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Permanent magnet (PM) electrical machine has far-reaching impacts in aviation electrification due to the continuous development in high power density and high efficiency electrical drives. The primary barrier to acceptance of permanent magnet machines for safety-critical starter-generator systems is its low fault-tolerance capability and low reliability (for the conventional designs). This article investigates a modular triple three-phase PM starter-generator comprehensively, including the tradeoff of fault-tolerant topology, optimization design process, analysis of electromagnetic (highlight the post-fault analysis) and thermal behavior, respectively. The triple three-phase segmented topology proposed meet the fault-tolerant requirement along with complete electrical, magnetic, and thermal isolation. There would be cost penalty on the proposed topology, but it gets offset by the ease of manufacturing of coils and their insertion.

Evolving of aircraft design towards further electrification requires safe and fault-free operation of all the components. More electric aircraft are increasingly utilizing electro-mechanical actuators (EMA). EMAs are prone to jamming and subsequent failure due to large forces on the shaft. Large forces are generated due to the high reflected inertia of the electric machine rotor. To limit the force acting on the shaft, a torque limiting device is connected to the power train which can separate the rotating mass of the electric machine from the power train. In this paper, a concept of integration of torque limiter and the electric machine rotor is presented to reduce overall volume and mass. It is connected closely with the rotor, within the motor envelope. A commercially available torque limiter and an electric machine designed for actuator application are used to demonstrate the concept. While essential for safety, the torque limiter adds to the mass and size of the overall EMA.

Bleeding in engines is essential to mitigate the unmatched air massflow between low and High Pressure (HP) compressors at low speed settings, thus avoiding unstable operation due to surge and phenomena. However, by emerging the More Electric Aircraft (MEA) the engine is equipped with electrical machines on both high and Low Pressure (LP) spools which enables transfer of power electrically from one spool to another and hence provides the opportunity to operate engine core components closer to their optimum design point at off-design conditions. At lower power setting of the engine, HPC speed can be increased by taking power from LP shaft and feeding it to HP shaft which can lead to the removal of the bleeding system which in turn reduces weight and fuel consumption and help to overcome engine instability issues. Fuel consumption can be decreased by decreasing inconsistent thrust with the aircraft mission for flight and ground idle settings.

Vector based control strategies have been extensively employed for drive systems, and in recent times to the More Electric Aircraft (MEA) generator based systems. The control schemes should maintain the bus voltage and adhere to the generator system voltage and current limits throughout a wide speed range. Typically, the current limit is prioritised first due to ease of implementation and simple control structure. As a result, the voltage limit can be exceeded due to change in operating conditions or disturbance factors. In flux weakening regions, this may affect the controllability of the power converter and lead to generator system instability. In this paper, an alternative control strategy has been investigated to address this drawback. The proposed control scheme refers to the modulation index limit which is the ratio between the power converter input and output voltages as the voltage limit.

The solid state power controller (SSPC) is one of the most important power electronic components of the aircraft electrical power distribution (EPS) systems. This paper presents an architecture of the DC SSPC and provides the mitigation techniques for transient voltage overshoot during its turn-off. The high source side inductance carries breaking current (9xnominal current) just before turnoff and induces large voltage transient across the semiconductor devices. Therefore, the stored inductive energy needs to be dissipated in order to prevent semiconductor switches from over-voltage/thermal breakdown. Three different transient voltage suppression (TVS) devices to reduce voltage stress across switches are included in the paper for detail study. The comprehensive comparison of the TVS devices is presented. In addition, the thermal impact of the TVS devices on the semiconductor switches is also analyzed.

Studies of diesel system deposits continue to be the subject of interest and publications worldwide. The introduction of high pressure common rail systems resulting in high fuel temperatures in the system with the concomitant use of fuels of varying solubilizing ability (e.g. ULSD and FAME blends) have seen deposits formed at the tip of the injector and on various internal injector components. Though deposit control additives (DCAs) have been successfully deployed to mitigate the deposit formation, work is still required to understand the nature and composition of these deposits. The study of both tip and internal diesel injector deposits (IDID) has seen the development of a number of bench techniques in an attempt to mimic field injector deposits in the laboratory. One of the most used of these is the Jet Fuel Thermal Oxidation Tester or JFTOT (ASTM D3241).

High power density for aerospace motor drives is a key factor in the successful realization of the More Electric Aircraft (MEA) concept. An integrated system design approach offers optimization opportunities, which could lead to further improvements in power density. However this requires multi-disciplinary modelling and the handling of a complex optimization problem that is discrete and nonlinear in nature. This paper proposes a multi-level approach towards applying random heuristic optimization to the integrated motor design problem. Integrated optimizations are performed independently and sequentially at different levels assigned according to the 4-level modelling paradigm for electric systems. This paper also details a motor drive sizing procedure, which poses as the optimization problem to solve here. Finally, results comparing the proposed multi-level approach with a more traditional single-level approach is presented for a 2.5 kW actuator motor drive design.

This paper presents novel development of a reconfigurable assembly cell which assembles multiple aerostructure products. Most aerostructure assembly systems are designed to produce one variant only. For multiple variants, each assembly typically has a dedicated assembly cell, despite most assemblies requiring a process of drilling and fastening to similar tolerances. Assembly systems that produce more than one variant do exist but have long changeover or involve extensive retrofitting. Quick assembly of multiple products using one assembly system offers significant cost savings from reductions in capital expenditure and lead time. Recent trends advocate Reconfigurable Assembly Systems (RAS) as a solution; designed to have exactly the functionality necessary to produce a group of similar components. A state-of-the-art review finds significant benefits in deploying RAS for a group of aerostructures variants.

This paper proposes an integrated system’s approach towards design of aero-engine subsystems - seals, bearing chamber, generator and power system. In a conventional design approach, the design of the overall system is typically broken-down into subsystems. Therefore, the focus is not on the mutual interaction between different components or subsystems, resulting in a lack of characterization of the overall system performance at the design phase. A systems design approach adopts a much broader outlook, focusing on the overall optimization of the system performance. This paper is divided into two parts. The first part presents an integrated approach for modelling the electrical, mechanical and hydraulic subsystems of aero engines, in order to analyze the fluid dynamics interactions and reduce the transversal shaft vibrations. For this, an in-line starter/generator and an air-riding seal are studied.

The paper will deal with the problem of establishing a desirable power sharing in multi-feed electric power system for future more-electric aircraft (MEA) platforms. The MEA is one of the major trends in modern aerospace engineering aiming for reduction of the overall aircraft weight, operation cost and environmental impact. Electrical systems are employed to replace existing hydraulic, pneumatic and mechanical loads. Hence the onboard installed electrical power increases significantly and this results in challenges in the design of electrical power systems (EPS). One of the key paradigms for future MEA EPS architectures assumes high-voltage dc distribution with multiple sources, possibly of different physical nature, feeding the same bus(es). In our study we investigate control approaches to guarantee that the total electric load is shared between the sources in a desirable manner. A novel communication channel based secondary control method is proposed in this paper.

The future aircraft electrical power system is expected to be more efficient, safer, simpler in servicing and easier in maintenance. As a result, many existing hydraulic and pneumatic power driven systems are being replaced by their electrical counterparts. This trend is known as a move towards the More-Electric Aircraft (MEA). As a result, a large number of new electrical loads have been introduced in order to power many primary functions including actuation, de-icing, cabin air-conditioning, and engine start. Therefore electric power generation systems have a key role in supporting this technological trend. Advances in modern power electronics allow the concept of starter/generator (S/G) which enables electrical engine start and power generation using the same electrical machine. This results in substantial improvements in power density and reduced overall weight.

There is a growing demand for composites to be utilised in the production of large-scale components within the aerospace industry. In particular the demand to increase production rates indicates that traditional manual methods are no longer sufficient, and automated solutions must be sought. This typically leads to automated forming processes where there are a limited number of effective options. The need for forming typically arises from the inability of layup methods to produce complex geometries of structural components. This paper reviews the current state of the art in automated forming processes, their limitations and variables that affect performance in the production of large scale components. In particular the paper will focus on the application of force and heat within secondary forming processes. It will then review the effects of these variables against the structure of the required composite component and identify viability of the technology.

Current assembly systems that deal with large, complex structures present a number of challenges with regard to improving operational performance. Specifically, aerospace assembly systems comprise a vast array of interrelated elements interacting in a myriad of ways, resulting in a deeply complex process that requires a multi-disciplined team of engineers. The current approach to ramp-up production rate involves building additional main assembly fixtures which require large investment and lead times up to 24 months. Within Airbus Operations Ltd there is a requirement to improve the capacity and flexibility of assembly systems, thereby reducing non-recurring costs and time-to-market. Recent trends to improve manufacturing agility advocate Reconfigurable Assembly Systems (RAS) as a viable solution. Yet, adding reconfigurability to assembly systems further increases both the operational and design complexity.

In this paper, the load sharing principles in dc-distribution electric power systems (EPS) for future more-electric aircraft (MEA) are investigated. The study is conducted using a potential MEA EPS architecture with multiple sources feeding into the main dc bus. Corresponding reduced-order EPS models are established. The influence of the cable impedance on the load sharing accuracy is analyzed and sharing error is quantized in mathematical equations. In addition, source/load impedance of the droop-controlled system has been derived leading to the discussion of the stability issues in multi-feed dc EPS under different droop control strategies. The influence of load sharing ratio on the EPS stability margins has been investigated. The theoretical findings were supported by time-domain simulations in Matlab/SimPower.

This paper considers the electromechanical interconnection between the electrical power system of the More Electric Aircraft (MEA) and the shaft connecting the engine to the generator. In order to investigate the coupling between these two systems the effect of an electric load impact on the mechanical system of the MEA will be analysed. In the MEA, many functions traditionally powered by pneumatic, hydraulic and mechanical systems will be replaced by the electrical systems. Thus the electrical power rating will be considerably higher than that of a traditional aircraft. With the increase of electrical power, the impact of electrical load on the mechanical system, especially the engine shaft, will become significant. This paper focuses on the study of the interaction between the electrical and mechanical system.

This paper aims to develop a general functional model of multi-pulse Auto-Transformer Rectifier Units (ATRUs) for More-Electric Aircraft (MEA) applications. The ATRU is seen as the most reliable way readily to be applied in the MEA. Interestingly, there is no model of ATRUs suitable for unbalanced or faulty conditions at the moment. This paper is aimed to fill this gap and develop functional models suitable for both balanced and unbalanced conditions. Using the fact that the DC voltage and current are strongly related to the voltage and current vectors at the AC terminals of ATRUs, a generic functional model has been developed for both symmetric and asymmetric ATRUs. The developed functional models are validated through simulation and experiment. The efficiency of the developed model is also demonstrated by comparing with corresponding detailed switching models. The developed functional model shows significant improvement of simulation efficiency, especially under balanced conditions.

Aerospace assembly systems comprise a vast array of interrelated elements interacting in a myriad of ways. Consequently, aerospace assembly system design is a deeply complex process that requires a multi-disciplined team of engineers. Recent trends to improve manufacturing agility suggest reconfigurability as a solution to the increasing demand for improved flexibility, time-to-market and overall reduction in non-recurring costs. Yet, adding reconfigurability to assembly systems further increases operational complexity and design complexity. Despite the increase in complexity for reconfigurable assembly, few formal methodologies or frameworks exist specifically to support the design of Reconfigurable Assembly Systems (RAS). This paper presents a novel reconfigurable assembly system design framework (RASDF) that can be applied to wing structure assembly as well as many other RAS design problems.

Many More-Electric Aircraft (MEA) Electric Power System (EPS) architecture paradigms are being studied in order to provide the on-board electrical loads with high-quality supply and to ensure safe operation. EPS with dc distribution appears to be more promising due to higher efficiency, higher reliability, easier integration and lower overall weight. Another advantage of dc systems is the ease of sources paralleling, together with the optimization of load power sharing, this can lead to further EPS weight reduction. The DC bus can be fed by multiple sources such as generators, batteries and other energy storage devices. Many loads in MEA EPS are tightly controlled by power electronic converters and often behave as constant power loads (CPL). These are known as main contributors to the degradation of EPS stability margins. Therefore, stability study is one of the key topics in the assessment of potential EPS architecture candidates.

A 45kW, switched reluctance type, starter-generator, having a 1:4 constant power speed range has been designed as a possible candidate for a regional jet application. In the first section of this paper, a review of the major starter-generator topologies considered for the aerospace application is provided, highlighting the advantages of choosing the Switched reluctance topology for such a safety critical application. Following this, the required torque speed characteristic of the machine, along with the imposed physical constraints, in terms of cooling and outer dimensions, are also detailed. Section III provides a description of the Electromagnetic design, and challenges encountered in meeting both the low speed, peak torque node, at 8000rpm, and the high speed, high power node, at 32000rpm. The induced mechanical stresses in the rotor at such high speeds have also been evaluated and used as a material selection criterion for such a design as presented in section III.

The nature of internal diesel injector deposits (IDID) continues to be of importance to the industry, with field problems such as injector sticking, loss of power, increased emissions and fuel consumption being found. The deposits have their origins in the changes in emission regulations that have seen increasingly severe conditions experienced by fuels because of high temperatures and high pressures of modern common rail systems and the introduction of low sulphur fuels. Furthermore, the effect of these deposits is amplified by the tight engineering tolerances of the moving parts of such systems. The nature and thus understanding of such deposits is necessary to both minimising their formation and the development of effective diesel deposit control additives (DCA).