Review of advanced low-emission technologies for sustainable aviation

Energy 188 (2019) 115945

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Review of advanced low-emission technologies for sustainable aviation Kavindu Ranasinghe, Kai Guan, Alessandro Gardi, Roberto Sabatini* RMIT University, Melbourne, Victoria, 3001, Australia

a r t i c l e i n f o

a b s t r a c t

Article history: Received 22 February 2019 Received in revised form 8 August 2019 Accepted 13 August 2019 Available online 15 August 2019

Turbofan engines are the most widely used propulsion technology in commercial transport aircraft and are directly involved in many of the environmental impacts of aviation. Advancements in turbofan technology have thus a very significant potential in reducing aviation impacts on the environment. This article reviews the main technological advances currently being pursued in low-emission aircraft propulsion including: combustion and thermofluidic enhancements, gearbox technology, lightweight materials and intelligent engine health management systems. The paper discusses some key opportunities and challenges for these new technologies, with a particular outlook on the historical emission trends. Particularly, historical records from the ICAO aircraft engine emissions databank are used to extrapolate current trends and the progress against the ambitious targets set by international bodies. The analysis highlights that the sustained investments made by the aviation industry have yielded progressively diminishing returns and that the emission objectives will not be achieved at the current pace. Disruptive technological advances will therefore be required to significantly improve the fuel efficiency and mitigate environmental impact of commercial transport aircraft in the future. © 2019 Elsevier Ltd. All rights reserved.

Keywords: Turbofan Emission reduction Fuel efficiency Sustainable aviation Aircraft propulsion

1. Introduction The aviation sector is responsible for approximately 2e3% of Carbon Dioxide (CO2) emissions worldwide, and based on most current growth forecasts, this figure is expected to double by 2050 [1]. The predicted increase is primarily due to the fact that the demand for air transport is growing faster than advancements in technology and operational procedures towards the reduction of emissions can offset [1]. Taking this significant growth rate into account, the projection of the contribution of aviation to total CO2 emission will continually increase unless steps are taken to significantly reduce the environmental impacts of aviation. Moreover, other aircraft pollutant emissions, including Nitrogen Oxides (NOX), Unburned Hydrocarbons (HC), Carbon Monoxide (CO) and Sulphur Oxides (SOX) are also a concern as they are dispersed into the atmosphere at all levels, with some directly or indirectly prompting stratospheric ozone (O3) depletion, others generating harmful tropospheric O3 and others inducing noxious effects to the health of humans and of the ecosystem.

* Corresponding author. School of Engineering, RMIT University, Melbourne. E-mail address: [email protected] (R. Sabatini). https://doi.org/10.1016/j.energy.2019.115945 0360-5442/© 2019 Elsevier Ltd. All rights reserved.

Engines are the main source of these environmental emissions from aircraft, thereby attracting the attention of researchers worldwide to explore the possible improvements to current engine technology because they have a strong potential in moderating these adverse environmental impacts. At present, most airline transport aircraft are equipped with turbofan engines, which have been in widespread use for decades and have evolved considerably in terms of their efficiency and stability. In a conventional turbofan engine, the air inflow is divided after the fan into a core flow and a bypass flow. The bypass airflow is exploited to produce most of the thrust, whereas the core inlet airflow participates in an exergonic thermodynamic cycle involving compression, combustion and turbine expansion. Most of the currently operated turbofan engines feature two coaxial shafts, respectively connecting High Pressure (HP) compressor to HP turbine and Low Pressure (LP) compressor, turbine and fan. In recent years, many technological advances were proposed to improve the efficiency and environmental sustainability of turbofan engines. Due to the physics involved, some of these improvements resulted in trade-offs between the various emissions and other factors, therefore it is essential to formulate a multidisciplinary and multi-objective design optimization problem. Aerospace industry experts have predicted that by the year 2050,

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advancements in technology would lead to a 40e50% improvement in average fuel efficiency with fuel efficiency taking priority and a 30e40% improvement in fuel efficiency with NOX reduction taking priority respectively [2]. There are nonetheless several recognised challenges associated with every new technology, with various direct and indirect implications to the design, manufacturing, operation and maintenance phases. This paper reviews some of the most promising technological advances and outlines the main benefits and challenges identified in the literature to assess their overall viability to advance the turbofan engine technology. 1.1. Structure and layout of the article The article attempts to provide a comprehensive and broadscope review of recent progresses and trends in aircraft propulsion technology, focusing on some of the most distinctive trends. Where different approaches have been undertaken by major manufacturers, the article provides an assessment of the performance improvements and drawbacks for comparable technologies. The review starts with a discussion of turbofan emissions (Section 2) followed by the important subjects of increasing bypass ratio (Section 3) and the progresses in the use of composite materials (Section 4). Hence, the article discusses low-emission combustion technologies (Section 5) and casing treatments (Section 6). Successively, some of the relevant thermofluidic improvements (Section 7), combined cycle and interstage combustion concepts (Section 8), integrated health monitoring and engine management systems (Section 9) are discussed. Section 10 gives an overview of the current trends in terms of environmental impacts and Section 11 summarises the conclusions. 2. Turbofan emissions in aviation Turbofan engines generate a variety of gaseous and particulate material in their exhaust during normal operating conditions due to the combustion of fuel. The exhaust comprises about 70% carbon dioxide CO2 and about 30% water vapour (H2O) which are unwanted emissions but are unfortunately unavoidable provided that fossil fuels are used for propulsion [3]. However, the total quantity of exhaust emissions can be reduced by improving the overall efficiency of the aircraft, including improvements to the thermodynamic cycle of the engine as well as aerodynamic efficiency. The rest of the constituents comprise less than 1% of the exhaust and include nitrogen oxides (NOX), oxides of sulphur (SOX), carbon monoxide (CO), partially combusted or unburned hydrocarbons (HC), particulate matter (PM), and other trace compounds [4]. Nitrogen oxides (NOX) are unwanted by-products of combustion that are produced when nitrogen and oxygen combine together when air is subject to extremely high pressure and temperature conditions during the combustion process. Four main types mechanisms for the formation of nitrogen oxides have been identified in the literature [5], however thermal NOX was found to be the most dominant form. Sulphur oxides (SOX) are formed when small quantities of sulphur, inherently present in all petroleum fuel, combine with oxygen from the air during combustion. The formation of sulphur oxides also contributes to the secondary particulate matter (PM) formation. Hydrocarbons (HC) are formed as result of a poor fuel-to air (FAR) ratio and incomplete combustion. This occurs at low temperature and pressure regions and often leads to the formation of carbon monoxide (CO) due to the lack of oxygen required to complete the combustion reaction and form CO2. CO can also be produced in high temperature regions due to the dissociation of CO2 molecules, however this phenomenon is negligible as the CO is later burned out in the air-rich regions

downstream of the combustion chamber [3]. Particulate matter (PM) consist of finely distributed particles of soot that form as a result of incomplete combustion. The characteristics of the PM emissions are largely influenced by the type of engine, the thrust level (power) at which it is operated as well as the properties of the fuel itself [6]. 2.1. Environmental and health impacts of aircraft emissions Fig. 1 shows a representation of the breakdown of emissions at different flight stages during aircraft operation. The main contributors to the greenhouse effect, and thereby global warming, are CO2 and H2O (in the form of contrails) which traps infrared radiation from the surface of the earth and prevents heat from emitted into outer space. In particular, contrails and aircraft-induced cirrus clouds are amongst the largest contributors to aviation-induced global warming, possibly with a greater impact compared to CO2 for aircraft flying at high altitudes [8]. While aircraft emissions produced in the cruise phase cause pollution issues on a global scale, those that are produced in proximity to airport areas have a more direct effect on human health [9]. The products of incomplete combustion such as hydrocarbons (HC) and PM are the main contributors to urban smog and can be readily inhaled by humans, leading to increased health risks related to respiratory, cardiovascular and neurological systems [10]. Additionally, these particles remain in the atmosphere for extended periods of time, absorbing and reflecting sunlight. Environmental protection organisations are particularly concerned about NOX emissions from aircraft, mostly because they have a direct impact on the troposphere layer of the atmosphere and ozone depletion [11]. NOX emissions have an enhanced effect on the environment as they are released in concentrated regions near airports and at higher altitudes in the atmosphere. Tropospheric NOX emissions from aircraft trigger a chain reaction of chemical processes in the presence of ultraviolet (UV) light to form ozone (O3) in the troposphere. The formation of tropospheric O3 is associated with a positive radiative forcing, which is an indication of a forcing agent's influence on change in global mean surface temperature [12]. Additionally, at this proximity to the ground, ozone is a pollutant which is harmful to breathe for living beings and damages crops, trees and other vegetation. The production of tropospheric O3 is magnified due to the fact that this chemical process is self-sufficient as NO2 is recycled in the process, with each molecule of NO2 producing ozone multiple times. The only limiting factor is essentially the amount of volatile organic compounds (VOC), hydrocarbons (HC), and CO in the atmosphere. On the other hand, NOX emissions from aircraft in the stratosphere destroys O3 rather than producing it [1], leading to the depletion of the protective ozone layer naturally occurring in the stratosphere and thereby increasing risk of skin cancer, cataracts, and impaired immune systems in humans from overexposure to UV radiation. 2.2. Emission reduction targets Several organisations are implementing a systematic approach to define the emission reduction targets in both the short and long term and to track the progress achieved. The Advisory Council for Aviation Research and Innovation in Europe (ACARE), a publicprivate partnership between the European Commission and the aerospace sector, has set specific goals to address environmental sustainability of aviation, namely Vision 2020 and Flightpath 2050, to be achieved within 20 and 50 years relative to year 2000 technology respectively. The Vision 2020 and Flightpath 2050 goals relative to the environmental impact of aviation are recalled in the following table.

K. Ranasinghe et al. / Energy 188 (2019) 115945

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Fig. 1. Repartition by flight phases of the environmental impacts associated with aircraft operation [7].

Table 1 Emission reduction goals defined by ACARE [13]. Goal

Noise

CO2

 Halve perceived noise ¼ 10 dB per operation  65 LDEN at airport boundaries (no one impacted outside airport boundaries) Flightpath Reduce by 65% perceived noise ¼ 15 dB 2050 per operation

Vision 2020

Other emissions

Green life cycles

Alternative fuels

Progress in reducing the environmental None Decrease Decrease NOX by 80% CO2 by 50% (equivalent 60%/CAEP6) impact of the lifecycle of the aircraft

Decrease  Decrease NOX by 90% Air vehicles are recyclable CO2 by 75%  Europe at the forefront of atmospheric research  Emission-free taxiing

These goals help to create the vision for the scientific research and industry development programmes and ACARE continually provides support to the aviation research industry to guide it along a path towards an emission-free future. Similarly to ACARE, other aviation organisations and modernisation programmes worldwide such as ICAO, Clean Sky and NASA Environmentally Responsible Aviation (ERA) [14e16] have set ambitious environmental objectives to reverse the adverse effects of air transport on the environment. Most organisations are attempting the achievement of the environmental targets by taking a holistic approach, hence encompassing the design, manufacturing, operation and lifecycle management of aircraft.

and of the exhausts relative to the engine, m_ 0 is the mass flow rate of influx air, m_ e is the mass flow rate of exhaust, m_ c is the core inlet mass flow rate, m_ b is the bypass mass flow rate and finally m_ f is the fuel mass flow rate. Since bypass ratio (BPR) can be calculated as b ¼ m_ b =m_ c , Equation (1) can be rewritten as:

F ¼ ð1 þ bÞm_ c ðVe  V0 Þ þ m_ f Ve


F ¼ m_ e Ve  m_ 0 V0 ¼ m_ c þ m_ b þ m_ f Ve  ðm_ c þ m_ b Þ Ve ¼ ðm_ c þ m_ b Þ ðVe  V0 Þ þ m_ f Ve

(2)

Equation (3) gives the thrust contribution by the bypass flow, Fb , as a function of the fan diameter d and speed relative to ambient air, where r is air density and the velocity gain is DVbðVe  V0 Þ. The equation shows that at constant density, inflow speed and velocity gain the thrust generated by the fan increases quadratically with fan diameter.

3. Increasing engine bypass ratio The bypass flow of a turbofan engine is the main contributor to the total thrust, and at the same time, contributes to cooling the engine core. Assuming that exhausts are allowed to expand to ambient pressure, the thrust generated by a turbofan engine can be expressed as [17]:

Europe centre of excellence on sustainable alternative fuels

Fb ¼ m_ b ðVe  V0 Þzr Ab V0 DV ¼ r

pd2 4

V0 DV

(3)

Therefore, larger bypass airflows can greatly contribute to increasing thrust or alternatively to reduce the engine core sizes and required velocity gains at constant thrust. 3.1. Impacts of increasing fan size

(1)

where V0 and Ve are the velocities respectively of the ambient air

As evidenced by the equations, larger fans and fan inlets are beneficial to turbofan engine efficiency as they directly increase the airflow mass ratio and yield higher overall engine BPR. As BPR is

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related to the propulsive efficiency, increasing it has a direct impact on improving the thrust specific fuel consumption (TSFC), with added benefits in terms of noise reduction. Fig. 2 illustrates the historical trends pertaining to bypass and overall pressure ratios of aircraft engines. It is evident that there is an initial increase in BPR in the 1970s due to the introduction of high BPR turbofans. However, this is followed by a long period of time with no significant improvement in bypass ratio, which can be attributed to the several implications and engineering aspects that had to be overcome to increase the bypass ratio of a turbofan engine. On the other hand, the trend in OPR exhibits a continuous growth during this period owing to the improvements in compressor technology, materials being used and the introduction of multi-spool engines. In addition to geometric size limitations for ground clearance and transportation, larger fans and fan cases increase the total empty weight and cruise drag of the aircraft as both aerodynamic form factor and structural loads are worsened. Furthermore, more powerful LP Turbines (LPT) are required, further increasing engine weight. Another known issue relates to the tangential velocity of fan blade tips: the overall increase in the magnitude of the vector sum of the incoming flow velocity and the fan blade tip velocity can lead to degradations in aerodynamic efficiency and vibration characteristics of the fan. These, in turn, can amplify the oscillatory torsional loads on the shaft, degrading its fatigue life characteristics and requiring structural strengthening with its associated weight penalties. Fig. 3 shows the variation of weight, drag, noise, TSFC and fuel burn with increasing fan diameter for the CFM LEAP-1B turbofan engine used to power the Boeing 737 MAX [19]. From the figure, it is evident that increasing the fan diameter has both positive and negative impacts on the total fuel burn of a turbofan engine, and it is important for designers to take all these factors into consideration to find the optimum fan diameter to minimize fuel burn for a specific application. In terms of fan blade aerodynamics, the significant drag penalties incurred at high tangential speeds are correlated to the local Mach number M, defined as the ratio between the local tangential velocity as a function of the radial coordinate vðrÞ ¼ r ufan and the local speed of sound vsound , as:

MðrÞ ¼

r ufan vsound

(4)

The drag divergence Mach number is defined as the Mach number at which the aerodynamic drag affecting an aerofoil begins to increase significantly with respect to a small increase in Mach number. This is primarily due to the formation of shockwaves which induce flow separation and adverse pressure gradients over

Fig. 3. Optimization of fan diameter for the CFM LEAP-1B [19].

the aerofoil. An approximate method for estimating the transonic performance of aerofoils is using the Korn equation:

MDD þ

CL t þ ¼ kA 10 c

(5)

where MDD is the drag divergence Mach number, cL is the lift coefficient, t/c is the aerofoil thickness to chord ratio and kA is the aerofoil technology factor, with 0.87 assumed for conventional and 0.95 for supercritical aerofoils respectively. As shown in several studies, the value of MDD for conventional aerofoils is approximately 0.6e0.7 [20,21]. These velocities are easily exceeded when increasing the fan diameter, leading to severe degradations in aerodynamic efficiency. The most restrictive conditions are attained at altitude, as the speed of sound reduces with the square root of the temperature, which decreases linearly with altitude in the troposphere [22]. For instance, an average temperature of 56.5  C (216.65 K) and a reference speed of sound of 295 m/s encountered at 11,000 m are to be assumed as worst-case conditions. If the maximum tangential Mach number is limited to 0.6 for optimal drag, the blade tip velocity of the fan is vr  206:5m=s. Thus, assuming a fan diameter (D) of 2.66 m as an example e representative of the General Electric (GE) aviation engine GEnx-2B e the optimal angular velocity is u ¼ 155:26 rad=s, corresponding to about 1500 RPM. This value is substantially lower than optimal LP shaft rotation speeds. There are two main strategies being pursued to increasing the fan size and hence the BPR: the use of gearbox technology and the addition of extra compressor and turbine stages. Since the latter approach can easily lead to excessive weight and volumes, a viable strategy is to add an additional shaft. The recent increase in BPR of

Fig. 2. Historical trends of BPR and OPR in aircraft engines [18].

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aircraft engines from the 2000s onwards as shown in Fig. 2 can be attributed to such strategies.

3.2. Geared turbofan engines In conventional turbofans, the fan and LP Compressors (LPC) are driven by the same shaft and are thus rotating at the same angular velocity. Relatively high rotation speeds are required to maximise the efficiency of compressor stages, but since the LP shaft speed is limited by the restriction on the tangential tip velocity of fan blades, the efficiency of the LPC stages are significantly lower than the HP Compressor (HPC) stages. The adoption of a gearbox between fan and LP shaft has been proposed to solve the conflicting optimal rotation speeds of fan and LP compressor and turbine stages, increasing the efficiencies of both and allowing greater BPR to be achieved [23]. The increased efficiency potentially allows for the total number of compressor stages to be reduced, thereby also reducing overall engine weight. As an additional advantage, having the fan run at a low rotation speed reduces the aerodynamic noise generated by the engine and the resulting decrease in structural load and stress provides the opportunity to reduce the blade thickness and thereby reduce the total engine weight [24]. Although the use of gearboxes in turbofan engines was proposed as early as the 1970s, stability and fatigue lifecycle drawbacks discouraged their adoption. Furthermore, the addition of the gearbox has negative impacts on the manufactory and maintenance processes, with changes required to production lines and additional maintenance personnel training requirements [25], and with the introduction of new parts and changed sizes, the engine weight distribution is notably affected and this requires new airframe integration studies [26]. Despite these drawbacks, major aircraft engine manufactures such as Pratt and Whitney (P&W) have invested significant resources to improve fan drive gear system (FDGS) technology in order to produce cleaner, quieter, and more efficient high bypass ratio turbofan engines to their customers. Over recent years, significant improvements to FDGS design has enabled greater and more efficient power transmission, higher reliability and longer life-span of gear systems in turbofan engines. This section explores the innovative solutions brought forward by the P&W research program to overcome the drawbacks of traditional gear systems in turbofan engines.

3.2.1. Gear design The preliminary factor that determines the nature of the gear design is the gear ratio that matches the desired rotation speeds of the fan and the LPC. With increasing fan bypass ratio (BPR), the required gear reduction ratio also increases. P&W has patented the process of selecting this gear ratio [27] to cover all the ranges of possible gear ratios and how they interact with the turbine and fan. In their study, they found that typically, a BPR greater than 12:1 can be optimized to run at a fan tip speed below 350 ms1 and a pressure ratio less than 1.45. There are two different arrangements that can be utilized to achieve the desired reduction ratio: star systems and planet systems. These two epicyclic gear systems are illustrated in Fig. 4. In both configurations, the centre gear, called the sun gear, is driven by the input torque from the LP turbine shaft. The sun gear then drives an intermediate gear (star or planet), which then engages a ring gear. When the carrier is fixed in a star system, the star gears drive a ring gear which in turn drives the fan shaft, whereas the planet system utilizes a fixed ring gear and allows the carrier to rotate the fan shaft. The gear reduction ratio of epicyclic gear systems can be calculated using:

ur  uc Ns ¼  us  uc Nr

5

(6)

where ur, us and uc are the angular velocity of the ring, sun and carrier respectively, and Nr and Ns are the number of teeth of the ring and sun gear respectively. With the carrier fixed in star systems (uc ¼ 0 rad/s), the reduction ratio is equal to -Ns/Nr with Ns < Nr. Fixing the ring in planet systems (ur ¼ 0 rad/s) and using the planetary carrier as the output allows a higher gear reduction ratio of 1þNr/Ns. This difference makes the star system more suitable for applications requiring low gear reduction ratios and with counter rotation desired between the fan and the LP turbine. On the other hand, planet systems are better suited for high gear reduction ratios and when the fan is rotated in conjunction with the LP turbine.

3.2.2. Cooling and lubrication system for FDGS A major drawback of using a gear system in a large turbofan engine is the significant amount of additional heat generated by the FDGS. This puts further pressure on the thermal management system to prevent the temperature from reaching critical levels, with limited amounts of space available to fit extra heat exchangers and other cooling system components within the engine nacelle. In addition to this, the extremely high load capacity of a geared turbofan engine induces a considerable stress on the FDGS gears and bearings through continuous cyclic loading, leading to rapid wear and tear. Consequently, the maintenance and repair costs of the engine increase, with a reduction in reliability over time. To tackle the challenge of heat dissipation from the FDGS, P&W brought forward a unique baffle and gutter system, injecting lubricant at the sun/star mesh and carrying it through the gears and bearings until it exits the system via a static gutter that surrounds the gear system. It is important that the correct amount of oil is used for this process, as churning excess oil contributes to a large amount of heat loss and using insufficient oil would not provide adequate lubrication for the gears. The scavenge oil that enters the gutter is taken to an internal storage tank, and is then cooled and filtered so that it can be re-used in the cycle in a closed-loop system. Fig. 5 illustrates the schematic of the cooling and lubrication system of the FDGS in a typical geared turbofan engine. There are several other innovative features that allow this unique lubrication scheme to protect the gears and journal bearings under all types of circumstances and flight conditions, leading to near-infinite lifespan [28]. In their research program, P&W found that the leading factor for failure of gears and bearings was not due to under-sizing of FDGS components, but to the misalignment between the gears and bearings with the surrounding engine components, leading to additional loading and stress. In order to prevent this, the isolation of the gear system from the surrounding engine case was suggested. This involved introducing flexible couplings of the star gear mounted on a flexible support so that it can move freely under the driving action of the ring gear and the sun gear mounted on the input shaft from the LP turbine to compliment it. Details of the load isolation system can be found in Refs. [29,30].

3.3. Technological comparison of different industrial approaches to high BPR turbofan engines This section provides an outline of the different development pathways undertaken by the leading turbofan manufacturers globally. The section substantiates the state-of-the-art in the domain and serves as a baseline for subsequent sections.

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Fig. 4. Fundamental arrangements for epicyclic gear systems a) Star system b) Planet system (adapted from Ref. [27]).

Fig. 5. Unique FDGS lubrication system (adapted from Ref. [27]).

3.3.1. Pratt & Whitney 1000G PurePower series Beginning with the PW8000 in 1998, P&W have been the pioneers of developing geared turbofan engines. The PW8000 was targeted for the 25000e35000 lbf (110e160 kN) range and had a bypass ratio of 11:1. A decade afterwards, P&W decided to launch the PW1000G Pure Power engine product line, to bring cleaner, quieter and more efficient high bypass ratio turbofan engines to the market. Utilizing gearbox technology, the PW-1100G turbofan engine is able to operate with 3 LPC and 8 HPC stages, which is a significantly lower number of stages than that of other comparable engines, hence contributing to a sharp reduction of compressor weight [31] and achieving a high bypass ratio of about 12:1. This BPR is significantly larger than comparable non-geared turbofan engines, making it more fuel efficient [31]. In 2010, P&W announced the development of an ultra-high bypass variant, with a BPR significantly higher than the PW1100G's 12:1, to improve fuel consumption by 20% compared to a CFM56-7 and reduce noise relative to the FAA's Stage 4 by 25 dB [32].

The PW1000G series of engines, and especially the PW1100G family, is known for having high bypass ratios compared to other large commercial turbofan engines. The trend seems to suggest that P&W aim to push the limit of bypass ratio higher and higher in their future engines. The resulting low speed of the fans contributes to lower noise, yielding better passenger comfort and reduced flightrelated fatigue. The only downsides are related to technological immaturity as both the manufacturer and the operators have limited experience with large geared turbofan engines, casting initial issues for dispatch reliability. 3.3.2. Rolls-Royce Advance and ultrafan Rolls-Royce (RR) is a key partner in ACARE and has produced the Trent family of engines with continuous improvements in each successive variant to progress towards the emission reduction goals (Table 1). The Trent XWB is the last member of the family hitting the market and integrates a number of enhancements enabling it to comfortably meet current noise and combustion emissions

K. Ranasinghe et al. / Energy 188 (2019) 115945

requirements. However, this variant is only an intermediate step in the path towards achieving the mainline ACARE emission reduction goals. Based on continuous engagement with airframe manufacturers, RR has structured their development plans in two stages: Advance and Ultrafan, currently predicted for launch in 2020 and 2025 respectively [33]. The evolutionary pathway is outlined in Fig. 6. In particular, starting from the RB211, RR has distinctively adapted a three-shaft direct drive architecture for their large commercial turbofans, with three turbine stages to drive the fan, the Intermediate Pressure (IP) and HP compressors. This layout has allowed the Trent XWB to reach an overall pressure ratio (OPR) of 50:1, however studies have found that for even higher OPR's, a redistributed workload between the IP and HP shafts is more suitable, as it allows increased overall compressor and turbine system efficiencies, improved internal air system and lower cooling requirements, and reduced temperatures at principal internal structures and bearing chambers. The Advance concept capitalizes on these advantages to yield at least 20% better fuel burn and CO2 emissions than the Trent 772B [34], and deliver the highest overall pressure ratio of any commercial turbofan engine ever-made. Furthermore, the Advance core combines new component technologies such as hybrid ceramic bearings, dynamic sealing, smart adaptive cooling systems, lightweight high-efficiency compressors and turbines and an advanced low-NOX combustor to achieve higher temperatures and pressure ratios, yielding substantial improvements to thermal efficiency. The only major drawback is that the Technology Readiness Level (TRL) of these technologies is low, and they would require thorough validation and certification processes for them to mature and be brought forward to the market. The RR UltraFan, on the other hand, takes the evolution of Advance even further, featuring all the same technology and more to deliver a step-change increase in bypass ratio. One such addition is the introduction of a speed reduction power gearbox between the low-speed CarboneTitanium (CTi) composite fan and the intermediate pressure turbine to accommodate the ultra-high OPR of 70:1 and BPR of 15:1. Additional stages of IP turbine will be added, allowing to eliminate the LP turbine altogether and making the overall design smaller, lighter and more compact and therefore more efficient. RR also intend to take the CTi fan technology from the Advance one step further and make the fan blade pitch variable

7

in all phases of flight, thereby eliminating the need for a thrust reverser and enabling a slim design with minimal drag. 3.3.3. CFM international LEAP-X CFM International (CFM) is a 50-50 joint venture company between GE Aviation in America and Safran Aircraft Engines (previously known as Snecma) in France. Instead of opting for the geared turbofan or the three-shaft turbofan approaches to increase bypass ratio, CFM stuck to the conventional engine design of a twin spool engine, with a low speed LP shaft connecting the LP compressor stages and the fan and the HP compressor stages running at a higher speed, to produce the CFM56 series of high-bypass turbofan aircraft engines. In 2016, CFM brought forward the LEAP (“Leading Edge Aviation Propulsion”) as a new engine design to replace the CFM56 series, bringing with it 16% efficiency savings by using more composite materials and achieving a higher bypass ratio of 11:1 [35]. CFM's main goal with the LEAP engine was to simultaneously increase propulsive and thermal efficiency, by ensuring high compression ratios, high core temperatures and using a Twin Annular Pre-Swirl (TAPS) combustor to provide lean burn. This combustor is similar to a double-annular combustor in that it has two combustion zones, however the flow is swirled before it enters, creating an ideal fueleair mixture, and enabling the combustor to generate much less NOX emissions. Tests on a CFM56-7B engine demonstrated an improvement of 46% over single-annular combustors and 22% over double-annular combustors [36]. Advanced cooling techniques are employed to enable the turbine blades to withstand the extremely high temperatures, and a ceramic matrix composite (CMC) is utilized to provide additional thermal resistance for the blades. The main advantage the LEAP brings over its competitors is the proven reliability of the engine design, which a geared turbofan engine of comparable size is yet to demonstrate. The high TRL and reliability of this high-bypass ratio turbofan engine makes it attractive to many commercial aircraft manufacturers, such as the Airbus A320NEO family and the Boeing 737MAX family. Fig. 7 outlines some of the features of the CFM LEAP that help maintain its low fuel burn, high durability and low noise characteristics. 3.3.4. General Electric GEnx GE Aviation has been focusing on the development and use of

Fig. 6. Evolution of the Rolls-Royce Advance and UltraFan concepts (adapted from Ref. [33]).

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Fig. 7. CFM International LEAP component technologies (adapted from Ref. [37]).

Table 2 Key parameters of modern turbofan engines (Data obtained from respective engine datasheets). PW1100G

CFM LEAP

RR Trent 900 and successors

GEnx

Aircraft

Airbus A320neo

Airbus A320neo Boeing 737MAX

Boeing 787/747-8

Rated takeoff thrust [kN]

30/33G: 147.28 27G: 120.43 24/22G: 107.82

1A: 143.05 1B: 130.41

Bypass Ratio

12:1

Overall Pressure Ratio Fan Diameter [in]

38:1 81

1A: 11:1 1B: 9:1 40:1 1A: 78 1B: 69.4

Compressor stages (Fan-LP-HP) Turbine stages (HP-IP-LP)

1-3-8 2-0-3

1-3-10 1A: 2-0-7 1B: 2-0-5

Engine weight [kg]

2857.6

1A: 2990e3153 1B: 2780

Thrust-to-weight ratio

3.85e5.26

4.62e4.88

A380 (Trent 900) B787 (Trent 1000) A350 (Trent XWB) A330neo (Trent 7000) 900: 310e340 1000: 265.3e360.4 7000: 324 XWB: 430 1000/7000: 10:1 XWB: 9.6:1 (37e40):1 900: 116 1000/7000: 112 XWB: 118 1-8-6 XWB: 1-2-6 1000/7000: 1-1-6 900: 1-1-5 900: 6246 1000: 5936e6120 7000: 6445 XWB: 7550 5.13e6.01

advanced, lightweight and durable composite materials and specialized coatings in their jet engines, produced a dual rotor, axial flow, high-bypass turbofan jet engine named the GEnx (General Electric Next-generation). Derived from the GE90 family, it was introduced in 2007 for the Boeing 747-8 and 787 Dreamliner as a replacement for the CF6 in their product line. Carrying on the composite technology from the GE90 family, the GEnx employs 18 composite fan blades, a composite fan case and titanium aluminide stage 6 and 7 low-pressure turbine blades. The reduction of weight due to the utilization of lightweight composite materials yielded an improvement in specific fuel consumption of 15% comparative to the CF6 [38]. The GEnx is able to achieve a high BPR of 9.0:1 and a massive OPR of 58.1:1 thanks to a 10 stage HP compressor helped by larger, more efficient fan blades. 3.3.5. Summary Table 2 summarises the key characteristics of the latest turbofan

339

9:1 58.1:1 111.1

1-4-10 2-0-7

6147

5.62

engine variants hitting the market from the leading engine manufacturers. From the table, it is evident that in terms of achieving high bypass ratios amongst turbofan engines in current market, the adoption of gearbox technology has demonstrated the best results, with the PW1100g achieving the highest bypass ratio and the lowest number of total compressor stages. 4. Carbon fibre composites The fan blades of traditional turbofan engines are mostly made of Titanium or Aluminium alloys since these materials have good intensity and are lighter than other metallic materials. However, studies have shown that using composite material as alternatives could help reduce a considerable amount of weight while maintaining structural integrity. For the next generation of commercial engines, researchers are looking to composites made with carbon fibres coated with carbon nanotubes, called carbon fibre

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composites (CFRP). In particular, carbon fibre composite fan blades are 10% lighter than Titanium alloy structures of the same strength [39]. However, when considering the strength of composite parts, it is misleading to directly compare properties of carbon fibre versus aluminium or titanium, as these metals are homogeneous and have isotropic properties. By comparison, in a carbon fibre composite, the strength resides along the axis of the fibres, and thus fibre properties and orientation greatly impact mechanical properties. Table 3 shows the physical characteristics of a unidirectional lamella of 0 carbon fibres in ideal conditions. The table highlights the anisotropic behaviour of carbon-fibre composites as the strength in the 0 direction is significantly higher than 90 . Thus, carbon fibres need to be used compositely to withstand shear stress and other non-axial loads. In order to create a composite part, the carbon fibres, which have excellent stiffness properties in tension and compression, need a stable matrix to reside in and preserve their form. Epoxy resin is an excellent plastic with good compressive and shear properties, and is often used to form this matrix, whereby the carbon fibres provide the reinforcement. It is typical in engineering structural design to measure the benefit of a material in terms of strength to weight ratio and stiffness to weight ratio. Therefore, it is also important to consider the density of viable materials to understand their feasibility for different applications. Table 4 lists some representative densities of jet engine materials as reviewed by NASA. It can be noted that Aluminium is nearly half as dense as Titanium, whereas the density of carbon-fibre composites, which is around 2000 kg/m3, is even lower than Aluminium [41]. Using lightweight composite materials has a direct impact on the fuel efficiency of the engine as the overall weight would be substantially reduced. This also allows room for the fan diameter to be increased to enhance the bypass ratio and thereby further increase engine efficiency. Additionally, when a metallic fan blade fractures, the sharp metallic debris can easily damage other engine components, potentially leading to uncontained failures, whereas carbon-fibre composite blades do not have sharp edges when fractured and cannot withstand high temperatures, so their debris will be burned off when passing through the compressor and combustion chamber, thereby protecting engine core components from secondary damage. This allows the engine to meet blade-out test requirements without the manufacturing, assembly and weight of a separate containment ring. Moreover, because carbon-fibre composites have a better fatigue cycle, the total maintenance efforts could be reduced [42]. Assuming the General Electric Next-generation (GEnx) engine that employs a carbon-fibre composite fan as an example, the overall weight of carbon-fibre composite fan blades is 66% lower than Titanium ones but they are 100% stronger [43]. At the same time, because of the increased strength of carbon-fibre composites, the fan blades can be designed thinner and have a larger area than

Table 3 Typical mechanical properties of a flat carbon-fibre sheet [40]. Property

Value 3

Density (kg/m ) 0 Tensile strength (MPa) 0 Tensile modulus (GPa) 0 Compression strength (MPa) 0 Compression modulus (GPa) 90 Tensile strength (MPa) 90 Tensile modulus (GPa) 90 Compression strength (MPa) 90 Compression modulus (GPa)

1550 1400 123 850 100 18 8.3 96 8.4

9

Table 4 Representative densities of some commonly considered materials for jet engines. Material

T Lim [K]

Density [kg/m3]

Aluminium alloy Titanium alloy Stainless steel Nickel alloy Nickel crystal Ceramic

500 833 1111 1388 1666 1666

2726 4693 7633 8252 8252 2630

Titanium ones. Thus, the GEnx engine achieved a reduction in the count of fan blades from 22 to 18, further reducing engine weight. Furthermore, as the number of blades is reduced, the noise level is also reduced, allowing further increase to the diameter with no noise penalty [44]. However, due to their limited thermal resistance, they cannot be used to replace core engine components, and this characteristic introduces new risks to the engine. For example, in the event of accidental exposure to high temperatures as a result of an uncontrolled fire, the carbon-fibre composite fan and fan case may incur substantial damage. Carbon fibre composites cannot sustain high temperatures as the mass loss will significantly increase when temperatures are higher than 490  C for 3 min. Another major concern of using carbon fibre composites is the issue with their recyclability at their end-of-life stage. Most metals used for industrial and manufacturing purposes have a proven track record for recycling and minimizing waste, whereas the reclamation and repurposing of CFRP is a relatively novel concept, with further research required into understanding the processes and cost of closing the life-cycle loop of composite parts. 4.1. Evolution of composite materials in turbofans Although CFRP is a relatively new material technology with many refinements and upgrades as more research is carried out over the recent decades, the use of composite materials in the aviation industry dates back to the 1950s, when Rolls-Royce actually used glass fibre/epoxy in its RB108 engine compressor blades and casing. This was followed by a carbon fibre-reinforced epoxy called Hyfil in its RB211-22B turbofan blades in the 1970s. However, RR had to abandon the development of these composite blades, owing to the fact that they failed catastrophically during bird-strike tests [45]. At this stage, designers had limited understanding of composite materials and were therefore reluctant in using polymer composite materials in safety critical engine components. Over the next few decades, advanced research and analysis into the structure and failure mechanisms of composite materials at a micro-scale level was enabled by the development of techniques such as composite fractography and infrared spectroscopy. Consequently, the understanding of damage and failure modes of composites was greatly improved, and the design and manufacture of composites adapted to impede the causes of failure and maximise damage containment [45]. This paved the way for the reintroduction of CFRP in turbofan blades, beginning with the GE Aviation's GE90 in 1995, and even the use of CFRP in the front fan case as in the case of the GEnx turbofan in 2006. The driving factor behind the increased use of CFRP in turbofan engines was the desire by engine manufacturers to achieve higher BPRs for their engines, which resulted in better fuel efficiency and weight-savings. In addition to investigation techniques, there were also significant technological advancements with respect to the manufacturing methods used in the production of composite components [46]. For example, the injection of the epoxy resin into the polymer matrix and the curing process was combined into one

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operation called resin transfer moulding (RTM). Weaving methods that were previously carried out by manually were soon replaced by weaving machines able to produce 3D woven textile preforms featuring complex geometry with much more precision and accuracy at a much faster rate of production. Employing these efficient manufacturing methods, CFM International has managed to achieve an extremely high production rate of composite fan blades for the manufacture of their LEAP engines. After their unsuccessful attempt of introducing composite materials to turbofan blades in the 1970s, RR invested heavily on research into new blade design with composite materials incorporated. The aim was to deliver light fan blades with strong structural integrity while retaining aerodynamic performance. The outcome of their efforts was the development of the CTi blade, referring to the lightweight carbon composite structure and the aerodynamically efficient titanium leading edge, which was first tested in 2014 on a Trent 1000 “donor” engine with promising results. Currently, RR utilize composite materials in their Trent family of engines, with the Trent XWB having a CFRP fan case, fan track liners, bifurcation linings and anti-fluid panels. However, CTi fan blade technology is a major step forward for the design of the RR's next generation of turbofan engines, the RR Advance and Ultrafan, allowing the former to save up to 680 kg in total engine weight and achieve a 20% better fuel efficiency and reduction in carbon emissions when compared to the early engines in the Trent family [46]. The future of component manufacturing for turbofan engines seems promising with highly automated composites manufacturing robots being able output complex 3D geometries with high levels of accuracy and precision at a rapid production rate. This provides room for growth into the next stage of manufacturing technology, “adaptable manufacturing”, where the level of automation allows machines to react to changes in manufacturing processes without requiring human intervention.

5. Low emissions combustion technologies The design of low emissions combustion technologies is a complex multi-disciplinary optimization (MDO) design challenge because although the focus is on the reduction of emissions, there are several other factors and consequences to be considered that could affect other attributes of the aircraft. The fundamental challenge is to increase engine cycle efficiency, while keeping emissions at the lowest possible levels; however, it is important that characteristics such as performance, safety and operability are not compromised. Some of the design considerations taken into account when reviewing low emission combustion technologies in this section are given below:  Combustion Efficiency - how effectively the stored chemical energy in fuel is converted to useful heat energy  Combustion Stability - how vulnerable the process is to small perturbations altering the combustion characteristics  Pressure Drop - the magnitude of the pressure loss due to the increase in temperature  Smooth Ignition - the likelihood of controlled ignition being achieved without auto-ignition/flashback risks  Combustor Length/Weight - space and weight savings directly correlate to fuel savings  Temperature Distribution - the ratio of the difference between the average and peak temperatures in the combustion zone  Liner Life - the durability of the inner wall of the combustion chamber  Structural Integrity - the ability to support the required load during operating conditions with appropriate safety margin

5.1. Rich-burn quick-quench lean-burn combustion The first stage of conventional combustion is the “rich burn” stage, which involves a fuel-rich environment with stable combustion due to the high concentration of energetic hydrogen and hydrocarbon radical species at a relatively low temperature and concentration of oxygen. Therefore, the amount of nitrogen oxides formed at this stage is very low, but this produces a high concentration of combustible substances such as CO, the unburned hydrocarbons (HC), hydrogen (H2), and also carbon (soot). These substances cannot be freely released into the atmosphere; therefore, a quench section is employed downstream to dilute the mixture. In conventional combustors, there is a gradual and continuous admission of air to reduce the fuel-to-air ratio (f (FAR)), which results in the rapid formation of NOX, particularly close to when the reaction achieves its stoichiometric point, as shown in Fig. 8. Therefore, from the point of view of reducing NOX emissions, the combustion chamber of the engine should operate far from the stoichiometric point in either air enriched or depleted zones. The idea of quick-quench is to minimize the time taken to reduce the FAR, and rapidly input the jet air through holes in the walls of the flame tube to dilute the mixture and reach the lean burn zone, where the low FAR promotes the oxidation of the unburnt substances. This is the approach taken in Rich-burn, Quick-quench and Lean-burn (RQL) combustion, depicted in Fig. 9. Ideally, the exhaust gas leaving the combustion chamber should have an acceptable composition of the products of complete combustion (CO2 and H2O) along with N2 and O2. The main focus and the problem of such technology is to ensure fast and qualitative mixing of the gas stream in the intermediate stage (Quick-Mix) in order to prevent the formation of a mixture of stoichiometric composition and minimize NOX formation. If this process is not achieved effectively, it can lead to a sudden increase in flow temperature with undesirable consequences, both in terms of harmful emissions and damage to structural elements. Due to the critical nature of the quick quench process for RQL combustion, several measures can be taken to ensure its effectiveness [47]:  Closer spacing between primary and quench jets e increasing the proximity of the dilution holes to the primary rich burn zone where the primary air is injected boosts the strength of mixing

Fig. 8. Formation of nitrogen oxides and working principle of RQL (adapted from Ref. [3]).

K. Ranasinghe et al. / Energy 188 (2019) 115945

11

Fig. 9. Engine cross-section schematic illustrating RQL combustion.

and accelerates the process by which the combustion reaction moves from rich burn to the lean burn zone  Minimize time spent by the mixture in the quench zone e designing the combustor geometry to optimize the volume and distribution of combustor area to reduce the residence time of the mixture in the quench zone  Adding local cooling air e the NOx formation rate is highest at gas temperature peaks in the quench zone, therefore injecting air from additional cooling holes downstream of the primary zone leads to further reduction in NOX  Optimization of jet hole pattern spacing, sizing and geometry e these characteristics of the dilution holes play an important role in structure of the flow field distribution and therefore several research studies and experiments have been carried out to determine the optimum setup There is concern that the most widely used RQL technology also has limited potential to further significantly decrease NOX emissions whilst maintaining the other emissions criteria and satisfying all operability requirements, however it has been refined over the years by the world's largest engine manufacturers for their combustor designs. Pratt & Whitney developed the TALON (Technology for Advanced Low NOX) family of RQL combustors for commercial aircraft engines with reductions in NOX emissions with each progressive model, the latest being the TALON X. More detailed information on the design of the TALON family of combustors can be found in Ref. [48]. Other industrial approaches include GE's Low Emission Combustor (LEC) RQL technology which includes swirling devices to optimize fuel-air mixtures in the different regions, and RR's Phase 5 RQL combustor which was installed in the Trent family of engines. 5.2. Double Annular Combustor The idea behind using annular combustors is to eliminate the separation of the combustion zones and instead have a continuous liner and casing in a ring configuration. The simpler design and smaller size of this combustor greatly reduces the total weight and promotes more uniform combustion, leading to uniform gas exit temperatures. Double Annular Combustors (DAC) is a variant of this concept with two combustion zones around a ring: the pilot zone and the main zone, as shown in Fig. 10. The pilot zone (the outer annulus) acts like that of a single annular combustor (SAC), and typically operates at low power settings to raise the FAR to increase the combustion efficiency and promote complete combustion. At high power settings, the main zone (the inner annulus) is used as well, increasing air and mass flow through the combustor to achieve lean burn for minimal NOX and smoke formation. There is also an intermediate setting where part of the main zone is fuelled and ignited. The demand for different fuel supplies to different

Fig. 10. Double Annular Combustor cross-section [49].

combustion zones based on power setting demands a more complex fuel injection system. Similar to SAC, the radial arrangement enables the stable, efficient and low emission combustion to be achieved while shortening the length of the combustion chamber and thereby reducing the weight. The utilization of separate combustion zones at different power settings also allows more flexibility and control when compared to a SAC, with the potential to have lower NOX and smoke emissions at all power levels. However, the increase in surface area due to the addition of another annulus requires a larger amount of cooling, and with the injectors arranged in a radial configuration, insufficient cooling could lead to uneven distribution of temperature in “hotspots”. The radial temperature profile also has negative impacts on the durability of the turbine and its ability to extract work done, thereby reducing the fuel burn efficiency. Another major challenge arising from having the pilot and main stage in parallel with each other is to obtain the desired performance during the intermediate power settings when both the annular chambers are outside their optimum conditions. This provides significant challenges for designers to ensure that the benefits of an additional annulus outweigh the drawbacks in terms of fuel burn efficiency. CFM adopted the DAC design for their CFM56 series of engines to power the Airbus A320 and A321 aircraft. GE also extended this technology for large thrust category applications as is the case for the GE90 series which powers the Boeing 777 aircraft. Emission tests on these engines reveal that the margin between the NOX emission level and the CAEP/6 is generally wider than that of RQL combustors; however the margin tends to decrease with increasing OPR of the engine [50]. The drawbacks of DAC technology, in particular the cooling requirement, seem to limit the OPR that the engine can operate at with acceptable levels of NOX emissions.

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Therefore, significant improvements would have to be made to keep this technology viable in the future of engine designs with continually higher OPRs. 5.3. Axially staged combustors Similar to DAC, Axially Staged Combustors (ASC) have a pilot and a main zone of combustion with similar working principles, however instead of a radial configuration, the fuel injection zones are arranged in the axial direction with the main stage downstream of the pilot stage as shown in Fig. 11. With this configuration, the pilot zone continuously delivers higher upstream gas temperatures for more reliable and prompter ignition of fuel in the main stage, as well as more stable, efficient and complete combustion. Due to the arrangement of the stages in an axial manner, NOX emissions can be kept low because of the short residence time at the high temperatures experienced at high power settings, as shown in Fig. 12. The sharp drop in NOx formation when the power setting is increased can be attributed to the transition from part load to full load when the main stage comes online. On the other hand, the extremely high downstream temperatures have some negative side effects that have to be considered. The higher the temperature, the more susceptible the fuel becomes to the process of coking, in which a solid residue (coke) is created when fuel undergoes severe oxidative and thermal breakdown at extreme engine temperatures. The separation of the pilot and main fuel injection systems eliminates the cooling effect of the pilot fuel that is present in other two zone combustors, and also requires separated penetration of the combustor casing which calls for higher strength and stiffness of the casing material for structural integrity. Furthermore, engine manufacturers have to take extra precaution to ensure the in-line alignment of the pilot and main zone is perfect, otherwise the imperfect flow path would result in substantial pressure losses. Despite the concept of ASC arising in the same timeframe as DAC in the 1970s, the only major application of this technology is in the Pratt & Whitney IAE V2500 series of two-shaft high-bypass turbofan engines, primarily used to power the Airbus A320 family. Other engine manufacturers have opted to follow different routes when investing resources in the development of combustor technology, thereby the TRL of ASC is still fairly low comparative to the other types of combustor discussed in this paper. However, P&W recently carried out further development of this technology, renaming it to Axially Controlled Stoichiometry (ACS) under the NASA ERA program, and successfully demonstrated the lowest levels of NOX emissions tested (88% margin to CAEP/6 LTO NOX regulations for an advanced Nþ2 GTF cycle), shown in Fig. 13,

Fig. 11. Pratt & Whitney V2500-A5 ASC axial staged combustor cross-section [51].

Fig. 12. NOX emissions vs. power for conventional and axially staged combustors (adapted from Ref. [52]).

Fig. 13. Comparison of rig test results of the TALON X and ACS at NASA and UTRC for an advanced Nþ2 GTF cycle (adapted from Ref. [53]).

providing the most margin for development compared to the other concepts tested [53].

5.4. Lean direct injection (LDI) combustion In LDI combustors, fuel is injected directly into the combustion chamber without external premixing. The injector helps to produce a toroidal recirculating air flow with the fuel which can help the fuel and airflow to mix optimally. Since a large proportion of air is used in this process, the combustion takes place at lower overall FAR compared to conventional combustors, and the absence of a separate dilution zone means that the FAR is kept constant throughout the combustion chamber. An advantage of direct fuel injection into the combustion zone is that the risk of auto-ignition and flashback is eliminated, and provided that the fuel-air mixture is well mixed uniformly and rapidly to achieve constant lean burn, an extremely low level of NOX emissions can be achieved. On the other hand, constant lean burn leads to a lower maximum temperature during combustion reaction, which although significantly improves the liner life, has negative impacts on the stability and efficiency of combustion, leading the production of CO, HC and soot. RR has been investigating Lean-Direct-Injection (LDI) concepts of premixed and partially pre-vaporised lean combustion in the framework of the NEWAC (NEW Aero Engine Core) concepts, with the goal of achieving substantial NOX reduction for high OPR (>30) aero-engine applications. Test results show that the RR LDI combustor was able to achieve NOX reduction of 60% versus the

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CAEP/6 LTO NOX regulation for an engine with 39:1 OPR. Though the TRL for this technology is approaching 7 for emissions and operability, further LDI development will include further optimization in terms of weight and system complexity and better understanding of the required fuel staging functionality. 5.4.1. Multipoint injection integrated LDI The high dependency on the quality of mixing performance for LDI technology suggested the use of multiple injection points with several stages of LDI combustion. A study by Heath [54] can be adopted to assess the combustion efficiency of this technology. The study included a multiple LDI combustion consisting of 3 single LDI modules, the aspect of which is depicted in Fig. 14. In the experiments, the length of combustor was set as a constant and maintained for all combustions, however, 20 different designs were tested by varying certain parameters of the LDI modules. These parameters include the LDI module diameter X1 (varying between 1.810 cm and 3.016 cm), vane angle X2 (varying from 35.0 to 85.0 ), venture angle X3 (varying from 25.0 to 55.0 ) and number of vanes X4 (ranging from 4 to 8). The combustor air flow pressure drop (DP) is one of the key values that need to be determined to assess the combustion efficiency. To measure the effect of lean burn technology in reducing NOX emissions, the EINOx also needs to be considered. It is very important that the combustion has steady characteristics to make the engine and measured data reliable. Therefore, the nonuniformity U(x) has to be taken into consideration when assessing the results of this study. This can be calculated using Equation (7).

UðxÞ ¼

n 1X varðxi Þ n i¼1 mi

(7)

where x is the local temperature extracted from each computational solution and m is the average temperature. At the same time, the average Sauter mean diameter (D32) also needs to be calculated through Equation (8). The s is fuel surface tension (N/m), mL is fuel dynamic viscosity (Pa.s), mL is fuel injector mass flow rate (kg/s), PL is injector fuel flow pressure (Pa) and rA is injector air density (kg/ m3). Using these calculations,

D32 ¼ 2:25s0:25 m0:25 P 0:5 r0:25 m_ 0:25 L L L A

(8)

The exit temperature (T4) could also be used to assess the lean burn combustion non-uniformity. The target exit temperature

Fig. 14. Illustration of single LDI module (adapted from Ref. [55]).

13

when burning fuel is 1735.5K, however it was observed that some of the lean burn models could not reach this temperature, which suggests that an additional source of heat is required to increase the exit temperature up to the target value. The variation of the combustor air flow pressure drop DP with increasing vane angle in designs was also tested. It was found that DP varies linearly with the vane angle X2. When the vane angle decreases, the airflow travels through the combustion chamber in a shorter period of time and the direction of the airflow velocity is more aligned with the combustion chamber. Thus, the pressure drop will become lower which makes the power drop also lower. However, the models with 8 vanes exhibited a large pressure drop when compared to the models with 4 vanes. It can be deduced that increasing the number of vanes will also cause the pressure drop to increase at the same time. Therefore, the pressure drop is related to both the vane angle and the number of vanes. When considering the non-uniformity U(x) of every model, it was observed that most of the models outside the midrange vane angle (56 e72 ) have low non-uniformity, indicating that steady lean burn combustion could be achieved in these regions. Therefore, it can be inferred that the efficiency of lean burn in MLDI combustion is closely related to the vane angle and the number of vanes. The other two factors could affect the combustion efficiency, but do not have disciplinary characteristics that could be analysed. Therefore, the study only focuses on the vane angle and the number of vanes. High combustion uniformity supports steady conditions and thus efficiency and repeatability. At the same time, a lower pressure drop could also lead to better combustion efficiency. Lowering EINOx is the main advantage of using MLDI combustion. 5.5. Twin annular premixing swirler combustion Evolving primarily from other staged combustion concepts, Twin Annular Premixing Swirler (TAPS) technology has many resemblances to conventional SACs, with concentrically mounted pilot and main stages. However, a unique feature of TAPS combustors is the application of premixed combustion, achieved through a multi-swirler arrangement producing two co-annular swirling flow streams produced by the pilot and a main mixer respectively. The pilot uses a simplex atomizer surrounded by two co-rotating swirlers to spray fuel onto a pre-film lip where it is atomised to make it suitable for engine start-up and low-power operations. At higher power settings, the main mixer, consisting of a cyclone or radial inflow swirler, is utilized. A mixing layer is formed where the flow streams from the two mixers interact with each other, and this helps to stabilize the main flame and ensure clean combustion. Almost all the air entering the combustor is directed to the pilot and main swirlers, with the remaining air being used for cooling the combustor dome and liners. This can be attributed to the elimination of the need for dilution air during the latter stages of combustion, since the fuel-air mixing process is already significantly more thorough due to the TAPS mixers. This leads to better structural integrity of the combustor and shorter overall length due to the absence of dilution holes. A major advantage of having an internally staged configuration is that the required exit temperature of the TAPS combustor can be easily achieved and maintained, unlike that of conventional DACs. The improved distribution of exit temperature provides longer turbine life as well as lower fuel burn due to better thermal efficiency. The durability could be further improved through the use of advanced composite materials in the liner to resist higher temperature, and this is a critical parameter for the next generation of ultra-high OPR engines being developed. As with other premixed combustion techniques, there is always

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a risk of auto-ignition and flashback which has to be handled carefully through rapid and thorough mixing. Therefore the quality and effectiveness of the mixing is of paramount importance to the success of TAPS technology, and this requires detailed research into flow dynamics and heat release analysis using techniques such as computational fluid dynamics (CFD) tools and particle image velocimetry (PIV) [56]. GE has significantly invested in to developing TAPS technology for the turbofan engines, beginning with DAC versions in the late 1990s and transitioning into SAC TAPS combustors in the production of their CFM56 7-B engine. Based on the experience gained on fuel staging in these annular combustors, GE developed a TAPS combustor for the GEnx turbofan engine, which was able to achieve ultra-low NOX emissions (60% and 40% margins relative to CAEP/ 6 at 35 and 47 OPR respectively [47]) and good combustion efficiency. The success and potential for further emission reduction of TAPS has led to continuous improvement and scaling of the technology, termed TAPS II and TAPS III by GE, and has a very promising future in low emission combustion technology.

5.6. Lean premixed pre-vaporised combustors The underlying working principle of the Lean Premixed Prevaporised (LPP) concept is to supply a homogenous fuel-air mixture of low FAR into the combustion zone, and then carry out the combustion phase at an equivalence ratio very close to the lean blowout limit. The probability of NOX formation is significantly reduced due to the low flame temperature and absence of hot spots in the combustion zone. This also benefits liner durability due to low flame radiation. The schematic concept of this is shown in Fig. 15, with the LPP split into 3 major areas: the fuel preparation duct for premixing and pre-vaporising, the combustion zone and the dilution zone. Inside the premixer, the air passes through axial swirlers and is mixed with an evaporating liquid fuel spray from the fuel injection nozzles. These processes of mixing, fuel drop injection and evaporation have to be optimized to ensure the suitable conditions for extremely lean burn when the resulting mixture is supplied to the combustion zone. The dilution zone further reduces the FAR in a similar manner to conventional combustors. In order to optimize the preparation process, it is tempting to increase the residence time of the mixture in the fuel preparation duct, however there are other implication of this that need to be considered. In addition to a longer duct and thereby larger engine weight, increase in the residence time also makes the fuel-mixture more susceptible to auto-ignition. Flashback is another major concern of LPP combustors, which occurs when the flame speed is higher than magnitude of the velocity of the fuel-air mixture exiting the preparation duct. These risks are only amplified at

higher engine OPRs, which calls for a new technique or modification of this technology to deal with them for further application in future engines. 5.7. Flameless combustion Flameless combustion was developed to burners in heating industrial furnaces using preheated combustion air; however its ability to suppress the formation of thermal NOX even at extremely high temperatures makes it suitable to be adapted for aero-engine applications, especially with the demand for higher OPR engines. In conventional combustion, the combustible mixture of fuel and air is ignited to develop a flame. The stability and control of combustion is usually determined by the flame stabilisation, where a stable flame front provides a constant and controlled reaction and is also used as an indicator for flame supervision. However, this is not the only way of producing controlled combustion within a combustion chamber. The principle idea behind FC is the elimination of the flame front, which can be achieved in extremely high chamber temperatures (>850  C) with the help of preheated air. This type of reaction is only able to take place at temperatures above the self-ignition temperature, in a large dispersed volume as opposed to a restricted flame front. There are numerous advantages of this type of reaction, including very low levels of CO and NOX formation and the elimination of noise. The requirement of flame supervision is also eliminated as there is no danger of the reaction extinguishing, and thereby no risk of explosion. The flameless oxidation mode allows a large reduction of NOX emissions by avoiding localized peaks of temperature and significantly reducing the concentration of oxygen available for NOX formation. As shown in Fig. 16, the feeding of oxidising air and fuel gas is performed separately at high injection speeds. This creates large internal recirculation flows which, along with the high temperature of combustion products, initiates the flames mode of combustion and distributes the reaction throughout the volume of the combustion chamber. It is important to note the homogeneity in the temperature of the entire process, as opposed the existence of a peak in the temperature at the reaction zone of conventional combustion. The use of flameless combustion has been largely restricted to industrial applications such as power generation and steel and gas manufacturing industries. Unlike these ground-based applications, aero-engine combustors have additional restrictions in terms of volume, weight and high-pressure ratios, making it difficult to implement flameless combustion. Gas turbine engines also operate at very fuel-lean conditions in order to reduce their SFC, and the combustor liner and turbine blades have a limited resistance to extremely high gas temperatures, with limited space to add heat exchangers within the engine nacelle. This also means a large portion of oxygen remains unconsumed after combustion which makes it difficult to effectively lower the oxygen concentration, which is an important factor for the low NOX emissions. Therefore, there is certainly a gap in the knowledge of applying flameless combustion technology in an aero-engine scenario. 6. Casing treatments

Fig. 15. Schematic representation of LPP working principle (adapted from Ref. [47]).

One way to improve the propulsive efficiency of an aircraft is to increase the engines' thrust-to-weight ratio. Normally, commercial aircraft engines have about 4 stages of LP compressors and 9e10 stages of HP compressors which are responsible for a considerable portion of the engine's weight. Thus, increasing the pressure ratio of compressor stages can help reducing the number of the stages and engines' weight. However, increasing the pressure ratio of individual stages equates to an increase of the compressor blade

K. Ranasinghe et al. / Energy 188 (2019) 115945

15

Fig. 16. Comparison of the two combustion modes (adapted from Ref. [57]).

loading. A higher blade loading is more subject to stall which can compromise the engines' thrust or even lead to catastrophic engine failures [58]. Therefore, the stall margin is a crucial parameter to be observed when investigating evolutionary compressor technologies. One of the factors prompting a compressor stage to stall is the unsteady flow at the tips of the compressor blades. Because of the gaps between the blade tips and the compressor case, the airflow becomes highly unsteady in proximity of these gaps. Therefore, carefully designed casing treatments can help control the stability of the air flow at higher stage loadings and prevent the onset of stall [59]. Based on Boyce's analysis [60], Equation (9) could be derived, where a2 and a4 are the angles of inflow outflow of compressor blades. Based on this equation, pressure ratios could be increased by increasing the compressor blade angles. However, this also increases the stall point and thus the blade angle cannot be too large. Casing treatments could be used to potentially relieve this limitation and allow more room for increasing pressure ratio.

 g gþ1 P2 UVz ¼ ðtana2  tana4 Þ þ 1 P1 gc cp Tin

(9)

6.1. Stall precursor suppressed casing treatment The compressor stall precursor wave has been recently identified as a fundamentally important factor and can explain how compressor stall originates. Thus, increasing the stall margin translates into how to mitigate the stall precursor wave completely to avoid compressor stall [61]. The Stall Precursor Suppressed (SPS) casing treatment is one way that could help to mitigate the precursor wave. With respect to SPS casing treatment, the unsteady tip flow is drawn into an annular backchamber towards the front of the compressor stage rotor. Fig. 17 illustrates the mechanism of SPS casing treatment. A study on SPS casing treatment was carried out by Sun et al. (2014) [63] in which a low-speed TA36 fan equipped with controllable bleed valves, which can accurately change the operating conditions near and away from stall, was tested to explore the effect of SPS casing treatment on compressor stability. From the study, it was concluded that the SPS casing treatment helps to improve the compressor stall margin by around 10% by improving its stall point. The pressure efficiency is nearly the same with or without casing treatment, however, because the SPS casing treatment could improve the stall margin, it could indirectly improve the compressor efficiency. The compressor blade shape could be designed to sustain more loading without compromising stall margin, and the pressure ratio could be increased by using more

efficient compressor blades, thereby indirectly assisting in increasing the compressor efficiency. Furthermore, it could be deduced from the results of the study that with SPS casing treatment, the compressor could support more steady power even at the near stall point. However, because the SPS casing treatment adds extra cases on the compressor casing, the compressor needs to be fully redesigned. At the same time, adding extra case outside the core inlet may affect the fan inlet and cooling system. Thus, the structural integrity of the new compressor case needs to be tested to verify whether the new case is structurally reliable.

6.2. Recirculating casing treatment Tip leakage flow (TLF) is another important factor that is known to have negative impacts on the performance and operability of most modern axial flow compressors. Slot-type or groove casing treatments were suggested to provide compressor stability improvements by manipulating the TLF for tip-critical compressors, however there were drawbacks to these strategies in terms of loss in compressor efficiency [64]. A study by Guinet et al. (2014) [65] concluded that the benefits of using casing treatments were more prominent in compressors with higher tip clearance, but on the other hand larger tip clearance leads to the formation of a strong tip leakage vortex which has detrimental effects to compressor stability. In order to provide compressor stability without compromising for efficiency, a Recirculating Casing Treatment (RCT) was brought forward as a potential solution. RCT involves extracting highpressure air from locations downstream of the compressor and injecting it at the tip of the compressor blades. This tip injection process provides the range extension which is essential to negate the adverse effects of having a large tip clearance. Fig. 18 illustrates the method by which RCT can be implemented. A comprehensive study carried out by Wang et al. (2016) included both a numerical and experimental investigation into the effects of RCT in an axial flow compressor. The results including the variation of total pressure ratio and adiabatic efficiency with the mass flow rate (F) through the compressor are plotted in Fig. 19. From the results, it is evident that RCT has a positive impact on the stall margin of the compressor, with an improvement in performance and efficiency. The difference in the simulated and experimental pressure and adiabatic efficiency parameter at the near-stall point can be attributed to the non-uniformity of the rotor tip clearance and the incapability of the turbulence model; however, the mass flow rates at the simulated and true stall points are very similar, with both showing a 40% range extension in the stall margin for RCT. There are clear benefits in terms of stall range extension by

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Fig. 17. SPS casing treatment (adapted from Ref. [62]).

using RCT, with no predicted loss in efficiency or pressure increase capability at design speed conditions, however the accuracy of the design and placement of the injection and bleed ports is an important factor in ensuring proper recirculation of air, otherwise misalignments of the jets or insufficient fluid injection could have adverse effects on the compressor performance. 7. Interstage combustion and combined cycle technologies When considering fuel efficient technologies in engines, it is important to analyse the conversion of thermal energy into mechanical energy in its thermodynamic cycle. A simple Brayton-Joule cycle of a gas turbine engine is depicted in Fig. 20. The ideal thermodynamic cycle includes four stages correlating to their respective engine components: two isentropic (compressor and turbine) and two isobaric (combustion chamber and exhaust). The thermal efficiency hth of the cycle is related to the difference between the heat added (Q) in the combustion chamber (Stage 2e3) and the heat released in the exhaust (Stage 4-1) according to Equation (10).

hth ¼ Fig. 18. Mechanism of RCT (adapted from Ref. [64]).

Q23  Q41 Q23

Fig. 19. Characteristics of axial flow compressor for smooth casing (SC) and recirculating casing treatment (RCT) (adapted from Ref. [64]).

(10)

K. Ranasinghe et al. / Energy 188 (2019) 115945

17

Fig. 20. Ideal Brayton-Joule cycle for a gas turbine engine.

In simple terms, the area enclosed by the curves is an indicator of the thermal efficiency of the entire process, with a larger area correlating to a cycle that is able to achieve a better conversion of thermal energy to useful work done. Improving the efficiency of this conversion using thermodynamic optimization methods is key to improving the overall system performance and reducing pollutant emissions [66]. This section explores technologies that attempt to manipulate the thermodynamic cycle of aircraft gas turbine engines in the hopes of achieving better fuel efficiency with minimal loss in thrust and overall pressure ratio of the engine. 7.1. Interstage Turbine Burner (ITB) Afterburners were designed to energise the exhaust flow in specific flight stages to obtain extra thrust without increasing the size and weight of core engine components. However, this approach is highly inefficient as the afterburner is installed downstream of the HP and LP turbines, therefore the burning products are directly discharged through the nozzle into the atmosphere with very minimal thermal recovery. Conversely, some of the combustion takes place outside the engine nozzle and hence does not produce any propulsion for the aircraft. The afterburner technology could instead be positioned immediately after the HP turbine to exploit the thermal recovery offered by the LP turbine while still yielding significant extra thrust. Based on this principle, the Interstage Turbine Burner (ITB) has been conceived to be installed between HP turbine and LP turbines, yielding extra power for the LP shaft and fan. Thus, the engine could achieve extra thrust and use less fuel compared to afterburners. At the same time, because the ITB is positioned after HP turbine, the airflow temperature is reduced, thereby limiting NOX emissions [67]. In terms of thermodynamics, Fig. 21 illustrates the effect of introducing ITB in the thermodynamic cycle of a turbofan engine. To estimate the theoretical advantages offered by ITB, a preliminary estimation method are presented in this section. Equation (11) is used to calculate the uninstalled thrust-specific fuel consumption S where f0 is the air-fuel ratio, F is the uninstalled thrust and m_ 0 is the engine mass flow rate. The thermal efficiency hth is used to assess the fuel usage efficiency of the new technology and it could be calculated as per Equation (12). E_ kinetic; gain is the net rate of the kinetic energy gain, m_ f is the fan mass flow and hPR is the low

heating value of fuel. For detailed derivation of these equations below, the reader is referred to Ref. [67].

S ¼ f0 =ðF=m_ 0 Þ

hth ¼

E_ kinetic; gain m_ f $hPR

(11)

(12)

Nevertheless, ITB technology requires substantial design changes to the turbofan engine in relation to the additional fuel injectors between HPT and LPT. In particular, the design of fan and LP compressors and turbines is also significantly affected, and as a result the engine becomes more complex with operational and economic penalties. In particular, interstage burning increases the airflow temperature through the LPT. Furthermore, with fuel burning in close proximity or through LPT, substantial carbon deposition will likely happen on LPT blades, therefore, the maintenance work will be increased, and the carbon deposition may also affect the LPT blades’ lifetime. In addition, because ITB use the air which has already been burned in combustion, the residual available O2 is limited and non-homogeneous, and this can potentially cause severe incomplete/unstable combustion issues, producing more CO and HC. In addition, the ITB turbofan engine is less efficient when the interstage combustion is not engaged due to the structural redesigns introduced by the technology. 7.2. Intercooled and recuperated aero-engines (IRA) The concept of intercooling is aimed at decreasing the net work input of the compressor required for a given pressure ratio by cooling down the air temperature between the outlet of the LP compressor and the inlet of the HP compressor while maintaining constant pressure. This results in an effective increase in the net work output of the engine, however this is done at the expense of a reduction in thermal efficiency, as heat is removed from the process [68]. In order to compensate for this, intercooling must be used in conjunction with a reheating process. This can be easily achieved by introducing a heat exchanging system that utilizes the hot exhaust gas from the LP turbine to increase the temperature of the air entering the combustion chamber. The schematic of this combined process and the associated modification in thermodynamic cycle of

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Fig. 21. Thermodynamic cycles of a turbofan with ITB [67].

the engine is shown in Fig. 22. An additional benefit of this cycle is its significantly lower typical OPR, which supports the utilization of ultra-low NOX emission combustors such as DAC and LPP combustors, which otherwise have difficulty being implemented to high OPR engines. In terms of the aviation industry however, engine manufacturers are hesitant to incorporate intercooling and recuperation cycles in their engines mainly due to the added size and weight of the additional components required. Therefore, despite the potential of achieving better fuel consumption and lower emissions in aircraft, the use of this technology has been widely restricted to groundbased gas turbine power plants. In order to successfully implement this technology onto aircraft, an innovative design and method of integration which minimizes weight and size is required. Nevertheless, an EU-funded campaign EEAFAE (Efficient and Environmentally Friendly Aero Engine) have already invested heavily in the testing and validation of the IRA cycle with the help of the CLEAN (Component Validator for Environmentally-Friendly Aero-Engine) demonstrator, due to the potential improvements in fuel consumption and emission reduction that it could achieve. Based on the results of these tests that were completed in 2005, NEWAC (New Aero Engine Core Concepts) have continuously developed the concept further in more detail, taking into account optimum components to utilize, their arrangement and integration with the engine core. The “Intercooled Recuperative Core Concept”

is included in their list of subprojects that consist of core engine technologies with the best potential to achieve ACARE 2020 objectives. Significant advancements have been made in terms of the associated components technologies of IRA such as a profiledtubed recuperator (designed by MTU Aero Engines), centrifugal HP compressor and an advanced LPP combustor, and NEWAC's objective is to investigate and optimize these technologies and address their potential drawbacks in order raise their TRL to a level acceptable for the application of IRA in commercial aircraft.

8. Thermofluidic improvements The performance of gas turbine engine is dictated by the flow rate of air through it, thereby any restrictions or losses that inhibit the smooth flow of air will have a direct impact on the aerodynamic efficiency of the engine. Optimizing the path of the air flow through a gas turbine engine, in particular by addressing profile losses, tip clearance losses and end-wall losses, can contribute to a significant improvement in overall engine performance. Aerodynamic performance benefits can also be gained by optimizing the flow characteristics at the engine inlet with the use of morphing engine structures [70], and the engine outlet by manipulating the bypass nozzle after-body design [71]. Several studies have been performed to improve thermofluidic losses in different components of the engine; however the breadth and width of these studies is beyond

Fig. 22. Thermodynamic cycle of IRA (adapted from Ref. [69]).

K. Ranasinghe et al. / Energy 188 (2019) 115945

the scope of this review as it would require a dedicated article to address them. In this section, a top-level overview of some of the notable technologies in this area is presented. 8.1. Air tabs The majority of the noise pollution produced by aircraft can be attributed to the engines during take-off, and one of the main contributions is from the turbulent mixing of the hot core outflow with the colder bypass flow at the nozzle. An optimal mixing of these airflows can help reduce the fan case noise significantly, therefore researchers have been investigating the optimal mix these two airflows [72,73]. However, many of the traditional mixing solutions negatively affect the thrust due to the blockage effect, with the total thrust lost being about 10% with no external airflow and about 13e20% exploiting subsonic external airflow respectively [74]. The adoption of air tabs was proposed to help mix the airflow with minimum thrust losses. In a numerical simulation model study by Gu et al. (2014), six air tabs were placed on the splitter plate between the fan inlet and the core outlet to investigate the effectiveness of connecting the bypass and core outlet flows to allow optimal mixing. Fig. 23 illustrates the dimensions of the experimental set up. Since the airflow in the bypass outlet has a higher pressure compared to that of the core outlet, the colder flow passes through the tabs towards the core [73]. Preliminary investigations highlighted that the high temperature area of an engine with air tabs is of considerably smaller size than normal ones. In particular, the region where temperatures exceed 650 K is reduced from 29.2% to 5.2% when adding air tabs [73]. An additional benefit is obtained in terms of reduction in peak temperatures at the core outlet, protecting the nozzle materials from overheating and potentially enabling the adoption of lighter materials for the nozzle. According to the law of conservation of energy, a lower average temperature means more internal energy is recovered quasi-adiabatically into mechanical energy that results in additional thrust. On the other hand, since the flow through the air tabs produces vorticity, the losses associated with this additional vorticity have to be estimated. The non-dimensional streamwise vorticity stv can be calculated using:

stv ¼

D
z$v ve

jvj

(13)

where D is the shroud diameter at nozzle entrance, ve is the nozzle exit velocity, v is the velocity vector and z is the vorticity vector. Since the airflow at different positions has different velocities, stv will also vary accordingly. As ve is decided by the air-tabs position, the two variables of the size of air-tabs (d) and the air-flow ratio (u) were kept constant to aid in the analysis of the stv. It was observed that when the airflow goes further inside the nozzle, the airflow continues to collapse increasingly deeper. However, the vorticities which are created by the airflow going through the air-tabs become weaker as the airflow crosses the nozzle. The vorticities and the

19

collapsing effects were shown to increase with an increase in u. The increase in vorticity is more sensitive at the nozzle intake than the nozzle exit, but the collapsing effect shows the completely opposite sensitivity. The vorticities are similar with different size of air-tabs, but the collapsing effect becomes deeper with smaller air-tabs. To fully understand the benefits of using air-tabs, the thermal mixing efficiency hT also needs to be taken into consideration. Equation (14) shows how the mixing efficiency is calculated and Tmix is calculated through Equation (15).

ð

hT ¼ 1 

_ ðT  Tmix Þ2 dðmÞ

T 2core m_ core þ T 2core m_ fan þ T 2airtab m_ airtab  T 2mix m_ mix (14)

Tmix ¼

Tcore m_ core þ Tfan m_ fan þ Tairtab m_ airtab m_ core þ m_ fan þ m_ airtab

(15)

T is static temperature and m_ is the mass flow rate. Since both the flow ratio (u) and the size of air-tab (d) could influence the mixing flow and mixing temperature, the control variables methodology needed to be used to fully analyse the air-tab efficiency. From the results, it was evident that the presence of air-tabs is ineffective when x =D < 0:2; however, there was a sharp increase in efficiency afterwards. The mixing efficiency increased with the increase of u up to the value of u ¼ 0.76%, after which the mixing efficiency started to decrease. Thus, it could be proven that the collapsing ratio is not the only variable to measure the mixing efficiency. Similarly, it was evident that the mixing efficiency would increase with the increase of the air-tab tube inner diameter d up to the value of d ¼ 0.0036D, after which the mixing efficiency started to decrease. It is important to note that the mixing of the exhaust airflow lead to a loss in pressure, therefore, the total pressure recovery coefficient sairtab was taken into consideration to assess the feasibility of using air tabs. Equation (16) shows the calculation of sairtab, and it was found that the pressure recovery coefficient for all the models is over 0.98.

ð

sairtab ¼ ð

P mix dm_ mix

ð ð P core dm_ core þ P fan dm_ fan þ P airtab dm_ airtab

(16)

The best thermal mixing efficiency is attained with the lowest temperatures at the core air tab outlet. Thus, to get the lowest temperature, the best air tab size needs to be found. The air tabs mixing efficiency was modelled as a function of flow ratio and inner tube diameter respectively, and it was found that a higher u and d could achieve higher efficiency, except for the highest value tested in both cases. On the other hand, the air tab will produce vorticities which will lead to the pressure ratio decrease. Therefore, to get the largest benefits, the best design would balance the thermal mixing efficiency and pressure recovery gains with the

Fig. 23. Turbofan nozzle with air tabs (adapted from Ref. [73]).

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losses associated with the vorticity. Some potential drawbacks of air tabs technology include the weakening of the structural integrity of the nozzle due to the addition of holes, thereby causing it to suffer more structural and thermal loadings. Additionally, the air tabs need to be cleaned frequently to prevent blockage, thereby increasing the maintenance workload. No data is currently available to ascertain how often the air tabs need to be cleaned. Furthermore, air tabs may incur water deposition inside the hole, and if undetected, this can lead to corrosion inside air tabs, potentially compromising the structural integrity of the nozzle.

8.2. Cavity tips of turbine blades The turbine blades and casings are designed to feature a tip clearance to prevent rubbing. However, this small gap affects the stability of the airflow at blade tips. The unsteady airflow at the blade tips causes the high temperature and pressure upstream flow to leak downstream, inducing considerable turbulence and instability in the flow and reducing the turbine stage efficiency [75]. The problem becomes more serious for HP turbines because the airflow has a higher temperature and pressure, thus this leak and the associated turbulent phenomena become more severe. Advanced technologies that reduce the turbulence at the turbine blade tips can help to increase the engine efficiency significantly. The conventional strategy to reduce the turbulence is using flat tips. With flat tips, the airflow separates into two flows and one of the flows mixes with the airflow coming from the gap to reduce its turbulence [76]. A new advanced tip technology consists of using cavity tips which have has two squealers to help reduce the tip turbulence directly [77]. The new cavity tips yield lower thrust losses and are more efficient than flat tips [78,79]. Because of the internal cavity, the blades will have better temperature transfer characteristics, however to fully analyse the differences between cavity tips and flat tips, both stationary and moving endwalls need to be investigated for a more accurate assessment. The work relative to the endwall velocity that is used to measure the loss coefficient could be calculated with Equation (17) where the Fendwall is the resultant of the viscous forces and Vendwall is the velocity of the endwall. _ of the Because the Vendwall for the stationary endwall is 0, the W stationary endwall is also 0. Thus, the total loss coefficient Yp, total and profile loss Yp, profile could be calculated through Equations (18) and (19) respectively. The V is the average mass velocity and the subscript is means isentropic condition.

_ ¼F W endwall $Vendwall

Yp;total ¼

1m 2

Yp;profile ¼

(17)


_ V 2outlet;is  V 2outlet  W 1 mV 2 outlet;is 2

V 22profile;is  V 22

(19)

V 22outlet;is

Finally, the tip leakage loss Yp, equation (20).

Yp;tip ¼ Yp;total  Yp;profile

(18)

tip

tip, the moving endwall is 10.3% better than the stationary endwall. Thus, it could be proved that the cavity tip with moving endwall model has the lowest loss coefficient. Overall, the heat transfer coefficient is nearly the same across the various models. However, the cavity tips perform about 9.5% better than flat tips. Therefore, cavity tips are a better solution yielding lower thrust losses and better heat transfer. At the same time, a beneficial weight reduction can be obtained because the cavity tip is hollow inside. However, the structural integrity may be negatively impacted compared to the use of flat tip, therefore more studies need to be undertaken to analyse the design optimization of cavity tips, taking into account structure, reliability and manufacturing complexity. For example, in-depth investigations should determine the corrosion that can occur inside the cavity tip and outline new inspection methodologies to support line maintenance. 8.3. Microjets Efficient turbines help the engine produce more thrust and get higher pressure ratios. The low-pressure turbine (LPT) is used to power the fan and low pressure compressor; therefore it plays a significant role in providing maximum power during take-off and landing. However, since the LPT is operated in low Reynolds number regions, especially at high altitudes, the efficiency ratio is poor. When the Reynolds number is less than 100,000, the airflow at the surface of the turbine blades becomes laminar and will increase the turbine's losses [80]. As the turbine blades have a considerable amount of weight, reducing the number of blades could help to increase the engine efficiency and save fuel cost. At the same time, reducing the turbine blades count could help to make the maintenance work easier and increase the turbine life circle. However, to negate the loss in power due to fewer turbine blades, the turbine blade-to-blade space would be larger and the subsequent blade loading would increase. These changes will make the air flow at the LPT more unsteady and the turbine blades will fatigue more easily when exposed to high loading factors. Therefore, understanding how to control the laminar flow in low Reynolds number regions could help to reduce turbine losses and thereby increase the turbine efficiency. In order to achieve this, the concept of microjets was proposed. Microjets are small holes on the turbine blade ideally located in close proximity to the point at which the airfoil profile experiences non-reattaching separation of flow. They blow air perpendicular to the airfoil surface in order to control the flow over it, in an attempt to mitigate the flow separation leading to the loss of lift. Through the use of microjets, the non-reattaching separation of flow could be reduced to increase the turbine blades’ efficiency [81]. In a study carried out by Kumar et al. (2009) [82], the improvement of the flow field structure due to the effect of microjets placed at the 60% axial chord length (Cx) of the turbine blades was analysed. To quantify the improvement, the pressure coefficient Cp was calculated through Equation (21) where PT and PS are the total and static pressure respectively. The variation of pressure coefficient at different Reynolds numbers (Re) without the effect of microjets was calculated and used as a baseline for comparison.

could be calculated through

Cp ¼ (20)

In the study done by Zhou et al. [75] it could be seen that, the cavity tip models all have lower loss coefficient than the conventional flat tip models. In stationary models, the cavity tip has a 13% smaller loss coefficient than flat tips, and the loss coefficient reduction is 12% in moving cases. On the other hand, for the cavity

PT; inlet  PS; local PT; inlet  PS;inlet

(21)

The microjet velocity ratio B is the ratio between microjet exit velocity and local freestream velocity at 53% Cx as shown by Equation (22). From the study, it was evident that the variation in Cp is significantly altered with the use of microjets, and the magnitude of highest value of Cp depends on the value of B. The pressure coefficients increase compared to their corresponding baseline values

K. Ranasinghe et al. / Energy 188 (2019) 115945

with increase in B, however past B ¼ 4.5, a drop in Cp values is observed, with some locations being worse than the baseline model.

B¼

Ujet Ulocal

(22)

With microjet technology, the turbine efficiency could be increased which could help to increase the thrust of the fan and LP compressors. Microjets with B ¼ 4.5 show the maximum benefit at low Reynolds numbers and it could help to increase Cp by about 16% in total when Re z 20,000. When Re z 40,000, the B ¼ 12 and B ¼ 4 models could reach the highest Cp at about 60% Cx, with an improvement of about 10.1% compared with the baseline model. Therefore, it could be proved that the optimum value of B is between 4 and 4.5 to get the highest Cp in low Reynolds numbers. However, this new technology will require holes in the turbine blades and would therefore affect the structural integrity of the turbine blades. The fan blades may not sustain high loading as reliably as non-microjets blades, and they may be more prone to crack propagation. At the same time, the fatigue characteristics of turbine blades may also change, making them more susceptible to breaking off along the line of microjet holes. Furthermore, corrosion inside the microjet could potentially be a serious problem, because of the difficulty in detecting corrosion in tiny microjet holes as well as the removal of trapped water. These drawbacks of microjet technology may reduce the turbine fan blades lifetime and may increase the maintenance workload.

9. Integrated health monitoring and engine management systems The thrust control of modern turbofan engines is managed by the full authority digital engine control (FADEC) system on-board the aircraft. Essentially, the FADEC is responsible for interpreting the pilot power request input from the throttle setting or power level angle (PLA) and adjusting the fuel flow to the combustion chamber accordingly, and while doing so, ensuring that the engine is able to operate within safety limits. This process is important for the smooth and stable operation of the engines as uncontrolled adjustments to engine parameters could lead to compressor surge,

21

causing significant damage to the engine and potential loss of human life. Fig. 24 shows a simplified version of the control logic implemented by the FADEC to control engine thrust. Since the engine's thrust cannot be measured directly, the engine fan speed (N1) or the engine pressure ratio (EPR) can be used as the primarily control variable correlated to the thrust. Other variables that are measured by sensors in the engine include the core compressor spool rotation speed (N2), pressure and temperature measurements at the fan inlet, LP compressor and HP compressor exits, and an exhaust gas temperature (EGT) sensor that measures the gas temperature at the exit of the LP turbine. Despite providing guaranteed performance and safe operation throughout the engine operating life, current engine control architectures are unable to dynamically adapt to changes in the engine's operating environment, because their design is based on clean operating conditions [84]. The set of control gains used in traditional engine control logic is determined using aero-thermal mathematical models of the engine based on mass, energy and momentum conservation laws [85]. However, during operation, engines are exposed to several physical conditions such as erosion, corrosion and the build-up of dirt and carbon (soot) in sensitive areas. Consequently, important parameters such as flow capacity, effective nozzle areas and engine efficiency are affected negatively, thereby altering the performance characteristics of the engine [86]. The problems associated with this gradual deterioration become noticeable over a given period of time, with several pilots stating that they are required to manually make adjustments to the throttle in order to achieve the desired output setting, thereby increasing their workload and stress [87]. In order to implement dynamic adaptation of the control logic, a model-based control architecture which takes data from engine sensors to perform fault diagnostics and recalibration of control gains in real-time has been proposed. An engine monitoring unit (EMU) can be installed on the engine to assist the FADEC with thrust control. The associated modified control architecture is illustrated in Fig. 25. The on-board model of the engine adapts to the changes in engine condition and makes adjustments to the fan speed command signal accordingly. This not only alleviates workload from pilots, but also helps with the maximization of fuel efficiency, reduction of harmful emissions and increasing engine life. An

Fig. 24. Process diagram of simple engine control system (Adapted from Ref. [83]).

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Fig. 25. Block diagram of modified engine control loop (Adapted from Ref. [87]).

additional benefit includes better control and operability of the aircraft when flying outside design conditions, for example in an emergency scenario, as the model makes dynamic changes to the command inputs based on real-time data from sensors. However, in order to safely allow full autonomy in intelligent engine management, control and diagnostics systems, it is essential that the realtime model used to describe the behaviour of the turbofan engine has a high confidence level. Examples of potential models that can be utilized for this application include STORM (Self Tuning OnBoard Real-time Model) by P&W and ALPVM (Adaptive Linear Parameter Varying Model) [88]. Aside from real-time applications, engine health monitoring and management systems can also increase the safety and reliability of engine operation in the long term, by early detection of performance degradation, continuous tracking of engine health parameters and accurate prediction of remaining useful life (RUL) [89]. As of this point, the only factor involved in the control loop which modifies the operating conditions of the engine is the fuel flow; however, this technology could potentially pave the way to other factors such as variable pitch turbofan blades and variable reduction ratio gearbox technology being involved. The TRL of this technology is still relatively low and would require more development, testing and validation to reach the level of technical maturity required for commercial use by engine manufacturers. 10. Aircraft engine emissions trends Unsurprisingly, modern turbofan engines exploiting some of the emerging technologies reviewed in this paper are consistently more fuel efficient and produce fewer emissions. An analysis was carried out to extrapolate trends in the levels of thrust specific HC, CO and NOX emissions as well as the total fuel burn per unit thrust respectively for an LTO cycle of 584 aircraft engines over the past decades with certified data taken from the ICAO Aircraft Engine Emissions Databank Issue 25A (30th May 2018). The Emission Index (EI) of CO and HC (g/kg of fuel) were taken for the aircraft idle phase and the EI of NOX (g/kg of fuel) was taken for the aircraft take-off phase. The amount of fuel consumed in a reference ICAOdefined LTO cycle [90] was used as the most suitable indicator to obtain the thrust specific level of emissions for each engine. The calculation of fuel burned per unit thrust F (kg/kN) for each aircraft engine is shown in Equation (23).

!   FFtakeoff FF FLTO ¼ ttakeoff $ þ tclimb $ climb Ttakeoff Tclimb     FFdescent FFtaxi þ ttaxi $ þ tdescent $ Tdescent Ttaxi

(23)

where t is the time (sec), FF is the fuel flow (kg/s) and T is the thrust (kN) at each corresponding phase. The details of an ICAO-defined reference LTO cycle is given in Table 5. The results of this analysis are plotted in Figs. 26e29. Additionally, the variation of HC, CO, NOX and CO2 emission indexes as well as the TSFC of these engines with increasing thrust setting are plotted in Figs. 30e33 and are categorized by the decade of introduction to the aircraft engine market. The data points for each thrust setting were obtained from their corresponding flight phase as detailed in Table 5. A nonlinear symbolic regression method was used obtain the best fit trends of the datasets, using exponential models for HC and CO and 2nd order polynomial models for NOX, CO2 and TSFC. The fitting functions and the associated coefficients of determination R2 are detailed in Table 6. From Figs. 26 and 27, it is evident that the general trend of HC and CO emissions from aircraft engines has closely followed a negative exponential decay, with a few exceptions arising from 1985 to 1995. The LTO cycle TSFC depicted in Figs. 29 and 34 has experienced a steady decline since the 1970s; however, while apparently linear, the trend is more closely captured by a negative exponential fitting (higher R2), so the trend is flattening out progressively. Due to the significant correlation between fuel consumption and CO2 emissions, a similar trend is observed in CO2 emissions in Fig. 33. While economic factors (fuel consumption) rather than climate

Table 5 Reference Landing and Takeoff (LTO) cycle [90]. Phase

Thrust (T)

Time (t)

Takeoff Climb Descent Taxi

100% 85% 30% 7%

0.7 min 2.2 min 4.0 min 26 min

K. Ranasinghe et al. / Energy 188 (2019) 115945

23

Fig. 29. Variation of total fuel burn per unit thrust for a reference LTO cycle of aircraft turbofan engines. Fig. 26. Variation of thrust specific HC emissions for a reference LTO cycle of aircraft engines.

Fig. 27. Variation of thrust specific CO emissions for a reference LTO cycle of aircraft turbofan engines.

Fig. 30. Empirical fits of fuel-specific HC emissions as a function of the throttle of aircraft turbofan engines.

Fig. 28. Variation of thrust specific NOX emissions for a reference LTO cycle of aircraft turbofan engines.

concerns played and still play a prominent role in the decrease of CO2 emissions, reductions in CO and HC emissions were partly justified by the increase in efficiency and partly by the mitigation of air quality and health hazards posed by these emissions. However, these pollutant emission reductions were mostly afforded by ordinary technological advancements. Moreover, as Figs. 30 and 31 show, these emissions are still very significant in the most critical conditions (low temperature/throttle). Even more concerning is the historical evolution of specific NOX emissions shown in Figs. 28 and 32, which follows a completely different trend, with a notable

Fig. 31. Empirical fits of fuel-specific CO emissions as a function of the throttle of aircraft turbofan engines.

increase up to 1980s and a mostly steady level up to 2010s. This trend is mostly due to the increase in engine core temperatures to

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Fig. 32. Empirical fits of fuel-specific NOX emissions as a function of the throttle of aircraft turbofan engines.

Fig. 34. Empirical fits of TSFC as a function of the throttle of aircraft turbofan engines.

carbon emission reduction policies may lead to assume that declining trend in emissions could continue, however Figs. 26, 27 and 30, 31 very clearly indicate that the magnitude of reduction achieved decreased each decade despite sustained investments. In other words, technological advancements have yielded diminishing returns and will not significantly impact overall emissions in the future. This issue is well acknowledged in the scientific community and was also brought to the attention of policy makers because disruptive technologies which could afford a shift of gear will require policy incentives and substantial research funding. From this point on, changes to other limiting factors such as type of fuel, propulsion technology and aircraft configuration will be required to significantly improve the fuel efficiency and mitigate environmental impact of commercial transport aircraft. This puts a greater emphasis on research into areas such as alternative aviation fuels [91,92], intelligent engine control systems and novel propulsion methods such as hybrid-electric distributed propulsion, in the effort of producing ultra-efficient and environmental-friendly aircraft [93e95]. Fig. 33. Empirical fits of thrust-specific CO2 emissions as a function of the throttle of aircraft turbofan engines.

achieve better thermal efficiency, which counterbalanced the decreases in NOX formation afforded by higher bypass ratios and associated fuel efficiency. Because of this undesirable trend emerging both in aviation and in the road transport sector, the concern for NOX emissions has recently grown to widespread public attention and significant improvements are now mandated, particularly with the development of lean combustion methods. Advancements in low-emissions technologies e particularly including the ones reviewed in this paper e coupled with stricter

10.1. Hybrid-electric distributed propulsion As illustrated in the previous section, the advancements in aircraft engine systems to improve efficiency and reduce emissions are beginning to yield lower gains for greater effort, whilst more improvements are still required to achieve sustainability. The next step in the development of aircraft design for improved efficiency is considered to be the integration of the subsystems of the aircraft to gain synergistic benefits through positive interactions between these systems [96]. A proposed method of achieving this is using distributed propulsion, in which the thrust force of an aircraft is distributed about the airframe with multiple smaller engines as

Table 6 Fitting functions and the associated coefficients of determination of the empirical fits. Year

EICO

EIHC eð9:75tþ4:72Þ

EINOx eð15tþ4:13Þ

1970e1980

EICO ¼

1980e1990

EICO ¼ eð9:95tþ4:10Þ þ 0:769

EIHC ¼ eð13tþ2:40Þ þ 0:166

1990e2000

EICO ¼ eð11:1tþ3:97Þ þ 0:541

EIHC ¼ eð11:2tþ1:79Þ þ 0:082

2000e2010

EICO ¼ eð9:27tþ3:83Þ þ 0:380

EIHC ¼ eð17:2tþ1:54Þ þ 0:0224

2010þ

EICO ¼ eð10:4tþ3:71Þ þ 0:293

EIHC ¼ eð16:3tþ0:666Þ þ 0:031

þ 0:570

EIHC ¼

þ 0:447

TSFC ¼ :9996

TSFC ¼ :0319t2  :0409t þ :0226R2 ¼ :9232

EINOX ¼ 6:27t2 þ 19:5t þ 2:79R2 ¼ :9974

TSFC ¼ :0228t2  :0290t þ :0180R2 ¼ :9274

EINOX ¼ 5:18t2 þ 19:2t þ 3:66R2 ¼ :9910

TSFC ¼ :0194t2  :0245t þ :0156R2 ¼ :9142

EINOX ¼ 12:2t2 þ 10:9t þ 5:54R2 ¼ :9799

TSFC ¼ :0196t2  :0247t þ :0150R2 ¼ :9086

EINOX ¼ 17:2t2 þ 5:99t þ 5:58R2 ¼ :9602

TSFC ¼ :0179t2  :0224t þ :0133R2 ¼ :9064

EINOX ¼

13:4t2

þ 15:2t þ

2:03R2

K. Ranasinghe et al. / Energy 188 (2019) 115945

opposed to the conventional propulsion configuration of using large engines producing concentrated thrust vectors. In particular, this concept has a good synergy with emerging hybrid-electric technologies as an alternative to solely using internal combustion engines. Both concepts can combine and provide major benefits in terms of noise reduction, better reliability, shorter take-off and landing distances, better specific fuel consumption and improved flight stability [94]. Hybrid power systems have been proposed by many to bridge the gap between energy densities of electrochemical cells and liquid chemical fuels used in internal combustion engines. 11. Conclusions Turbofan engines are the most widely adopted aircraft propulsion technology at the moment and they are likely to maintain a dominant role at least for a couple of decades. Significant expectations are put on advanced turbofan technologies to revert the currently unsustainable impacts of aviation on the environment in spite of the steady growth of traffic. The advanced technologies reviewed in this paper improve turbofan engines in different ways, with some addressing weight reductions, pressure and bypass ratio increases and thrust and thermal efficiency increases. Many of the novel turbofan technologies discussed in this paper contribute to achieving this objective, however, significant challenges are still faced when considering a widespread adoption of these technologies. In addition to the abnormally rapid wearing recently encountered by several production models, the analysis of current trends in aircraft emissions points to the fact that recent advancements in turbofan engine technology have yielded diminishing returns despite sustained investment and that the reduction in emissions is progressively flattening. This calls for urgent disruptive technological advances such as turboelectric propulsion, alternative aviation fuels and innovative aerodynamic configurations. Due to the significant risk associated to these technologies, effective policy incentives and targeted research funding initiatives will be essential to spur a new generation of more sustainable aircraft technologies eventually meeting the emission goals set by international organisations. References [1] Penner E, Lister DH, Griggs DJ, Dokken DJ, McFarland M. Aviation and the global atmosphere - a special report of IPCC working groups I and III. 2000. [2] Baughcum SL, Begin JJ, Franco F, Greene DL, Lee DS, McLaren M-L, et al. Aircraft emissions: current inventories and future scenarios. 1999. [3] Wulff A, Hourmouziadis J. Technology review of aeroengine pollutant emissions. Aero Sci Technol 1997;1:557e72. [4] Aviation Administration Federal. Aviation emissions, impacts & mitigation: a primer. 2015. [5] Nicol D, Malte PC, Lai J, Marinov NN, Pratt DT, Corr RA. NOx sensitivities for gas turbine engines operated on lean-premixed combustion and conventional diffusion flames. 1992. https://doi.org/10.1115/92-GT-115. V003T06A012. € nenberger D, Siegerist F, Fischer A, et al. [6] Elser M, Brem BT, Durdina L, Scho Chemical composition and radiative properties of nascent particulate matter emitted by an aircraft turbofan burning conventional and alternative fuels. Atmos Chem Phys 2019;19:6809e20. https://doi.org/10.5194/acp-19-68092019. [7] Gardi A, Sabatini R, Ramasamy S. Multi-objective optimisation of aircraft flight trajectories in the ATM and avionics context. Prog Aerosp Sci 2016;83:1e36. https://doi.org/10.1016/j.paerosci.2015.11.006. [8] Lim Y, Gardi A, Sabatini R. Modelling and evaluation of aircraft contrails for 4dimensional trajectory optimisation, vol. 124; 2015. [9] Zaporozhets O, Synylo K. Improvements on aircraft engine emission and emission inventory asesessment inside the airport area. Energy 2017;140: 1350e7. https://doi.org/10.1016/j.energy.2017.07.178. [10] Froines JR. Ultrafine Particle Health Effects. URL: http://www.aqmd.gov/docs/ default-source/technology-research/ultrafine-particles-conference/preconference_2_froines.pdf?sfvrsn¼2. €hret Y, Kıncay O, Karakoç TH. An environment-friendly engine selection [11] S¸o methodology for aerial vehicles. Int J Green Energy 2018;15:145e50. https:// doi.org/10.1080/15435075.2017.1324788.

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