed in a supersonic airstream, the shock wave which is formed detaches from the leading edge of the object. The wave will also detach when a wedge shaped or a cone shaped angle exceeds a certain "critical" value. Figure 19-11 below and on the next page shows the formation of a shock wave for differently shaped leading edges.

Figure 19-11 Shock Wave Formations

When a shock wave is made perpendicular to the free stream flow it is called a "normal shock wave" (normal = at right angles). The low-pressure area immediately behind the shock wave is subsonic. The flow immediately behind a normal shock wave is always subsonic, regardless of the free stream MN. In fact, the higher is MFS, the lower is the subsonic MN behind the shock wave. As an example, when MFS is 1.50, the airflow immediately behind the shock wave is M 0.70. If MFS is 2.6, the subsonic flow is M0.50.

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! A normal shock wave forms immediately in front of any relatively blunt object in a supersonic air stream. This slows the air stream to subsonic speed so that the air stream is influenced by the blunt object and will flow around it. Once the air stream is past the object it may remain subsonic or accelerate to supersonic speed. Whether this happens or not will depend on the particular shape of the object and the MN of the free air stream. ! A normal shock wave may also be formed when no object is in the free air stream. This will occur when an aircraft travelling at supersonic speed slows to subsonic speed without turning. ! A normal shock wave forms at the boundary between the supersonic and subsonic regions. This explains why aircraft can encounter compressibility effects before reaching LSS. ! As the local supersonic waves move aft, a normal shock wave forms and the flow may return to subsonic speed and rejoin the subsonic free stream without discontinuity. The transition to subsonic is smooth and there are no shock waves if the transition is made gradually with a smooth surface. The transition of flow from supersonic to subsonic without a change in direction always forms a shock wave. A supersonic air stream passing through a shock wave always experiences certain changes: ! The air stream is slowed to subsonic. The local MN behind the wave is approximately equal to the reciprocal of the MN ahead of the wave, for example, when the MN ahead of the wave is 1.25, behind the wave it will be ± 0.80 ! The airflow direction immediately behind the wave does not change ! The static pressure behind the wave increases greatly ! The static temperature behind the wave increases greatly and so the LSS increases also ! The density of the air stream behind the wave is also greatly increased ! The available energy of the air stream (this is given by the sum of the dynamic & static pressures) is greatly reduced. ! The normal shock wave therefore wastes energy. EXPANSION WAVE If a supersonic wave is turned away from the preceding flow, an "expansion wave" is formed. The flow around the corners, shown in Figure 19-12 causes sudden sharp changes only at the corners themselves. The flow is therefore not a true shock wave. A supersonic air stream passing through an expansion wave experiences certain changes: ! The supersonic air stream is accelerated. ! The velocity and MN behind the wave are greater ! The static pressure behind the wave decreases ! The static temperature behind the wave decreases, thus reducing LSS ! Since the flow change is gradual, no shock occurs nor is energy lost into the air stream. ! The expansion wave does not dissipate air stream energy. ! The flow direction is changed provided separation does not occur.

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Figure 19-12 Expansion Wave Formations

THE EXPANSION WAVE IN THREE DIMENSIONS The expansion wave in three dimensions is somewhat different. The principal difference is that there is a tendency for the static pressure to start increasing past the shock wave (following the initial decrease in pressure). Figure 19-13 gives a summary of the characteristics of the three principal waveforms encountered with supersonic airflow. Type of Wave

Flow Direction Change

Effect on Velocity and Mach Number

Effect on Static pressure Static Temp And Density

Effect on Available Energy

Flow Into Centre

Decreased But still Supersonic

Increase

Decrease

No Change

Decreased to Subsonic

Great increase

Great Decrease

Flow Around a corner

Increased To Higher Supersonic

Decrease

No Change (No shock)

Expansion Wave

Figure 19-13 The three principal wave supersonic forms

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DRAG AT SUPERSONIC SPEEDS As speed increases, drag increases also. Along with the rise in drag will come: " buffeting " trim and stability changes. There will also be a decrease in the affectivity of the controls. Conventional aileron, rudder and elevator surfaces, which are subjected to the buffeting produced by the changes, may make "buzzing" noises and unwanted stick forces. Operating outside normal limits can easily cause structural damage. When airflow separation occurs on the wing as a result of shock formation a loss of lift and subsequently, a loss of downwash occurs aft of the affected area. If there is a sideslip, or the wings are differently shaped, there may be a differential between the shock waves on the wings which can lead to a rolling moment and wing drop. If the shock induced separation occurs near the wing root, a decrease in downwash behind the area will be a natural consequence of the loss of lift. A decrease in downwash on the tailplane produces a diving moment and the aircraft can "tuck under". The tail will normally have down loads at low aircraft angles of attack (in cruise) and up loads at high angles of attack (low speed flight) "Tuck under" is not the same as control reversal. Although the tendency to nose down with a speed increase in the transonic regime is opposite to the nose up tendency in the subsonic regime, a pull force on the control column will still raise the nose and a push force will still lower it. SWEPT WING PLANFORMS With a swept wing planform, the direction of pitch caused by shock induced flow separation depends where the shock first appears. If the shock first appears on the wing root, the wing root loses lift whilst there is still lift at the wing tips. This results in a nosedown tendency. If the shock first appears on the wing tips, the reverse applies and there is a nose-up tendency. DELAYING SHOCK INDUCED SEPARATION Since most of the problems of transonic flight are associated with shock induced flow separation any means of delaying the onset of shock waves, or lessening their effect, will improve the aerodynamic characteristics. An aircraft configuration can utilize thin surfaces of low aspect ratio as well as sweepback to delay the onset and to reduce the magnitude of transonic flight divergence. In addition, various methods of boundary layer control (e.g. injecting high-speed air into the boundary layer, high lift devices (leading edge slats, slots and trailing edge laps), vortex generators etc may be applied to improve transonic characteristics. As an example, vortex generators mounted on a surface can produce higher surface velocities and increase the kinetic energy of the surface layer. In this case a greater pressure gradient (a stronger shock wave) is necessary to produce the airflow separation. Copyright © 2012 EAA

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VORTEX GENERATORS

Vortex Generators Figure 19-14 Vortex generators

Vortex generators draw high-energy air from outside the boundary layer into the slower moving air close to the skin of the aircraft. They are usually small aerofoils of low aspect ratio, which are mounted in pairs at opposite angles of attack to each other and perpendicular to the aerodynamic surface they serve. They develop lift and very strong tip vortices. These vortices make the air flow outwards and inwards in circular paths around the ends of the aerofoils. (Figure 19-14). Vortex generators serve three quite different purposes. Their purpose determines their location: ! aileron mounted – rows of vortex generators located on the upper surface of the wing just upstream of the ailerons delay the onset of drag divergence at high speeds. ! rudder mounted – rows of vortex generators mounted on both sides of the vertical fin just upstream of the rudder prevent airflow separation over the rudder during extreme angles of yaw. This applies more to propeller multi-engined aircraft. ! elevator mounted -rows of vortex generators located on the underside of the horizontal stabiliser (and occasionally the upper surface) just upstream of the elevators prevent airflow separation over the elevators at very low speeds. Vortex generators on wing surfaces improve high-speed characteristics and when installed on the tail, they improve the low speed characteristics. STALLING SPEED AT HEIGHT The high thrust available from jet engines means that greatly increased altitudes are possible. Nevertheless, maximum altitude needs to be considered carefully. The TAS of a high incidence stall will increase with altitude, while the TAS of a shock stall will decrease from sea level to the base of the stratosphere. It will then remain constant. The situation is shown in Figure 19-15. A sea-level stall speed of 90 kt and a Mcrit of 0.80 are assumed. The shaded portions of the graph indicate where flight is not possible without stalling, either from a high angle of incidence or from

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Figure 19 – 15 Subsonic speed range (TAS)

This aircraft cannot fly above F780 without stalling, regardless of the power available and it can only maintain flight at this level at Vs. Slower speeds will produce a high incidence stall and higher speeds will produce a shock stall. This situation was often called "Coffin corner". The same situation is illustrated in Figure 19-15, but with IAS as datum. It can be seen that, regardless of engine power available, there is a limit to the maximum altitude for a specific aircraft and the operating speed range narrows considerably with increasing altitude.

Figure 19 – 16 Subsonic speed range (IAS)

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Manoeuvring will considerably reduce operating altitude, as the Vs from high incidence will increase with manoeuvres. The shock stalling speed will also decrease. Illustrative figures are given in Figure 19-17.

90 kts

180 kts

360 kts

380 kts

Increase of High Incidence Stalling Speed

400 kts

570 kts

M = 1.0 Sea Level

M = 0.7 40 000 ft

At 4g, this may be Reduced to, say, 380 kts

M = 1.0 40 000 ft

Shock Stall (True Air Speed)

Further increase at high M To, say

Stall at 4g, at 40 000 ft

Normal Stall at 40 000 ft

Normal Stall, Sea Level

High Incidence Stall (True Air Speed)

661 kts

Decrease of Shock Stalling Speed

Figure 19 – 17 Vs At Different Speeds

THE EFFECTS OF INCREASING MACH NUMBER Figure 19-18 is a very informative diagram, which shows the formation of shock waves and pressure distribution about a symmetrical bi-convex aerofoil section (at an angle of attack of 2°) with increasing MN. Pressure on the top and bottom surfaces is shown with the shock pattern illustrated below: ! The change of pressure distribution and shock pattern is shown progressively for MNs 0.75, 0.81, 0.89, 0.98 & 1.4. ! The solid line represents the pressure distribution on the upper surface of the wing and the broken line shows the pressure distribution below. ! Decreased pressure is shown upwards and increased pressure below, just as on the wing itself. ! As on a wing, the lower pressure above the wing is mainly responsible for lift and any decrease in pressure below the wing acts against lift. ! The difference between the two lines will indicate the efficiency of a particular part of the wing section. If the broken line is above the solid line, lift is actually negative. ! Total lift is shown by the area between the two lines and ! The centre of pressure is shown by the CG of the area. ! An increase in speed is shown upwards.

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A detailed analysis of the diagram shows the following.

Figure 19 – 18 Effect Of Increasing MN

At (a), separation is already apparent at the trailing edge and there is little lift over the rearmost 30% of the aerofoil. The centre of pressure is well forward and (as shown in Figure 18-18) the lift coefficient is rising steadily. In (b), the incipient shock wave has appeared on the top surface with a sudden increase of pressure (shown by the falling line) and decrease of speed at the shock wave. The centre of pressure has moved back slightly but the area is large (i.e. lift is good) as is shown in Figure 18-18 and is rising steadily. The CD is rising rapidly. In (c) the pressure distribution shows clearly why there is a sudden drop before the whole aerofoil reaches LSS. On the rear portion of the wing the lift is negative because the reduced pressure (suction) on the top of the wing has been spoilt by the shock wave. On the lower surface there is still good suction and high-speed flow. On the front portion there is nearly as much suction below as above. The centre of pressure has moved well forward again and drag is increasing rapidly (Figure 19-19). At (d) the picture is interesting as it shows the important results of the shock waves moving to the trailing edge of the wing. The shock waves no longer spoil the suction or cause separation. The speed of the flow over the surfaces is nearly 100% supersonic and the centre of pressure has moved back to approximately 50% MAC. As the pressure over most of the upper wing surface is reduced and rather less on the lower surface, the CL has actually increased (Figure 19-19). The CD has almost reached maximum with the position being almost at the "barrier" At (e) the bow wave has appeared and the barrier has been "broken". For the first time the speed of the airflow over some 50% of both surfaces is less than the MN of the aerofoil as a whole. The CL has fallen again as the pressures on both surfaces are nearly equal. For the first time since Mcrit near the approach to the "barrier", the aircraft is now in the "transonic" regime.

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Figure 19-19 shows the CL and CD corresponding to Figure 19-18.

Figure 19 – 19 CL with increasing Mach number

Figure 19 – 19 CD with increasing Mach number

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FILL IN THE GAPS Fill in the gaps (or circle the correct answer) in the following exercise by referring to the text in the previous chapter. 1.

The speed of sound is defined as the speed at which a very small ________ ___________ is propagated in a fluid under specified conditions.

2.

The speed of sound varies with _______ and ___________

3.

The speed of sound can be calculated from the formula: ___ = k1T.

4.

At low speed the air around a wing is termed "______________".

5.

As speed increases, it _______ __ with the pressure waves moving ahead of it.

6.

As an aircraft approaches M1.0, a ___________ ____ forms at the leading edge.

7.

Pressure disturbances can/cannot be propagated faster than the speed of sound.

8.

Compressibility effects depend mainly on Mach number/airspeed.

9.

The "Free Stream Mach Number" is the MN of the airflow near the wing, which is affected/unaffected by the passage of the wing.

10.

Mcrit occurs when the free air stream around an aerospace first reaches the _____ __ _____.

11.

The speed of propagation of a pressure wave is equal to the ___.

12.

Compressibility effects can occur/cannot occur at speeds below LSS.

13.

A shock wave normal to the free stream flow is called a "______ shock wave".

14.

The flow immediately behind a normal shock wave is always ________, regardless of the free stream MN.

15.

The transition of flow from supersonic to subsonic without a change in direction always forms a _____ ____.

16.

The LSS, air density & temperature behind the shock wave always ________.

17.

When a supersonic wave is turned away from the preceding flow, an "_________ ____" is formed.

18.

The rise in pressure at supersonic speeds causes an ________ in drag.

19.

At supersonic speeds, the effectivity of the controls will ________.

20.

"Tuck under" is not the same/is the same as control reversal.

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ANSWERS: 1. pressure disturbance 4. incompressible 7. cannot 10. speed of sound 13. normal 16. increase 19. decrease

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2. 5. 8. 11. 14. 17. 20.

density\temperature catches up 6. Mach number LSS subsonic. "expansion wave" is not the same.

3.

LSS compression wave 9. unaffected 12. can occur 15. shock wave. 18. increase

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CHAPTER TWENTY – HYDROPLANING

CONTENTS OF THIS CHAPTER: !

The dangers of hydroplaning

!

Types of hydroplaning

!

Viscous hydroplaning

!

Dynamic hydroplaning

!

Hydroplaning speeds

!

Factors reducing hydroplaning

!

Summary of hydroplaning types

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HYDROPLANING INTRODUCTION Hydroplaning (which is often referred to as "aquaplaning") may occur to a certain degree whenever a tyre contacts a wet runway. When a tyre touches a (dry) runway, it normally spins up straightaway to the aircraft groundspeed from zero. If it later moves into a wet part of the runway, it may spin down or even stop rotating completely. This happens because when a tyre is rolling along a wet runway it continuously squeezes water from underneath the tread. The result is that pressures can be produced which will actually raise the tyre from the runway surface. This reaction then reduces the amount of friction normally developed by the tyre. Clearly, any reduction in friction between the tyre and the runway will reduce the braking ability of the aircraft – it will increase the braking distance required. In fact a tyre impacting on a water-covered runway may not even spin up at all because of hydroplaning. It will therefore require a much greater available stopping distance, as the brakes will not be effective until the wheel leaves the hydroplaning section of the runway and is in proper contact with the dry runway surface. It could even be said that whenever a tyre moves over a wet surface, it is in fact – technically speaking – hydroplaning. The friction which is normally produced will attenuate as speed reduces. However, the time (i.e.0 distance) required to slow down and stop may well mean that the aircraft will overrun the far end of the runway. Every year a number of over-run accidents take place which are attributable to the primary factor of reduced runway friction. When hydroplaning takes place stopping distances increase and directional control reduces. THE DANGERS OF HYDROPLANING The dangers of hydroplaning are obvious and the effects are serious. Braking efficiency is reduced – or can disappear completely; steering is difficult – particularly with crosswind conditions. The most important result is that landing distance required is greater than normal, sometimes by a very high factor. In these cases, stopping distances required can be even 250% greater than the normal dry runway stopping distance. TYPES OF HYDROPLANING There are three types of hydroplaning : ! Viscous hydroplaning ! Dynamic hydroplaning ! Reverted rubber hydroplaning. Viscous Hydroplaning Viscous hydroplaning is the most common form of hydroplaning and it can occur on any wet runway. Viscous hydroplaning is the technical term to describe the lubricating action or slipperiness of water.

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A thin water film on the runway surface acts like a lubricant and hydroplaning can take place at speeds lower than with "dynamic" hydroplaning (see later). Although viscous hydroplaning conditions may exist whenever the runway is wet, most runway surfaces have a coarse textured "finish". This coarseness will break up the surface sufficiently and dissipate the thin water film which can produce viscous hydroplaning. When viscous hydroplaning is combined with the thin rubber coating left by aircraft tyres, which is often found in the touchdown zone of a runway, braking efficiency can be reduced by 65%. With ice or rain present, braking efficiency can be reduced by as much as 100%. This translates into a serious stopping problem. Factors contributing to viscous hydroplaning are : ! damp or wet runway ! medium to high speed ! poor runway (pavement) texture ! worn tyre tread.

Figure 20-1 Types of Hydroplaning

Dynamic Hydroplaning Dynamic hydroplaning occurs when a tyre lifts completely off the runway surface and "planes" on a wedge of water in the same way that a water ski rides on the surface. The depth of the wedge can be as small as 0.025 mm. This is "total dynamic hydroplaning". Conditions which lead to this type of hydroplaning are: ! high speed ! deep standing water (greater than 0.25 mm) ! poor surface friction ("macro texture"). Any one of these conditions must be present for dynamic hydroplaning to take place. If one condition disappears as the tyre is hydroplaning, the tyre will fall off the wedge of water, or hydroplaning will only affect a small portion of the tyre footprint. The condition is actually quite rare, but when it does occur, the tyre, as it is raised from the runway surface, loses friction and therefore wheel braking is non-existent.

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Other factors contributing to dynamic hydroplaning are : ! low tyre pressures ! worn tyre tread. Reverted Rubber Hydroplaning Reverted rubber hydroplaning occurs when a locked tyre skids along an icy or very slippery wet runway. It can occur at any speed above 20 kt. The locked tyre does not experience friction strong enough to rotate, although there may be enough friction to produce a "flat spot" on the tyre. On this flat spot, the rubber in the tyre can actually melt and the heat generated can produce steam or super-heated steam at temperatures as high as 300°C. This steam will produce "blisters" on the tyre. It often leaves white markings which can be seen on the runway surface. This type of hydroplaning is the most serious which is likely to be met, as there is hardly any braking friction while the wheels are locked. Factors which contribute to reverted rubber hydroplaning are : ! wet or flooded runway ! high speed ! poor runway texture ! deficient braking system. The most common form of hydroplaning occurs with the combined effects of viscous and dynamic hydroplaning. Reverted rubber hydroplaning appears to develop only in an aircraft with tyre pressure greater than 24 psi (165 kPa). HYDROPLANING SPEEDS Hydroplaning speeds are related directly to tyre pressures: hydroplaning speed (Vp) = 9 p where Vp is the speed in kt and p is pressure in psi. This formula is valid for a rotating tyre which subsequently enters a flooded section of the runway and then spins down or stops rotating. When a tyre is not rotating, the hydroplaning speed is slower : Vp = 7.7 p. This speed (or, of more interest – pressure) is of greater relevance when considering the actual speeds at which hydroplaning can take place. It should be remembered that this equation is in fact an approximation of hydroplaning speed. The actual speed will vary in practice with the texture of the runway surface, water depth and tyre tread depth.

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It may be thought a good idea to increase tyre pressures and thereby reduce the chance of hydroplaning. However, increased tyre pressure will mean a smaller tyre "footprint" (the area of the tyre in contact with the braking surface). This, in turn, therefore means less traction and greater dry and wet stopping distances. The best method is to follow the tyre manufacturer's recommendations. Recommended pressure will usually be a compromise between the various requirements of tyres used in different environments, all of which are demanding. FACTORS REDUCING HYDROPLANING Several factors can reduce the possibility of hydroplaning: ! runway grooving (longitudinal and transverse) ! high tyre pressures ! runway re-surfacing (coarse texture) ! effective anti-skid systems ! good tyre tread design ! runway crowning ! removal of surface rubber build-up on touchdown areas ! strict adherence to the calculated approach and touchdown speeds ! correct descent profiles avoiding an extended flare ! correct use of spoilers and reverse thrust ! correct braking techniques (auto-braking where possible) ! use of maximum landing flap possible. Perhaps the most important of all is to avoid excess speed. High speed obviously increases the stopping distance required. Excess speed on the approach is a contributory cause in almost every over-run situation. It produces "floating" and "re-bound". For example, a figure of 10 kts extra can typically mean an extra 2 000 ft of stopping distance. Once on the ground, an extra 10 kts requires 200 ft more (dry) and 500 ft more (wet runway). A "moderately firm" touchdown, then lowering the nose wheel as soon as the main wheels touch down is the Boeing recommendation. "Greasers" are definitely out in hydroplaning conditions. Boeing also recommends that crosswind landings should be avoided in the same situation. Turning off the runway can also be dangerous, since a rubber build-up can be anticipated at the far end of the runway. It's best to complete most braking in the initial and middle sections of the runway and anticipate poor braking conditions at the far end. Remember that reverse thrust is most efficient at high speed and should be used as soon as possible after touchdown.

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SUMMARY OF HYDROPLANING TYPES

Action

Factors which enhance the case

Factors which alleviate the case

VISCOUS

DYNAMIC

Normal slipperiness on wet/damp runway. A thin water film acts as a lubricant. Rubber deposits also can make tyres skid. Most common form.

Tyre "lifted" from runway by a wedge of water rather like a water ski. Known technically as total aquaplaning – wheel does not "spin-up".

Damp or wet runway Medium/high speed Poor runway texture Worn tyre tread Runway micro texture Runway grooving Good tread design

Flooded pavement High speed Low tyre pressure Worn tyre tread Runway macro texture Runway grooving Higher tyre pressure

REVERTED RUBBER Tyre is locked and then skids on smoother icy surface. Heat produced by friction on "flat spot" can generate steam and melt the tyre tread. Wet/flooded runway High speed Poor runway texture Deficient brakes Good runway texture Runway grooving Improved anti-skid Good tread design

References : Runway stopping (Professional Pilot) Hydroplaning (Boeing Aircraft Company) Stopping safely (World Air News) Aircraft Tyre Safety (ALPA-SA)

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CHAPTER TWENTY-ONE – FIRE PROTECTION SYSTEMS

CONTENTS OF THIS CHAPTER: !

Fire Detection

!

Fire Extinguishers

!

Classification of Fires

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FIRE PROTECTION SYSTEM INTRODUCTION When designing aircraft fire systems the focus is first to prevent fire and secondly to provide adequate fire protection. The most common causes of aircraft fire are: ! Fuel leaks in the vicinity of hot equipment ! Hot gas leaks from engines or ducting coming into contact with inflammable materials ! Electrical or mechanical malfunctions in equipment FIRE DETECTION Fire detection and overheat systems sense the presence of fire or excessive heat and consist of ‘area detectors’ in large fire zones and ‘spot detectors’ for individual pieces of equipment. Some types of detectors are: ! Rate of temperature rise indicators ! combustible mixture detectors ! Smoke detectors ! Overheat detectors ! Flame detectors ! Radiation sensing detectors In the case of fire, overheat or smoke the detection systems provide visual (red light) and audible alarms to the crew in order to identify the area in which the problem exists and then to complete electrical safety circuits within the fire extinguisher systems to permit necessary system operation by the crew. Three basic principles of operation are used in detectors,Simple electrical switches activated by the different thermal expansions of dissimilar metals (Fig. 21 – 1). At normal temperatures the contacts attached to the assembly remain open. As the temperature rises the cylinder expands allowing the contacts to spring together which closes the electrical circuit causing the light to go on and the alarm to sound.

Fig. 21 – 1 Typical Heat Sensitive, SelfResetting Thermal Switch

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Sensors – used in the majority of ‘area fire detectors’. The sensing element is in the form of a wire about 2 mm diameter and is known as the continuous wire detection element. (Fig.21 – 2) A continuous wire is routed through the various compartments and because it is vulnerable to damage caused by vibration it is protected by insulation known as ‘thermister’. A weak electrical current carried by the wire completes the circuit when any section along the wire exceeds the prescribed temperature. Stainless Steel Tube

Filling Material

Central Electrode Fig. 21 – 2 Continuous wire detection element

Crash switches are usually installed in the undercarriage bays or inside the belly of the aircraft. If subjected to excessive horizontal deceleration a bank of electrical contacts in the fire extinguisher circuits is actuated. FIRE EXTINGUISHERS Engine Fire Extinguishers Permanently installed fire extinguisher systems are provided to suppress fires in the engine section, ancillary section and main gear. The systems are designed to deliver predetermined volumes of extinguishing agent from the fire bottles to designated areas of the engine installation. These extinguishers are filled with inert gas or halocarbon agents (Halon which is a type of Freon). These are used due to their rapid knockdown effect. When released from the system the gas blankets the fire purging oxygen away from it. Although fire extinguishers vary in construction, they generally comprise of two sections, the steel or copper container, and the discharge or operating head. The container has an externally threaded neck, onto which the discharge head is screwed and then soldered. Container

Extinguishant Fig. 21 – 3 Example of an engine fire extinguisher

Discharge Head

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The head comprises the cartridge unit, which contains the charge. The purpose of the charge is to dislodge the plug which keeps the extinguishant in the cylinder. The charge is electronically actuated, by any one of the switches previously described.

Charge Explosive Charge Sleeve Cartridge Unit Plug Outlet Adapters

Detachable cup Charge Spigot

Fig. 21 – 4 Example of a discharge head

Extinguishant released into the system

Charge spigot

Plug Fig. 21 – 5 Operation of the fire extinguisher

As the system is actuated, the charge which is directed toward the plug, dislodges this plug, which in turn strikes the charge spigot. The charge spigot is propelled to the end of the head, where it strikes the mechanical indicator (Indicating that the system has been actuated), the plug is also positioned at the bottom of the head after actuation. This action releases the extinguishant into the extinguisher system. Cabin Fire Extinguishers Portable Fire Extinguishers (hand held) are installed in various sections of the cabin areas of the aircraft. These extinguishers contain either carbon dioxide or a water/glycol mixture. The extinguishers are prevented from expelling gas at high ambient temperature by a pressure relief valve.

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Copyright © 2012 EAA

Eltanin Aerospace Academy

EAA-CPL-MAN-010v1

CLASSIFICATION OF FIRES Fires are classified according to the material that is being burned. The four classes of fires, with the American and International symbols, are as follows: Class A:

Ordinary Combustibles – cloth, wood, paper, rubber, many plastics.

Extinguisher: (1) Pressurized water suitable for use on Class A only. (2) Dry Chemical, ammonium phosphate, rated for Class A, B, and C fires. Class B:

Flammable Liquids – Gasoline, Oil, Oil-based paint, Cooking Oil.

Extinguisher: (1). Carbon dioxide, expelled as a gas and non-corrosive. Must be used very aggressively because it dissipates quickly. (2). Dry Chemical, Ammonium phosphate, rated for Class A, B, and C fires. Works very well on Class B, however, it produces quite a mess. (3) Dry chemical, sodium bicarbonate and potassium bicarbonate, urea-based potassium rate for Class B and C, preferred for extinguishing cooking oil fires – also quite messy. (4) Halon is expelled as a gas and is non-corrosive. Halon extinguishers are being phased out due to environmental issues. Class C:

Energized electrical equipment, including appliances, wiring, circuit breakers, and fuse boxes.

Extinguisher: (1) Carbon dioxide, expelled as a gas and non-corrosive. Must be used very aggressively because it dissipates quickly. (2) Dry Chemical, Ammonium phosphate, rated for Class A, B, and C fires. Works very well on Class B, however, it produces quite a mess. (3) Dry chemical, sodium bicarbonate and potassium bicarbonate, urea-based potassium rate for Class B and C, preferred for extinguishing cooking oil fires – also quite messy. (4) Halon is expelled as a gas and is non-corrosive. Class D:

Combustible metals such as magnesium, sodium, lithium, hafnium, powdered aluminium, etc.

Extinguisher: Extinguishing agents for Class D fires must match the type of metal involved. Extinguishers rated for Class D fires have a label which lists the types of metals on which the extinguisher may be used.

Copyright © 2012 EAA

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