The growth in air transport and the ambitious targets in emission reductions set by advisory agencies are some of the driving factors behind research towards new fuels for aviation. Liquefied Natural Gas (LNG) could be both environmentally and economically beneficial. However, its implementation in aviation has technical challenges that needs to be quantified. This paper assesses the application of LNG in civil aviation using an integrated simulation and design framework, including Cranfield University’s aircraft performance tool, Orion, and engine performance simulation tool Turbomatch, integrated with an LNG tank sizing module and an aircraft weight estimation module. Changes in tank design, natural gas composition, airframe changes, and propulsion system performance are assessed. The performance benefits are quantified against a Boeing 737–800 aircraft. Overall, LNG conversion leads to a slightly heavier aircraft in terms of the operating weight empty (OWE) and maximum take-off weight (MTOW). The converted aircraft has a slightly reduced range compared to the conventional aircraft when the maximum payload is considered. Compared to a conventional aircraft, the results indicate that although the energy consumption is increased in the case of LNG, the mission fuel mass is decreased and CO₂ emissions are reduced by more than 15%. These benefits come with a significant reduction in fuel cost per passenger, highlighting the potential benefits of adopting LNG for aviation.

Liquid hydrogen (LH₂) has long been seen as a technically feasible fuel for a fully sustainable greener aviation future. The low density of the cryogenic fuel would dictate the redesign of commercial aircraft to accommodate the large tanks, which are unlikely to be integrated within the whole internal volume of the wing. In the ENABLEH2 project, the morphological aspects of a LH₂ aircraft design are discussed and a methodology for rapid concept comparative assessment is proposed. An exercise is then carried on to down-select short-to-medium range (SMR) and long-range (LR) concepts, able to carry 200 passengers for 3000 nmi and 414 passengers for 7500 nmi, respectively. The down-selection process was split into two phases with the first considering 31 potential airframe architectures and 21 propulsion-system arrangements. The second phase made the final down-selections from a short-list of nine integrated design concepts that were ranked according to 34 criteria, relating to operating cost, revenue, noise, and safety. Upon completion of the process, a tube and wing design with the tanks integrated into extended wing roots, and a blended-wing-body design were selected as the best candidates for the SMR and LR applications, respectively. Both concepts feature distributed propulsion to maximize synergies from integrating the airframe and propulsion systems.

Geared contra-rotating open rotors have the potential to further reduce fuel consumption relative to geared turbofans, but require a more-complex speed-reducing transmission system to drive the propellers. Hitherto, the preliminary design for such transmission systems has been reported independently of the overall engine modeling or has been limited by many predetermined engine constraints. This has restricted the feasible design space of the transmission systems. A simple transmissions preliminary design approach is needed that does not involve complicated mechanical assessments and can be integrated with engine preliminary design studies. This paper presents a novel design framework for sizing a double-helical differential planetary gearbox for a contra-rotating open rotor. An up-to-date methodology is proposed for the design of maximum load capacity gears for this application based on the power-transmitted, durability, and space envelope requirements. The methodology was validated with published data, demonstrating only very minor differences in geometry dimensions. Parametric analyses have been made to assess the impact of the design assumptions on gearbox dimensions. The framework also enables the identification of feasible torque ratios between the output shafts of the speed-reducer and the contra-rotating propellers driven by them. The impact of the torque ratio on the size of the gearbox has been analyzed for equal propeller rotational speeds and different speed ratios between the output shafts. This study shows that potential torque ratios lie between 1.1 and 1.33, with the higher ratios enabling more compact gearboxes having four or five planet gears.

Conventional propulsion systems are typically represented as uninstalled systems to suit the simple separation between airframe and engine in a podded configuration. However, boundary layer ingesting systems are inherently integrated, and require a different perspective for performance analysis. Simulations of boundary layer ingesting propulsions systems must represent the change in inlet flow characteristics, which result from different local flow conditions. In addition, a suitable accounting system is required to split the airframe forces from the propulsion system forces. The research assesses the performance of a conceptual vehicle, which applies a boundary layer ingesting propulsion system—NASA's N3-X blended wing body aircraft—as a case study. The performance of the aircraft's distributed propulsor array is assessed using a performance method, which accounts for installation terms resulting from the boundary layer ingesting nature of the system. A “thrust split” option is considered, which splits the source of thrust between the aircraft's main turbojet engines and the distributed propulsor array. An optimum thrust split (TS) for a specific fuel consumption at design point (DP) is found to occur for a TS value of 94.1%. In comparison, the optimum TS with respect to fuel consumption for the design 7500 nmi mission is found to be 93.6%, leading to a 1.5% fuel saving for the configuration considered.

Demands for rotorcraft with increased flight speed, improved operational performance and reduced environmental impact have led to a drive in research and development of alternative concepts. Compound rotorcraft overcome the flight speed limitations of conventional helicopters with additional lifting and propulsive components. Further to operational benefits, these augmentations provide additional flight control parameters, resulting in control redundancy. This work aims to investigate the impact of optimal control strategies for a generic coaxial compound rotorcraft, equipped with turboshaft engines, targeting the minimization of mission fuel burn and gaseous emissions. The direct redundant controls considered are: (a) main rotor speed, (b) propeller speed, and (c), fuselage pitch attitude. A simulation tool for coaxial compound rotorcraft analysis has been developed and coupled to a zero-dimensional engine performance model and a stirred-reactor combustor model. Firstly, experimental and flight test data were used to provide extensive validation of the developed models. A parametric analysis was then carried out to gain insight into the effect of the redundant controls. This was followed by the derivation of a generalized set of optimal redundant control allocations using a surrogate-assisted genetic algorithm. Application of the optimal redundant control allocations during realistic operational scenarios has demonstrated reductions in fuel burn and NO_X of up to 6.93% and 8.74% respectively. The developed method constitutes a rigorous approach to guide the design of control systems for future advanced rotorcraft.

Designing commercial aircraft to use Liquid Hydrogen (LH~2~) is one of very few ways to substantially diminish life-cycle CO~2~ emissions from aviation. The merits of hydrogen as an aviation fuel have long been recognized. However, apart from very short-range applications, its integration into commercial aircraft is only viable when it is liquefied, due to the heavy weight and