Recent research suggests that large battery-electric aircraft can achieve greater ranges than previously assumed and can therefore be a promising solution to decarbonize the aviation sector on short distances. However, several technical challenges need to be investigated for such aircraft to become technically and commercially viable. This paper summarizes the findings of a series of research projects investigating these technical challenges, which are then incorporated in the conceptual redesign of a 90-seater battery-electric aircraft. The top-level aircraft requirements are revisited and the main configuration trade-offs are discussed. The resulting design presents a maximum take-off mass of 82.5 t and achieves a battery-only range of 750 km for a pack energy density of 320 Wh/kg, while reserves are covered by a dedicated fuel-based reserve energy system (RES). The impact of using the RES for range extension is also investigated. A comparison to conventional aircraft configurations in terms of CO2-equivalent and operating costs per passenger-kilometer demonstrates that such aircraft can become a cost-competitive solution to mitigate the climate impact of aviation on short ranges.

Distributed propulsion systems, characterized by multiple propellers, represent a promising approach for full-electric aircrafts, offering several advantages but also introducing technical challenges. The main objective of this paper is to quantify how the propeller performance and noise emissions of an eight-propeller full-electric aircraft configuration compare to that of a conventional fuel-based turboprop. In both cases, the key parameters driving the trade-off between noise emissions and aerodynamic performances are analysed as well as the benefits of each configuration. The propeller noise emissions are analysed in terms of the perceived noise emissions at the three certification points: approach, take-off, and flyover. Optimizations are performed as a function of blade count to investigate the performance and noise trends for different propeller configurations. The results show a promising performance for the battery-electric aircraft with distributed propulsion, achieving a propeller efficiency between 83% and 88% in cruise without incurring a major noise penalty compared to the reference turboprop aircraft, despite the large increase in aircraft size and weight.

The preliminary design of the thermal management system for the electric powertrain of a 90-seater battery-electric aircraft is addressed in this study. The main heat loads of the powertrain are associated with the battery pack and the electronic power unit, comprising an electric motor, an inverter, and optionally a gearbox. The main operating points of the thermal management system throughout the nominal flight mission are identified as take-off, top of climb, cruise, and top of descent. Two candidate architectures are assessed and compared in terms of key performance metrics, i.e., weight, drag penalty, and electric power consumption. Moreover, the equivalent battery mass required to power and carry the thermal management system throughout the flight mission is computed. The results show that the architecture exploiting the difference in the temperature levels of the heat sources in the powertrain features superior performance than the baseline configuration. In particular, the total equivalent battery mass is reduced by 388 kg. Overall, this translates into an increase in the usable range of 11 km prior to snowball effects, highlighting the importance of the thermal management system in the design of large battery-electric aircraft.

Battery-electric aviation is commonly believed to be limited to small aircraft and is therefore expected have a negligible impact on the decarbonization of the aviation sector. In this paper we argue that, with the correct choice of design parameters and top-level aircraft requirements, the addressable market is actually substantial. To demonstrate this, the Class-II sizing of a battery-electric 90-seater is performed, and the environmental impact is assessed in terms of well-to-wake CO2-equivalent emissions per passenger-kilometer. The resulting 76-ton aircraft achieves a battery-powered useful range of 800 km for a pack-level energy density of 360 Wh/kg. For this range, it has an energy consumption of 167 Wh per passenger-kilometer and an environmental impact well below that of kerosene, eSAF, or hydrogen-based aircraft alternatives and comparable to land-based modes of transport. These results indicate that, to successfully reduce the climate impact of the aviation sector, battery-electric aircraft should not be designed as a niche product operating from small airfields but as commercial transport aircraft competing with fuel-based regional and narrowbody aircraft.

Thus far, battery-electric propulsion has not been considered a promising pathway to climate-neutral aviation. Given current and expected battery technology, in most literature battery electric aircraft are only considered feasible for short ranges (< 400 km) and small payloads (< 19 pax). As a result, battery-electric aircraft development focuses on new aviation segments such as regional and urban air mobility. However, little effort has been made to develop battery-electric aircraft that can replace existing larger aircraft. This paper re-examines the assumptions that lead to the conclusion of limited applicability of battery-electric aircraft. Starting from the range equation, this paper assesses the drivers of two key parameters: the ratio between energy mass and maximum take-off mass, and the maximum lift-to-drag ratio. This assessment, based on Class-I mass and aerodynamic-efficiency estimates, shows that there is a design space where these two parameters can reach significantly higher values than often assumed in the open literature. Based on this finding, several parametric aircraft designs are evaluated, relying on Class-II mass and aerodynamics methods. These parametric studies validate the conclusion from the Class-I assessment. This implies that battery-electric passenger aircraft can play a larger role in climate-neutral aviation than was previously envisioned.

Optimized power system architectures and lighter weight are enabling considerations for the successful development of all-electric aircraft (AEA). In this article, a cross-redundant connection architecture and weight reduction solutions are investigated for a 90-seater full battery-electric aircraft from the perspective of high-power aviation cable. Design criteria of the power system architecture are introduced. Material selection, sizing, and weight estimation methods of cable for AEA are discussed by combining ground cable standards with aviation requirements. The influence of the conductor materials, voltage level, current, battery pack quantity, and operating temperature on cable evaluation is thoroughly discussed and analyzed. Weight comparison under two controversial voltage level options (800V and 3kV) is conducted. Comparison results show that the utilization of an aluminum conductor, PTFE insulator, and a voltage level of 3kV proves to be a preferable selection for current AEA medium and high voltage cables. Increasing the rating operation temperature to 120°C is a conservative and secure option. The layout of battery packs consistent with the quantity of distributed electric motors is preferable to achieve the lightest cabling system. This study provides a guideline for the cable sizing methods of high-power aviation cables and an optimized design solution for the power system architecture of AEA from the perspective of cable layout and weight assessment.