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As part of the More Electric Initiative, there is a significant interest in designing energy-optimized More Electric Aircraft, where electric power meets all non-propulsive power requirements. To achieve this goal, the aircraft subsystems must be analyzed much earlier than in the traditional design process. This means that the designer must be able to compare competing subsystem solutions with only limited knowledge regarding aircraft geometry and other design characteristics. The methodology presented in this work allows such tradeoffs to be performed and is driven by subsystem requirements definition, component modeling and simulation, identification of critical or constraining flight conditions, and evaluation of competing architectures at the vehicle and mission level. The methodology is applied to the flight control actuation system, where electric control surface actuators are likely to replace conventional centralized hydraulics in future More Electric Aircraft. While the potential benefits of electric actuation are generally accepted, there is considerable debate regarding the most suitable electric actuator -electrohydrostatic or electromechanical. These two actuator types form the basis of the competing solutions analyzed in this work, which focuses on a small narrowbody aircraft such as the Boeing 737-800. The competing architectures are compared at both the vehicle and mission levels, using as metrics subsystem weight and mission fuel burn, respectively. As shown in this work, the use of this methodology aids the decision-making process by allowing the designer to rapidly evaluate the significance of any performance advantage between the competing solutions.

The aerospace industry is currently transitioning to More Electric subsystem architectures due to steadily improving electric technologies and the technology saturation of established conventional architectures. For aircraft with such unconventional architectures, the lack of historical information and the presence of increased inter-subsystem interactions create a significant design challenge. These necessitate a greater focus on subsystems design earlier in the design process than typically seen for aircraft with conventional subsystem architectures. At the same time, however, to be suitable for the early design phases, the subsystem analyses must be computationally inexpensive and not require detailed aircraft definition. This work presents an integrated, modular, and tool-independent approach to the sizing and performance analysis of the aircraft and its subsystems, in which inter-dependencies are established between relevant aircraft and subsystem level parameters. The approach allows the assessment of subsystem architectures using vehicle and mission level metrics for a fixed vehicle design, and also the amplified effect of re-sizing the vehicle in accordance with a pre-defined rule-set. The proposed approach was demonstrated through a comparative assessment of a predominantly electric subsystem architecture and a conventional one for a representative single-aisle aircraft. In this assessment, the impacts of changing the form of secondary power extraction, the change in vehicle empty weight, and re-sizing of the vehicle were successively identified.

This paper presents a methodology for the sizing and synthesis of power generation and distribution (PG&D) subsystems. The PG&D subsystem models developed in a previous work done by the authors were applied within a parallel hybrid electric propulsion architecture using the Dornier 328 as the baseline aircraft. The hybridization took place only during the cruise segment. Analyses were performed in Pacelab SysArc, a system architecture design tool, to assess the impact of different hybrid electric propulsion architectures and changing PG&D subsystem characteristics at aircraft and mission levels. To this end, sensitivity analysis was conducted to reveal the sensitivity to the subsystem level characteristics. Moreover, six different architectures were compared in terms of their mission level performance. These architectures included the PG&D subsystems with current state of the art technology, NASA 15-year technology goals and a more advanced battery technology. Although neither the current state of the art PG&D subsystems nor NASA 15-year technology goals were advanced enough to match the design range requirement of the baseline aircraft, some of the competing architectures met the practical range target while enjoying substantial amount of fuel reductions. Finally, it was observed that in order to reach a break-even point in terms of the design mission range, a battery specific energy of 5 kWh/kg was necessary for a 50% level of hybridization during cruise. In this work the Dornier 328 was used as a testbed, however the methodology can be generalized for all parallel hybrid electric propulsion applications.

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