Civil Aero-engines Research Articles

Heat Load Development and Heat Map Sensitivity Analysis for Civil Aero-Engines

The design complexity of the new generation of civil aero-engines results in higher demands on engines’ components, higher component temperatures, higher heat generation, and, finally, critical thermal management issues. This paper will propose a methodological approach to creating physics-based models for heat loads developed by sources, as well as a systematic sensitivity analysis to identify the effects of design parameters on the thermal behavior of civil aero-engines. The ranges and levels of heat loads generated by heat sources (e.g., accessory gearbox, bearing, pumps, etc.) and the heat absorption capacity of heat sinks (e.g., engine fuel, oil, and air) are discussed systematically. The practical research challenges for thermal management system design and development for the new and next generation of turbofan engines will then be addressed through a sensitivity analysis of the heat load values as well as the heat sink flow rates. The potential solutions for thermal performance enhancements of propulsion systems will be proposed and discussed accordingly.

Influence of Hot Streak and Swirl Clocking Position on Aerothermal Performance of High-Pressure Turbine

In modern civil aeroengines, the hot streak and swirl at the exit of the combustor have a significant impact on the aerothermal performance of the high-pressure turbine (HPT). Due to the different design purposes of the combustor and the turbine, hot streak (HS) and swirl (SW) have different spatial distributions at the turbine inlet. This paper conducts a transient simulation of the GE E3 first-stage HPT, considering the swirl and hot streak facing the middle of the passage and the leading edge of the nozzle guide vane, respectively, and also explores the impact of positive and negative swirl. The results show that different clocking positions and swirl directions will change the incident angle and streamline distribution of the vane, thereby affecting the migration of the hot streak, the temperature and the Nusselt number distribution on the stator surface. In positive cases, the hot streak gathers in the upper part of the passage, and in negative cases, it is in the lower part. In middle cases, high-temperature areas appear in both vanes, and the distributions are opposite. Affected by the swirl, when facing the passage center, the pressure side stagnation lines of the two vanes are also different, so the Nusselt number distribution is opposite. When facing the leading edge, only one vane appears. Due to the insensitive interference of the rotor–stator, the transient migration of the hot streak in the rotor is mainly affected by the inherent secondary flow and the temperature at the inlet of the rotor (especially the conditions facing the leading edge), while the upstream residual swirl is less affected. Unlike the middle case, in leading edge cases, the hot streak is separated and needs to be re-mixed before entering the blade passage, so the temperature change in the blade cascade is relatively gentle. Based on this, the Nusselt number distribution on the surface of the blade is similar. In order to obtain the most favorable operating conditions for the engine, the turbine efficiency is used to compare the aerothermal performance under different conditions. Ultimately, it was found that the turbine with the hot streak and positive swirl directly facing the leading edge was the most efficient.

Design optimisation of non-axisymmetric exhausts for installed civil aero-engines

Future civil aero-engines are likely to operate with higher bypass-ratios (BPR) than current power-plants to improve propulsive efficiency and reduce specific thrust. This will probably be accompanied by an increase of fan diameter and size of the power plant. Consequently, future configurations are likely to require more close-coupled installations with the airframe due to structural and ground clearance requirements. This tendency may lead to an increase in the adverse installation effects which could be mitigated with non-axisymmetric exhausts. However, due to the prohibitive computational cost, limited regions of the design space have been studied. For this reason, a relatively low-cost design approach for the integrated system is required. The aim of this work is to establish a method to map the non-axisymmetric exhaust design space where the effects of the propulsion system installation are taken into account. The methodology relies on the generation of a design database using inviscid computational fluid dynamics (CFD) methods. This is used to characterise the design space, identify the dominant design parameters and build response surface models for optimisation. The candidate designs that arise from the optimisation are assessed with viscous CFD simulations to assess the aerodynamic mechanisms and performance characteristics. The result is a set of design recommendations for installed configurations with non-axisymmetric exhausts. The method is an enabler for the optimisation of installed propulsion systems and has provided an exhaust design with a 0.7% improvement on net vehicle force relative to an axisymmetric exhaust, for a close coupled configuration where the fan cowl is overlapped with the wing. A reduction in net vehicle force is expected to lead to a similar reduction in cruise fuel burn.

Propulsion integration study of civil aero-engine nacelles

AbstractIt is envisaged that future civil aero-engines will operate with ultra-high bypass ratios to reduce the specific fuel consumption. To achieve the expected benefits from the new engine cycles, these new powerplants may mount compact nacelles. For these new configurations the aerodynamic coupling between the powerplant and the airframe may increase. For this reason, it is required to quantify and further understand the effects of aircraft integration for compact aero-engine nacelles. This study provides an insight of the changes in flow aerodynamics as well as quantification of the most relevant performance metrics of the powerplant, airframe and the combined aircraft system across a range of different installation positions. Relative to a conventional architecture, there is an aerodynamic benefit in net vehicle force of about 1.2% for a compact powerplant when installed in forward positions. This is the same improvement that was identified when the aero-engine nacelles were in isolation. However, for close-coupled installation positions, the aerodynamic benefit in net vehicle force erodes to 0.44% due to the larger effects of aircraft integration on compact nacelles.

Effects of swirl and hot streak on thermal performances of a high-pressure turbine

Advanced civil aero-engines tend to adopt lean burn combustors to meet emission requirements. The exit of a lean burn combustor experiences highly non-uniformities in both temperature (Hot Streak, HS) and flow (swirl). This paper presents a numerical investigation on the behaviors of a High-Pressure (HP) turbine under a combined effect of swirl and hot streak. The investigation was conducted on a GE-E3 HP turbine with unsteady numerical simulations, which considered the realistic clocking position of the HP Nozzle Guide Vane (NGV) relative to the combustor. The influences of swirl orientations on the HS migration and thermal performances on the blade surface were examined. Results indicate that, inside the NGV passage, the swirl’s induced incidence angle effect dominates the HS radial migration. The transversal movement of HS follows the cross flow and thus makes itself approach the Suction Side (SS) and keep away from the Pressure Side (PS) as passing through the NGV, so that HS near the SS is more influenced by the incidence angle effect than that near the PS. As for the heat transfer, swirl affects the Heat Transfer Coefficient (HTC) on the NGV’s PS and SS mainly through the incidence angle effect. Different from the NGV, the inlet swirl and HS have limited effect on the HTC on the rotor blade’s PS, while on the rotor blade’s SS, the original vortex system dominates; therefore, the inlet non-uniformities merely enhance the HTC on the SS rather than alter its distribution characteristics.

Application of an Automatic Mesh Convergence Procedure for the Large Eddy Simulation of a Multipoint Injection System

AbstractLean combustion is an attractive alternative to limit pollutants levels in order to meet the imposed limitations for the next generation of civil aero-engines. However, its implementation involves important technological questions related to the augmentation of the air dedicated to the combustion process. An effort on the injection system design is required and large eddy simulation (LES) can be a useful tool in order to explore the design of novel concepts. At the state of the art, the validation of LES in high-pressure reactive conditions and in presence of the liquid phase is still limited. This shrinks the understanding and optimization of lean devices. The industrial project PERCEVAL, between Safran Tech and the CORIA laboratory, aims at extending the actual knowhow on lean combustion. Novel optical experimental techniques have been developed at CORIA to gain detailed information on industrial injection systems at high-pressure conditions. Within PERCEVAL, Safran Tech is in charge of the assessment of LES by using the experimental dataset collected at CORIA. In this framework, a novel automatic mesh convergence (AMC) procedure, based on adaptive mesh refinement, has been developed in the YALES2 platform to speed-up the calculation process. In the present paper, the AMC framework is described and then applied on the lean injection system designed at Safran Tech and tested during PERCEVAL. An analysis is carried out to evaluate the interest and gains offered by the AMC framework.

Towards the design and optimisation of future compact aero-engines: intake/fancowl trade-off investigation

PurposeRelative to in-service aero-engines, the bypass ratio of future civil architectures may increase further. If traditional design rules are applied to these new configurations and the housing components are scaled, then it is expected that the overall weight, nacelle drag and the effects of aircraft integration will increase. For this reason, the next generation of civil turbofan engines may use compact nacelles to maximise the benefits from the new engine cycles. The purpose of this paper is to present a multi-level design and optimisation process for future civil aero-engines.Design/methodology/approachAn initial set of multi-point, multi-objective optimisations for axisymmetric configurations are carried out to identify the trade-off between intake and fancowl bulk parameters of highlight radius and nacelle length on nacelle drag. Having identified the likely optimal part of the design space, a set of computationally expensive optimisations for three-dimensional non-axisymmetric configurations is performed. The process includes cruise- and windmilling-type operating conditions to ensure aerodynamic robustness of the downselected configurations.FindingsRelative to a conventional aero-engine nacelle, the developed process yielded a compact aero-engine configuration with mid-cruise drag reduction of approximately 1.6% of the nominal standard net thrust.Originality/valueThe multi-point, multi-objective optimisation is carried out with a mixture of regression and classification functions to ensure aerodynamic robustness of the downselected configurations. The developed computational approach enables the optimisation of future civil aero-engine nacelles that target a reduction of the overall fuel consumption.

Influence of Wake Intensity on the Unsteady Flow Characteristics of the Integrated Aggressive Interturbine Duct

The interturbine transition duct (ITD), located between the high-pressure (HP) and low-pressure (LP) turbines of aeroengines, tends to be designed as an aggressive ITD integrated with wide-chord struts to meet the requirements of civil aeroengines for high bypass ratios and thrust–weight ratios. This paper presents a detailed unsteady numerical investigation of the effects of the HP rotor trailing-edge radius on the unsteady flow characteristics in the integrated aggressive interturbine transition duct (AITD), including the transport and dissipation of HP rotor wakes, the control mechanism of HP rotor wakes on flow separation and the influence of wake parameters. A sweeping rod, with a nondimensional diameter ranging from d/s = 0.056~0.143 (based on the pitch (s) of wide-chord struts at the midspan) and a reduced frequency (f) of 1.07, is used to simulate the HP rotor wake to decouple its influence from other secondary flows. Using the k-ω SST turbulence model and gamma–theta transition model, a structured grid with 6.3 million nodes can achieve similar global results. The wake in the lower part of the AITD channel dissipates rapidly because of the stretching between its own circumferential motion and the radial upward secondary flow, especially for a small d/s. Only the residual wake in the upper part can reach wide-chord struts in the case with large d/s. A sweeping rod with a large d/s can reduce the radial pressure gradient in the AITD, inhibit the internal secondary flow to a certain extent, reduce the dissipation rate of the wake, enhance its suppression effect on flow separation on a wide-chord strut, and decrease the flow loss. However, the wake can also enhance the passage vortex due to the increasing circumferential pressure gradient in the wide-chord strut channel, resulting in increasing blade profile loss. In the scope of this study, the aerodynamic gain of the wake is still not enough to compensate for its loss increment (including its own dissipation loss). Therefore, selecting a small trailing-edge radius of the HP rotor is conducive to improving the aerodynamic performance of the integrated AITD.

Impact and quasi-static bending tests for a novel composite structure of ceramic and polymeric matrix composites

To resolve the dilemma between impact resistance and structural stiffness of fan containment casing of modern civil aeroengines, a novel composite structure (CS) is proposed. It combines the hard and stiff ceramic which layered at the impact side and the lightweight and tough polymeric matrix composites (PMC) on the back side. The new structure was tested under ballistic impact loadings and quasi-static four-point bending. The results indicate that the impact resistance of the new structure is improved, compared with the PMC-only samples. Meanwhile, the bending stiffness of the new structure is also substantially larger at the cost of much lower bending strength.

A Weakly Supervised Gas-Path Anomaly Detection Method for Civil Aero-Engines Based on Mapping Relationship Mining of Gas-Path Parameters and Improved Density Peak Clustering.

Gas-path anomalies account for more than 90% of all civil aero-engine anomalies. It is essential to develop accurate gas-path anomaly detection methods. Therefore, a weakly supervised gas-path anomaly detection method for civil aero-engines based on mapping relationship mining of gas-path parameters and improved density peak clustering is proposed. First, the encoder-decoder, composed of an attention mechanism and a long short-term memory neural network, is used to construct the mapping relationship mining model among gas-path parameters. The predicted values of gas-path parameters under the restriction of mapping relationships are obtained. The deviation degree from the original values to the predicted values is regarded as the feature. To force the extracted features to better reflect the anomalies and make full use of weakly supervised labels, a weakly supervised cross-entropy loss function under extreme class imbalance is deployed. This loss function can be combined with a simple classifier to significantly improve the feature extraction results, in which anomaly samples are more different from normal samples and do not reduce the mining precision. Finally, an anomaly detection method is deployed based on improved density peak clustering and a weakly supervised clustering parameter adjustment strategy. In the improved density peak clustering method, the local density is enhanced by K-nearest neighbors, and the clustering effect is improved by a new outlier threshold determination method and a new outlier treatment method. Through these settings, the accuracy of dividing outliers and clustering can be improved, and the influence of outliers on the clustering process reduced. By introducing weakly supervised label information and automatically iterating according to clustering and anomaly detection results to update the hyperparameter settings, a weakly supervised anomaly detection method without complex parameter adjustment processes can be implemented. The experimental results demonstrate the superiority of the proposed method.

Non-axisymmetric aero-engine nacelle design by surrogate-based methods

For many aerodynamic design tasks, a key challenge is the balance between the non-linearity of the transonic flow, the inherent 3D nature of the geometry and flow field, the computational cost and the level of accuracy. Within an optimisation process this is further compounded by the high degrees of freedom to define the geometry and the requirement for 3D computations at both design and off-design conditions. An example of this is the design of compact nacelles for future civil aero-engines which will have larger bypass ratios than current in-service architectures. This paper presents an approach for the design and optimisation of 3D drooped and scarfed non-axisymmetric nacelles. To reduce the computational expense, a range of surrogate-based adapted methods are investigated. Relative to the conventional approach of full numerical simulations in the optimisation loop, the adapted method identifies an acceptable design space with a 65% reduction in the total computational cost. Overall, this demonstrates a useful approach for reducing the time and cost of high-dimensional, aerodynamic design problems.

Fuzzy Controller Structures Investigation for Future Gas Turbine Aero-Engines

The Advisory Council for Aeronautics Research in Europe (ACARE) Flight Path 2050 focuses on ambitious and severe targets for the next generation of air travel systems (e.g., 75% reduction in CO2 emissions per passenger kilometer, a 90% reduction in NOx emissions, and 65% reduction in noise emission of flying aircraft relative to the capabilities of typical new aircraft in 2000). In order to meet these requirements, aircraft engines should work very close to their operating limits. Therefore, the importance of advanced control strategies to satisfy all engine control modes simultaneously while protecting them from malfunctions and physical damages is being more crucial these days. In the last three decades, fuzzy controllers (FCs) have been proposed as a high potential solution for performance improvement of the next generation of aircraft engines. Based on an analytic review, this paper divides the trend of FCs design into two main lines including pure FCs (PFC) and min–max FCs (MMFC). These two main architectures are then designed, implemented on hardware, and applied in a case study to analyze the advantages and disadvantages of each structure. The analysis of hardware-in-the-loop (HIL) simulation results shows that the pure FC structure would be a high potential candidate for maneuverability and response time indices improvement (e.g., military applications); while min–max FC architecture has a great potential for future civil aero-engines where the fuel consumption and steady-state responses are more important. The simulation results are also compared with those of industrial min–max controllers to confirm the feasibility and reliability of the fuzzy controllers for real-world application. The results of this paper propose a general roadmap for fuzzy controllers’ structure selection for new and next generation of aircraft engines.

Civil turbofan propulsion aerodynamics: Thrust-drag accounting and impact of engine installation position

It is envisaged that the next generation of civil aero-engines will employ high bypass ratios to lower specific thrust and improve propulsive efficiency. This trend is likely to be accompanied with the integration of compact nacelle and exhausts in podded under-wing installation positions that are close coupled to the airframe. This leads to the requirement for a comprehensive methodology able to predict aerodynamic performance for combined airframe-engine architectures. This paper presents a novel thrust and drag accounting approach for the aerodynamic analysis of integrated airframe-engine systems. An integral metric is synthesised based on the concept of net vehicle force. This is accomplished through the consolidation of aerodynamic coefficients, combined with the engine cycle characteristics obtained from a thermodynamic matching model. The developed approach is coupled with an in-house tool for the aerodynamic design and analysis of installed aero-engines. This framework is deployed to quantify the impact of engine installation position on the aerodynamic performance of a future large turbofan installed on a commercial wide-body airframe. The governing flow mechanisms are identified and their influence is decomposed in terms of the impact on airframe, nacelle, and exhaust performance. It is shown that it is essential to include the impact of installation on the exhaust for the correct determination of overall airframe-engine performance. The difference in net vehicle force for a close coupled position can reach up to -0.70% of nominal standard net thrust relative to a representative baseline engine location.

Predictions of NOx and CO emissions from a low-emission concentric staged combustor for civil aeroengines

This study is aimed to establish a detailed chemical reactor network model based on the analysis of complex reaction flowfield structures in aeroengine combustors, so that the emissions of nitrogen oxides and carbon monoxide from advanced civil aeroengines can be predicted quickly and accurately. In this study, a low-emission concentric staged combustor with three axial swirlers is designed for civil aeroengines, and numerical simulations of the three-dimensional reaction flowfields of the combustor during four load phases of takeoff, climb, approach, and idle, are conducted. Based on the numerical results, a simple chemical reactor network model with seven perfectly stirred reactors and a detailed chemical reactor network model using up to 15 perfectly stirred reactors are established. Using the developed chemical reactor network models and the detailed JP10 chemical reaction mechanism—composed of 374 step elementary reactions and 82 species, the emission variations of nitrogen oxides and carbon monoxide are predicted and compared with those estimated using an empirical formula and with the numerical results as a function of the combustion load. Using a combined chemical reactor network–computational fluid dynamics analysis method, the variations of the formation path, the mechanism, and the amounts of nitrogen oxides in the combustor and in the perfectly stirred reactors, are analyzed as a function of the combustion load. In addition, the effects of fuel and air pilot-to-total ratio on nitrogen oxides emissions for the 100% load condition are also analyzed. It is found that at high loads, the production rate of the thermal NO is the highest, while at low loads, the production rate of the prompt NO is the highest. The nitrogen oxide is mainly produced in the pilot zone and the recirculation zone, while its production in the outer main stage zone is low. The results show that the NOx emissions predicted by the complex chemical reactor network model are most consistent with those elicited using the empirical formula.

Study of Noise Reduction Based on Optimal Fan Outer Pressure Ratio and Thermodynamic Performance for Turbofan Engines at Conceptual Design Stage

With the development of civil aero-engines, the noise emission is now an essential issue to be considered during engine design. In this study, the preliminary design for a model fan, which was used for the fan noise estimation, was carried out. The fans with different bypass ratios, different tip speed as well as different aspect ratio were designed and the rough dimensions of the fan were obtained. Based on the dimension and performance of the fan, the effect of fan bypass ratio, tip speed, rotor blade numbers, and rotor–stator spacing on noise generation was investigated. The results indicate that the fan noise can be reduced as high as 10 dB by the increase of bypass ratios, and fewer rotor blade numbers and larger rotor–stator spacing are proved to be benefit to the fan noise reduction as well. However, lower tip speed does not achieve the noise reduction as expected. This is because more rotor blades are added to maintain the constant fan pressure ratio for the fan with lower tip speed. The study demonstrates that trade-off study should be carried out when considering fan noise reduction during engine design.

Surrogate-based aerodynamic optimisation of compact nacelle aero-engines

Genetic algorithms are a powerful optimisation technique for the design of complex engineering systems. Although computing power continuously grows, methods purely based on expensive numerical simulations are still challenging for the optimisation of aerodynamic components at an early stage of the design process. For this reason, response surface models are typically employed as a driver of the genetic algorithm. This reduces considerably the total overhead computational cost but at the expense of an inherent prediction uncertainty. Aero-engine nacelle design is a complex multi-objective optimisation problem due to the nonlinearity of transonic flow aerodynamics. This research develops a new framework, that combines surrogate modelling and numerical simulations, for the multi-objective optimisation of aero-engine nacelles. The method initially employs numerical simulations to guide the genetic algorithm through generations and uses a combination of higher fidelity results along with evolving surrogate models to identify a set of optimum designs. This new approach has been applied to the multi-objective optimisation of civil aero-engines which are representative of future turbofan configurations. Compared to the conventional CFD in-the-loop optimisation method, the proposed algorithm successfully identified the same set of optimum nacelle designs at a 25% reduction in the computational cost. Within the context of preliminary design, the method meets the typical 5% acceptability criterion with a 65% reduction in computational cost.

Civil turbofan engine exhaust aerodynamics: Impact of fan exit flow characteristics

It is envisaged that future civil aero-engines will operate with greater bypass ratios compared to contemporary configurations to lower specific thrust and improve propulsive efficiency. This trend is likely to be accompanied with the implementation of a shorter nacelle and bypass duct for larger engines. However, a short bypass duct may result in an aerodynamic coupling between the exit flow conditions of the fan Outlet Guide Vanes (OGVs) and the exhaust system. Thus, it is imperative that the design of the exhaust is carried out in combination with the fan exit profile. A parabolic definition is used to parameterise and control the circumferentially-averaged radial profiles of stagnation pressure and temperature at the fan OGV exit. The developed formulation is coupled with a parametric exhaust design approach, an automatic computational mesh generator, and a compressible flow solution method. A global optimisation strategy is devised comprising methods for Design of Experiment (DOE), Response Surface Modelling (RSM), and genetic optimisation.A combined Design Space Exploration (DSE) comprising both geometric, as well as fan exit profile variables, is performed to optimise the exhaust geometry in conjunction with the fan exit profile. The developed approach is used to derive optimum exhaust geometries for a tip, mid, and hub-biased fan blade loading distribution. It is shown that the proposed formulation can ameliorate adverse transonic flow characteristics on the core after-body due to a non-uniform bypass inflow. The hub-loaded profile was found to be most penalising in terms of exhaust performance compared to the mid and tip-loaded variants. It is demonstrated that the combined fan exit profile and exhaust geometry optimisation offers significant performance improvement compared to the fixed inflow cases. The predicted performance benefits can reach up to 0.19% in terms of exhaust velocity coefficient, depending on fan loading characteristics. A notable improvement is also noted in terms of bypass nozzle discharge coefficient. This suggests that the combined optimisation can lead to an exhaust design that can satisfy the engine mass-flow rate demand with a reduced geometric throat area, thus potentially offering further exhaust size and weight benefits.

A comparative study of coupled and decoupled fan flutter prediction methods under variation of mass ratio and blade stiffness

Flutter is a long standing issue for fan blades of civil aero-engines and becomes of further concern for modern light-weight designs with increasing fan diameters to reach ultra-high bypass ratios. Accurate flutter prediction is therefore of prime consideration in the design process in order to avoid catastrophic blade failure in operation or expensive redesign iterations if spotted in ground or flight tests. The traditional energy method, which is based on the assumption of negligible aeroelastic coupling, has been used to great extend to predict flutter of turbomachinery components including aero-engine fan blades, and is today by far the most widely applied technique. The underlying assumption of fluid–structure decoupling, however, has to be questioned for large fan blades that are characterized by low mass ratios and low stiffness. Implications of the violation of the system decoupling assumption on the prediction capabilities of the energy method are important to understand for the fan designer in order to allow an informed decision on the flutter prediction tool to use. In this work a comprehensive comparative study is presented in which the energy method is contrasted to the predictions of a strongly coupled fluid–structure interaction method for varying values of mass ratio and blade stiffness of a transonic three-dimensional fan rotor. The strength of aeroelastic coupling is evaluated in terms of the aeroelastic frequency shift and its impact on the prediction accuracy of the energy method is investigated. The results show the capability of the energy method to accurately predict flutter for a wide range of mass ratio and stiffness configurations, but its prediction accuracy is reduced for combined low mass ratio and low stiffness blades. Mechanisms governing the aeroelastic frequency shift are explained to allow a better understanding of the effect and a method for its prediction based on results of a decoupled analysis is shown.

Thermal Management Systems for Civil Aircraft Engines: Review, Challenges and Exploring the Future

This paper examines and analytically reviews the thermal management systems proposed over the past six decades for gas turbine civil aero engines. The objective is to establish the evident system shortcomings and to identify the remaining research questions that need to be addressed to enable this important technology to be adopted by next generation of aero engines with complicated designs. Future gas turbine aero engines will be more efficient, compact and will have more electric parts. As a result, more heat will be generated by the different electrical components and avionics. Consequently, alternative methods should be used to dissipate this extra heat as the current thermal management systems are already working on their limits. For this purpose, different structures and ideas in this field are stated in terms of considering engines architecture, the improved engine efficiency, the reduced emission level and the improved fuel economy. This is followed by a historical coverage of the proposed concepts dating back to 1958. Possible thermal management systems development concepts are then classified into four distinct classes: classic, centralized, revolutionary and cost-effective; and critically reviewed from challenges and implementation considerations points of view. Based on this analysis, the potential solutions for dealing with future challenges are proposed including combination of centralized and revolutionary developments and combination of classic and cost-effective developments. The effectiveness of the proposed solutions is also discussed with a complexity-impact correlation analysis.

Design optimisation of separate-jet exhausts for the next generation of civil aero-engines

ABSTRACTThe next generation of civil large aero-engines will employ greater bypass ratios compared with contemporary architectures. This results in higher exchange rates between exhaust performance and specific fuel consumption (SFC). Concurrently, the aerodynamic design of the exhaust is expected to play a key role in the success of future turbofans. This paper presents the development of a computational framework for the aerodynamic design of separate-jet exhaust systems for civil aero-engines. A mathematical approach is synthesised based on class-shape transformation (CST) functions for the parametric geometry definition of gas-turbine exhaust components such as annular ducts and nozzles. This geometry formulation is coupled with an automated viscous and compressible flow solution method and a cost-effective design space exploration (DSE) approach. The framework is deployed to optimise the performance of a separate-jet exhaust for very-high-bypass ratio (VHBR) turbofan engine. The optimisations carried out suggest the potential to increase the engine’s net propulsive force compared with a baseline architecture, through optimum exhaust re-design. The proposed method is able to identify and alleviate adverse flow-features that may deteriorate the aerodynamic behaviour of the exhaust system.