Components and Subsystems of Gas Turbine Engines. A Detailed Analysis

Academic Paper, 2021

113 Pages, Grade: A




Gas turbine engine - introduction

Centrifugal-flow compressor
Impeller (rotor)
Guide vanes
Diffuser vanes and casing
Axial-flow compressor
Multi-stage axial compressors
Stators and rotors
Engine pressure ratio
Pressure and temperature rise through the compressor
Airflow through the compressor
Stage power and work
Velocity triangles
Benefits vs drawbacks of axial compressors and centrifugal compressors

Gas turbine engine fan modules
High bypass turbofans
Low bypass turbofans
Turbofan fan blades - function and design
Turbofan fan disc - function and design
Turbofan fan casing - function and design

Turbine definition

Types of turbine
Impulse turbine
Reaction turbine
Impulse-reaction turbine

Basic turbine components

Exploring centrifugal and axial flow compressor thermo-fluid and aerodynamic principles
Boundary layer primary losses
Boundary layer secondary losses
Pressure loss coefficient

Axial-flow and centrifugal-flow compressors

Comparison of between impulse and reaction turbines

Compressor characteristics and performance
Multi-stage compressor stage matching
Overall pressure ratio against inlet mass flow and surge line

Compressor issues outside the operating limits
Compressor stall/surge
Blade stall and flutter
How compressor stall arises
Preventing stall and surge
Variations in absolute velocity
Velocity triangles for an axial-flow turbine stage
Fuel mass flow rate
Characteristic map tuning and efficiency contours
Turbine blades and nozzle guide vanes

Gas turbine disc and blade cooling

Nozzle guide vane (NGV) cooling methods

Safety improvement for compressors

Safety improvement for gas turbine fans

Safety improvement for turbines

Task 2

Aerodynamic performance of air intake sections

Intake types
Circular subsonic intake
Supersonic intakes
Axisymmetric supersonic intake
Rectangular supersonic intake (variable/fixed geometry)

Intake section design and performance
Intake section throat area in subsonic high bypass fan
Intake section throat area in supersonic high bypass fan

Gas turbine exhaust system
Exhaust nozzle
Thrust reversers
Convergent nozzle
Convergent-divergent nozzle

Thrust control and augmentation

Combustion systems

Combustor types and design
Multiple combustion chamber
Annular combustion chamber
Tubo-annular combustion chamber

Fuel nozzles

Fuel atomisation

Combustor performance
Diffuser performance
Combustion losses and efficiencies - performance criteria
Combustion efficiency
System pressure losses
Outlet temperature distribution
Combustion stability and light-up limits

Flame stabilisation - stability performance
Stability factors
Static stability
Dynamic stability
The reasons for instabilities occurring
Fuel types
Fuel-air ratio
Gas pressure, temperature and velocity variation
Airflow pressure variation
Airflow velocity variation
Airflow temperature variation
Flame-holder shape and size

Improving air intake performance

Improving combustion chamber performance

Improving exhaust performance




The thermodynamic and aerodynamic analyses of gas turbine engines is a very important topic in aerospace engineering. Engineers are constantly trying to achieve higher gas turbine efficiencies, by implementing various design configurations, experimenting with new materials, different material combinations, etc.

This assignment will cover some of the common and relevant criteria that is considered by engineers when designing gas turbine engines. The assignment will start off with thermodynamic analysis of gas turbine engines in task 1, by describing and explaining, in detail, the gas turbine engine components, namely, compressors and their types and components, engine fan modules and their components, turbines and their components, as well as fan casing.

Furthermore, system losses will also be covered, which covers the critical evaluation aspect of gas turbine performance, namely, boundary layer primary and secondary losses, pressure losses, etc. Subsequently, the gas turbine components will be examined and evaluated in detail, with different types of each component being evaluated, e.g., axial flow compressors and centrifugal flow compressors.

Moreover, comparisons between these components will also be covered, i.e., comparison between the compressor types, turbine types, etc. Some other system behaviour parameters that will be covered include compressor stall, blade stall and flutter, and relevant preventive measures against these phenomena will also be covered.

Additionally, velocity variations and velocity triangles will also be covered, as will the fuel mass flow rate, in addition to a few more mathematical analysis methods. Turbine and nozzle cooling methods will also be looked at in sufficient detail, as will the safety improvements for compressors, gas turbine fans, and turbines.

After that, i.e., in task 2, the aerodynamic analysis of gas turbine engines will be covered, namely, the air inlet types, i.e., for use in subsonic and supersonic flight regimes. The gas turbine exhaust system will also be examined, with further details entailing exhaust nozzles, thrust reversers, afterburners, and nozzle types, as well as thrust augmentation, and combustor design and their various types.

Furthermore, fuel atomisation and combustion performance and losses will also be covered, in addition to flame stability performance, fuel types, fuel-air ratio, and the variation of airflow velocity, pressure, and temperature, as well as flame-holders.

Finally, improvement methods for air inlets, combustion chambers, and exhaust nozzles, will also be explored in sufficient detail, and relevant images, diagrams, and graphical data will be provided where necessary to aid in the explanation of the gas turbine principles.

Gas turbine engine - introduction


Gas turbine engines incorporate a compressor section that serves the function of increasing the incoming air pressure prior to the airflow entering the combustor. The combustor section is important because it’s the main section, or component, of a gas turbine engine that directly influences the total engine performance. As shown in the diagram below, there are two types of compressors, namely centrifugal and axial compressors. (, 2015)

As a brief introduction, the centrifugal compressor causes the airflow to propagate perpendicularly to the axis of rotation, whereas the axial compressor causes the airflow to propagate parallel to the rotational axis. Centrifugal compressors aren’t utilised for large gas turbine engines anymore; they’ve been replaced by axial flow compressors. However, they’re still used on small turbojet and turboshaft engines, as well as pumps on rocket engines.

The main reason that axial flow compressors are used rather than centrifugal compressors is due to the fact that axial flow compressors achieve a higher pressure increase; centrifugal compressors increase pressure by a factor of 4, whereas axial flow compressors increase pressure by a factor of 4.3. (, 2015)

One of the primary purposes of an engine’s compressor section is supplying sufficient amounts of air to the combustion chambers so that the air-fuel mixture undergoes complete combustion; complete combustion produces carbon dioxide as a by-product, whereas incomplete combustion produces toxic carbon monoxide. The compressor section, in simple terms, increases the incoming air’s pressure, after which the air is discharged to the combustor section burners while meeting the exact pressure and quantity requirements.

The compressor section is also responsible for ensuring a constant supply of bleed air to be used by the engine for its numerous requirements. Compressors come in two types, as explained below.

Centrifugal-flow compressor

A typical centrifugal-flow compressor consists of a rotor and stator at the start of the compressor section, which are also known as impeller and diffuser, respectively. These two components fit onto the compressor manifold, which basically serves the function of holding the impeller and diffuser in place. The impeller and diffuser make up the compressor stages; the reason that there are two stages in a centrifugal-flow compressor is due to the fact that this configuration provides the ultimate efficiency for the compressor section.

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Figure 2 Compressor section with an impeller/rotor (left), diffuser/stator (centre), and compressor manifold (right)

Centrifugal compressors have a high pressure increase per stage, typically at an 8:1 ratio. Despite the fact that the impeller is a standalone component that is tightly bolted to the compressor manifold, the entire structure is often referred to as the diffuser. For explanation purposes, these are treated individually, i.e., as separate assemblies.

There are two kinds of impellers: single-entry and double-entry. The primary distinctions between the two varieties are their respective sizes and ducting arrangements. As for the single-entry impeller, its ducting configuration facilitates ducting of the air directly to the impeller eye, as opposed to the double-entry type, which has a smaller diameter but makes up for that by its operability at higher rotational speeds, which ensures sufficient airflow through the compressor section. (, n.d.)

Impeller (rotor)

The single-entry impeller receives airflow more efficiently, albeit at a compromise of a larger diameter in order for it to deliver a similar amount of air as the double-entry impeller. What this means in simple terms is that single-entry impellers require the overall engine diameter to be larger.

Double-entry compressor engines contain something known as a plenum chamber, which ensures that the incoming airflow enters the engine at angles close to 90° to the engine axis. Consequently, in order for the incoming air to induce a positive flow, it must encapsulate the engine compressor at a relatively positive pressure prior to entering the compressor. Some plenum chambers contain what’s known as blow-in doors, or more commonly known as air intake doors, which admit air to the engine compartment during ground operations to make up for the reduced air intake capability of the engine on ground. (, n.d.)

Guide vanes

As the image here shows, inlet guides are of two types: fixed and variable. They’re usually situated ahead of each impeller eye. The operating principle behind these guide vanes is that they efficiently guide the incoming airflow into the impeller eye. Most process compressors utilise fixed guide vanes due to the fact that vanes are typically prone to corrosion, as well as jamming. Air compressors typically utilise variable guide vanes on the first stage.

Shown in the diagram below is a depiction of adjustable inlet guide vanes, which alter the airflow’s angle into and through the impeller eye. They’re utilised for altering the slope of the compressor performance curve (shown below). For instance, the guide vanes which add pre-swirl counter to the impeller rotation will experience increased head and reduced curve gradient, whereas guide vanes which add the pre-swirl along with the rotation of the impeller will experience decreased head and increased curve gradient. (Stewart, 2019)

Adjustable inlet guide vanes

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Figure 4 Centrifugal compressor cross-section depicting adjustable guide vanes (left), and graph showing the effect of inlet guide vanes

Diffuser vanes and casing

The diffuser is an annular chamber with many vanes that form a series of passages diverging into the manifold. The diffuser vanes responsible for directing the air flow to the manifold from the impeller at a certain angle that maximises the amount of energy transmitted by the impeller. Diffuser vanes are also responsible for supplying air to the manifold at a pressure and velocity suitable for utilisation by the combustion chambers.

Figure 2 depicts the compressor manifold, which directs airflow from the diffuser through to the combustion chambers. Each chamber has its respective manifold outlet port, which divides the air equally among the chambers. A compressor outlet elbow is also bolted to each outlet terminal. The air outlets are made in the shape of ducts, and their role is to alter the orientation of the airflow from radial to axial, whereby the diffusion of air is completed after the turn. To improve the performance of this operation, turning vanes may be installed in the outlet elbows. Air pressure losses are reduced by these vanes, namely, by the smoother turning surface of the turning vanes. (, n.d.)

The diffuser has an increasing internal diameter for a decreased velocity and increased air static pressure.

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Figure 5 Diffuser diagram with bleed ports labelled (aircraftengineering,

The air then leaves the compressor then moves through the diffuser section. As the moving air flows through the diffuser, it’s being prepared for the combustion section at a low airflow velocity so as to enable proper mixing of the air with the fuel that’s being inlet into the combustion chamber.

The diffuser has bleed ports built into its structure, and it is through these ports that the discharge air from the compressors is bled off from the aircraft engine. Bleed air for the diffuser's operational functions is also drawn from additional ports situated between the low-pressure compressor and high-pressure compressors on aircraft engines with dual compressors. In the case of a high-pressure compressor, the bleed air may also be taken at intermediate stages. During such engine operating conditions, this air is bled from a way out of the main airflow on conventional engines with overboard venting to keep the 2011) compressor from surging. (aircraftengineering, 2011)

Axial-flow compressor

Axial-flow compressors typically consist of two main components: stator and rotor. The rotor consists of fixed blades on a rotating spindle, and these blades carry out the task of impelling the incoming air towards the rear in the same manner as a propeller, because of their angle and aerofoil contour. The rotor spins at high speeds and, at the compressor inlet, takes in air, which is then impelled through a series of stages. The air then flows along an axial path from the inlet all the way to the exit, after which it gets compressed.

The rotor motion serves to increase the air compression at each subsequent stage, accelerating the air backward over multiple stages. As a result of the increased velocity, energy is transferred to the air from the compressor in the form of kinetic energy. The stator blades act as diffusers at each stage, partially converting high velocity to pressure.

Each pressure stage consists of a pair of stator and rotor blades in a sequence of pairs. The volume of air and overall pressure increase needed determine the number of stages, i.e., the number of rows of blades. As the number of compression stages increases, the compressor pressure ratio also increases accordingly, with most gas turbine engines utilising up to 16 stages and more. (, n.d.)

Multi-stage axial compressors

A multi-stage axial compressor typically consists of numerous alternating rows of rotors and stators, as shown in the image below. The first row of stators is generally known as the inlet guide vanes, and each subsequent pair of rotors and stators is called a compression stage. Consequently, compressors with numerous blade rows are called multi-stage compressors.

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Figure 6 Typical multi-stage axial flow compressor (Rolls-Royce gas turbine engine) (, n.d.)

A good way to understand the functionality of compressors is by considering energy exchanges, which can be approximated by the Bernoulli equation, shown below.

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Pt is stagnation pressure, which is “a measure of the total energy that the airflow carries", and p is the static pressure, which is “a measure of the internal energy’. (, n.d.) The velocity terms are each a measure of the kinetic energy that’s correlated to their individual components; u, v and w, i.e., radial, tangential, and axial velocities, respectively.

The rotor adds swirl to the flow, thereby resulting in an increase in the total kinetic energy carried in the airflow due to the increasing angular momentum. The stator then proceeds to remove the swirl from the flow, but due to the fact that it’s stationary, it doesn’t add any net energy to the airflow, instead converting the kinetic energy from the swirl to internal energy, which raises the airflow static pressure. (, n.d.)

Stators and rotors

The stator is made up of a series of vanes that are connected to an enclosing case. The stator vanes are fixed and project radially towards the rotor axis, fitting snugly , ,r ... on either side of the rotor blades at This figure has been removed for copyright reasons. each point. In certain cases, the compressor casing, which houses the stator vanes, is split horizontally into two halves. When carrying out component inspection or maintenance procedures, either of the halves, i.e., the lower or upper half, may be removed, thus enabling access to the stator and rotor blades.

The stator vanes serve the function of receiving air from the air inlet duct, or alternatively, from each of the preceding stages, after which the stator vanes increase the air pressure, thereby delivering it to the subsequent stage at the correct pressure and velocity. Stator vanes also carry out the task of regulating the airflow direction to each rotor stage, for obtaining the maximum possible efficiency of the Figure 7 Stator and rotor elements of a typical axial-flow compressor (, n.d.) compressor blades. The image above shows the typical axial-flow compressor stator and rotor elements. Just like centrifugal-flow compressors, the first stage rotor blades of an axial flow compressor can be preceded by a fixed or variable inlet guide vane assembly.

The rotor blades are typically composed of stainless steel, with titanium sometimes used in the later stages as well. The attachment of the blades to the rotor disk rims differs based on compressor design, although they’re generally fitted into disks by “fir-tree” or “bulb-type” methods., as shown in the image below.

Consequently, the blades are locked into place by various methods. The blades differ in length from entry to discharge due to the annular operating space, which starts at the drum and ends at the casing, being progressively reduced by the reduced casing diameter as you go rearward. (, n.d.)

This figure has been removed for copyright reasons.

Figure 8 Typical compressor blade attachment to rotor disk design (, n.d.)

Engine pressure ratio

The engine pressure ratio (EPR) of a gas turbine engine is “the ratio of the turbine discharge pressure divided by the compressor inlet pressure”. (, 2017)

The engine pressure ratio is used for measuring the quantity of thrust generated by a gas turbine engine. Due to the fact that there’s a finite limit as to the total pressure that This figure has been removed for copyright reasons . a certain engine is designed to create, the EPR value is useful for providing the pilot with feedback as the thrust lever is deflected. Alternatively, the EPR can also be displayed on the Full Authority Digital Engine Control (FADEC) so as to make sure that the limitations imposed upon the engine aren’t exceeded.

The EPR value is determined by the aircraft’s pressure measuring probes which are Fi9ure 9 ECAM display with EPR indicator (, mounted in the engine inlet as well as at the 2017) turbine exhaust. A differential pressure transducer then receives this pressure data, after which the transducer indicates the ratio of the two pressures on a flight deck EPR gauge. The EPR system automatically compensates for the effects imposed by the aircraft’s airspeed and altitude. (, 2017)

Pressure and temperature rise through the compressor

Combustion of the fuel-air mixture at the typical atmospheric pressure wouldn’t generate the energy required for producing useful work. As a result, the energy that the combustion process releases is proportional to the mass and pressure of the air that’s burned in the combustion chamber. Consequently, higher pressures are required so as to achieve reasonable efficiency for the combustion cycle.

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Figure 10 Gas turbine engine compressor cross-section diagram

The incoming air is delivered to the compressor face from the air inlet duct in axial-flow compressors, after which it flows through the inlet guide vanes, as explained previously. As air approaches the multi-stage axial-flow compressor's first set of revolving blades, it is deflected in the direction of rotation, after which it is "arrested" and rotated as it passes over the stator vanes. This completes the airflow through one of the stages; after that, this process is repeated each time another stage is entered and the next set of rotating blades move the air along, thus making the air flow thorough the compressor’s multiple stages.

As a result, the air pressure rises every time the air moves through a set of stators and rotors. To achieve this result, the blades are made with aerofoil shapes, thus rendering the blades as lifting surfaces akin to propeller blades or wings. The cascade effect is a primary consideration in the determination of the angle of attack, aerofoil section, and the spacing from blade to blade. The blades must be able to withstand high aerodynamic loads and centrifugal stresses. (aircraftengineering, 2011)

Airflow through the compressor

At the compressor’s discharge end, the stator vanes are configured such that they straighten the airflow for eliminating turbulence from the airflow. These stator vanes are termed straightening vanes and are also known as the outlet vane assembly. Axial-flow compressor casings also support the stator vanes, as well as providing the means for the extraction of the compressor air so as to accomplish its various purposes.

These vanes are typically made of steel with resistance to erosion and corrosion, and they may also be encapsulated by a band of suitable materials to simplified fastening processes. The vanes are welded onto the encapsulation shrouds, with radial reinforcing screws securing the outer shroud to the compressor housing inner surface.

As mentioned previously, the drum-type compressor rotor differs in operating volume as you move rearward, which gives a reasonably constant velocity through the compressor, thereby maintaining constant airflow. The rotor contains either disc-type or drum-type construction, as shown below.

This figure has been removed for copyright reasons.

Figure 11 Drum-type compressor rotor (left), and disc-type compressor rotor (, n.d.)

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Drum-type rotors consist of flanged rings which fit against one another, whereby they’re capable of supporting the whole assembly by bolts. Low-speed compressors typically utilise this type of rotor construction due to the relatively low centrifugal stresses. On the other hand, disc-type rotors are made up of a series of aluminium forging-machined discs which are then shrunk over a steel shaft with the rotor blades dove-tailed into the disc rims. (, n.d.)

Stage power and work

The Euler turbine equation compares the power applied or extracted from a flow to the spinning blade row's characteristics. The Euler equation is based on the principles of energy conservation and angular momentum conservation. The image below shows an application of control volume around the compressor and turbine stages in a simplified model.

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Figure 12 Control volume for Euler turbine equation (, n.d.)

When conservation of linear momentum is applied, the torque T must be equal to the rate of change of time of angular momentum in a stream-tube flowing through the device.

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This relation is true regardless of whether the blade row is rotating or not. The torque and angular momentum values can be either positive or negative depending on direction. In the example model depicted above, the torque is positive when Vtan out > Vtan in, which is the same sense as the angular velocity. In case the blade row is in motion, then work is carried out, either by the fluid or on the fluid.

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In case angular velocity and torque have the same signs, work is being done on the fluid; this is the case in the compressor stage. The opposite is true for the turbine stage. Based on the steady-flow energy equation, the following expression is obtained:

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When the above expression is equated to the angular momentum expression, the following equation is reached:

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For a perfect gas with Cp — constant, this expression becomes:

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This is the Euler turbine equation, relating the temperature ratio, and in turn the pressure ratio, across a compressor or turbine to the velocity of rotation, as well as the variation of momentum per unit mass. In the compressor section, Trc > Tfb, and, as mentioned previously, work is done on the fluid. (, n.d.)

Velocity triangles

Velocity triangles are generally utilised for relating blade design parameters and flow properties in the relative frame - rotating along with the blades - to the absolute frame, also known as the stationary frame. The operating principle of axial-flow compressors is putting work into the incoming air through the processes of diffusion and acceleration.

Air then subsequently enters the rotor with an absolute velocity V and angle of attack ai, which is combined vertically with the blade’s tangential velocity U, which generates the resultant relative velocity Wi at an angle of «2. Airflow thorough the passages created by the rotor blades is given a relative velocity Wi at angle «4, which is smaller than a.2 due to each blade’s camber.

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Figure 13 Typical velocity triangles for an axial-flow compressor (Boyce, 2011)

Due to an increase in the passage width as the blades become thinner towards the trailing edges, Wi is less than Wi. Consequently, the air in the stage’s rotor section undergoes diffusion. The combination of blade velocity with the relative exit velocity of air results in an absolute velocity Vi at the rotor exit. After that, the air passes through the stator, whereby it’s turned at a certain angle so as to direct the air into the next stage rotor with minimal angle of incidence. The air entering the rotor section has an axial component at absolute velocity Vzi, as well as a tangential component Vei. (Boyce, 2011)

Benefits vs drawbacks of axial compressors and centrifugal compressors

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Table 1 Advantages and disadvantages of axial and centrifugal flow compressors (aircraftengineering, 2011)

Gas turbine engine fan modules

High bypass turbofans

In the transonic range, i.e., Mach 0.75 to 0.9, the most commonly used engine is the turbofan. The images below show examples of an ultrahigh bypass turbofan and a high bypass turbofan, respectively. The gas horsepower provided by the core is only partially extracted in a turbofan to power the propulsor, which normally consists of a single low pressure ratio shrouded turbo compression level.

The fan is normally positioned in front of the core inlet, allowing air to penetrate the core to flow through it first before being partly compressed by it. The majority of the airflow, on the other hand, bypasses the heart and goes straight to the exhaust nozzle. The central stream, which still has some gas horsepower, then goes straight to its own exhaust nozzle. (Baxter & Ehrich, 2015)

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Figure 14 Ultrahigh bypass turbofan engine with geared fan and thrust reversal variable-pitch blading

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Figure 15 High bypass turbofan engine with a 2-spool core and mixed-flow jet (Baxter & Ehrich, 2015)

The bypass ratio is a crucial factor for the classification of turbofans. It is defined as “the ratio of the mass flow rate of the bypass stream to the mass flow rate entering the core”. Even though the highest propulsion efficiency values are achieved by the engines which have the largest bypass ratios, the bypass ratio varies based on the type of engine, hence the different categories: medium bypass (BPR = 2 to 4), “high bypass engines” (BPR = 5 to 8), and “ultrahigh bypass engines” (BPR = 9 to 15 or higher).

In addition to these turbojet categories, there’s also low bypass turbojet engines. The first generation of turbojet engines, also known as zero bypass engines, have been largely replaced by low and medium bypass engines. The medium and high bypass turbofan engines replaced the low and medium bypass turbofan engines.

The higher the bypass ratio, the bigger the fan diameter, which means heavier components. At a certain threshold, it becomes more difficult to mount the engine onto an aircraft wing while maintaining adequate ground clearance. As well as the actual fan section itself, the bypass stream reversal apparatus is also required to be more complex and heavier, and this increases with the bypass ratio.

Ultrahigh bypass turbofan engines may incorporate a gearbox in-between the fan and drive turbine, which simplifies the design of small-diameter turbines with no compromise whatsoever to the large-diameter fan’s performance. In such ultrahigh bypass turbofans, variable-pitch fan blades are typically needed for thrust reversal, whereas in medium and high bypass turbofans, it’s possible to achieve a slight increase in the propulsive efficiency by mixing the hot core’s airstream with the cold bypass airstream prior to the total airstream entering a single jet nozzle. (Baxter & Ehrich, 2015)

Low bypass turbofans

In the subsequent aircraft flight velocity regime, i.e., Mach 1 - 3, it’s better to utilise turbojet engines (lack of bypass stream), or low bypass turbofan engines, with a BRP value up to 2. An example of a low bypass turbofan is depicted by the image below. Even though low bypass turbofans resemble higher bypass turbofans, some of their characteristics are unique only to low BPR turbofans.

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Figure 16 Low bypass turbofan engine with afterburner (Baxter & Ehrich, 2015)

For equivalent energy quantities available from the drive turbine, the lower overall airflow in the fan usually entails a higher fan pressure ratio, hence a fan with these characteristics is typically composed of 2 or 3 turbo-compressor stages, rather than one stage.

Engines designed for operating in the lower supersonic range typically provide inadequate thrust in other flight regimes where they're required to operate for short periods of time, such as when accelerating into the transonic speed range, or take-off from airports situated at high altitudes, at conditions such as high gross weight and enormously high temperatures, or during combat manoeuvres in the supersonic flight range.

For such applications, it’s more effective to add an afterburner for turbofans used in the supersonic range, rather than installing larger engines. Although the afterburner isn’t as efficient as a non-afterburning turbofan, it is relatively lightweight and simple in design. The afterburner typically requires the moderately cool bypass air to be mixed with the hot core stream, and for this reason, a mixer is incorporated into the afterburner.

When the afterburner is in operation, the exhaust nozzle of an afterburning engine should have a variable throat area for accommodating the significant fluctuations in volumetric flow rate between the afterburner’s incredibly high-heat exhaust stream and the colder airstream discharged from the engine when the afterburner is turned off. These engines typically have much lower compression pressure ratios compared to high bypass engines. (Baxter & Ehrich, 2015)

Turbofan fan blades -function and design

Figure 17 Rolls-Royce engine fan blades (, 2015)

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A turbine blade is the individual component comprising a gas turbine engine’s turbine section. Turbine fan blades, also known as turbine rotor blades, extract energy from the combustor section’s pressurised gas at high temperature. Additionally, fan blades carry out the function of accelerating the incoming air into the engine, which is the mechanism behind propulsion. Turbine blades also incorporate thermal barrier coating technology, which has resulted in improved oxidation and corrosion resistance.

Historically, fan blades used to be manufactured completely from titanium. Nowadays, the blades are made of hollow titanium, with web or honeycomb structures serving as internal stiffeners. This development was introduced by Rolls-Royce. On the other hand, General Electric utilised carbon fibre reinforced polymers (CFRP) for the GE90 blade design, which was later incorporated into the GEnx blade design.

Pratt & Whitney has developed an aluminium-based wide chord fan to be incorporated into their GTF engine series. These fan blades are manufactured with a hollow core and aluminium foam filler, as well as aluminium sheets bonded around the core. (Fehrm, 2016)

Turbofan fan disc-function and design

A fan disc is the central component of turbofan engine fan modules. The fan disc is the round piece where the fan blades are attached to the fan disk, which in turn is rotated by a turbine-driven shaft. Most of the thrust required for propelling the aircraft forward is generated by the fans. The fans are in turn driven by the engine’s gas turbine.

In order to reduce mechanical stress, fan discs are attached to a shaft that’s driven by a multi-stage Low Pressure Turbine (LPT). The incoming airstream flows through the engine’s frontal section, which is where the air pressure is increased by the fan rotating within a duct. As mentioned in the high bypass turbofan section, most of the pressurised air is then exhausted through the engine’s rear end, whereby the air expands, and its flow velocity increases. (, n.d.)

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Figure 18 Gas turbine engine fan rotor disc (bottom), and fan blade (top) (, n.d.)

A fan disc must be able to withstand very large centrifugal forces subjected to them by the attached fan blades during rotation. Due to their relatively large sizes, broken fan discs can result in severe damages to the aircraft structure.

Furthermore, the fan disc of a turboprop undergoes increased aerodynamic loading while the tips of the fan blades are moving at speeds higher than that of sound. To put it into perspective, a 15-pound fan blade can experience as much as 60 tons of centrifugal force. Fan discs are typically made of titanium alloys. (, n.d.)

Turbofan fan casing -function and design

A fan casing is simply the outer cylindrical piece which is installed around the inlet section. Fan casings must be strong in order to be able to withstand unexpected stresses in case of obstruction, blade failure or spin-off, etc. The optimal solution to this issue is to utilise braided composites in constructing and manufacturing the fan casings for turbofan engines. Not only does this bring the benefit of stronger material characteristics, but it also reduces the overall weight of turbofan engines.

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Figure 19 Gas turbine fan casing diagram (, 2006)

This fan casing incorporates tri-axial carbon braid as an alternative to aluminium and other similar metals and metal alloys. The braided composite casing has a higher toughness than aluminium fan casings, and coupled with its lighter weight, reduces the turbofan engines’ overall fuel consumption. (, 2006)

The fan casing serves the function of enclosing the fan section, as well as shielding the passengers from damages, as mentioned previously. The image below shows the tri-axial composite structure used in the fan casing; it’s not known which materials the composites are made of, since that is a General Electric trade secret. The tri-axial structure image shown below is the result of visual inspection done on the GEnx fan casing. (Fromm, 2009)

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Figure 20 Tri-axial composite structure of the GEnx fan casing

A crucial design consideration in fan casings is a blade-off scenario, as mentioned previously; therefore, the fan casing must be strong enough to endure the repercussions from blade damage.

The image below depicts the course of events in case of a blade-off. On the left is a rigid casing diagram, while on the right is a flexible casing diagram. The rigid case results in what’s known as the domino effect of blade failures since the broken blade will impede the movement of the other blades after it.

On the other hand, the flexible casing ensures that only the blade that’s already damaged faces the structural repercussions. Even though both cases aren’t desired in normal scenarios, the second scenario would still be somewhat more desirable compared to the first, whereby the flexible casing creates a buffer area for the damaged blade to be wedged into, so that the other blades can keep on moving with no impedance whatsoever. (Fromm, 2009)

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Figure 21 Undesirable blade damage, 2 possible scenarios (Fromm, 2009)

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Figure 22 Desired basing-blade interaction (Fromm, 2009)

On the other hand, the diagram on the left depicts a desirable scenario in the event of fan blade failure, whereby the fan casing is made of a honeycomb structure, and the damaged blade then penetrates the thin steel layer before entering the honeycomb structure, which ensures that the broken blade is fully clear of the remaining blades, while still regaining the structural integrity of the fan casing.

The honeycomb structure is surrounded with Kevlar for enhanced strength, toughness, and durability. The Kevlar layer carries out the function of absorbing the remainder of the rouge blade’s kinetic energy. Once this happens, the affected engine is immediately shut off and the aircraft is brought to the ground. (Fromm, 2009)

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Figure 23 Fan casing cross-section diagram (Fromm, 2009)

Turbine definition

A turbine in simple terms is a machine which transforms a fluid’s thermal and kinetic energy into mechanical work. The turbine is essentially a set of blades attached to a rotor acting as a shaft, which produces mechanical energy as it rotates, and this mechanical energy is then converted to electrical energy or used for generating thrust. In aircraft gas turbine engines, turbines are used to generate the thrust required for propelling the aircraft forward.

This figure has been removed for copyright reaons.

Figure 24 Illustrated diagram of a typical multi-stage turbine (Abhishek, 2018) (

Types of turbine

Impulse turbine

In impulse turbines, a quick-moving liquid is fired from a small nozzle at the turbine blades, causing them to rotate around. An impulse turbine’s blades are normally “bucket” shaped, trapping the fluid and then redirecting it at an angle of up to 180 degrees (reflection). In impulse turbines, the fluid hits the turbine at high speeds.

A simple example that shows the mechanisms of an impulse turbine is the waterwheel, which works on a similar principle. A Pelton waterwheel spins as one or more high-pressure water jets from the blue component, controlled and regulated by the (green) valve, fire into the buckets around the (red) wheel’s edge. The wheel then turns, and each subsequent bucket is pushed on by the high-pressure water stream.

Based on the law of conservation of energy, we know that each time the wheel is struck by the water stream, the energy gained by the wheel is equivalent to the energy lost by the water stream, hence the reflected water stream travels slower than the incident water stream. Moreover, based on Newton’s 2nd law of motion, we know that the momentum the wheel gains as it gets hit by the incident water stream is equivalent to the momentum lost by the water stream, hence the larger the impact from the incident water stream, the more momentum it transfers to the waterwheel. (Woodford, 2020) (

This figure has been removed for copyright reaons.

Figure 25 Pelton waterwheel (left), and impulse turbine (right) (Woodford, 2020) (

In an impulse turbine, the bucket-shaped turbine blades must be shaped in such a way that the action of one incident jet stream doesn’t affect the subsequent turbine blade. (Woodford, 2020) (, )

Reaction turbine

A reaction turbine’s blades are exposed to a much larger amount of fluid. The blades of a reaction turbine rotate along with the fluid flowing past them. Unlike impulse turbines, reaction turbines don’t alter the direction of the incident fluid stream in as drastic a manner as that of impulse turbines; rather, they simply rotate as they’re pushed past by the fluid.

A wind turbine is an example of a reaction turbine. The image on the left below shows a typical reaction turbine used in geothermal power plants. The fluid used as a propulsion medium for this is water or steam. When the turbine is operational, the water flows past the angled blades, which pushes against the blades and rotates the central shaft attached to the blades. In turn, the shaft drives a generator that produces electric power. On the other hand, an aircraft gas reaction turbine shown below on the right, is akin to a propeller in that it partakes in generating thrust.


Excerpt out of 113 pages


Components and Subsystems of Gas Turbine Engines. A Detailed Analysis
Aircraft Gas Turbine Engine Design and Performance
Catalog Number
ISBN (Book)
thermodynamics, aerospace, gas turbine engine, components
Quote paper
Abdusselam Šabić (Author), 2021, Components and Subsystems of Gas Turbine Engines. A Detailed Analysis, Munich, GRIN Verlag,


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