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Within aircraft electrical network designs, energy storage systems (ESS) provide a means of decoupling the electrical-mechanical interactions between the aircraft electrical power system and the aircraft engine, meeting peak load demand and maintaining power quality during network disturbances and variable load conditions. Within the literature to date, control and management strategies of ESSs for such applications has primarily focused on normal network operation with only limited coverage on the behavior of such technologies under abnormal conditions and the subsequent impact on the operation of the wider power system. Through modeling and simulation of a generic aircraft electrical system, this paper will highlight the potential risks of the inherent, sub-optimal operation of certain existing control strategies during fault conditions.

Mass and efficiency are key performance indicators for the development and design of future electric power systems (EPS) for more-electric aircraft (MEA). However, to enable consideration of high-level EPS architecture design trades, there is a requirement for modelling and simulation based analysis to support this activity. The predominant focus to date has been towards the more detailed aspects of analysis, however there is also a significant requirement to be able to perform rapid high-level trades of candidate architectures and technologies. Such a capability facilitates a better appreciation of the conflicting desires to maximize availability and efficiency in candidate MEA architectures, whilst minimizing the overall system mass. It also provides a highly valuable and quantitative assessment of the systemic impact of new enabling technologies being considered for MEA applications.

There is a well-recognised need for robust simulation tools to support the design and evaluation of future More-Electric Engine and Aircraft (MEE/MEA) design concepts. Design options for these systems are increasingly complex, and normally include multiple power electronics converter topologies and machine drive units. In order to identify the most promising set of system configurations, a large number of technology variants need to be rapidly evaluated. This paper will describe a method of MEE/MEA system design with the use of a newly developed transient modeling, simulation and testing tool aimed at accelerating the identification process of optimal components, testing novel technologies and finding key solutions at an early development stage. The developed tool is a Matlab/Simulink library consisting of functional sub-system units, which can be rapidly integrated to build complex system architecture models.

Turbo-electric distributed propulsion (TeDP) for aircraft allows for the complete redesign of the airframe so that greater overall fuel burn and emissions benefits can be achieved. Whilst conventional electrical power systems may be used for smaller aircraft, large aircraft (~300 pax) are likely to require the use of superconducting electrical power systems to enable the required whole system power density and efficiency levels to be achieved. The TeDP concept requires an effective electrical fault management and protection system. However, the fault response of a superconducting TeDP power system and its components has not been well studied to date, limiting the effective capture of associated protection requirements. For example, with superconducting systems it is possible that a hotspot is formed on one of the components, such as a cable. This can result in one subsection, rather than all, of a cable quenching.

In addition to providing thrust, the engines on conventional civil jet airliners generate power for on-board systems and ancillary loads in the form of pneumatic, hydraulic and electrical power. Reduced fuel-burn and efficiency targets have driven the move towards More Electric Aircraft (MEA) technology which seeks to replace hydraulic and pneumatic loads with electrical equivalents. This technological shift, in conjunction with a growing electrical power load per passenger in general, has greatly increased the electrical power demands of aircraft in recent years - over 1 MVA for the Boeing 787 for example. With increasing fuel prices, there is a growing need to optimise efficiency of power extraction from the aircraft engines for the electrical system and loads. In particular, the utilisation of multi-shaft power off-takes, interconnected generation and power sharing between shafts is thought to offer potentially significant engine operability and fuel efficiency benefits.

NASA has compiled a set of research goals for five year periods starting 2015, 2020 and 2025 for three classes of future subsonic aircraft, N+1 (2015), N+2 (2020) and N+3 (2025). With the intention of progressively making reductions in noise emissions, greenhouse gas emissions, fuel burn and energy consumption at each of these points to achieve Technology Readiness Levels (TRL's) of between 4 and 6. In the last few years much progress has been made towards achieving these goals through the development of new technologies and designs. This paper assesses how the current More Electric Aircraft (MEA) design concepts are advancing to allow the near term, N+1 goals of reducing 32 dB of noise emissions, 60% of the landing and take-off (LTO) NOx emissions, 55% of cruise emissions and 33% saving of fuel burn and energy consumption, relative to single aisle B737-800, could be met and eventually surpassed.

The aviation industry has witnessed a technological shift towards the More Electric Aircraft (MEA) concept. This shift has been driven by a number of perceived benefits including performance optimization and reduced life-cycle costs. Increased electrification within MEA has made aircraft electrical networks larger and more complex and this necessitates an increased electrical power offtake from the engine. The paralleling of multiple generation sources across the aircraft is one potential design approach which could help improve engine operability and fuel efficiency within more-electric aircraft platforms. Accordingly, this paper will investigate options for the realization of paralleled generation systems within the context of current design and certification rules. The paper first illustrates, through simulation, that MIL-STD-704F voltage envelopes may be breached for some interconnected electrical architectures under fault conditions.

Direct current (DC) for primary power distribution is a promising solution that is being explored by aircraft system integrators for MEA applications to enable the paralleling of non-synchronized engine off-take generators, and to enable the reduction of energy conversion stages required to supply electronically actuated loads. However, a significant challenge in the use of DC systems is the reliable detection of arc faults. Arcing presents a significant fire risk to aircraft and their presence can result in critical system damage and potentially fatal conditions. Series arc faults in DC systems are particularly challenging to detect as the associated reduction in system current eliminates the use of conventional overcurrent and current differential methods for fault detection. This paper provides an overview of series arc faults in DC systems and presents both simulation and hardware results to illustrate key trends, characteristics and discriminating features.

The increasing electrical demand in commercial and military aircraft justifies a growing need for higher voltage DC primary distribution systems. A DC system offers reduced power losses and space savings, which is of major importance for aircraft manufacturers. At present, challenges associated with DC systems include reliable fast acting short circuit protection. Solid State Contactors (SSC) have gained wide acceptance in traditional 28 VDC secondary systems for DC fault interruption. However, the reliable operation at higher operating voltages and currents requires further technology maturation. This paper examines a supporting method to SSC for more reliable fault mitigation by investigating bidirectional AC/DC converter topology with DC fault current blocking capability. Replacement of semiconductor switches with full bridge cells allows instant reversal of voltage polarities to limit rapid capacitor discharge and machine inductive currents.

Distributed electrical propulsion for aircraft, also known as turbo-electric distributed propulsion (TeDP), will require a complex electrical power system which can deliver power to multiple propulsor motors from gas turbine driven generators. To ensure that high enough power densities are reached, it has been proposed that such power systems are superconducting. Key to the development of these systems is the understanding of how faults propagate in the network, which enables possible protection strategies to be considered and following that, the development of an appropriate protection strategy to enable a robust electrical power system with fault ride-through capability. This paper investigates possible DC protection strategies for a radial DC architecture for a TeDP power system, in terms of their ability to respond appropriately to a DC fault and their impact on overall system weight and efficiency.

A number of concepts have been proposed to meet future aircraft performance goals. One such concept under consideration is Turboelectric Distributed Propulsion (TeDP) featuring a large number of superconducting motors powered by two superconducting generators placed on each wingtip and connected through a DC distribution network. A key aspect in any design concept is the ability to prove that the system will exhibit a satisfactory reliability for all intended operating conditions. A common tool to support the calculation of failure rates and reliability is Fault Tree Analysis (FTA), and this will be utilized within this paper. The paper undertakes an architectural level FTA on a NASA proposed TeDP architecture to identify any significant factors contributing to the failure rate of key functionalities within the network.

Contemporary trends are leading towards the electrification of aircraft for urban mobility applications. Accordingly, there is a high demand for advancements in light-weight, high voltage technologies to realize these new aircraft types. Driven by recent developments in the automotive industry, hybrid Pyrofuse protection devices have emerged as one such new candidate technology. Pyrofuses offer rapid clearance of fault currents, reduced cost and weight when compared to conventional mechanical breakers. In addition, Pyrofuses have the ability to tune the time-current curve to fit the application’s fault response characteristics. However, Pyrofuses are non-resettable devices whose exclusive use for electrical protection could present potential operational hazards and certification challenges in aerospace applications. Model-based analysis will be critical in supporting this evaluation.