Basic Jet Engine
Principle of Jet Propulsion
Jet propulsion is a practical application of Sir Issac Newton’s third law of motion which states that for every force acting on the body there is an opposite and equal reaction for aircraft propulsion, the aerobody is atmosphere air that is caused to accelerate as it passes through the engine. The force required to give this acceleration has an equal effect in the opposite direction acting on the apparatus producing the acceleration.
The jet engine produces thrust in a similar way to the propeller/engine combination, but where as the propeller gives a small acceleration to a large weight of air, the jet engines gives a large acceleration to small weight of air.
The familiar whirling garden sprinkler is a more practical example of this principle, for the mechanism rotates by virtue of the reaction to the water jets. Jet reaction is definitely an internal phenomenon and does not, as is frequently assumed, result from the pressure of the jet on the atmosphere. In fact the jet propulsion engine, whether rocket, athodydes or turbojet is a piece of apparatus designed to accelerate a large stream of air and to expel it at an exceptionally high velocity.
There are of course a number of ways of doing this but in all instances the resultant reaction or thrust exerted on the engine is proportional to the mass of air expelled by the engine and to the velocity change imparted to it. In other words, the same thrust can be provided either by giving a large mass of air a little extra velocity or a small mass of air, a large extra velocity.
Method of Jet Propulsion
The types of jet engines, whether ram jet, pulse jet, rocket gas turbine, turbo/ram jet or turbo-rocket differ only in the way in which the thrust provider or engine, supplies and converts the energy into power for flight.
A Ram jet engine is an athodyde or aero thermodynamic duct to give it its full name. It has no major rotating parts and consists of a duct with a divergent entry and a convergent or convergent-divergent exit. When forward motion is imparted to it from an external source, air is forced into the air intake where it loses velocity or kinetic energy and increases its pressure energy as it passes through the divergent duct. The total energy therefore, increased by the combustion of fuel, and the expanding gases accelerate to atmosphere through the outlet duct. A ram jet is often the power plant for missiles and target vehicles, but is unsuitable as an aircraft power plant because it requires forward motion imparting to it before any thrust is produced.
A pulse jet engine uses the principle of intermittent combustion and unlike the ram jet it can be run at a static condition. The engine is formed by an aerodynamic duct similar to the ram jet but due to higher pressure involved, it is of more robust construction. The duct inlet has a series of inlet valves that are spring loaded into the open position. Air drawn through the open valves passes into the combustion chamber and is heated by the burning of fuel injected into the chamber. The resulting expansion causes the rise in pressure, forcing the valves to close, and the expanding gases are therefore ejected rearwards. A depression created by the exhausting gases allows the valve to open and repeat the cycle.
Pulse jets have been designed for helicopter rotor propulsion and some dispense with inlet valves by careful design of the ducting to control the changing pressure of the resonating cycle. The pulse jet is unsuitable as an aircraft power plant because it has a high fuel consumption and is unable to equal the performance of the modern gas turbine engine.
Types of Jet Engines (Gas Turbine)
The gas turbine engine is basically of simple construction although the thermal and aerodynamic problem associated with its design are somewhat complex.
There are no reciprocating components in the main assembly and the engine is therefore essentially free from vibration. Power is produced in continuous cycle by compressing the intake air and passing it to the combustion chamber where fuel is added and burnt to provide heat. The expansion of gasses rearwards through the turbine produces the power necessary to drive, the compressor, the residual energy being used to provide jet thrust or in the case of turboprop engines, to drive a propeller.
Propeller efficiency falls off rapidly above approximately 350 knots so that turboprop engines are normally used to power comparatively low speed aircraft faster aircraft use turbojet engines, by pass or turbo fan engines being favoured for high subsonic speeds because of their fuel economy and low noise level. After burning i.e. the burning of fuel in the jet pipe to provide additional thrust, is normally used only in military aircraft due to large quantities of fuel consumed, but it may be used on civil aircraft for take off and acceleration to supersonic flight.
TYPES OF JET ENGINES
Variation in Gas Turbine Engine
Turbo Jet Engine
The Turbo Jet Engine consist of a compressor section which is connected to the turbine section via a shaft. It can consist of one or two compressor and turbine section, one of which is high pressure and other is the low pressure.
TURBO JET ENGINE
Turbo Fan Engine
The Turbo Fan Engine is similar in construction to the Turbo Jet Engine except that a fan is attached in the font on the intake section which is connected to the low turbine section. It can be two or three section turbine. The last turbine driving the fan and the other can be the low and high pressure section of the compressor and turbine.
The Turbo Fan Engine can be high bypass or low bypass engine.
TURBO FAN ENGINE
The Turboprop Engine is based on the Jet Engine with a variation of a propeller in the front which is connected to the compressor with a Gear box.
TURBO PROP ENGINE
Turbo Shaft Engine
The Turbo Shaft engine consist of a free turbine called as power turbine which is connected to the power shaft.
TURBO PROP ENGINE
Construction of Gas Turbine
In many types of turbine engines, it is not possible to list all the major components and have the list apply to all engines. There are several components and common to most turbine engines.
The operation of any gas-turbine engine requires that provision be made for three principal functions:
The compression of air
The expansion of the air by burning fuel and
The extraction of power from the jet stream of the engine for driving the compressor and accessories.
Thus, we may say that gas-turbine engine comprises three main sections the compressor section, the combustion section and the turbine section.
JET ENGINE CONSTRUCTION
In addition to these three main section, there are also component which serve to provide transition from one main functional section to another. The following is a list of all the major components, arranged as they would appear from the front of the turbine to the rear.
Inlet duct and guide vanes
The diffuser, with or without air adaptor
The combustion chamber
The nozzle diaphragm
The exhaust cone
The afterburners (if the engine is so equipped)
The accessory section (which may be located at the front of the engine or further to the rear)
The Inlet Duct And Guide Vanes
Turbine engine inlet duct must furnish a relatively distortion free and high energy supply of air on the required quantity to the compressor, the uniform and steady air flow is necessary to avoid compressor stall and excessively high engine temperature at the turbines. The high energy enables the engine to produce an optimum amount of thrust.
The air inlet duct is considered to be an air frame part. Inlet ducts has following functions
It must be able to recover, as much air as possible and deliver this pressure to the front of the engine with minimum pressure loss.
The duct must uniformly deliver air to compressor inlet with as little turbulence and pressure variations as possible
There are two basic types of duct
This is the simplest and most effective because of the duct inlet is located directly ahead of the engine and the aircraft is in such a position that it scoops the undisturbed air. The duct can be built strong and straight with relatively gentle curvatures. In single engine aircraft installation the duct is necessarily is relatively curved and hence some pressure drop is possible by the long duct, but the condition is offset by smooth air flow characteristic, although a short inlet duct results in minimum pressure drop, the engine often suffer from inner turbulence specially at low air speed and high angle of attack.
Divided Entry Duct
The requirements of high speed single engine aircraft in which pilot seat is low in the fuselage and close to the nose render it a difficulty to employ a single entrance duct. Some types of divided duct which takes air from either side of the fuselage may be required.
The divided ducts can be following types
Wing root entrances
Variable Geometry Duct
A supersonic inlet duct progressively decreases in area in the down streams, again it will follow the general configuration until the velocity of the incoming air is reduced to match 1 and below. The aft section of the duct will then commence to increase in area since this part must act as a subsonic diffuser.
For very high speed aircraft, the inside area or configuration of duct will be changed by a mechanical device as the speed of the aircraft increases or decreases. The duct of this type is usually known as variable geometry inlet duct.
Two main methods used to diffuse the inlet air and reduce the inlet air velocity, at supersonic flight speed, one method is to vary the area of geometry of inlet duct either by using a movable restriction such as wedge inside the duct, another method is some short of variable air flow bypass arrangement which extracts part of inlet air flow from the duct ahead of the engine. Another method is by using a shock wave in the air stream, a shock wave is a thin region of discontinuity in a flow of air or gas during which velocity, density temperature of gas undergo a sudden change.
A shock wave is set up in a supersonic flow of air entering the duct by means of some restriction which automatically protrudes into the duct in high flight machine. The shock waves results in diffusion of this airflow which reduces the velocity of air. In some cases both shock waves method and variable geometry method of casing diffusion are used in combination.
Bell Mouth Air Inlet
Bellmouth air inlet are convergent in shape and are used on helicopter and slow moving aircraft which flies below ram recovery speed. This type of inlet reduces a large brake factor but drag is outweighed by high degree of Aerodynamic efficiency.
Engines being calibrated on ground run test, also utilizes bell mouth anti ingestion screen. Duct loss is slight in this design that it is considered to be zero. Engine performance data such as engine trimming while using a bellmouth engine inlet, aerodynamic efficiency and duct loss. The rounded L.E. allows air stream to make use of total inlet cross section where as effective diameter of sharp edge orifice is greatly reduced.
Air Inlet Vortex Destroyer
When the jet engine operating on ground, the engine air inlet vortex can sometimes be formed between the air inlet and ground. This vortex can cause strong suction force capable of lifting foreign objects from the ground into the engine causing serious damage. To minimize the ingestion of debris, an inlet vortex destroyer is used. This destroyer is nothing but a small jet stream directed downwards from the lower L.E. of the nose cowling to the ground to destroy the vortex base. Bleed air from the engine is used as the vortex destroying stream, it is controlled by a valve located in the nose cowl. The control valve is a two position valve which is opened by a L/G weight switch. The valve closes when the aircraft leaves the ground and weight of aircraft is removed from the L/G. The valve opens when touches ground and when weight switch contact is made.
Foreign Object Damage
One of the major problems encountered in the operation of axial flow engines is foreign object damage. Rocks drawn into the air inlet during taxing cause considerable damage because they nick or scratch the compressor and turbine blades as they pass through the engines, which can lead to fatigue failure with the result that the engine may throw a blade in flight. This could result in loss of the aircraft or serious damage to the engine.
To prevent foreign object damage, the air inlet on the engine are screened. These screens are effective in removing large objects from the air stream, but they will not prevent small rocks from entering the engine. Small rocks, sand and grass can do a great amount of damage to the engine.
Air Inlet Icing
The air screen at the inlet of an axial flow engine is subject to icing, with the result that the engine may stop. The engine nose cowling nose dome and inlet guide vanes are subject to icing, however, and it is necessary to incorporate provisions in the engine nose cowlings to prevent the formation of ice. Jet engine anti icing system normally make use of a high temperature air from the diffuser section.
There are two types of compressor,
Both types are driven by the engine turbine and are usually coupled direct to the turbine shaft.
The centrifugal flow compressor is a single or two stage unit employing an impeller to accelerate the air and a diffuser to produce the required pressure rise.
The axial flow compressor is a multi stage unit employing alternate rows of rotating (rotor) blades and stationary (stator) blades to accelerate and diffuse the air until the required pressure rise is obtained.
TYPES OF COMPRESSORS
With regards to the advantages and disadvantages of two types, the centrifugal compressor is usually more robust than the axial compressor and is also easier to develop and manufacture. The axial compressor, however, compresses more than a centrifugal compressor of the same frontal area and can also be designed for high pressure ratios much more easily. Since the air flow is an important factor in determining the amount of thrust, this means that the axial compressor engine will also give more thrust for the same frontal area.
The Centrifugal Flow Compressor
Have a single or double sided impeller and occasionally a two-stage, single sided impeller is used as on the Roll’s Royce Dart. The impeller is supported in a casing that also contains a ring of diffuser blades. If a double entry impeller is used, the airflow to the rear side is reversed in direction and a plenum chamber is required.
TYPES OF CENTRIFUGAL COMPRESSORS
Principles Of Operation
The impeller is rotated at high speed by the turbine and air is continuously induced into the centre of the impeller. Centrifugal action causes it to flow radially outwards along the vanes to the impeller tip, thus accelerating the air and also causing a slight rise in pressure to occur. The engine intake duct may contain vanes that provides an initial whirl to the air entering the compressor.
The air on having the impeller, passes into the diffuser section where it passages from the divergent nozzles and converts most of the kinetic energy into pressure energy. In practice, it is usual to design the compressor so that about half of the pressure rise occurs in the impeller and half in the diffuser.
The air mass flow through the compressor and the pressure rise depend on the rotational speed of the impeller, therefore impellers are designed to operate at tip speed of up to 1600 ft per second. By operating at such high tip speeds, the air velocity from the impeller is increased. So that greater energy is available for conversion to pressure. Another factor influencing the pressure rise is the inlet air temperature, for the lower temperature of air entering the impeller the greater the pressure, the pressure rise for a given amount of work put into the air by the compressor, is a measure of the increase in the total heat of the air passing through the compressor.
To maintain the efficiency of the compressor, it is necessary to prevent excessive air leakage between the impeller and the casing, this is achieved by keeping their clearances as small as possible.
CENTRIFUGAL FLOW COMPRESSOR
The construction of the compressor centres around the impeller diffuser and air intake systems. The impeller shaft rotates in ball and roller bearings and is either common to the turbine shaft or split in the centre and connected by a coupling, which is usually designed for case of detachments.
The impeller consists of a forged disc with integral, radially disposed vanes on one or both sides forming divergent passages. The vanes may be swept back but for ease of manufacture straight radial vanes are usually employed. To ease the change of air flow from the axial to the radial direction, the vanes in the centre of the impeller are curved in the direction of rotation. The curved sections may be integral with the radial vanes, or formed separately for easier and more accurate manufacture.
The choice of impeller is determined by the engine design requirements, but it is claimed that the single entry ducting allows the air to be fed into the compressor at the best all round efficiency. It is also claimed that the single entry ducting minimizes the chances of surging at altitude, because it makes more efficient use of the ram effect than, does the double entry ducting. A small amount of heating also occurs on the double entry ducting.
The diffuser assembly may be an integral part of the compressor casing or a separately attached assembly. In each instance it consists of a number of vanes formed tangential to the impeller, The vanes passages are divergent to convert the kinetic energy into pressure energy and inner edges of the vanes are in line with the direction of the resultant airflow from the impeller. The clearance between the impeller and the diffuser is an important factor, as too small a clearance will set up aerodynamic impulses that could be transferred to the impeller and create an unsteady airflow and vibration.
DIFFUSER VANE AND IMPELLER
Axial Flow Compressor
An axial flow compressor consists of one or more rotor assemblies that carry blades of aero foil section and are mounted between bearings in the casings in which are located the stator blades. The compressor is a multi stage unit as the amount of work done (pressure increase) in each stage is small, a stage consists of a row of rotating blades followed by a row of stator blades. Some compressors have an additional row of stator blades, known as intake or inlet guide vanes, to guide the air on to the first row of rotor blades. The angular setting of the vanes can be automatically controlled to suit the airflow requirements at various operating conditions.
From the front to the rear of the compressor, i.e. from the low to high pressure end, there is gradual reduction of the air annulus area between the rotor shaft and the stator casing. This is necessary to maintain the axial velocity of the air constant as the density increases through the length of the compressor. The convergence of the air annulus is achieved by the tapering of the casing or rotor. A combination of both is also possible, with arrangement being influenced by manufacturing problems and other mechanical design factors.
A single spool compressor consists of one rotor assembly and stators with as many stages as necessary to achieve the designed pressure ratio, and all the airflow flow from the intake passes through the compressor. The multi spool compressor consists of two or more rotor assemblies, each driven by their own turbine at an optimum speed to achieve higher pressure ratio and to give greater operating flexibility.
Although a twin spool compressor can be used for a pure jet engine, it is most suitable for the by pass type of engine where the front or low pressure compressor is designed to handle a larger mass airflow than the high pressure compressor. Only a percentage of the air from the low pressure compressor passes into a high pressure compressor, the remainder of the air, the bypass flow is ducted around the high pressure compressor. Both flows mix in the exhaust system before passing to the propelling nozzle.
A fan may be fitted to the front of a single or twin spool compressor and on these types of engines the fan is driven at the same speed as the compressor to which it is fitted. On engines of the triple spool type, the fan is in fact the low pressure compressor and is driven by its own turbine separately from the intermediate pressure compressor and the high pressure compressor. The low pressure compressor has large rotor (fan) blades and stator blades is designed to handle a far larger mass airflow and the other two compressor, each of which has several stages of rotor blades. A large proportion of air from the lower part of the fan and known as the cold stream, by passes the other two compressors and is ducted to atmosphere through the cold stream nozzle. The smaller airflow, from the inner part of the fan and known as hot stream passes through the intermediate and high pressure compressor when it is further compressed before passing into the combustion system.
Principles Of Operation
During operation, the rotor is turned at high speed by the turbine, so that air is continuously induced into the compressor, where it is accelerated by the rotating blades and swept rearwards on the adjacent row of stator blades.
The pressure rise in the airflow results from the diffusion process in the rotor blade passages and from a similar process in the stator blade passages; the latter also serves to correct the deflection given to the air by the rotor blades and to present the air at the correct angle to the next stage of rotor blades. The last row of stator blades usually act as “air straightener” to remove the whirl from the air so that it enters the combustion system at a fairly uniform axial velocity.
The changes in pressure and velocity occur in the airflow through the compressor. These changes are accompanied by a progressive increase in air temperature as density increases.
Across each stage, the ratio of the total pressures of the out going air and inlet air is quite small, being between 1:1 and 1:2. The reason for the small pressure increase through each stage is that the rate of diffusion and the deflection angles of the blades must be limited if losses due to air break away at the blades, and subsequent blade stall are to be avoided.
The small pressure rise through each stage together with the smooth flow path of the air, does much to contribute to high efficiency of the axial flow compressor. For instance, the maximum air velocity through the axial compressor corresponds to a Mach number of about 0.9 and the flow is almost of thorough. On the other hand, the velocity through a centrifugal compressor is super sonic in places, reaching a Mach number of 1.2; the flow in this instance is tortuous culminating in a right angle bend at the outlet to the combustion chamber.
Because an axial flow compressor requires a large number of stages to produce a high compressor ratio, as the number of stages increases it becomes more difficult to ensure that each stage will operate efficiently over a engine speed range.
An automatic system of airflow control is sometimes necessary to maintain compressor efficiency, but a more flexibly operated engine can be achieved by having more than one compressor with each compressor being an independent system, driven by separate turbine assemblies through coaxial shafts. The compressor, therefore be designed to operate more efficiently and with greater flexibility over a wide speed range.
A by pass engine invariably has a spool compressor with the low pressure compressor supplying sufficient air for both the by pass system and the high pressure compressor. Still greater flexibility can be obtained and higher maximum compression ratios reached by using an automatic airflow control system for high pressure compressor, this method is used on the Rolls Royce spey series of engines. Although an engine may have a front or an aft fan, the front fan is favored by most manufactures as giving greater reliability, due to the fan operating in the cold section of the engine.
The fan can have one or more stages of large blades, both rotor and stator. The rotor blades can be fitted to the front of a compressor or be part of a complete compressor driven by its own turbine. The air accelerated by the outer portion of the blades forms a by pass or secondary airflow that is ducted to atmosphere, the main airflow from the inner portion of blade passes through the remainder of the compressor and into the combustion systems. Only one stage of blades is used on the fan of triple spool engines, because the blades are designed to operate at transonic tip speeds. This permits the desired compression ratio to be achieved and not only reduces the weight of engine but also its noise level.
The construction of the compressor centers around the rotor assembly and casings. The rotor shaft is supported in ball and roller bearings and is coupled to the turbine shaft. The casing assembly consists of a number of cylindrical casings some of which are in two halves to facilitate engine assembly and inspection, these are bolted together to completely house the rotor.
The rotor assembly may be of a disc construction or of drum, or a combination of both types may be used. The drum type rotor consists of a one or two piece forging on to which are secured the rotor blades. The disc type rotor has the rotor blades attached to separate discs, which are then splined to the rotor shaft and separated by integral or individual spacer rings. In the former type axial thrust and radial load both are taken by the drums where as in disc type radial load is taken by the disc and axial thrust by the black platform and spacer rings. The accumulated end thrust is taken by the end of the or the end discs.
The rotor blades are of aero foil section and are usually designed to give a pressure gradient their length to ensure that the air maintains a fairly uniform axial velocity. The higher pressure towards the tip balances out the centrifugal action of the rotor on the airstream. To obtain thrust condition, it is necessary to twist the blade from root to tip to give the correct angle of incidence at each point. The length of the blades varies from front to rear, the front or low pressure blades being the longest.
The stator blades are again of aero foil section and are secured into the compressor casing or into stator blade retaining ring, which are themselves secured to the casings. The blades are often mounted in packs in the front stages and may be shrouded at their tips to minimize the vibrational effect of flow variation on the longer blades. It is also necessary to lock the stator blades in such a manner that they will not rotate around the casings.
Each stage of a multi-stage compressor processes certain airflow characteristics that are dissimilar from those of its neighbor thus, to design a workable and efficient compressor, the characteristics of each stage must be carefully matched. This is a relatively simple process to carry out for one set of conditions (design mass flow, pressure ratio and rotational speed), but is much more difficult if reasonable matching is to be retained when the compressor is operating over a wide range of conditions such as an aircraft engine encounters.
Outside the design conditions, the flow around the blade tends to degenerate into violent turbulence when the smooth flow of air through the compressor is disturbed. Although the two terms ‘stall' and ‘surge' are often used synonymously, there is a difference which is mainly a matter of degree.
A stall may affect only one stage or even a group of stages, but a compressor surge generally refers to a complete flow breakdown through the compressor.
Compressor blades are designed to produce a given pressure rise and velocity increase over the engine speed range.
If something should disturb the pressure, velocity, rotational speed relationship the airflow across the blade profile will break away and create eddies until eventually the blade ‘ stalls ‘. This could occur if the airflow was reduced due to icing or a flight maneuver, or if the fuel system scheduled too high a fuel flow; damage due to ingestion could, of course, create a similar condition.
If the stall condition of a stage or group of stages continues until all stages are stalled, then the compressor will surge. The transition from a stall to a surge could be so rapid as to be unnoticed; on the other hand, a stall may be so weak as to produce only a slight vibration or poor acceleration or deceleration characteristics.
At low engine speeds or 'off design' speeds, a slight degree of blade stalling invariably occurs in the front stages of the compressor, even though a system of airflow control may be used. This condition is not harmful or noticeable on engine operation.
A more severe compressor stall is indicated by a rise in turbine gas temperature, vibration or ‘coughing’ of the compressor. A surge is evident by a bang of varying severity from the engine and a rise in turbine gas temperature.
The value of airflow and pressure ratio at which a surge occurs is termed the ‘surge point’. This point is a characteristic of each compressor speed, and a line which joins all the surge points called the ‘surge line’ defines the minimum stable airflow that can be obtained at any rotational speed. A compressor is designed to have a good safety margin between the airflow and the compression ratio at which it will normally be operated and the airflow and compression ratio at which a surge will occur.
AXIAL FLOW COMPRESSOR
AXIAL FLOW COMPRESSOR WORKING
The Diffuser and Air Adapter
The function of the diffuser assembly is to direct air from the compressor to the combustion chambers and to change air pressure and velocity as required for best fuel combustion. The air discharge from a centrifugal impeller enters equally spaced diffuser passages, and at the end of each is a Wrist type of elbow containing four vanes which turn the air 90 deg. into the compressor discharge. The diffuser has boxed type of single casting, with elbows and turning vanes cast integrally.
The diffuser for an axial-flow engine serves to carry the air from the compressor to the combustion chambers. For an engine equipped with individual “can” type combustion chambers, the diffuser must have a separate outlet shaped to fit the inlet of each combustion chamber.
In some axial-flow engines, the diffuser section is called the midframe. It not only contains the diffuser but also provides support for the mid bearing and mountings for the fuel-nozzle assemblies.
Air adapters on a centrifugal-flow engine carry the air from the diffuser to the combustion chambers. They also provide attachment for the fuel nozzles, domes or end caps of the combustion chambers, air adapters aid in slowing the air velocity and increasing the pressure as is desirable at this point of the thermodynamic cycle.
On axial-flow engines, the air adapter is actually the outlet of the diffuser section. Usually this portion of the assembly is not even named as the air adapter.
The amount of fuel added to the air will depend upon the maximum temperature rise required and, as this is limited by the materials from which the turbine blades and nozzles are made, the rise must be in the range of 700 to 1200º C because the air is already heated by the work done during compressor, the temperature rise required at the combustion chamber may be between 500 and 800ºC.
Since the gas temperature required at the turbine varies with engine speed and in the case of turbo-prop engine upon the power required, the comb. Chamber must also be capable of maintaining stable and sufficient combustion over a wide range of engine operating condition.
Air from the engine compressor enters the combustion chamber at a very high velocity which is further diffused and static pressure increased in the combustion chamber, this is done because the burning of kerosene at normal mixture ratio is only few fts/sec, and any fuel which is burnt at high velocity will be blown off or away. A region of low axial velocity has therefore to be created in the chamber so that the flame will remain alight throughout the range of energy separating conditions.
In normal operation the over all air/fuel ratio of a combustion chamber can vary between 45% and 30% kerosene however will only burn efficiently at or close to a ratio of 15:1 so the fuel must be burnt with only part of the air entering the chamber, in what is called a primary combustion zone. This is achieved by means of a flame tube (combustion liner) that has various devices for metering the airflow distribution along the chamber.
Approx. 18% of the air mass flow is taken in by the snout or entry section. Immediately downstream of the snout, are swirl vanes and a perforated flare, through which air passes into the primary combustion zone. The swirling air induces a flow upstream of the center of the flame tube and promotes the desired recirculation. The air not picked up by the snout flows into the annular space between the flame tube and the air casing.
Through the wall of the flame tube body, adjacent to the combustion zone, are a selected number of holes through which a further 10 to 15% of the main flow of air passes into the primary zone. The air from the swirl vanes and that from the primary air holes inter acts and creates a region of low velocity recirculation. This takes a form of a toroidal vortex similar to a smoke ring and has the effect of stabilizing and anchoring the flame. The recirculating gases hasten the burning of freshly injected fuel droplets by rapidly bringing them to ignition temperature.
It is arranged so that the conical fuel spray from the burner, intersects the recirculation vortex at the center. This action, together with the general turbulence in the primary zone, greatly assists in bringing up the fuel and mixing it with the incoming air.
The temperature of the combustion gases released by the combustion zone is about 1,800 to 2000º C which is far too hot for entry to the nozzle guide vanes of the turbine. The air not used for combustion, which is therefore introduced progressively into the flame tube.
Approx. half of this is used to lower the gas temperature before it enters the turbine and the other half is used for cooling the walls of the flame tube. Combustion should be completed before the dilution air enters the flame tube, otherwise the incoming air will cool the flame and incomplete combustion will result.
An electric spark from an igniter plug initiates combustion and the flame is then self sustained.
COMBUSTION CHAMBER WORKING
Types Of Combustion Chamber
There are three main types of combustion chamber in use, they are multiple chamber, the turbo annular chamber and the annular chamber.
Multiple Combustion Chamber
This type of combustion chamber is used on centrifugal compressor engines and the earlier types of axial flow compressor engines. It is a direct development of the earlier type of whittle combustion chamber. The major difference is that whittle chamber had a reverse flow, but as this created a considerable pressure loss, the straight through multiple chamber was developed by Joseph Lucas ltd.
The chambers are disposed around the engine and compressor delivery air is directed by ducts to pass into the individual chamber. Each chamber has an inner flame tube around which, there is an air casing. The air passes through the flame tube snout, and also between the tube and the outer casing.
The separate flame tubes are all interconnected. This allows each tube to operate at the same pressure and also allows combustion to propagate around the flame tubes during engine starting.
Turbo Annular Combustion Chamber
The turbo annular C.C is a combination of the multiple and annular types. A number of flame tubes are fitted inside a common air casing. The airflow is similar to that already described and this arrangement, embodies the case of overhaul and testing of the multiple system with the compactness of annular system.
Annular Combustion Chamber
This type of combustion chamber consists of a single flame tube, completely annular in form, which is contained in an inner and outer casing. The air flow through the flame tube is similar to that previously, described, the chamber being open at the front to the compressor and at the rear to the turbine nozzles.
The main advantage of annular chamber is that, for the same power output, the length of the chamber is only 75% of that of a turbo annular system for an engine of the same diameter, resulting in considerable saving of weight and production cost. Another advantage is that because inter connection are not required the propagation of combustion is improved.
In comparison with a turbo annular combustion system, the wall area of a comparable annular chamber is much less; consequently the amount of cooling air required to prevent the burning of the flame tube wall is less, by approx. 15%. This reduction in cooling air raises the combustion efficiently to virtually eliminate unburnt fuel, and oxidizes the carbon monoxides to non toxic carbon dioxide, thus reducing air pollution.
A high by pass ratio engine will also reduce air pollution since for a given thrust the engine burns less fuel.
TYPES OF COMBUSTION CHAMBER
Turbine Nozzle And Nozzle Diaphragm
This diaphragm consists of a group of nozzle vanes welded, between two shroud rings. In the typical nozzle diaphragm, the inner and outer bands contains punched holes to receive the ends of nozzle vanes. The nozzle vanes are usually constructed of high temperature alloy, and they must be highly heat resistant.
In many engines the nozzle vanes are hollow and are formed from stainless steel sheet. They are then welded and ground smooth before being installed between the shroud rings. When there is more than one turbine wheel, additional nozzle diaphragms are installed to direct the hot gases from one wheel to the next.
Second third and fourth stage nozzle vanes are often constructed of solid steel alloy. These may be either forged or precision cast.
The purpose of nozzle diaphragm is two fold :
It increases the velocity of the hot gases flowing past this point and
It directs the flow of gases to strike the turbine buckets at the desired angle.
The gases flowing through the nozzle diaphragm attain their highest velocity at this point. The blades or vanes in the nozzle diaphragm are of aero foil design. In the pure reaction type design, the nozzle blades resemble turbine blades closely. This is particularly true of the nozzle or stator blades (vanes) between the turbines wheels in multi stage turbine assemblies.
The turbine has the task of providing the power to drive the compressor and accessories and, in the case of engines which do not make use solely of jet for propulsion, of providing shaft power for a propeller or rotor. It does this by extracting energy from the hot gases released from the combustion system and expanding them to a lower pressure and temperature.
High stresses are involved in this process, and for efficient operation, the turbine blade tip may rotate at speeds up to 1300 feet per second. To produce the driving torque, the turbine may consists of several stages, each employing one row of stationary nozzle guide vanes and one row of moving blades.
The no. 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.
The number of shafts varies with the types of engine, high compression ratio engines usually have two shafts, driving high and low pressure compressors.
On high bypass ratio fan engines that feature an intermediate pressure system, another turbine is interposed between the high and low pressure turbine thus forming a triple spool system.
On some propeller or shaft engines, driving torque is derived from a free power turbine. The shaft driving the propeller or the output shaft to the rotor blade of a helicopter, through a reduction gear, may be mechanically independent of other turbine and compressor shafts.
The bypass 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, there by obtaining important thermal efficiency and power/weight ratio.
The design of the nozzle guide vanes and turbine blade passages is based broadly on aerodynamic considerations, and to obtain optimum efficiency compatible with compressor and combustion design, the nozzle guide vanes and turbine blades are of a basic aerofoil shape.
The relationship and juxtaposition of these shapes are such that the turbine functions partly under impulse and partly under reaction condition, that is to say, the turbine blades experience an impulse force caused by the initial impact of the gas on the blades and a reaction force resulting from the expansion and acceleration of the gas through the blade passages.
Normally gas turbine engines do not use either pure impulse or pure reaction turbine blades. With an impulse turbine, the total pressure drop across each stage occurs in the fixed nozzle guide vanes and the effect on the turbine, blades is one of momentum only ; where as with a reaction turbine, the total pressure drop occurs through the turbine blade passages.
The proportion of each principle incorporated in the design of a turbine is therefore largely dependent on the type of engine in which the turbine is to operate, but in general it is about 50% impulse and 50% reaction. Impulse type turbines are used for cartridges and air starters.
IMPULSE VS REACTION TURBINE
The basic components of the turbine are the combustion discharge nozzles, the nozzle guide vanes, the turbine discs and turbine blades. The rotating assembly is carried on roller bearings mounted to a compressor shaft or connected to it by a self aligning coupling.
Nozzle Guide Vanes
Are of aerofoil shape, the passages between adjacent vanes forming a convergent duct. The vanes are located in the turbine casing in a manner that allows for expansion.
The nozzle guide vanes are usually of hollow form and be cooled by passing compressor delivery air through them to reduce the effects of high thermal stresses and gas load.
The turbine disc is a machine forging with an integral shaft or with a flange on to which the shaft may be bolted. The disc also has around its perimeter provision for the attachment to the turbine blades.
To limit the effect of heat conduction from the turbine blades to the disc a flow of cooling air is passed across both sides of each disc.
The turbine blades are of an aerofoil shape. The main air is to provide passages between adjacent blades that gives a steady acceleration of the flow up to the throat where the area is smallest and the velocity reaches that required at exit to produce the required degree of reaction.
High efficiency demands thin trailing edges to the sections but a compromise has to be made so as to prevent the blades cracking due to temperature changes during engine starting and stopping.
The method of attaching the blades to the turbine disc is of considerable important, since the stress in the disc around the fixing or in the blade root has an important bearing on limiting rim speed.
Various methods of blade attachment are
Fir tree Root (with locking plate)
Fir tree Root (with shank seals)
De laval By it Root (with locking screw)
B.M.W.hollow blade (with retaining pins)
IMPULSE VS REACTION TURBINE
Now a days majority of gas turbine use fir tree root type attaching method.
To reduce the loss of efficiency due to gas leakage, across the blade tips, a shroud is often fitted this is formed by forging small segment at the tip of each blade, so that when all the blades are fitted to the disc, the segment form a peripheral ring around the blade tips.
On a fan engine, where the fan is aft (rear) mounted the blade forming the fan present an additional thermal problem. This is because the outer portion of each blade operation is in a duct through which passes a cool air stream, while the inner portion operates in the normal gas stream to extract the energy for accelerating the fan airflow.
The exhaust cone is located directly behind the turbine wheel and its main function is to collect discharge gases from the turbine wheel and expel then at the correct velocity. The exhaust cone consists of a stainless steel outer shell and central cone supported from the shell by four streamlined struts or fins is to straighten out the airflow from approx. 45% to an axial direction. Air flowing through this section decreases in velocity and increases in pressure.
The outer surface of the exhaust cone is insulated in most installations and many different types of insulation are used. A typical arrangement consists of four layers of A1 foil, each separated from the next by a layer of bronze screening.
The insulation reduces the heat losses that would normally escape through the exhaust cone. The insulation also protects adjacent aircraft structures and equipment from damage caused by heat.
When the gas-turbine engine is delivered by the manufacturer to the ultimate consumer, the exhaust cone is the terminating component of the basic engine. In order to operate the engine and obtain the required performance, however, it is necessary to use a tailpipe and exhaust nozzle.
The length of the exhaust pipe or tail pipe varies with each airplane installation, and therefore, by necessity, the pipe must be manufactured by the air-frame manufacturer.
IMPULSE VS REACTION TURBINE
Jet engines installed in jet airlines are equipped with thrust reversers to provide a braking action after the airplane has landed. The thrust reverser blocks gas flow to the rear and directs it forward to produce reverse thrust up to 5,000 lb. or more. The reverse thrust is produced when the air baffle doors or " clamshells " are moved into the gas stream by actuating cylinders controlled by the reverse thrust lever in the cockpit.
Early type of engines usually had the accessory drive sections located at the nose of the engine, where it had the effect of limiting the available area for air intake. This was particularly true on axial flow engines.
Present designs place the main accessory drive section at the bottom of the engine, and fitted closely against the case. The accessory drive is geared to the shift of the high pressure compressor.
Engine Driven Accessories
Includes the accessories needed for engine operation and also those required for the operation of a/c. The following is a list of accessories commonly driven by the engine and mounted on the accessory section.
Engine starter (connected to the engine during starting).
Emergency fuel pump
After burner fuel pump
Fuel control unit
Air bleed governor
Oil pump and scavenge pump
Some gas-turbine engines have one accessory power section, while others may have more. For example, one modern engine drives most of the accessories from a power takeoff at the bottom of the engine but also has an auxiliary accessory driven at the front of the engine. Two or three of the smaller accessories are driven from this front accessory section.
After burning (or reheat) is a method of augmenting the basic thrust of an engine to improve the aircraft-take off, climb and (for military aircraft) combat performance. The increased power could be obtained by the use of a larger engine, but as this would increase the weight, frontal area and specific fuel consumption, after burning provides the best method of thrust augmentation for short periods.
After burning consists of the introduction and burning of fuel between the engine turbine and the jet pipe propelling nozzle utilizing the unburned oxygen in the exhaust gas to support combustion. The resultant increase in the temperature of the exhaust gas gives an increased velocity of the jet leaving the propelling nozzle and therefore increases the engine thrust.
As the temperature of the after burning flame can be in excess of 1700 deg. C., the burners are usually arranged so that the flame is concentrated around the axis of the jet pipe. This allows a proportion of the turbine discharge gas to flow along the wall of the jet pipe and thus maintain the wall temperature at a safe value.
The area of the after burning jet pipe is large than a normal jet pipe would be for the same engine, to obtain a reduced velocity gas stream. To provide for operation under all conditions, an afterburning jet pipe is fitted with either a two position or a variable-area propelling nozzle. The nozzle is closed during non-afterburning operation, but when after burning is selected the gas temperature increases and the nozzle opens to give an exit area suitable for the resultant increase in the volume of the gas stream. This prevents any increase in pressure occurring that would affect the functioning of the engine and enables after burning to be used over a wide range of engine speeds.
The thrust of an after burning engine, without after burning in operation, is slightly less than that of a similar engine not fitted with after burning equipment; this is due to the added restrictions in the jet pipe. The overall weight of the power plant is also increased because of the heavier jet pipe and after burning equipment.
After burning is achieved on bypass engines by mixing the bypass and turbine streams before the afterburner fuel injection and stabilizer system is reached so that the combustion takes place in the mixed exhaust stream. An alternative method is to inject the fuel and stabilize the flame in the individual bypass and turbine streams, burning the available gassed up to a common exit temperature at the final nozzle. In this method, the fuel injection is scheduled separately to the individual streams and it is normal to provide some form of interconnection between the flame stabilizers in the hot and cold streams to assist the combustion processes in the cold bypass air.
Operation of Afterburning
The gas stream from the engine turbine enters the jet pipe at a velocity of 750 to 1200 feet per seconds, but flow is diffused before it enters the afterburner combustion zone, i.e. the flow velocity is reduced and the pressure is increased. However, as the speed of burning kerosene at normal mixture ratios is only a few feet per second., any fuel lit even in the diffused air stream would be blown away. A form of flame stabilizers is, therefore, located downstream of the fuel burners to provide a region in which turbulent eddies are formed to assist combustion in operation.
An atomized fuel spray is fed into the jet pipe through a number of burners, which are so arranged as to distribute the fuel evenly over the flame area. Combustion is then initiated by a catalytic igniter which creates a flame as a result of the chemical reaction of fuel/air mixture being sprayed on to a platinum-based element by an igniter plug adjacent to the burner, or by a hot streak of flame that originates in the engine combustion chamber; this latter method known as hot shot ignition. Once combustion is initiated the gas temperature increases and the expanding gases accelerated through the enlarged area propelling nozzle to provide the additional thrust.
In view of the high temperature of the gases entering the jet pipe from the turbine, it might be assumed that the mixture would ignite spontaneously. This is not so, for although cool flames form at temperatures up to 700 deg. C., combustion will not take place below 800 deg C. If , however, the conditions were such that spontaneous ignition could be effected at sea level, it is unlikely that it could be affected at altitude where the atmospheric pressure is low. The spark or flame that initiates combustion must be of such intensity that a light-up can be obtained at considerable altitudes.
For smooth functioning of the system, a stable flame that will burn steadily over a wide range of mixture strengths and gas flows is required. The mixture must also be easy to ignite under all conditions of flight, and combustion must be maintained with the minimum loss of pressure.
The increase in thrust due to after burning depends solely upon the ratio of the absolute jet pipe temperatures before and after extra fuel is burnt. For example, neglecting small losses due to the afterburner equipment and gas flow momentum changes, the thrust increase may be calculated as follows.
Assuming a gas temperature before after burning of 640 deg. C (913 deg K) and with after burning of 1269 deg. C (1542 deg. K) then the temperature ratio = (1542/913 ) = 1.69 The velocity of the jet stream increases as the square root of the temperature ratio. Therefore, the jet velocity 1.69 = 1.3. Thus, the jet stream velocity is increased by 30 percent and the increase in static thrust, in its instance, is also 30%.
Static thrust increases of up to 70% are obtainable from bypass engines fitted with after burning equipment, and at high forward speeds several times this amount of thrust boost can be obtained.
High thrust boosts can be achieved on by pass engines because of the large amount of exhaust oxygen in the gas stream and the low initial temperature of the exhaust gases. It is not possible, however, to go on increasing the amount of fuel that is burnt in the jet pipe so that all the available oxygen is used, because the jet pipe would not withstand the high temperatures that would be incurred.
Advantage and Disadvantage of Gas Turbine Engines
In its present forms, the gas turbine, either propeller-drive or pure jet, has numerous attractive features that have been mentioned frequently. These may be reviewed briefly as follows.
Advantages of Gas Turbine Engine
Freedom From Vibration
This permits lighter propeller sections and mounting structure. Vibration is reduced by the elimination of reciprocating parts such as connecting rods and pistons.
Simplicity Of Control
Only one lever is required for controlling the speed and power of the unit. Such surfaces add weight and drag. Very small coolers for lubricating oil are used on large jet engines and turboprop engines.
Negligible Cooling Air Required
Conventional engines require from five to eight times as much air for cooling as is required for power production. The acceleration of this air to airplane speed represents an appreciable loss of power, particularly in climb, even though much of this may be recovered later by the use of carefully designed radiators.
No Spark Plugs Required Except For Starting
After combustion is once established, it is self supporting.
Hence there is no carburettor icing and no mixture controls. There is some question as to this advantage, since large gas turbines do require very complex fuel-control units. The automatic features of these units compensate for their complexity, however.
Available Supply Of Compressed Air
This air is used for driving cabin superchargers and small turbines and for anti-icing purposes.
Decreased Fire Hazard
Fuels used for gas turbines are usually less volatile than the high-octain fuels used in reciprocating engines.
Lower Specific Weight
A gas turbine may develop several times as much power as a reciprocating engine of the same weight.
Lower Oil Consumption
Disadvantages of Gas Turbine Engine
High Specific Fuel Consumption at Low Airspeeds
This applies chiefly to pure jet engines. Turboprop engines have performance comparable to reciprocating engines in some instances, since they have attained specific fuel consumption as low as 0.40 lb per hp per hr.
Inefficient Operation At Low Power Levels
Slow Acceleration From Minimum Speed
This condition applies chiefly to turbojet engines. Turboprop and turbofan engines are able to accelerate quite rapidly.
High Starting Power Requirements
Starting large gas-turbine engines has been a problem in the past; however, starters have been developed within the past few years which make it relatively simple.
High Cost Of Manufacture
Although the gas turbine is much simpler in operation than the piston engine, the special materials and manufacturing processes needed make the cost of the gas turbine much higher.
Susceptibility To Damage By Foreign Material
Such material is rapidly drawn into the air inlet.
Basic Difference Between Gas Turbine and Reciprocating Engine
The basic differences between the gas-turbine engine and the reciprocating engine may be classified into five main groups.
Smaller nacelles possible; negligible cooling power required; high speed jet is a more efficient propulsive means than propeller at high flight speeds.
A high speed jet is a less efficient propulsive means than a propeller at lower flight speeds and during takeoff. The development of turboprop and turbofan engines has made it possible to combine the advantages of the turbine engines with the efficiency of the propeller for lower speeds and for takeoff.
Turbine engines are considerably lighter than reciprocating engines for the same power output. Turbojet engines have been developed with a weight/power ratio of less than 0.13 lb. per lb. of thrust. Turboprop engines have attained 0.39 lb. per equivalent shaft horsepower (eshp) in comparison with reciprocating engines which usually have a weight/power ratio of approximately 1.0 lb. per hp or more.
Fuel Consumption (Neutral Characteristics)
Best specific fuel consumption occurs near maximum output. (Best specific fuel consumption reciprocating engines occurs at about one-half maximum power.) At a given flight speed, specific fuel consumption of the turbojet engine tends to decrease with altitude. The specific fuel consumption for turboprop engines is comparable to that of the best reciprocating engines. It is likely that continued development will produce a turboprop engine with much better specific fuel consumption than any reciprocating engine. Turbojet engines have been developed to a point where specific fuel consumption is excellent, at operating speeds and altitudes, they are more efficient than of 1:2:1 and have brought about specific fuel-consumption figures of less than 0.70 lb. per pound of static thrust.
Best fuel consumption is in general is poorer for the turbojet engine; however this disadvantage appears to be decreasing rapidly.
Operation of the gas turbine engine at varying altitudes is somewhere between that of an unsupercharged and supercharged reciprocating engine. That is, the turbine engine is more adaptable to varying altitudes than the unsupercharged engine and perhaps a little less adaptable than the supercharged reciprocating engine.
Low power plant vibration, relatively constant speed over a wide range of output.
High engine speed (advantageous for generator drive)
A comparison of the performance of aircraft powered by reciprocating turboprop, and turbojet engines indicates that for a cruising altitude of 35,000 ft :
Aircraft with top speeds below 335 mph achieve their maximum range when powered by reciprocating engines.
Aircraft with top speeds above 610 mph achieve their maximum range when powered by turbojet engines.
Aircraft with speeds between those specified in the foregoing statements in general achieve their maximum range when powered by turboprop or turbofan engines.
Axial Flow V/S Centrifugal Flow
Within the classification of turbojet engines, some engines have advantages over other. At the present stage of development of these engines, the axial-flow type has the following advantages over the centrifugal-flow type:
Lower Specific Fuel Consumption
This advantage is quite important, since turbine powered aircraft are now designed to fly vast distances without refuelling. The lower fuel consumption is accomplished with the axial-flow engine because the axial compressor makes possible higher pressure ratio. This is particularly true of the turbofan engines.
Smaller Diameter Or Frontal Area
This characteristic makes the axial-flow engine more suitable for wing installation.
The following are some of the advantages of the centrifugal flow engine over the axial-flow type :
Simple Manufacture And Fewer Parts - This reduces initial cost and maintenance.
Lower Specific Weight - A centrifugal engine with an equivalent compression ratio may be lower in weight for the amount of thrust.
Faster Installation And Removal - No close fitting ducts to engine are necessary.
More Effective Water Injection - On this type of engine the water can be injected directly into the compressor, where as in an axial flow engine, water is injected into the C.C.
Faster Acceleration Of The Rotor Section - The advantages of higher compressor ratios possible with the axial flow engines will make this type of engine more desirable, especially for high performance a/c. Turbofan or by pass engines increase air mass flow by feeding additional air into the jet stream directly to the rear of the turbine. This results in thrust augmentation because a greater mass of air is accelerated than would be the case with the simple jet engine.