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.

Hybrid-electric propulsion has emerged as a promising technology to mitigate the adverse environmental impact of civil aviation. Boosting conventional gas turbines with electric power improves mission performance and operability. In this work the impact of electrification on pollutant emissions and direct operating cost of geared turbofan configurations is evaluated for an 150-passenger aircraft. A baseline two-and-a-half-shaft geared turbofan, representative of year 2035 entry-into-service technology, is employed. Parallel hybridization is implemented through coupling a battery-powered electric motor to the engine low-speed shaft. A multi-disciplinary design space exploration framework is employed comprising modelling methods for multi-point engine design, aircraft sizing, performance and pollutant emissions, mission and economic analysis. A probabilistic approach is developed considering uncertainties in the evaluation of direct operating cost. Sensitivities to electrical power system technology levels, as well as fuel price and emissions taxation are quantified at different time-frames. The benefits of lean direct injection are explored along short-, medium-, and long-range missions, demonstrating 32% NOx savings compared to traditional rich-burn, quick-mix, lean-burn technologies in short-range operations. The impact of electrification on the enhancement of lean direct injection benefits is investigated. For hybrid-electric powerplants, the take-off-to-cruise turbine entry temperature ratio is 2.5% lower than the baseline, extending the corresponding NOx reductions to the level of 46% in short-range missions. This work sheds light on the environmental and economic potential and limitations of a hybrid-electric propulsion concept towards a greener and sustainable civil aviation.

Hydrogen for propulsion could lead the industry to achieving the set environmental goals. This work performs a comparative cycle and engine design for a hydrogen-fired and conventional Jet-A burning configuration. Aircraft design and mission performance complete the conceptual design loop. A 19-passenger small commuter aircraft is investigated. A multi-disciplinary framework is developed for the study. A multipoint synthesis scheme is employed for conventional, and hydrogen powered engine assessment. Cryogenic hydrogen tanks and a thermal management system are integrated in the aircraft and their volumetric and gravimetric impact are assessed showing an expected 13% increase of aircraft mass. Direct combustion of hydrogen leads to more efficient and smaller turboprop engines due to the increased specific heat capacity of combustion products. Engine cycle design reveals that the optimum aircraft energy consumption lies at an overall pressure ratio at cruise of 15 and a combustor outlet temperature of 1,400 K. However, the hydrogen tanks and fuel conditioning system increase aircraft mass sufficiently to offset the fuel consumption benefits resulting in increased block energy demand between the H2 and Jet-A configurations by 3–4%. This work reveals a lower bound of H2 fuel system gravimetric performance due to aircraft weight certification and highlights the design guidelines for the engine, aircraft, and storage system.